Pilatus PC-12 Pilot s Information Manual

Pilot's Information

Manual

Revision

PILOT'S INFORMATION

MANUAL

Reprinted 1 November, 2003; Including revisions 1 through 11

PILOT'S OPERATING HANDBOOK AND

FOCA APPROVED AIRPLANE FLIGHT MANUAL

PC-1 2

MSN 101-320 AND 322-400

Report No. 01973-001

Manufacturer's Serial No. -

Registration No. -

FOCA Type Certificate No: F56-30 FAA Type Certificate No: A78EU

PILATUS AIRCRAFT LTD. CH-6370 STANS SWITZERLAND

FOCA APPROVED IN THE NORMAL CATEGORY BASED ON FAR 23 THROUGH AMENDMENT 42. THlS DOCUMENT MUST BE CARRIED IN THE AIRPLANE AT ALL

I TIMEs.

THlS HANDBOOK INCLUDES THE MATERIAL REQUIRED TO BE FURNISHED TO THE PILOT BY THE FEDERAL AVIATION REGULATIONS AND ADDITIONAL INFORMATION PROVIDED BY THE MANUFACTURER AND CONSTITUTES THE FOCA APPROVED AIRPLANE FLIGHT MANUAL.

This Handbook meets General Aviation Manufacturer's Association (GAMA) Specification No. 1, Specification For Pilot's Operating Handbook, issued 15 February 1975, revised 1 September 1984.

This Handbook is also FAA approved for U.S. registered aircraft in accordance with FAR 21.29.

Approved by - Federal Office for Civil Aviation (FOCA) of Switzerland

Signature -

Initial Issue - March 30. 1994

BPILATUS AIRCRAFT LIMITED 1994 All rights resewed

ePILATUSE PC12 LIST OF SERVICE BULLETINS INCORPORATED

Aircraft Serial No:

This list of Service Bulletins are those that are mentioned in the various Sections of the POH. It is not a complete list of SB's. The purpose is to show the modification status of the aircraft to / assist the pilot in the correct use of the Emergency and Normal Procedures.

I

Mark an X in the box when the Service Bulletin has been incorporated on the aircraft.

SB No.

21-001

21-002

21 -003

21-004

23-003

24-004

24-008

24-010

25-001

25-005

25-008

25-010 ,

Issued: February 14, 1994 Revision 10: September 1,2000

Report No: 01 973-001 LOSB-1

Title

Optional removal of the cabin outflow solenoid valve fro the pressurization system. lncorporated in build from MSN 141

Optional modification of the auxiliary heating system. lncorporated in build from MSN 181

Optional replacement of the ECS temperature control valve. lncorporated in build from MSN 181

Optional installation of pilot and copilot foot heaters

Optional replacement of 25 kHz VHF transceivers with 8.33 kHz

Optional replacement of the 30 second starter sequence relay for a 60 second relay. lncorporated in build on MSN 181-230. Not applicable from MSN 231 new starting system

Optional installation of an external power control unit. lncorporated in build from MSN 231

Optional replacement of Gen 2 and removes Gen 2 reset switch. lncorporated in build from MSN 231

Optional modification to the crew seats adjustment mechanisms. lncorporated in build from MSN 162

Optional installation of passenger seat luggage restraint bar

Optional installation of a passenger door thermal blanket

Optional installation of a larger baggage net

Incorporated

r PILATUSW LIST OF SERVICE BULLETINS INCORPORATED PC12

Mark an X in the box when the Sewice Bulletin has been incorporated on the aircratt.

build from MSN 321

Incorporated in build from MSN 181

Report No: 01 973-001 LOSB-2

24-017

25-022

Issued: February 14, 1994 Revision 1 1 : March 1,2003

Optional replacement of nickel cadmium batteries with lead acid batteries

Optional replacement of ELT Narco 910 with Kannad 406 AF

LIST OF SERVICE BULLETINS INCORPORATED


SB No.

25-024

25-026

31-005

32-013

33-007

33-009

35-003

79-005

Issued: February 14, 1994 Revision 1 1 : March 1.2003

Report No: 01973-001 LOSB-3

Title

Optional installation of a smaller ELT remote control panel

Optional installation of a pilots relief tube

Mandatory replacement of CAWS computer unit

Optional replacement of brake units, with carbon heat sinks for steel heat sinks

Optional installation of baggage compartment light

Optional installation of dual filament navigation lights

Optional installation of an oxygen mask in the toilet compartment

Optional installation of chip detector in engine accessory gearbox

Incorporated

SPILATUSS LIST OF SERVICE BULLETINS INCORPORATED PC12


Report No: 01973-001 LOSE-4

Issued: February 14,1994 Revision 10: September 1,2000

XPILATUSF PC12 LIST OF EFFECTIVE PAGES

Page No. Rev. No.

Title 1 LOSB-1 10 LOSE-2 and 3 11 LOSB-4 10 LOEP-1 and 2 11 LOTR 1 9 LOTRQ 10 LOTR-3 11 LOTR-4 10 LOR-1 & 2 1 LOR-3 thru 5 2 LOR-6and7 3 LOR-8 and 9 4 LOR-10 thru 13 5 LOR-14 thru 16 6 LOR-17 and 18 7 LOR-19 8 LOR-20 9 LOR-21 8 LOR-22 thru 24 9 LOR-25 thru 27 10 LOR-28 thru 30 11 CONTENTS-1 1 0-i 10 0-ii 1 0-1 1 0-2 and 0-3 10 0-4 1 1-i & 1-ii 8 1-1 1 1-2 2 1-3 and 1-4 1 1-5 5 1-6 2 1-7 5 1-8 1 1-9 7 1-10 4 1-1 1 thru 1-20 8 2-1 2 2-ii and 2-iii I I

LlST OF EFFECTIVE PAGES

Page No. Rev. No.

2-iv 4 2-1 7 2-2 2 2-3 3 2-4 5 2-5 9 2-6 10 2-7 3 2-8 11 2-9 7 2-10 2 2-1 1 3 2-12 9 2-1 3 6 2-14 thru 2-19 8 2-20 thru 2-22 9 2-23 thru 2-44 11 3-1 10 3-ii thru 3-iv 11 3-1 10 3-2 5 3-3 6 3-4 and 3-5 9 3-6 2 3-7 8 3-8 2 3-9 11 3-10 2 3-1 1 5 3-12 10 3-13 3 3-14thru3-16 11 3-17 2 3-18 thru 3-64 11 4-1 8 4-11 10 4- 1 9 4-2 6 4-3 and 4-4 8 4-5 and 4-6 10 4-7 thru 4-9 11

Page No. Rev. No.

4-10 8 4-1 1 10 4-12 and 4-13 8 4-14and4-15 11 4-1 6 6 4-17 8 4-1 8 9 4-1 9 and 4-20 6 4-21 10 4-22 9 4-23 and 4-24 6 4-25 9 4-26 thru 4-29 11 4-30 8 4-31 thru 4-41 6 4-42 thru 4-46 8 4-47 10 4-48 6 4-49 and 4-50 10 5-i 8 5-ii 5 5-iii 6 5-iv 2 5- 1 7 5-2 thru 5-9 2 5-10 and 5-1 1 5 5-12 2 5-13thru5-16 5 5-17 7 5-18 thru 5-23 10 5-24 5 5-25 thru 5-28 7 5-29 5 5-30 thru 5-39 7 5-40 thru 5-49 5

5-50 7 5-51 and 5-52 5 5-53 7 5-54 and 5-55 5 5-56 7 5-57 thru 5-66 5

Page No. Rev. No.

5-67 thru 5-69 7 5-70 thru 5-77 6 5-78 5 5-79 10 5-80 thru 5-84 6 6-i 11 6-ii 9 6- 1 9 6-2 and 6-3 11 6-4 9 6-5 and 6-6 11 6-6A and 6-6B 11 6-7 thru 6-26 9 6-27 11 6-28 thru 6-31 9 6-32 11 6-01-1 9 6-01 -2 11 6-01 -3 10 6-01-4 thru -6 9 6-02-1 9 6-02-2 and 3 11 6-02-4 thru -6 9 6-03-1 9 6-03-2 and -3 1 1 6-03-4 thru -6 9 6-04-1 9 6-04-2 and -3 11 6-04-4 thru -6 9 6-05-1 9 6-05-2 and -3 1 1 6-05-4 thru -6 9 6-06-1 9 6-06-2 and -3 11 6-06-4 thru -6 9

6-07- 1 9 6-07-2 and -3 11 6-07-4 t h r ~ -6 9 6-08-1 9 6-08-2 and -3 11 6-08-4 thru -6 '9

Issued: February 14. 1994 Revision 11: March 1,2003

Report No: 01973-001 LOEP-1

WPILATUSP LIST OF EFFECTIVE PAGES PC12

LlST OF EFFECTIVE PAGES

Page No. Rev. No.

7-i 6 7-ii 11 7-iii 8 7-iv and 7-v 11 7-vi 3 7-1 and 7-2 2 7-3 3 7-4 5 7-5 thru 7-7 11 7-7A and 7-76 11 7-8 2 7-9 5 7-10 9 7-1 1 2 7-12 6 7-1 3 10 7-14and7-15 2 7-16 6 7-17 3 7-18 9 7-1 9 10 7-20 8 7-21 11 7-22 2 7-23 9 7-24 3 7-25 and 7-26 11 7-27 5 7-28 4 7-29 2 7-30 5 7-31 and 7-32 2 7-33 6 7-34 thru 7-37 2 7-38 11 7-39 5 7-40 9 7-41 2 7-42 9 7-43 and 7-44 10 7-45 thru 7-48 9

Page No. Rev. No.

7-48A and 488 9 7-49 9 7-50 10 7-51 and 7-52 7 7-53 5 7-54 3 7-55 4 7-56 and 7-57 5 7-58 3 7-59 5 7-60 3 7-61 9 7-62 and 7-63 3 7-64 7 7-65 11 7-66 10 7-67 11 7-68 10 7-69 11 7-70 and 7-71 8 7-72 thru 7-74 11 7-75 10 7-76 8 7-76A 11 7-766 9 7-77 and 7-78 10 7-79 11 7-80 10 7-81 11 7-82 6 7-83 10 7-84 6 7-84A & 7-846 10 7-85 3 7-86 and 7-87 6 7-88 11 7-89 and 7-90 8 7-90A thru 7-90F 10 7-91 thru 7-93 3 7-94 and 7-95 5 7-96 3

Page No. Rev. No.

7-97 thru 7-100 5 7-101 10 7-102 11 7-103 thru 7-105 9 7-106 3 7-107 8 7-108and7-109 4 7-1 10 3 7-1 11 10 7-1 12 11 7-1 13 4 7-1 14 6 7-1 15 5 7-116 11 7-117 4 7-1 18 5 7-119 10 7-120 11 7-121 and 7-122 9 7-123 4 7-124 5 7-125 9 7-1 26 5 7- 127 11 7-128 and 7-129 10 7-130 thru 7-134 11 7-135 thru 7-151 3 7-152 6 7-153 thru 7-163 3 7-164and7-165 9 7-1 66 3 7- 167 6 7-168 and 7-169 5 7-170 4 7-171 thru 7-173 3 7-174 7 7-1 75 thru 7-178 3 8-i & 8-ii 10 8- 1 5 8-2 1 8-3 6

Page No. Rev. No.

8-4 10 8-5 1 8-6 2 8-7 11 8-7A and 8-78 1 1 8-8 and 8-9 1 8-10 5 8-1 1 8 8-12 5 8-13 8 8-14and8-15 1 8-16 thru 8-19 10 8-20 and 8-21 11 8-22 and 8-23 8 8-24 thru 8-32 1 9-i and 9-ii 11 9-00-1 and -2 11 104 10 10-ii 1 10-1 thru 10-3 6 10-4 thru 10-7 7 10-8 and 10-9 9 10-10 thru 10-12 10

Report No: 01 973-001 LOEP-2

Issued: February 14, 1994 Revision 1 1 : March 1, 2003

LOG OF TEMPORARY REVISIONS


The incorporation of Temporary Revisions into this manual are to be recorded on the sheet below. Instructions for the removal of Temporary Revisions will given in the Instruction Sheet issued wlh each regular revision.

Issued: February 14.1994 Revision 9: September 1, 1999

.

Report No: 01973-001 LOTR-1

NO.

1

2

3

4

St

S2

S3

5

6

7

8

9

10

TEMPORARY REVISION TITLE

8 SEAT EXECUTIVE

COMMUTER BENCH SEAT

AHRS FAILURE

FLAP ACTUATORS

SUPP 08 - COMMUTER LOADING FORM

SUPP 08 - COMM BENCH SEAT LOAD FORM

SUPP 0 8 - 8 SEAT EXEC LOADING FORM

EXECUTIVE BENCH SEAT

6 EXEC AND 2 STAND SEAT FIT

ENGINE FAILURE

FLAP RESET SWITCH

LOGO LIGHTS

ENGINE INSTRUMENT SYSTEM

DATE OF ISSUE

SEP 2/98

OCT 27/98

NOV 5/98

NOV 18/98

NOV 4/98

NOV 5/98

NOV 5/98

DEC 2/98

JAN 18/99

JAN 26/99

JAN 27/99

MAR 16/99

APR 1/99

CANCELLED BY

POH REV 9

POH REV 9

POH REV 9

AMM TR 04-02

POH REV 9

POH REV 9

POH REV 9

POH REV 9

POH REV 9

POH REV 9

POH REV 9

POH REV 9

POH REV 9

'FdPILATUSW LOG OF TEMPORARY REVISIONS K12

LOG OF TEMPORARY REVlSlONS

Report No: 01 973-001 LOTR-2


CANCELLED BY

POH REV 9

POH REV 9

POH REV 10

POH TR 21

POH REV 10

POH REV 10

POH REV 10

POH REV 10

POH REV 10

POH REV 10

SB 30-006

POH TR 25

POH REV 10

POH REV 10

POH REV 10

DATE OF ISSUE

APR 16/99

APR 19/99

DEC 22199

DEC 17/99

APR 3/00

JAN 21/00

MAR 1/00

MAR 2W00

MAR 23/00

MAR 28/00

MAY 18/00

MAY 23/00

MAY 1 1/00

MAY 18/00

JUN 27/00

NO.

11

12

13

14 -

15

16

17

18

19

20

21

22

23

24

25


GEN 1 AMPNDC INDICATOR

PREPAREDUNPAVEDSURFACES

CABIN INTERIOR BULKHEAD 8 CURTAIN

WINDSHIELD HEATER OPERATION

MODIFIED FLAP SYSTEM

LEAD ACID BATTERY INSTALLATION

400 AMP STARTERIGENERATOR

WEIGHT ON WHEELS SIGNAL

CABIN PLACARDS

EMERGENCY POWER SYSTEM (EPS)

WINDSHIELD HEATER OPERATION

MSN 321 8 UP HEATING SYSTEM

FUEL ANTI-ICING ADDITIVES

VARIOUS CHANGES (Not issued)

MODIFIED HEATING SYSTEM

r PILATUSW PC12 LOG OF TEMPORARY REVISIONS




NO.

26

27

28

29

30

31

32

33

34

35

36

37

DATE OF ISSUE

Oct 26/00

Dec 1 1 100

Dec 5/00

May 17/01

Not Issued

Dec 21/00

Jan 8/01

Oct 26/01

Jan 30102

Aug 19/02

Sep 27/02

Jun 6/03


DUAL LEAD ACID BATTERY INSTALLATION

NEW EIS INSTALLATION FOR MSN 321

NEW ELT INSTALLATION FOR MSN 321

NEW CAWS & OHP INST FOR MSN 321

NEW FLAP POWER DRIVE UNIT

PREVENTING OF FROZENlLOCKED BRAKES

FLAP LIMITATIONS

ELECTRICAL LOAD SHED PROCEDURE

ELECTRICAL LOAD SHED PROCEDURE

CREW OXYGEN MASK DONNING

AHRS LIMITATION

FUEL LIMITATIONS

CANCELLED BY

POH REV 10

NEW POH

NEW POH

NEW POH

POH REV 10

POH REV 10

SB 27-01 1

TR 34

POH REV 11

POH REV 11

POH REV 11

SB 28-01 1

€ PILATUS W LOG OF TEMPORARY REVISIONS PC12



NO.

Issued: February 14.1994 Revision 10: September 1,2000

TEMPORARY REVISION TITLE DATE OF ISSUE

CANCELLED BY

WARNING!

This PC-12 Pilot's lnformation Manual is published for familiarization and training purposes ONLY!

This Pilot's lnformation Manual does not meet FAA, FOCA or any other civil aviation authority regulations for operation of any aircraft!

This Pilot's lnformation Manual is a reproduction of a PC-12 Airplane Flight Manual, however it is not revised or updated in any manner after printing.

This Pilot's lnformation Manual does not reflect the configuration or operating parameters of any actual aircraft.

Only the Approved Airplane Flight ManuallPilot's Operating Handbook issued for a specific serial number aircraft may be used for actual operation of that serial number aircraft.

Pilatus Aircraft, Ltd. Stans, Switzerland

Pilatus Business Aircraft Limited Broomfield, Colorado USA



PllATUS AIRCRAFT LTD. CH-6370 STANS SWITZERLAND .

FOCA APPROVED IN THE NORMAL CATEGORY BASED ON FAR 23 THROUGH AMENDMENT 42. THlS DOCUMENT MUST BE CARRIED IN THE AIRPLANE AT ALL TIMES.

THlS HANDBOOK INCLUDES THE MATERIAL REQUIRED TO BE FURNISHED TO THE PILOT BY THE FEDERAL AVIATION REGULATIONS AND ADDITIONAL INFORMATION PROVIDED BY THE MANUFACTURER AND CONSTITUTES THE FOCA APPROVED AIRPLANE FLIGHT MANUAL

Thls tlandbook meels General Avlatlon Manufaclurer's AssodaUon (GAMA) Speclflcatlon No. 1, SpecificaUon For Pilot's Operating Handbook, Issued 15 February 1976, revlsed 1 September 1984.

Thls Hendbook Is also FAA approved for U.S. registered alrcraft In accordance wth FAR 21.29.

Slgnature -

BPILATUS AIRCRAFT LIMITED 1994 Afl rlgtlls resewed

31E PILATUS W SC XI1 LOG OF REVISIONS

LOG OF REVISIONS

Issued: February 14, 1994 Revision 1: June 10, 1994

Report NO: 01 973-001 LOR-1

Description

Reformatted whole document to A5 size. Section Table of Contents revised accordingly. Relocated LOEP, LOR, 8 POH Contents pages before Section 0. Revised page numbering accordingly. Added information to Symbols, Abbreviations, and Terminology. ~dcied pages.

Revised Airspeed Limits. Revised Misc. Instrument Matkings. Revised Weight Limits. Revised Center of Gravity Limits. Added Kinds of Operation. Added Equipment to the KOEL.

Revised Windshield Limitation. Added Autopilot Limitations. Added pages. Revised Placard List.

Revised Emergency Procedures Checklist.

Added pages. Added Autopilot Procedures.

Revised terminology and airspeed. Revised Normal Procedures Checklist.

Added pages. Added Autopilot Procedures.

Revised Performance Chart Titles.

Revised Peflomance Charts. Revised Performance Charts.

?

Revision Number

and Date

1

10 June 94

Page Number

All

Sect. 0

Sect. 1 1-17 thtu

1-20 2- 1 2-9 2-10 2-1 1 2-12

2-13 thru 2-16 2-18 2-20

2-25 thru 2-34

3-5, 3-7, 3-10,3-13 t h ~ 3-16, 3-18,3-19, 3-21 t h ~ 3-25, 3-29 thru 3-34 3-35 thru

3-52 4-2

444-7 , 46, 4-10 t h ~ 4-14, 4-16, 4-17, 4-19, 4-20 4-21 thnr

4-40 5-2 thru

5-6 5-7 8 5-8 5-14 thru

5-20

LOG OF REVISIONS

LOG OF REVISIONS (CONT.)

Report No: 01 973-001 LOR-2

Revision Number

and Date

Issued: Febfuafy 14, 1994 Revision 1 : June 10, 1994

9-02-4 O Federal Office 13 ,&Y /4294 for Civ i l Aviation 0

Page Number

5-21 thru 5-26

6-1 thru 6-3. 6-5,

6-9 6-26,6-28 6-35 thru

6-46 Sect. 7

7-131 thru 7-176 8- 1 8-4 8-5

8-1 0 8-27 thru

8-32 9-00-1 8 9-00-2

9-01-1 thru 9-0 1-4

9-02-1 thru

Description

Added pages. Revised Performance Charts.

Revised aircraft weighing procedure.

Revised Weight and Balance Determination for Right. Added pages. Revised Figure 7-8. Loading Form.

Misc. changes throughout for new CAWS panel, autostart, auto ignition, Heating 8 Cooling systems, AHRS switching, standby power system, cabin fire extinguisher, rear bulkhead safety net, copilot windshield, cargo door motor. Revised cockpit switch markings, circuit breaker panels, flap 8 landing gear system, stick pusher system, oxygen system, ECS, pressurization system, ground ops prop rpm warning, 8 torque limiter setting. Removed windshieM demist system and . maximum tie down weight recommendation. Added pages. Added optional MFD, Wx Radar, 8 Autopilot info.

Revised Identification Plate location. Revised tc w limits. Revised Figure 8-1. Revised leveling procedure. Added pages.

Revised page numbers.

Added Supplement No. 1, BendiKing KLN 90A GPS.

Added Supplement No. 2, BendidKing

LOG OF REVISIONS

LOG OF REVISIONS

Issued: February 14,1994 Revision 2: February 14, 1995

Report No: 01 973-00 1 LOR-3

Description

Revised Aircran dimensions Revised Fuel capacities Revised Weights and loadings for 4100 kg MTOW Revised ElectricaVAvionic Abbreviations

Revised Airspeed Limitations for 4100 kg MTOW Revised Airspeed Indicator KlAS value Revised Engine Operating Limits table heading Minor changes Minor changes Revised Weight Limits for 4100 kg MTOW Minor change Revised Fuel capacities Minor changes Revised Placard Revised Placards

Minor change Revised Weights and Speeds for 41 00 kg MTOW Revised Emergency Procedures

Revised Airspeeds for Normal Operations for 4100 kg MTOW

Revision Number

and Date

2 14 Feb 95

Page Number

Sect. 1 1 -2 1-5

1-6 8 1-7 1-17

Sect 2 2-1 2-3 2 4

2-6 & 2-7 2-9

2-10 & 2-1 1 2-12 2-17

2-19 & 2-21 2-22

2-24 lhru 2-27

Sect 3 3- 1 3-2

3 4 8 3-5 3-7 & 3-8

3-10 & 3-1 1 3-15 thN

3-2 1 3-25 8 3-26 3-29 8 3-30

3-32 t h ~ 3-35 3-40

3-43 8 3 4 4 3-46 & 3 4 7 3-50 thN

3-52 Sect 4

4-2 4-7 8 4-8 4-10 l h ~

4-12

LOG OF REVISIONS -- --


Report No: 01973-001 LOR-4

Issued: February 14, 1994 Revision 2: February 14, 1995

Description

Revised Performance Graphs

Added pages for new Performance Graphs

Minor change Added pages. Executive Interior Seat Locations Minor change Added Executive Interior Minor changes Added Executive lnterior Minor change Added Loading Limitation Added Executive lnterior Added Executive Interior Moment C:,arts Revised Fuel Moments Revised Example Loading Form Revised MRW and MTOW weights Added Executive lnterior Loading Form Revised MRW and MTOW weights Revised C of G Envelopes Revised pages throughout to reflect latest information and indude s assisted cargo door dosing option, revised fuel indcation system and radar altimeter. Effectivities of specifiic equipment added.

Revised IdentificationPlate information Revised parking information Minor change

Revis~on Number

and Date

Page Number

4-14 8 4-15 4-17 8 4-18

4-20 l h ~ 4-23

4-26 WIN 4-28 4-30 4-32

4-34 lhru 4-39

Sect 5 5-1 lhru

5-26 5-27 lhru

5-84 Sect 6

6-3 6-7 8 6-8

6-9 6-1 3

6-14 8 6-15 6-16 6-1 7 6-19 6-27

6-31 8 6-32 6-35 8 6-36

6-38 6-39 6-40 6-41

6-43 8 6-44 Sect 7

Sect 8 8- 1 8-6 8-16

PILATUS W ?c XI1 LOG OF REVISIONS

Issued: February 14,1994 Revlsion 2: February 14, 1995


Description

Revised Information

r

Revlslon Number and Date

Page Number

Sect 9 9-02-1 thru 942-3

LOG OF REVISIONS



Revision Number and Date

3 29 Sep 95


Page Number

Sect 1 1-5 1-7 1 - 18

Sect 2 2-1 2-3 2 4 2-6 2-8 2-7 2-1 1 2-1 5 2-1 7 2-18 2-19 2-20 2-21

2-26 thru 29 2-34 thru 36

Sect 3 3-iii 3-2 3-3 3-5 3-1 3 3-15 3- 16 3-21 3-32 3-33 3-34 3-35 3-37 3-38

Sect 4 4-5 and 4-6

4-7 4-10 4-12 4-24 4-36 440

Sect 5 5-31 and 65

Descriptron

Fuel conversion values standardized. Oil capacity corrected Emergency exit dimensions corrected Fuel conversion values standardized

Airspeed Limitations - flap degree values clarified Airspeed Indicator Markings - KlAS White Arc value changed Engine Operating Limits - min idle Ng % info changed Oil - tank capacity corrected, fuel anti-icing paragraph changed Chip detector - first sentence deleted Propeller - Made angles at station conected Center of Gravity limits - weight correction KOEL - AOA Deice added Fuel Limitations - conversion values standardized Seating Limits - corporate commuter and executive clarified Oxygen System - limits expanded Autopilot - limits corrected Cargo tie down straps limitation clarified Placards corrections and additions Placards corrections and additions. Pages added

Contents page update Airspeeds for Emergency Operations - 7060 Ibs added Rejected Takeoff - wheel and brake caution added Engine Failure in Flight - paragraph correction Propeller Underspeed - paragraph correction CockpitlCabin Fire, Smoke or Fumes - paragraph addition Smoke Evacuation - paragraph addition Forced Landing - paragraph mrreclion Battery Bus Failure - services inoperative addition Gen 1 Bus Failure - services inoperative addition Gen 2 Bus Failure - services inoperative addition Gen 1 Off - effectivity added and services inoperative addition Bus Tie - effectivity added N EsnU Bus - effectivity added and new paragraph added

PreRight Inspection - correction and addition Preflight Inspection - addition Engine Starting - correction Engine Starting - correction Oxygen control valve set to AUTO at 10,000 feet Autopilot Operation - rad alt caution wording changed Noise Level - values changed

Schedule and ~raph correctiom

SPLATUSW PC XI1 LOO OF REVISIONS


Issued: February 14,1994 Revision 3: September 29.1995

Revlslon Number and Date


Page Number

Sect6 6-6 6-8 6-9 6-15

6-38 thru 41 Sect 7

7-47 thru 7-178 Sect 8

8-1 and 8-R 8-4 8- 16 8-20

Sect 9 9-1

9-03-1 thru 9-034

944-1 thnr

Description

Corporate Commuter - illustration vlew change Paragraph deleted Fuel conversion values standardized Moment value corrected Minor conections Misc. changes throughout to ghre Information on allemn/~dder Interconnect, hydraulic system, seats, pass. oxygen masks stowage, CAWS and executive cabin features. Added pages for new EIS and changes to fuel and elecMcal systems

Contents pages updated Service Bulletin Info a d towing condition added Fuel capacity standardized Tire pressures updated

Contents page updated Added Supplement No. 3, Bendix/Klng KHF 950

Added Supplement No. 4. Bendixn<ing KLN 908 GPS

O Federal Office for Clvl l Avia!lon

A 3 f f o ~ . J995

BPLATUSW LOG OF REVISIONS ?c XI1

-- -- -- - -



I


4 24 Nov 95

Issued: February 14,1994 Revision 4: November 24,1995

Page Number

Sect 1 1-10 and

1-11 Sect 2 2-12 2-13

2-14 thru 2-17 2-19 2-20

2-25 2-26 2-27

Sect 3 3-2

3-35 and 3-36 3-45 3-46 3-47 348 3-49 3-50 3-51 3-56

Sect 4 4-2 4-5

4-1 1 and 4-12 4-14 4-15

4-1 6 and 4-17 4-17 4-18 4-19 4-20

4-21

4-22 4-24

Description

lcing Meteorological Terminology added

Flight into known icing conditions added lcing Limitations added lcing conditions added to the KOEL

Ice mode added to stall waminglstick pusher system limits lcing conditions added to pneumatic deice system limits and probe heat system limits added Flap extension limits in icing conditions placard added Change to magnetic compass placard Airplane approved for icing conditions added to Operational Limitations placard

Landing approach speeds for deice system equipment failures added Terminology change to AOA heater plates

Propeller deice failure added Boot deice failure added Inertial separator failure added LH windshield deice failure added AOA probe deke failure added Pitot and static probe delce failure added Pusher ice mode failure added Alt Loss for APR 3" ILS and 6' ILS interchanged

Airspeeds for normal operations in king conditions added Propeller deicing boots pre flight inspection changed Inertial separator procedure added

Pusher ice mode added to before taxiing test lcing conditions procedures added to before taxiing procedures lcing conditions procedures added to before takeoff procedures

Takeoff - power control lever setting pr0~edure Changed Flight into known icing cortdiions procedures added Climb - power control lever setting procedure changed Cruise - power control lever setting procedure changed, ice protection system procedure added. WSHLD contrd corrected Approach ice protection system and inertial separator procedure added Balked landing ice protection system procedure added Shutdown procedure changed -

LOG OF REVISIONS


Issued: Fel Revision 4:

Number page I

4-42 thru 4-45 4-46

Sect 5 5-40 thru

5-79 5-80 and

5-81 Sect 7

7-7 7-167 Sect 9

9-04-1 thru 9-04-12

9-05-1 thru 9-05-4

law 14.1994 ovember 24,

Description

Amplified procedure for flight in icing conditions added

Figure added for Ice Protection Systems CAWS Advisory

Maximum cruise power performance removed and figure numbers changed Flight in icing conditions added

Deicing systems description and operation updated to include all the new deicing equipment New triple trim indicator added Autopilot outputs to triple trim indicator lights added

Supplement No. 4, BendixlKing KLN 908 GPS revised

Added Supplement No. 5, Argus 5000 Electronic RMIhloving Map Display


LOG OF REVISIONS

Number

Sect 1

1-7 Sect 2

2-1

2-4

2-5 2-13 2-18

2-23 2-26 2-34

2-35 Sect 3

3-ii and 3-iii 3- 1 3-2 3-4 3-5 3-1 1 3-15 3-16 3-18 3-19

3-21 3-26 3-28 3-29

3-30 3-32 3-35 3-36 3-37 3-39 3-40 3-41 3-42 3-43 3-44

01973-001


I Description I

Total fuel capacity values standardized Aircraft config change for usable fuel given, oil quantity correction Correction to power loading figure

Additional wording added to general para Mathematical symbol for max. flap extended speed changed Note 9 added to takeoff condition Ax cont and max climb defined Note 9 explanation given Definition of Icing conditions given and further info added Total fuel capacity values standardized Aircraft config changes for total usable and unusable fuel given Aircraft config total usable capacity placards given Applicability added to Standby Compass placard Liters added to oil capacity placard, oil type deleted and note added to oil type placard Aircraft config fuel capacity placards given

Contents pages updated Aural Wamings description paragraph updated Landing app-oach speeds clarified, balked landing info added Engine failure after rotation - landing info clarified Airspeed changed Fuel control unit failure - landing distance clarified Editing and page layout changes Page layout change Windshield heat action changed Maximum Rate Descent procedure updated. Emergency Descent Profiles chart added Forced landing - note added ref silencing flap aural warning Circuit breaker (CB) location added Caution added ref flap asymmetry Para 3.13 renamed to include both inadvertent pusher and shake Para 3.13.1 allocated for inadvertent pusher operation Para 3.13.2 added for inadvertent shaker operation Note added ref EFlS CMPST switch Gen 1 Off procedure updated Gen 2 Off and Bat Off procedures updated CB location and Note added ref EFlS CMPST switch EIS failure procedure for MSN 101-1 11 expanded EIS failure procedure for MSN 112-999 added Page run on and applicability added to VIA switch Page run on and CB location added Page run on. CB location and new para 3.17.5 added Page re-arranged

Issued: February 14,199 R~\,isinn 5. May 10, 198

PC XI1 LOG OF REVISIONS



Page Number

3-45 3-46 3-47 3-48 3-49 3-50 3-51 3-53 3-54 3-55 3-56 3-57

Sect 4 4-2 4-7 4-8 4-10 4-1 1 4-12 4-14 4-15

4-16 4-17 4-18 4-19 4-20 4-21 4-22

4-26 4-38

4-42 thru 4-46

Sect 5 5-i thru 5-iii

5-1 5-10 and

5-1 1 5- 13 thru

5-77 5-78 thru

5-80 Sect 6

6-9 6-1 1 8 6-12

Description

Para title, warning and CB location clarified. Procedure expanded Para title, warning and CB location clarified. Procedure expanded Para title, warning and CB location clarified. Procedure expanded Para title and CB location clarified Para title, warning and CB location clarified. Procedure expanded Para title, warning and CB location clarified. Procedure expanded Para title, warning and CB location clarified. Procedure expanded Pax changed to passenger. AHRS failure procedure expanded Note added ref EFlS CMPST switch CAWS malfunction procedure added CB location added CB location added

Information expanded Control lock stowage clarified and 2nd AHRS procedure added Page run on Parameter change and note clarified Gen switching changed to separate steps and Ng speed change Note clarified and Gen switching changed to separate steps Pusher test updated Pusher test page run on, inertial separator check moved from If lcing Conditions expected checks to normal checks Page run on. Trim setting clarified Page run on. Reference to Section 5 charts given Page run on. Warning expanded Page run on. Flight into Known lcing Conditions para expanded Page run on. PCL setting reference to Section 5 charts given Before Landing checks updated with icing information Page run on. Crosswind reference updated. Balked Landing icing information added Oxygen abbreviations changed to wording Autooilot alideslooe info exoanded ~rn~iified>roced;re for ~l i$ht in lcing Conditions updated and expanded

Contents updated Additional general information added Graphs re-issued to a standardized format, effect of icing conditions info added to appropriate graphs and reverse thrust graphs revised and retitled to Short Field. Total Number of graphs reduced to 61 Flight in lcing Conditions, page numbers changed and information updated and expanded

Basic Empty Weight form usable fuel figures updated Weight and Balance Record forms made usable for PC-12 & 145

Issued: February 14, 1994 Revision 5: May 10, 1996

Report No: 01 973-001 LOR-11 I

EPILATUSI LOG OF REVISIONS ?c XI1


Report No: 01 973-001 I LOR-12

- Revision Number and Date

Issued: February 14,1994 Revision 5: May 10, 1996

Page Number

6-35 & 6-36 Sect 7

7-4 7-9

7-16 7-27 7-30 7-33 7-39 7-53 7-56 7-57

7-59 7-65 7-69

7-70 8 7-71

7-72 & 7-73 7-75 7-88

7-94 & 7-95 7-97 thru

7-100 7-102 7-1 12 7-1 14 7-1 15 7-1 16 7-1 18 7-1 19 7-121 7-1 22 7-1 24 7-126 7-1 30 7-132 7-168 7- 169 Sect 8

8-1 8-10 8-12 8-16

Description

"

Fuel Moment Charts updated

Trim description clarified Tire pressures deleted Landing Gear figure updated with aircraft configurations Contrd lock stowage point description added Inertial separator operation corrected Propeller Deice figure switch placarding corrected Engine Oil System figure editing changes Redundant paragraph deleted Propeller deice description updated Wing fuel capacity for aircraft configurations and SB 28-001 added Fuel System figure updated to include SB 28-001 Additional info added to the Power Supplies description Electrical Power System figure updated CB locations deleted and replaced with typical CB panel figures for MSN 101-120 Typical CB panel figures for MSN 121 and upwards included Typical Overhead Panel figure updated with MSN configs Heating System Operation additional info added Cabin Pressurization System updated to include SB 21-001 Cabin Pressurization System updated to indude SB 21-001

Oxygen Control Valve position corrected AOA DE ICE caption description clarified Pitot Static description updated Typical Pilot and Static figure updated with icing mods Stall WarningIStick Pusher icing description clarified Stall WarningIStick Pusher operation clarified Stall WarninglStick Pusher CAWS indications clarified Stall WarningIStick Pusher schematic updated with icing mods Stall WarninglStick Pusher figure CAWS readouts changed Pneumatic De Ice failure para wording change Pneumatic De Ice schematic updated with icing mods Typical Avionics schematic typing error correction 2nd AHRS and Yaw Rate Giro installations added Autopilot Mode Controller DNlUP control description clarified Page run on

Pilatus department, tel and fax numbers updated Caution added reference jacking the aircraft Jacking figure updated to show ballast attached to tail jack point Fuel quantities deleted

LOG OF REVISIONS


Issued: February 14, 1994 Revision 5: May 10. 1996

Report No: 01973-001 LOR-13 I

Description

Second type of hydraulic oil used in brake system, to avoid confusion hydraulic oil type deleted and reference made to refer to the AMM for complete information Supplements Table of Contents updated Supplements 9-07 Rev 1, 9-08 and 9-09 included

A/.%& sppm V C - ~

2 5 dud fY36,h.o.

Rev~sion Number and Date

Page Number

8-20

9-i

ePILATUS= LOG OF REVISIONS ?C %(I

I

LOG OF REVlSlONS (CONT.)

Description

Amp added to generator max. load figures PCL operation limitations made more specific Takeoff in falling and driving snow limitation deleted Oxygen system shut off handle added HIC info from Temp Rev included in this section New fuel, hand pump, ECS placard on center console added Post SB 28-002 fuel shut off lever placard added Standby compass placard changed Max freight load placard position corrected Tire inflation placards added

Contents pages updated Fuel shuton lever operation updated to include locking latch Fuel shuton lever operation updated to include locking latch Fuel shuton lever operation updated to include locking latch Warning added relerence number of air Start attempts Fuel shutofl lever operation updated to include locking latch Landing with immobil~zed horizontal stabilizer procedure added Editorial. Paragraphs re-numbered CAWS Hydr caution blinking on ground action deleted Paras numbered. New main or altern stab trim procedures added Inoperative services lists updated Total landing distance factor added Total landing distance factor added Air Gnd CAWS warning anected systems info updated post icing AHRS failure procedure updated Autopilot malfunction max altitude lnss figures corrected

Contents pages updated Maximum demonstrated crosswind limits changed Hydraulic hand pump stowage added to cockpit checks Oxygen shut off handle added Engine start procedure using battery updated for cold starting Additional info added to starter sequence. Engine start procedure using ext power updated for cold starting. Additional info added to starter sequence Page run on Page run on Page run on. Cabin pressure controller setting step changed Page run on. ECS switch position step changed. Note changed Pages run on Page run on. Cabin rate control setting step deleted Page run on Page run on. Balked Landing PCL setting step changed Page run on

Issued: February 14, 1994 Revision 6: Dec 6. 1996


6 6 Dec 96

Report No: 01 LOR- 14

Page Number

Sect 2 2-8 2-8 2- 13 2-20 2-22 2-25 2-25 2-26 2-29 2-36

Sect 3 ii thru iv

3-3 3-4 3-5 3-7 3-14 3-24 3-25 3-26 3-31

3-33 8 34 3-46

3-49. 50,51 3-52 3-53 3-56

Sect 4 4-i 8 4-ii

4-2 4-8 4-9 4-10 4-1 1 4- 12

4-13 4-16 4-17 4-18

4-19 thru 21 4-22 4-23 4-24 4-25

973-001

SPILATUS3. PC XI1 LOG OF REVISIONS


Issued: February 14.1994 Revision 6: Dec 6. 1996

Report No: 01973-001 LOR- 15

Description

Oxygen shut off handle added, steps renumbered Oxygen shut off handle added Pages run on Page run on. Crosswind operation para added Page run on. Para re-numbered. Snow added to deep slush runway operations. Takeoff in falling and driving snow para deleted. Pages run on

Contents page updated New charts issued New charts issued Takeoff reformance on slush and snow covered runways added. Icing landing distance increases changed to per cent factors Flight planning example added

Pax seat 112 position corrected. MSN 101 -1 71 added to figure New fig Pax seat 5 position changed from MSN 172 MSN101-171addedtofigure New fig Pax seats weight 8 moment change from MSN 172 MSN 101-171 added to figure MSN 101 -1 71 added to ligure New Ib-in moment chart from MSN 172 New mkg moment charl from MSN 172 Pax 5 seat arm for MSN 101 - 17 1 and from MSN 172 added

Contents page updated Flap position and rotation sensors description clarified Text run on from page 7-5 Hydraulic system description - pre standardization confii deleted Hydraulic system operating pressures corrected Landing Gear System figure updated with new hyd hand pump Wheels and tires description added Engine Controls figure updated with fuel shutoff lever locking latct- Starter re-engagement paragraph added Engine torque limiter Pre 8 Post SB 72-001 updated ECS temp control valve description changed Figures made pre SB 2 1-003 and 2 1-002 New figures added post SB 21-003.21-002 and MSN 181 8 UP Text updated to pre and post SB 21-003,21-002 8 MSN 181 8 UP Text updated to pre and post SB 35-001

BUS TIE CAWS caption description updated Reference to figure corrected EFlS control panel figure changed to updated panel Autopilot mode controller location corrected


Page Number

4-26 4-27

4-28 thru 42 4-43 4-44

4-45 thru 48 Sect 5

5-iii 5-17 to 23 5-67 to 77

5-79

5-80 to 84 Sect 6

6-6A 8 6B 6-7 6-16

6-16A 8 16B 6-31 6-32

6-32A 6-328 6-40

Sect 7 7-i 7-5 7-7 7-10 7-12 7-16 7-21 7-33 7-40 7-51 7-82

7-83 8 7-84 7-84A 8 848 7-86 thru 90 7-101 thru

7-103 7-1 11 7-1 14 7-152 7-167

LOG OF REVISIONS


# (0

1 7.M --

Report No: LOR- 16

Page Number

Sect 8 8-3

Sect 9 9-i

Sect 10 10-i 10-1 10-3


Description

CAT connector inlormation added

Contents page updated Supplement 9-06 Initial lssue Supplement 9-08 Rev 1 pages 1.2, 3,6, 12, 18-30. 79-89, 91. 94 Supplement 9-09 Rev 1 pages 1 lhru 4 Supplement 9- 10 Initial lssue

Contents page updated Flammable materials and Crosswind Operations info added Flammable materials, pressure vessels and equipment locations Figure added i

I

Issued: February 14.1994 Revision 6: Dec 6. 1996

LOG OF REVISIONS


Issued: February 14. 1994 Revision 7: July 1, 1997


7 1 July 97

Report No: 0 1973-00 1 LOR- 1 7

Page Number

Sect 1 1-9

1-12 Sect 2

2-1 2-9 2-14 2-1 5 2-16 2-17

2-19 2-20 2-20 2-20 2-22 2-22 2-22 2-23 2-24 2-25 2-26

2-27 - 2-29 2-30

2-31 - 2-34 2-35

2-36 - 2-38 Sect 3

3-ii 3-9

3-14 3-32

Sect 4 4- 10 4- 12 4- 18 4-21 4-43 4-44

4-45 - 4-47 Sect 5

5-1 5-17 - 5-23 5-25 - 5-28 5-30 - 5-37

Description

Vsl definition added Abbreviation MOR added to manual override terminology

Airspeed Limitations table note deleted Power Plant Instrument Markings for Ng Green Arc corrected New paragraph added to KOEL Page run on Propeller MOV requirement updated Oxygen system and pass oxy annunciator moved from pressurized flight section to mechanical systems section Flap indicator required for takeoff deleted Inertial separator must function all flights deleted Pneumatic deice system requirement changed Probe heat requirement changed Cargo Limitations changed Luggage Limitations added HIC paragraph deleted Intontionally blank page added Intentionally blank page added Page moved on Page moved on. Oxygen shutoff lever placard added Pages moved on Page moved on. Executive seat placards updated Pages moved on Page moved on. Terminology change Pages moved on

Contents page changed Chip detector indications updated Paragraph title changed Inoperative systems list updated

Note changed Note changed Torque limiter temp rev info added Power control lever selling inlo changed Noise level values changed Reference to new info in Sect t 0 para added Pages run on

Torque setting para changed Torque limiter temp rev graph changes included Graphs changed Graphs Changed

EPILATUSE LOG OF REVISIONS ?c %(I

LOG OF REVISIONS (CONT.) I

Report NO: 01973-001 LOR-18

Issued. February 14. 1994 Revls~on 7: ~ u l y I. 1997 1

Description

Torque limiter temp rev graph change included Torque limiter temp rev graph change included Torque limiter temp rev graph change included Torque limiler temp rev graph change included Torque limiter temp rev graph changes ~ncluded

Minor dimension changes Minor dimension changes Unusable fuel moment value corrected Cargo net statemenl added, attachment points clarified Seat 5 moment updated. Cargo net statement added Cargo net statement added, attachment points clarified Loading L~mitations updated Cargo tie down ligure updated Cargo and luggage net figure updated F~gure l~tle changed Figure title changed. Pax 5 data changed Figure title changed Combi interior loading figure updated Pass 5 data changed

Stab trim warning inlo updated Starter relay inlo updated. Cold starting info para added Torque limiter temp rev inlo added Ch~p detector info updated Fuel low level conversion figures changed Typlcal CB panel f~gures updated Ch~p detector inlo changed AP disengage caution info updated

Contents page updated Supplement 9-08 Rev 2 pages 1, 18 thru 30, 32 thru 37, 39 thru 48, 59, 62, 65, 68, 79 thru 81 Supplement 9- 11 Initial Issue 24 Mar 97. Rev 1 pages t , 2 .6

Contents page updated rmal n initial issue ,. '3

Revis~on Number and Date

for C ~ v l l Aviation &_-h.- j : L

A9 Set ) . 4 9 3 7

Page Number

5-38 - 5-39 5-50 5-53 5-56

5-67 - 5-69 Sect 6 6-6A 6-7 6-9 6-15 6-16

6-16A 6- 19 6-24 6-25

6-29 - 6-30 6-31 - 6-32

6-32A-6-328 6-37 6-40

Sect 7 7-7 7-40 7-51 7-52 7-64

7-72 - 7-73 7- 1 1 1 7- 1 74 Sect 9

9-i

Sect to 10-1

'EPILATUS W , K12 LOG OF REVISIONS


8 1 Sep 98

Issued: Feb~ary Revision 8: Seflember 1,1998 LOR- 19

Page Number

LOTR-1 LOTA-2 Sect 1

1 -i and 1 -ii 1-1 1

1-12 thru 1-20

Sect 2 2-ii and 2-iii

2-14

2-15 2-16 2-17 2-10 2-19 2-20

2-21 2-22 2-23 2-28

Sect 3 3-7 3-15 3-16 3-18 3-20 3-28 3-32 3-34 3-35 3-36 3-37 3-38 3-53

Sect 4 4-i & 4-ii

4-3 4-4

14,1994


Description

Log for recording of Temporary Revisions added

Contents pages updated Severe king Conditions added to Meteorology Terminology Pages moved on

Contents pages updated Wing inspection light operative added to Icing Limitations. Severe king Condiions information added Page moved on Second battery requirements added to Equipment List Wing inspection light requirement added to Equipment List Alternate flap control system requirement deleted Pages moved on Trim system limits changed Heated Windshield limits changed Flap system limits deleted Oxygen System limits changed Autopilot limits changed Page moved on Standby compass placard changed and ELT placard added

Second battery info added Step numbering corrected. Second battery info added lnfo on page moved on Second battery info added Second battery info added Flaps Failure - Alternate system info deleted Second battery info added Second battery info added lnfo added to the Note. Extra step added. Second battery info added Note added. Second battery info added Second battery info added Passenger door warning additional info added

Contents pages updated Larger capacity oxy system and hydraulic system info added The word Stabilator corrected to Stabilizer and steps re-numbere

Repon No: 01973-001

LOG OF REVISIONS


Issued: February 14,1994 Revision 9: September 1, 1999

-

Description

Nose oxygen service bay changed to a standard system. SB 35-001 applicability added SB 25-008 thermal blanket info added. Sequence changed and second battery info added. CB info changed Page moved on Second battery info added. Pass oxygen check added. Beacon lights info added. Second battery info added. Second note added External power staning sequence changed Ex1 pwr changed to ext pwr switch Beacon and recognition lights info added Second battery info added Recognition lights info added Recognition lights info added Beacon and recognition lights info added Beacon, recognition lights and second battery info added Oxygen SOV handle changed to shut-off lever Fig made applicable to standard oxygen system Duplicated autopilot preflight check deleted, already in mair procedure. New Fig for larger capacity oxygen System added Para re-numbered and a Note added lnfo from next page moved fotward New section title Amplified Procedures added Para title simplified, severe icing info added lnfo on page moved on Funher info added New page, severe icing conditions info added Blank page added

Figure titles corrected

Contents page updated Standard seat location figure updated to SB 25-006 Combi conversion figure updated to SB 25-006 Seat occupant moments updated to SB 25-006

Figures and pages info moved on

Loading forms updated to SB 25-006 Figures and pages info moved on, pages added


Page Number

4-5

4-7

4-0

4-9

4-10 4-1 1 & 4-12

4-13 4-15 4-17 4-18 4-22 4-25 4-26 4-27 4-28 4-29

4-30 4-42 4-43 4-44

4-45 8 4-46 4-47 4-49 4-50

Sect 5 5-i

Sect 6 6-i 6-6 6-15

6-29 thru 6-32

6-33 thru 6-41

6-42 & 6-43 6-44 thru

6-50

-PLATUS=

I ?c %I1 RECORD OF REVISIONS


Issued: Februaty 14,1994 Revision 8: September 1. 1998



Page Number

7-iii & 7-iv 7-5 7-6 7-7

7-19 & 7-20 7-23 7-38 740

7-49 li 7-50 7-65 thru 7-

68 7-69 thru 7-

74 7-75

7-76A 7-768

7-77 & 7-78 7-81

7-89 & 7-90 7-104 7-107 7-127 7-128 7-134 Sect 8 8-1 1 8-13

8-22 & 8-23 Sect 9

9-i

Description

Contents pages updated Flaps a!ternate system info deleted Flaps alternate system info deleted from system figure Flaps alternate system info deleted Probe deice info deleted. SB 25-003 info added Chip detector option added SB 80-001 info added EIS Cautions and Warnings limits minor changes Second battery and SB 24-010 info added. Minor text changes

Extra sheets added to figure. Circuit breaker panel figures updated Options and applicability added to figure New schematic for second battery added New fiilrre for second battery overhead panel added Second battery info added Optional beacon and recognition lights info added lnfo corrected Optional larger capacity oxygen system info added Flaps alternate system info deleted from cockpit figure SB 25-008 thermal blanket info added lnfo on page moved on Optional ELT remote control info added

Jacking information updated Battery sewicing info updated for second battefy Oxygen sewicing info updated for larger capacity system

Contents page updated Supp 04 Rev 2 pages 1 thru 12 Supp 08 Rev 3 pages l,9 and 10 Supp 12 Rev 1 pages lthru 4 Supp 13,14, 15 and 16 initial issue

+Federal Offic? for f lvil Aviation

4.g Sty/. AS 93

mPLATUSEC LOG OF REVISIONS P(12

I



Issued: February 14,1994 Revision 9: September 1. 1999

Description

Log of Temporary Revisions updated Log of Temporary Revisions updated Page header corrected

Contents pages updated Oil cooler thermal actuator SB 79-003 info added Pneumatic deicing boots system operation paras added Interior seating variations and code numbers added Pages run on Passenger seat lap balt extension info added Extendable baggage net and optional wardrobe limits Xkbd Placards layout changed to exterior, cockpit, cabin and placards required for each interior seating variation

Contents page updated Engine failure Temp Rev No. 7 included ECS pressurization terminology clarified Forced landing procedure updated Flaps failure Temp Rev No. 8 included Stick pusher control wheel force updated ESNTL BUS CAWS warning voltage value corrected Gen 2 reset Pre SB 24-010 statement added. AOA heater plates Inoperative effectivity statement added AV BUS CAWS warning voltage value corrected ECS pressurization and temperature terminology clarified AHRS failure Temp Rev No. 3 included Page run on

Introduction to Short Cheddit added Preflight oil quantity check info changed ECS pressurization terminology clarified Optional external lights standardized as external lighting. External lighting standardized, remaining steps re-numbered External lighting standardized External lighting standardized External lighting standardiied, remaining steps re-numbered External liihting standardized, remaining steps re-numbered Some info moved to preceding page

Contents page updated Complete section re-arranged. General information added. Seat location info removed from weighing procedure. Carno loadins


9 1 Sep 99

Page Number

LOTR-1 LOTR-2 LOR-20 Sect 2

2-ii 8 2-iii 2-5 2-12 2-20

2-21 I3 2-22 2-23 2-24

2-25 thru 2-43

Sect 3 3-i

3-4 I3 3-5 3-15 8 3-16

3-21 3-28 3-29 3-32 3-36

3-37 3-44 3-53 3-54

Sect 4 4-1 4-6 4-8 4-9 4-15 4-18 4-22 4-25 4-26 4-27

Sect 6 6-1

6-1 thru 6-32

LOG OF REVISIONS



I

Report No: 01 973-001 LOR-23

Page Number

Sect 7 7-5 7-6 7-10 7-13 7-18 7-23

7-38 7-40 7-42 7-43 7-44

7-45 thru 7-47 7-48 7-48A 7-488 7-49 7-50 7-61 7-68

7-72 & 73 7-75 7-76A 7.760 7-77 7-78 7-79 7-80

7-102


Description

and hazardous materials info added. Extendable baggage net Added to Fig 6-6. Weight and balance determination for flight procedures revised, passenger seat occupant moments removed, baggage area moments for extendable baggage net added, loading form standardized, one example and one general loading form given. Interior configurations info added. Annexes Added for each interior variation giving seat locations, permitted seats weights and moments, seat occupant moments. Temp Revs l,2,5 and 6 included

Flaps reset switch Temp Rev No. 8 info added Flaps reset switch added to Fig 7-1 MSN 231 and UP larger nitrogen pressure tank info added Fig 7-2 updated with larger nitrogen pressure tank Hydraulic control CB title corrected. KlAS airspeed corrected Extendable baggage net info added. Crew seats de~cription Updated. Optional three seat bench info added Oil level info updated Ignition Temp Rev No. 10 info added EIS Temp Rev No. 10 info added Page run on EIS Temp Rev No. 10 info added Fig 7-10 Sheet Nos increased

Fig 7-10 EIS backlight power supply added Fig 7-10 New Sheet added Post SB 77-002 and MSN 261-999 Blank page added EIS OAT info updated and moved from next page Corrections and Temp Rev No. 10 info added Fig 7-12 ref to €IS power supplies deleted. Given in Fig 7-10 Stby power para clarified. Gen I amp/voR Temp Rev 1 1 added Fig 7-13 Sheets 4 and 5 CB panels updated Fig 7-13 Sheet 7 Gen 1 amp/voR meter added Fig 7-13 Sheet 9 Power system schematic two batteries correctc Fig 7-13 Sheet 10 Overhead panel Gen 2 bus indic light added Page run on Gen 1 amphlt indicator Temp Rev 11 added Text from page 7-80 brought forward Lighting text brought forward from page 7-81. Exterior lighting beacon and logo lights Temp Rev 9 added Oxygen points for passenger seat interior variations added

EPLANSB LOG OF REVISIONS K12



Revisin Number and Date

Issued: February 14,1994 Revision 9: September 1, 199s

Page Number

7-103 thru 7- 105

7-1 16 7-120 7-121 7-122 7-125

7-127 8 128 7-132 7-164 7-165 Sect 8 8-20 8-21

Sect 9 9-i

Sect 10

10-i 10-8 Uiru

10-10

Description

Pages run on

Stick pusher actuator and capstan clutch forces changed Fig 7-20 Sheet 1 schematic updated with ice mode Fig 7-20 Sheet 2 schematic updated for MSN 261-999 Fig 7-20 Sheet 3 schematic updated wlh CAWS ice mode Fig 7-21 Sheet 1 Deicing system schematic improved Cabin Features section updated Second AHRS magnetic sensor unit location corrected Fig 7-31 EADl heading scale value corrected to 32 Fig 7-32 EADl heading scale value corrected to 32

Tires text change from type to size and ref to AMM added Hydraulic system ref MSN 231-999 added

Contents page updated Supp 08 Rev 4 pages 1,9,91 thru 95. Temp Revs 8-1 3-2 and 8- 3 included Supp 11 Rev 2 pages 1 and 6 Supp 17 Initial issue

Contents page updated Operations from prepared unpaved surfaces Temp Rev 12 added

C 0 2 1999

LOG OF REVISIONS

LOG OF REVISIONS (COW.)

Description

Log of Service Bulletins Incorporated added Log of Temporary Revisions updated New Log of Temporary Revisions pages Issued

Contents page updated Temporary Revisions added to Revision Procedure para List of Effective Pages para changed Caution ref revision incorporation responsibility added Temporary Revisbns para added

Temp Rev 23 Fuel Anti-icing Additives info added Luggage Limitations para (Ref. SB 25-019) changed Cockpit placard updated Standby compass placard updated for new heating system Temp Rev 19 Cabin Placards added. New heating system placard added Baggage placards updated (Ref. SB 25-019) Interior code and cabinet drawers placard weight limits updated. Placard near each executive seat wording corrected

Contents page para numbers corrected Urgency of landing definitions added Urgency of landing terminology aligned with definlion Urgency of landing terminology aliined with definition Ditching evacuation step clarified Flaps failure procedure changed to Pre SB 27-08 TR 15 Flaps failure procedure Post SB 27-08 new page added Intentionally left blank page added Stby att operative with EPS option added, land terminology Alternative flap control deleted Temp Rev 16 & 26 Lead Acld Batteries info Note added Urgency of landing terminology aligned with definition Urgency of landing terminology aligned with definlion Caution deleted AHRS failure with new heating system procedure step added Contents page corrected Engine and fuel drains and fuel filter terminology clarified Engine oil level and replenishment instruction expanded Note added ref EPU power Engine oil temperature excursions info added AOA fast slow pointer para deleted. Should changed to must Preventing of frozen brakes ~rocedure added

Report No: 01973-001


10 1 Sep 00

Issued: February Revision 10: September 1,2000 LOR-25

Page Number

LOSB-1 & 2 LOTR-2

LOTR3 (L 4 sect 0

0-1 0-2 0-3 0-3 0-3

Sect 2 2-6 2-24 2-30 2-33 2-36

2-37 2-40 & 2-41

Sect 3 3-1 3-1 3-9 3-12 3-25 3-28

3-28A 3-288 3-32 3-33 3-34 3-40 343 3-46 3-53 4-11

4-5 & 4-6 4-6 4-1 1 4-21 4-47

4-49 & 4-50 14,1994

=PILATUsw LOG OF REVISIONS K12

i




Page Number

Sect 5 5-18 thru

5-23 5-79

Sect 6 6-01-3 6-02-3 6-06-3 Sect 7

7-5 & 7-6 7-7A thru

7-7D 7-1 3 7-19 7-43 7-44 7-50 7-65

7-66 & 7-67 7-68 7-75

7-77 LL 7-78 7-79

7-83 8 7-84A 7-848 7-88

7-90A thw 7-90F 7-101 7-1 02 7-1 11 7-1 19 7-127 7-128 7-129

7-1 34 Sect 8

8-i and &ii 8-4

~escription

See Flight in lcing Conditions para added to figure

lcing conditions take off performance additional para added

Seat Nos. corrected Temp Rev 13 Bulkhead and Curtain info added Temp Rev 13 Bulkhead and Curtain info added

Flaps description and figure changed to Pre SB 27-08 Temp Rev 15 Flaps Post SB 27-08,27-011 and 27-013 info added Figure 7-2. Bleed air key deleted Temp Rev 18 WeigM on Wheels Signal info added EIS fault codes para changed Fuel quantity reset info changed DC volt Gen info corrected General para changed to add Temp Revs. 400 Amp StarterIGer and EPS option. Info ref external power disconnection added Pages run on MSN number changed Gen 1 indiitor info changed Pages run on from page 7-68 Temp Rev 16 Lead Acid Battery info added Figure 7-14. Terminology sensor to switch change Figure 7-14 Sheet 2. Footwarmer option added Heating System info effectivity added Temp Rev 25 New Heating System info addad

Oxygen general para changed Page run on Temp Rev 16 Lead Acid Battery info added Three last sentences of first paragraph deleted Commuter interior bulkhead and curtain (Post SB 25-017) Zdd8d Executive interior bulkhead and curtain (Post SB 25-017) added Optional second audio panel info added. Optional SB 23-003 info ref 8.33kHz channel spacing added ELT remote control switch info added

Contents pages updated Towing by lifting nose wheel restriction deleted

FPILATUSS

I PC12 LOG OF REVISIONS



Report No: 01 973-001 LOR-27

Description

Temp Rev 23 Fuel Anti-icing Additives info added Pages run on Temp Rev 23 Fuel Anti-icing Additives info added Page run on

Contents pages updated Pages 1, 25 thru 30 lcing Conditions statement added, page 91 lcing conditions take off performance additional paras added, pages 92 and 94 revision status corrected ECTM ref to SB updated, pages 1 and 2 PC-12 Registered in France - initial general issue GPS KLN 900 - initial general issue 2" Pitot Static Sys & Co-pilot Instruments - initial general issue Specific customer only - not general issue Emergency Power System - initial issue includes Temp Rev 20 Multifunction Display (KMD 850)

Contents updated Passenger briefings added


Rev 5

2 1

1

Page Number

8-1 6 8-17 & 8-18 8-19 & 8-20

8-21 Sect 9

9-i & 9-ii Supp 08

Supp 12 Supp 18 Supp 19 Supp 20 Supp 21 Supp 22 Supp 23 Sect 10

10-i 10-1 0 thru

10-12

I r PILATUSF

LOG OF REVISIONS PC12


Report No: 01 973-001 LOR-28


11 1 Mar 03

Issued: February 14,1994 Revision 1 1 : March 1,2003

Page Number

Title LOSB-2 & 3

LOTR-3 Sect 2

2-ii & 2-iii 2-8 2-23 2-24

2-25 thru 30 2-31

2-32 & 33 2-34

2-35 and 36 2-37 2-38

2-39 and 40 2-41 2-41

2-42 thru 44 Sect 3

3-ii thru 3-iv 3-9

3-14 thru 16

3-18 thru 21

3-22 3-23 thru 31

3-32

3-33 thru 38 3-39 3-40

3-41 thru 47 3-48 3-49

3-50 thru 64

Description

MSN effectivity added Log of Service Bulletins Incorporated updated Log of Temporary Revisions updated

Contents pages updated Gen 2 max load figures revised TR No. 36 AHRS limits included Executive cabin floodlights limitation added Pages run on Pre and Post SB 11-002 flap placards shown Pages run on Control wheel placards updated & ELT placard Post SB 25-020 Pages run on Placard MSN applicability corrected Max freight load placard Ibs value corrected Pages run on Post SB 35-003 lavatory oxygen mask placards added New executive seat placard added Pages run on, one extra page added

Contents pages updated Chip detector changed to plural Post SB 79-005 2m chip detector. In flight procedure changed. CAWS Oil Quantity warning info updated TR No. 35 Crew Oxygen Mask Donning procedure added Cockpitfcabin Fire, Smoke or Fumes procedure revised TR No. 34 Electrical Load Shed procedure info added TR No. 35 Crew Oxygen Mask Donning procedure added Maximum range descent procedures rewritten for single & dua battery installation Para re-numbered & TR No. 4 Crew Oxygen Mask Donning Pages run on Pre SB 27-008 info removed. Post SB 27-008 is now standard and moved to this page. Word practicable changed to practical Pages run on Gen 2 load, step 5 changed to reduce electrical load Gen 2 off, step 4 changed Pages run on TR No. 35 Crew Oxygen Mask Donning procedure added New ECS failure procedure added Pages run on, six extra pages added

LOG OF REVISIONS



Revision Number

and Date

Report No: 01 973-001 LOR-29

Page Number

Sect 4 4-7

4-8 4-9

4-14 & 15 4-26 4-27

4-28 & 29 Sect 6

6-1 6-2 6-3

6-5 & 6

6-6A & 68 6-27 6-32

6-01-2 6-02-2 6-02-3 6-03-2 6-03-3 6-04-2 6-04-3 6-05-2 6-05-3 6-06-2 6-06-3 6-07-2 6-07-3 6-08-2 6-08-3 Sect 7

7-11 7-iv & 7-v

7-5 thru 7-7

7A & 78 7C & 7D

7-21 7-25

Description

Over-wing emergency exit lock pin info added para 4.3.7 step 5. If installed added to para 4.3.8 step 7 Page run on Fuel pumps operation check step added Stick pusher test, PCL set to idle step added, steps re-numbered CAWS oil quantity check added to shutdown procedure Para 4.a. changed Oxygen Masks Connected added to table column heading

Contents page updated Page title changed Procedure changed to weighing with load plates Amm corrected to Arm in table heading. Decimal point removed from mm figures New weighing procedure with jacks and load cells added (kg) and (kg-m) removed from table heading New para ref two interior codes placard added Direction of flight indicator added to seat position info Direction of flight indicator added, seat positions clarified New bench seat Part Nos. added Forward and aft facing executive seat positions clarified New executive seat Part Nos. added Forward and aft facing executive seat positions clarified New executive seat Part Nos. added Forward and aft facing executive seat positions clarified New executive seat Part Nos. added Forward and aft facing seat positions clarified New executive seat Part Nos. added Forward and aft facing seat positions clarified New executive seat Part Nos. added Forward and aft facing seat positions clarified New executive seat Part Nos. added

Contents page updated Contents pages updated Flaps Pre SB 27-008 info removed. Post SB 27-008 is now the standard and info moved to these pages. More info added Figures on pages 7-C and 7D moved to these pages Pages deleted Post SB 32-013 steel brakes info added Aircraft security para added

WPILATUSW I LOG OF REVISIONS PC12 1


Report No: 01 973-001 LOR-30

Revision Number

and Date

Issued: February 14,1994 Revision 11: March 1, 2003

Page Number

7-26 7-38

7-65 7-67 7-69 7-72 7-73 7-74

7-76A 7-79

7-80 & 81 7-88

7-102 7-1 12 7-1 16 7-120

7-127 7-130 7-131 7-1 32 7-133 7-134

Sect 8 8-7

8-7A 8-78 8-20 8-21

Sect 9 9-i and 9-ii 9-00-1 & 2

Supp 6 Supp 8

Supp 22 Supp 24 Supp 29

Description

Figure 7-4 updated to show emergency exit locking pin Chip caution on conditions defined. SB 79-005 second oil c h l ~ detector info added Gen 2 amp rating figure changed Caution ref Gen 2 load figures changed Figure 7-13 Sheet 1 schematic corrected. Gen 2 amp changed Figure 7-13 Sheet 4 CB MSN number changed Figure 7-13 Sheet 5 re-issued previous type face difficult to read Figure 7-13 Sheet 6 115v AC bus removed from schematic Figure 7-13 Sheet 9 schematic corrected. Gen 2 amp changed Gen 2 amps figure changed, no discharge from battery added Additional baggage comp light and dual filament nav lights added CAWS ECS caution description corrected Post SB 35-003 lavatory oxygen mask info added CAWS ECS annunciator on sentence clarified Pusher disc corrected to pusher intr Figure 7-20 Sheet 1 airspeed input deleted, engine running switch input added Fire extinguisher info changed for different types Figure 7-22 Typical Avionics schematic updated LCR-92 continuous orbiting maneuvering info added Page run on ELT info brought forward and updated Post SB 25-022 Kannad ELT info added. SB 25-029 and 030 info added

Blanks and Covers updated and changed to two Figures New Blanks and Covers Flgure added Blank page added Post SB 32-013 steel brake wheel type info added. Brake wear indicator pin info clarified

Contents updated Supplements List of Effective Pages log added TCAS - Revision 1 issued PC-12/45 - Revision 6 issued EPS - Revision 1 issued EGPWS - Initial issue Pilot's Relief Tube - Initial issue

A

WPILATUSW I PCXII h CONTENTS

CONTENTS

Subject

INTRODUCTION

GENERAL

LIMITATIONS

EMERGENCY PROCEDURES

NORMAL PROCEDURES

PERFORMANCE

WEIGHT AND BALANCE

AIRPLANE AND SYSTEMS DESCRIPTION

HANDLING, SERVICING, AND MAINTENANCE

SUPPLEMENTS

SAFETY AND OPERATIONAL TIPS

Page

0-1

1.1

2-1

3-1

4-1

5-1

6-1

7-1

8-1

9-1

10-1

THE CONTENTS OF SECTIONS 2,3,4,5,6, AND 9 ARE FOCA APPROVED

Issued: February 14, 1994 Revision 1: June 10, 1994

1 PILOT'S OPERATING L HANDBOOK

L Introduction

SECTION 0 INTRODUCTION

SECTION 0

INTRODUCTION

TABLE OF CONTENTS

Subject

GENERAL

WARNINGS, CAUTIONS, AND NOTES

REVISION MARKINGS

REVISION PROCEDURE

Transmittal Letter Log of Revisions List of Effective Pages New or Revised Pages Temporary Revisions

Issued: February 14, 1994 Revision 10: September 1, 2000

Page

Report No: 01973-001 0-1


Report No: 01 973-001 I 0-1,

THIS PAGE INTENTIONALLY LEFT BLANK

Issued: Feb~ary 14, 1994 Revision 1: June 10, 1994


GENERAL

This Pilot's Operating Handbook (POH) is designed to provide the information requlred for the operation of the airplane. Each airplane 1s delivered with a POH that reflects the standard alrplane with all of the approved options plus any special equipment installed on an individual basis.

WARNINGS, CAUTIONS, AND NOTES

The following definitions apply to the warnings, cautions, and notes as used in this manual: .

I WARNING I ANY OPERATING PROCEDURE, PRACTICE, OR CONDITION WHICH, IF NOT STRICTLY COMPLIED WITH, MAY RESULT IN PERSONAL INJURY OR LOSS OF LIFE.

ANY OPERATING PROCEDURE, PRACTICE, OR CONDITION WHICH, IF NOT STRICTLY COMPLIED WITH, MAY RESULT IN DAMAGE TO THE AIRPLANE OR EQUIPMENT.

NOTE

Any operating procedure, practice, or condition that requires emphasis.

Issued: June 10, 1994 Revision 1: June 10, 1994

Report No: 01 973-001 0-1 I


REVISION MARKINGS

Additions, deletions, and revisions to existing POH material will be identified by a verlical revision bar (black line) in the outside margin of the applicable page, next to the change.

When a revision causes expansion or deletion of text or illustration which results in unchanged material appearing on a different page, that page will be identified by a revision bar in the outer margin next to the page number.

At the bottom of each page, opposite the page number, there will be the original issue date of the page. As the page is subsequently revised, the original issue date will be fobwed by the current revision number and date. If a new page is issued, it will be identified by having the same original issue and revision date and have a revision bar in the outer margin next to the page number.

The revision bar will only indicate the current change on each page. Physical relocation of material or the correction of typographical or grammatical errors, outside of the material revised, will not be identified by a revision bar.

REVISION PROCEDURE

To keep this POH current, revisions will be issued to latest registered owner of airplane. Revisions to this POH will consist of:

- Transmittal Letter

- Log of Revisions

- List of Effective Pages

- New or Revised Pages

I - Temporary Revisions

The Equipment List is not included in the Revision Procedure. The Equipment List is a separate report and was current at the time of license at the manufacturer and must be maintained by the airplane owner.

The Transmittal Letter will show the revision number and date. All POH pages affected by the Transmittal Letter will be listed along with instructions for incorporating the revision into the POH.

Report No: 01873-001 0-2

Issued: June 10,1994 Revision 10: September 1, 2000


LOG OF REVISIONS

The Log of Revisions provides a brief description of each revision.

LIST OF EFFECTIVE PAGES

The List of Effective Pages will list all of the current POH page numbers with the applicable revision number. I NEW OR REVISED PAGES

In accordance with the instructions of the Transmittal Letter, new or revised pages will be incorporated into the POH and superseded pages destroyed.

IT IS THE RESPONSIBILITY OF THE OWNER OR OPERATOR TO MAINTAIN THIS PILOT'S OPERATING HANDBOOK IN A CURRENT STATUS AND INCORPORATE SUCCESSIVE REVISIONS.

TEMPORARY REVISIONS I Temporary Revisions are issued when the POH must be revised between the regular formal revisions. They are issued on yellow paper and must be recorded on the Log Of Temporary Revisions. Temporary Revisions should normally be put at the front of the POH, apart from Section 9 Temporary Revisions which should be put in front of the applicable Supplement. Temporary Revisions must only be removed from the POH when instructed to do so by, the Transmittal Letter of the next issue of a formal revision, superseded by another temporary revision and sometimes by the incorporation of a Service Bulletin. The Log Of Temporary Revisions will be updated and issued with each formal revision.

issued: June 10,1994 Revision 10: September 1,2000

Report No: 01973-001 0-3



Issued: June 10,1994 Revision 1 : June 10, 1994

PC XII PILOT'S OPERATING

HANDBOOK

General

SECTION 1 GENERAL

SECTION 1

GENERAL

TABLE OF CONTENTS

Subject Page

GENERAL 1-1

INTRODUCTION 1-1

DESCRIPTIVE DATA

ENGINE PROPELLER FUEL OIL MAXIMUM WEIGHTS TYPICAL AIRPLANE WEIGHTS CABIN AND ENTRY DIMENSIONS SPECIFIC LOADINGS

SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY 1-8

GENERAL AIRSPEED TERMINOLOGY AND SYMBOLS 1-8 METEOROLOGICAL TERMINOLOGY 1-10 POWER TERMINOLOGY 1-12 ENGINE CONTROLS AND INSTRUMENTS TERMINOLOGY 1-13 AIRPLANE PERFORMANCE AND FLIGHT PLANNING TERMINOLOGY 1-14 WFIGHT AND BALANCE TERMINOLOGY 1-15 GENERAL ABBREVIATIONS AND SYMBOLS 1-17 ELECTRICAUAVIONIC ABBREVIATIONS 1-18

CONVERSION INFORMATION

GENERAL STANDARD TO METRIC METRIC TO STANDARD

Issued: February 14. 1994 Revision 8: September 1, 1998

Reporl No: 0 1 1973-00 1 -i

SECTION 1 GENERAL

Report No: 01 973-001 I 1-~ i


4

Issued: February 14.1994 Revision 8: September 1, 1998 4

=PLATUSW SECTION 1

?c XI1 GENERAL

GENERAL

This section contains basic data and information of general interest to the pilot. It also contains ' definitions and explanations of symbols, abbreviations, and terminology that is used throughout this POH.

INTRODUCTION

This POH includes the material required to be furnished by the Federal Aviation Regulations and additional information provided by the manufacturer and constitutes the FOCA Approved Airplane Flight Manual. This POH must be read, and thoroughly understood, by the owner and operator in order to achieve maximum utilization as an operating guide for the pilot.

This POH is divided into numbered sections which are separated by tabs. Section 3, Emergency Procedures, is further highlighted by the use of a red tab to facilitate quick recognition.

Pages that have been intentionally left blank will be so indicated by the statement 'THIS PAGE INTENTIONALLY LEFT BLANK'.

Issued: February 14,1994 Revision 1: June 10,1994

Report No: 01 973-001 1-1 I

SECTION 1 BPLATUSW SC XI1 I

GENERAL

Figure 1 -1. Airplane Three View and Dimensions

Report No: 01973001 1-2

Issued: February 14,1994 Revision 2: February 14, 1995

SECTION 1 GENERAL

Figure 1-2. Airplane Ground Turning Clearance

, Issued: Febtuary 14,1994 Revlskn 1: June 10,1994

SECTION 1 WPLATUSE GENERAL PC %I1

DESCRIPTIVE DATA

ENGINE

Number of Engines 1

Engine Manufacturer

Engne Model Number

Pratt 8 Whitney Canada

PT6A-678

Engine Type

This airplane incorporates a Iwin shaft turboprop engine with 4 axial and 1 centrifugal compressor stages, an annular combustion chamber, and a 3 stage turbne where one stage drives the compressor and two stages power the propeller.

Horsepower Rabng and Engine Speed

Takeoff Power 1,200 shp

Maximum ClimbIC~ise Power 1,000 shp

Compressor Turbine (No)

Speed (1 04%) 39,000 rpm

Propeller Speed (N,) 1,700 rpm

PROPELLER

Number of Propellers 1

Propeller Manufacturer

Propeller Model Number

Number of Blades 4

Propeller Diameter 105' (2.67 m)

Propeller Type

The propeller assembly consists of a hub unit and four metal blades, and is a hydraulically actuated, constant speed, full feathering and reversible type.

Report No: 01973-001 1 1-4

Issued: February 14, 1994 Revision 1: June 10. 1994

=PILATUSW SECTION 1 'PC %I1 GENERAL

FUEL

APPROVED FUELS

JET A, JET-A-1, JET 8, JP-4

Any other fuel which complies with the latest revision of Pratt & Whitney Service Bulletin 14004.

TOTAL CAPACITY

406.8 US gal, 2,736.5 Ib (1,540 liters, 1,241.3 kg)

USABLE FUEL

MSN 101-140 400.4 US gal, 2,693.3 Ib (1,515.7 liters, 1,221.7 kg)

MSN 141 AND UP 402 US gal, 2,703.6 Ib (1,521.5 liters, 1,226.4 kg)

ANTI-ICING ADDITIVE

Anti-Icing additive conforming to MIL-1-27686.

Anti-icing additives should be in compliance to Pratt & Whitney Service Bulletin 14004.

OIL

OIL GRADE OR SPECIFICATION

Any oil specified by brand name in the latest revision of Pratt & Whitney Service Bulletin 14001.

OIL QUANTITY

Total Oil Capacity 3.6 US gal (13.6 liters)

Drain and Refill Quantity 2.0 US gal (7.6 liters)

Oil Quantity Operating Range 1.0 US gal (3.8 liters)


Report No: 01973-001 1-5

GENERAL ?c XI1

MAXIMUM WEIGHTS

Maximum Ramp Weight

Maximum Takeoff Weight

Maximum Landing Weight

Maxlmum Zero Fuel Weight

Maximum Cargo Weight

Baggage Area

Cabin Area

9,083 Ib (4,120 kg)

9,039 Ib (4,1 W kg)

9,039 Ib (4,100 kg)

8,160 Ib (3,700 kg)

400 Ib (1 80 kg)

2205 Ib (1,000 kg)

TYPICAL AIRPLANE WEIGHTS

Empty Weight 5,439 Ib (2,467 kg) '

Useful Load 3,600 Ib (1,633 kg)

'Empty weight of standard airplane without 9 passenger seats and cabin floor covering.

CABIN AND ENTRY DIMENSIONS

Maximum Cabin Width

Cabin Floor Width

Maximum Cabin Length

Cabin Floor Length

Maximum Cabin Height

Forward Cabin Door

Width

Height

Cargo Door

Width

Height

Report No: 01 973-001 1-6

Issued: Febmaty 14,1994 Revision 2: February 14, 1995

SECTION 1 GENERAL

Overwing Emergency Exit

Width

Height

Compartment Volume

Baggage

Cabin

SPECIFIC LOADINGS

Wing Loading

Power Loading


32.5 Iblsq ft (158.8 kglsq m)

9.0 lblshp (4.1 kglshp)

Report No: 01 973-001 1-7

GENERAL

SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY

GENERAL AIRSPEED TERMINOLOGY AND SYMBOLS

CAS

IAS

TAS

Report No: 01973401 I 1-8

Calibrated airspeed means the indicated airspeed of an aircraft, corrected for position and instrument error. Calibrated airspeed is equal to true airspeed in standard atmosphere at sea level.

Ground speed is the speed of an airplane relative to the ground.

lndicated airspeed means the speed of an aircraft as shown on its pitot-static airspeed indicator uncorrected for airspeed system error. IAS values in this handbook assume zero instrument error.

Calibrated airspeed expressed in knots.

lndicated airspeed expressed in knots.

Means Mach number. Mach number is the ratio of true airspeed to the speed of sound.

Maximum operating limit speed is the speed limit that may not be deliberately exceeded in normal flight operations. M is expressed ill Mach number.

True airspeed means the airspeed of an airplane relative to undistutbed air which is the CAS corrected for altitude, temperature, and compressibility.

Maximum R a p extended speed is the highest speed permissible with wing flaps in a prescribed extended position.

Maximum landing gear extended speed is h e maximum speed at which an airplane can be safely flown with the landing gear extended.

Maximum landing gear operating speed is the maximum speed at which the landing gear can be safely extended or retracted.

Issued: Februaw 14, 1994 Revision 1: June 10, 1994

SECTION 1 GENERAL

VR

" s

"so

vsc

vx

"Y

Issued: February 14. 1994 Revision 7: July 1, 1997

Maximum operating speed is the speed limit that may not be exceed at any time. V is expressed in knots

Maximum Operating Maneuvering airspeed is the maximum speed at which application of full available aerodynamic control will riot overstress the airplane.

NOTE

Vo is defined in accordance with FAR 23 Amendment 45

Rotation speed used for takeoff.

Stalling speed or the minimum steady flight speed at which the airplane is controllable.

Stalling speed or the minimum steady flight speed at which the

airplane is controllable in the landing configuration at maximum gross weight.

Sta~l~ng speed or the minimum steady flight speed at which the

airplane is controllable in the specified configuration at the specified weight.

Best angle of climb speed is the airspeed which delivers the greatest gain 01 altitude in the shortest poss~ble horizontal distance.

Best rate of climb speed is the airspeed which delivers the greatest gain of altitude in the shortest possible time.

Report No- 01973 001 1-9

SECTION 1 GENERAL

METEOROLOGICAL TERMINOLOGY

The number actually read from an altlmeter when the barometrlc subscale has been Pressure set lo 29 92 In hg (1013 2 mbar)

Indicated Outslde Alr Temperature IS the temperature obtalned from an lndlcator and not corrected for ~nslrunient error and compressrb~l~ty effects

ISA International Standard Atmosphere ~n whlch

the alr IS a dry, perfect gas;

the temperature at sea level IS 59' F (15" C) ;

the pressure at sea level IS 29.92 in hg (1013.2 mbar);

the temperature gradient from sea level to the altltude at which the temperature is -69.7' F (-56.5' C) is - 0.003564" F (-0.00198" C) per foot and zero above that alt~tude.

OAT

Pressure Allrlude

Statlon Pressure

Wind

Report No: 01973-001 1-10

Outslde Air Temperature is the free air static temperature. obtained either from intlight temperature indications or ground meteorological sources, adjusted for instrument error and compressibrllty effects.

Pressure Altitude measured from standard sea level pressure (29.92 in hgl1013.2 mbar) by a pressure or barometrlc altlmeter. It IS the Indicated pressure alt~tude corrected for positlon and instrument error. In this AFM, altimeter instrument errors are assumed to be zero.

Actual atmospheric pressure at held elevation

The wlnd velocities recorded as variables on the charts of thls AFM are to be understood as the headwind or tailwlnd components of the reported wlnds.

Can exlst when the oulslde alr temperature (OAT) on the ground and for take-off, or total alr temperature (TAT) In fl~ght, IS 1OGC or colder, and vlslble molsture In any lorm 1s present (such as clouds, log or mlst w~th vlslbll~ty of one mlle or less. ram, snow, sleet and Ice crystals)

Issued: February 14, 1994 Revisron 4: November 24, 1995

SECTION 1 GENERAL

king Can exist when the OAT on the ground and for take-off is 10°C Conditions (Continued) or colder when operating on ramps, taxiways or runways, where

surface snow, ice, standing water, or slush may be ingested by the engine, or freeze on the engine, or the engine nacelle.

Can exist when there are visible signs of ice accretion on the aircraft

Severe Icing Conditions

Severe icing may result from environmental conditions during flight in freezing rain, freezing drizzle, or mixed icing conditions (supercooled liquid water and ice crystals) which may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming all of the protected surfaces.

Issued: Februaty 14, 1994 Revisbn 8: September 1, 1998

Report No: 0 1973-00 1 1 - 1 1

SECTION 1 GENERAL

POWER TERMINOLOGY

Cruise Climb Power

The power recommended to operate the airplane in a cruise climb (a continuous, gradual climb) profile.

Flight Idle Power The power required to run an engine, in flight, at the lowest speed that will ensure satisfactory engine and systems operation and airplane handling characteristics. Power setting is achieved with the Power Control Lever at the ldle Detent and the Condilion Lever in the Flight ldle psilion.

Ground ldle power

Maximum Cl~mb Power

Maximum Cruise Power

Reverse Thrust

Takeott Power

Zero Thrust

Report No: 01973-001 1 1-12

The power required to run an engine on the ground, as slowly as possible, yet sutficient to ensure satisfactory engine, engine accessory, and airplane operation with a minimum of thrust. Power setting is achieved with the Power Control Lever at or all ot the ldle Detent and the Condition Lever in the Ground ldle position.

The maximum power approved for climb.

The maximum power approved lor cruise.

The thrust of the propeller directed opposite the usual direction. thereby producing a braking action. Power setting is achieved with the Power Control Lever aft of the ldle Detent and the Condition Lever in Flight or Ground Idle.

The maximum power permissible for takeon (limited to 5 minutes).

The absence of appreciable thrust, in either direction.

Issued: February 14,1994 Revision 8: September 1. 1998

SECTION 1 GENERAL

ENOINE CONTROLS AND INSTRUMENTS TERMINOLOGY

Adjustable Minimum Prop Pitch in flight

Beta Range

Condiiion Lever

Constant Speed Range

ITT Gauge

The Power Control Lever position selects the minimum pitch in flight (6" to 12") when forward of the idle detent. This pitch can only be reached when the propeller is underspeeding at low power and low airspeed conditions.

The range of propeller pitch where the beta valve in the Constant Speed Unit (CSU) controls the pitch. Forward of the idle Detent only the minimum pitch is limited by the beta valve. In case of a propeller overspeed the CSU moves the propeller to a coarser pitch. Below flight regime. i.e. aft of the Idle Detent, the pneumatic section of the CSU limits the propeller speed to an underspeed condition and the beta valve, i.e. PCL .

position, directly controls the propeller pitch.

This lever selects the gas generator idle speed and fuel cutoff, and feathers the propeller when in the CUTOFFIFEATHER position.

The engine operating range where the propeller is out of Beta range and operating at a constant rpm, under control of the propeller governor.

A temperature measuring system that senses gas temperature in the turbine section of the engine.

Manual Override (MOR) The device that controls engine power in case of a pneumatic failure in the engine control systems. It can also control engine power in case of a power control lever failure.

Power Control Lever

The lever used to control engine power, from reverse (see Beta Range) to maximum power (see Constant Speed Range).

Propeller Feather This is a propeller pitch condition which produces minimum drag in a flight condition.

Propeller Governor The device that keeps propeller rpm constant by increasing or decreasing propeller pitch through a pitch change mechanism in the propeller hub. See Beta and constant speed Range.

Py Pressure


P3 pressure (after engine compressor) to limit fuel flow during engine acceleration in order to not cause compressor surges. The torque limiter and the Nf governor reduce P3 pressure (which is called Py pressure) to limit fuel flow so that the torque and NI limits are not exceeded.

Reporl No: 01973-001 1-13 1

SECTION 1 GENERAL

Tachometer

Torquemeter

Torque Limiter

An instrument that indicates rotational speed. Gas generator tachomelers measure speed as a percentage of the nominal maximum speed of the turbine(s), while propeller tachometers measure actual propeller rpm.

An indicating system that displays the output torque available on the propeller shaft. Torque is shown in reference terms, such as the oil pressure generated by the engine torquemeter piston.

A device which monitors torque pressure and adjusts the Py air pressure to the Fuel Control Unit to prevent an overtorque condition by limiting engine power.

AIRPLANE PERFORMANCE AND FLIGHT PLANNING TERMINOLOGY

Cl~mb Gradient The demonstrated ratio of the change in height during a pot i in of a climb, to the horizontal distance traversed in the same time nterval.

Demonstrated The demonstrated crosswind velocity is the velocity of the Crosswind Velocity crosswind component for which adequate control of the

airplane during takeotl and landing was actually demonstrated during certification tests. The value shown may or may not be limiting. Whether or not the value shown is limiting will be stated.

MEA

Route Segment

Reporl No: 01 973-001 1 1-14

Minimum Enroule IFR Altitude.

A pan of a route. Each end of that part is identified by: (1) a geographical location; or (2) a point at which a definite radii fix can be established.


SECTION 1 GENERAL

MIGHT AND BALANCE TERMINOLOGY

A.O.D. AR of Datum

Arm

Basic Empty Weight

The horizontal distance from the reference datum to the center of gravity (C.G.) of an item.

Standard empty weight plus optional equipment.

Center of The point at which an airplane would balance if suspended. Its Gravity (C.G.) distance from the reference datum is found by dividing the total

moment by the total weight of the airplane.

C.G. Arm

C.G. Limits

Datum

The arm obtained by adding the airplane's individual moments and dividing the sum by the tolal weight.

The extreme center of gravity locations within which the airplane must be operated at a given weight.

An imaginary vertical plane from which all horizontal distances are measured lor balance purposes.

Maximum Maximum weight approved for the landing touchdown. Landing Weight

Maximum Ramp Weight

Maximum weight approved tor ground maneuver. It includes weight of start, taxi, and run-up fuel.

Maximum Maximum weight approved for the start of the takeoff run. TakeoH Weight

Maximum Zero Maximum weight exclusive of usable fuel. Fuel Weight

Moment The product of the weight of an item multiplied by its arm. Moment divided by a constanl is used to simplify balance calculations by reducing the number of digits.

Payload Weight of occupants, cargo, and baggage

Standard Weight of a standard airplane including unusable fuel, lull Empty Weight operating fluids, and full oil.


Report No: 01 973-001 1-15 (

SECTION 1 GENERAL

A location along the airplane fuselage usually given in terms of ' distance from the reference datum. 4

Unusable Fuel Fuel which may not be considered usable for flight planning. 4 4

Usable Fuel

Uselul Load

Fuel available for flight planning. 1

D~tlerence between takeoff weight, or ramp weight if applicable, 1 and basic empty weight.

4

Repon No: 01973-001 ( 1-16

1

Issued: February 14. 1994 Revision 8: September 1, 1998 4

SECTION 1 GENERAL

GENERAL ABBREVIATIONS AND SYMBOLS

C Celsius

w Cubic

F Fahrenheit

FAA Federal Aviation Administration (U.S.A.)

FOCA Federal Office lor Civil Aviation (Switzerland)

fpm Feet per Minute

ft Feet

B Unit of acceleration measured against the force of gravity

gal Galbn (US)

hg Mercury

IFR Instrument Flight Rules

in Inches

kg Kilogram

KTAS Knots True Airspeed

Ib Pound (mass)

m Meter

MAC Mean Aerodynamic Chord

max Maximum

mbar Millibar


mkg Moment in meters~kilograms

min Minimum

mm Millimeters

nm Nautical Mile

NIA Not Applicable

psi Pounds per Square Inch

rpm Revolutions Per Minute

sec Second

ShP

sm

TED

TBO

VFR

0

Shan Horsepower

Statute Mile

To Be Determined

Time Between Overhauls

Visual Flight Rules

Degrees

Feet

Inches

Report No: 01973-001 1-17 1

SECTION 1 GENERAL

ELECTRICAL I AVIONIC ABBREVIATIONS

Alp Autopilot

ADF Automatic Direction Finder

ALT Autopilot Altitude hold mode

APR Aulop~lot Approach mode

ARC EFlS sectored display

AS1 Airspeed Indicator

ATC Transponder with altitude XPNDR reporting capability

ATT Autopilot Attitude hold mode

BC Back Course mode

CAT Cabin Air Temperature or Computer Aided Testing

CAWS Central Advisory and Warning System

CDI Course Deviation Indicator

CRS Course

CWS Control Wheel Steering

DU Display Unit

EADl Electronic Attitude Director lndicator

EFlS Electronic Flight Instrumentation System

EHSl Electronic Horizontal Situation lndicator

EIS Engine Instrument System

ELT Emergency Locator Transmitter

Report No: 01973-001 1 1-10

EIS

ELT

FD

FDWU

HDG

HF COM

IAS

Engine Instrument System

Emergency Locator Transmitter

Flight Director

Flap Drive Warning Unit

Heading

High Frequency Band Communication

Autopilot Airspeed hold mode

MFD Multi Function Display

MKR Marker Beacon

NAV Autopilot Navigation mode

OBS Ommi Bearing Selector

RMI Radii Magnetic lndi i tor

WA Radar Altimeter

SPWU Stick Pusher Warning Unit

SR Son ride mode

VIS Vertical Speed

VSI Vertical Speed Indicator

VHF VHF band communication radio COM

VHF VHF Navigation radii N AV

WX Weather


SECTION 1 GENERAL

CONVERSION INFORMATION

All numerical data contained in this AFM is shown in standard format with the metric equivalent immediately following in parenthesis, ex. T 3' (2.1 m). The lollowing formulas can be used to make required conversions.

GENERAL

Fahrenheit ("F) = ("C x 1.8) + 32

Celsius ("C) = ("F - 32) x 0.556

Statute Mile (sm) = Nautical Mile (nm) x 1.151

Nautical Mile (nm) = Statute Mile (sm) x 0.869

Jet Fuel (JET A) Standard Weights at 15" C (Relative Density 0.806)

One (1) Liier = 1.777 Ib

One (1) U.S. Gallon (US gal) = 6.73 Ib

One (1) Imperial Gallon (IMP gal) = 8.078 lb

STANDARD TO METRIC

Millimeters (mm) = Inches (in) x 25.4

Centimeters (cm) = lnches (in) x 2.54

Meters (m) = Feet (ft) x 0.305

Meters (m) = Yards (yd) x 0.914

Kilometers (km) = Statute Miles (sm) x 1.61

Kilometers (km) = Nautical Miles (nm) x 1.852

Liers = US Gallons (US gal) x 3.785

L i e n = lmperial Gallons (IMP gal) x 4.546

Kilograms (kg) = Pounds (Ib) x 0.454

Bar = psi x 0.069


Report No: 01973-001 1-19 I

SECTION 1 GENERAL

METRIC TO STANDARD

Inches (in) = Millimeters (mm) x 0.039

Inches (in) = Centimeters (cm) x 0.393

Feel (11) = Meters (m) x 3.281

Yards (yd) = Meters (m) x 1.094

Statute Miles (sm) = Kilometers (km) x 0.621

Nautical Miles (nm) = Kilometers (km) x 0.54

US Gallons (US gal) = Liters x 0.264

Imperial Gallons (IMP gal) = Liters x 0.22

Pounds (Ib) = Kilograms (kg) x 2.205

psi = Bar x 14.504

Report No: 01973-001 1 1-20 Issued: June 10,1994

Revision 8: September 1, 1998

1 PILOT'S OPERATING I HANDBOOK

Limitations

= PILATUS Z? PC XI1

SECTION 2 LIMITATIONS

Subject

GENERAL

AIRSPEED LIMITATIONS

SECTION 2

LIMITATIONS

TABLE OF CONTENTS

AIRSPEED INDICATOR MARKINGS

POWER PLANT LIMITATIONS

ENGINE ENGINE OPERATING LIMITS

OIL Oil Grade or Specification Oil Quantity

FUEL Approved Fuel Grades Anti-Icing Additive

PROPELLER STARTER GENERATOR POWER CONTROL LEVER OPERATION CHIP DETECTOR

POWER PLANT INSTRUMENT MARKINGS

MISCELLANEOUS INSTRUMENT MARKINGS

WEIGHT LIMITS


Page

Report No: 01973-001 2 -1 I


Page Subject

CENTER OF GRAVITY LlMlTS

MANEUVER LlMlTS

FLIGHT LOAD FACTOR LlMlTS

FLIGHT CREW LlMlTS

KINDS OF OPERATION

PNEUMATIC DEICING BOOT SYSTEM

ICING LIMITATIONS

SEVERE ICING CONDITIONS

KINDS OF OPERATION EQUIPMENT LIST

FUEL LIMITATIONS

MAXIMUM OPERATING ALTITUDE LlMlTS

OUTSIDE AIR TEMPERATURE LlMlTS

CABIN PRESSURIZATION LlMlTS

MAXIMUM PASSENGER SEATING LlMlTS

SYSTEMS AND EQUIPMENT LlMlTS

STALL WARNINGISTICK PUSHER SYSTEM TRIM SYSTEMS HEATED WINDSHIELD FIRE DETECTION SYSTEM ENGINE ICE PROTECTION OXYGEN SYSTEM EFlS PNEUMATIC DEICE SYSTEM PROBE HEAT ENGINE INSTRUMENT SYSTEM (EIS) FLAP SYSTEM CYCLE LlMlTS AUTOPILOT AHRS

Issued: June 10,1994 Revision 1 1 : March 1,2003

Repolt No: 01973-001 2-ii


Subject

OTHER LIMITATIONS

PASSENGER SEAT LAP BELT EXTENSION CARGO LIMITATIONS EXECUTIVE CABIN FLOODLIGHTS LUGGAGE LIMITATIONS STRUCTURAL LIMITATIONS SMOKING

PLACARDS

EXTERIOR COCKPIT CABIN SEATING VARIATIONS

Issued: June 10,1994 Revision 1 1 : March 1, 2003

Page

Report No: 01 973-001 2-iii


Report No: 01973-001 I 2-iv


Issued: February 14, 1994 Revision 4: Novenber 24,1995 '


GENERAL

This section contains the FOCA Approved operating limitations, instrument markings, color coding. and basic placards necessary lor the operation of the airplane, its engine, systems, and equipment. Compliance with approved limitations is mandatory.

Limitations associated with systems or equipment which require POH supplements are included in Section 9. Supplements.


Issued June 10. 1994 Rev~s~on 7 July 1, 1997

AIRSPEED

Maximum operating speed

- 'MO

- M~~

Maximum Operating Maneuvering Speed - V,

9039 Ib (4 100 kg)

8380 Ib (3800 kg)

7940 Ib (3600 kg)

7500 Ib (3400 kg)

7060 Ib (3200 kg)

661 0 Ib (3000 kg)

61 70 Ib (2800 kg)

5730 Ib (2600 kg)

Maximum flap extended speed - V,,

r 15" . 15"

Report No. 01973-001 2.1

SIGNIFICANCE

Do not exceed this speed in any operations. Refer to VMo schedule

for niaximum speed above 15,200 I t

(See Fig No. 2-1. Vmo Schedule)

Do not make lull or abrupt control movements above this speed.

Do not exceed this speed w~th flaps extended.

KCAS

240 .

0.48

154

148

144

140

136

132

127

123

165

130

KlAS

236

151

145

141

137

133

129

124

120

163

130


AIRSPEED LIMITATIONS (CONT'D)

MAXIMUM OPERATING SPEED

Figure 2- 1, VMo Schedule I

Do not retract or extend landlng gear above this speed.

Do not exceed thls speed WIUI landlng gear extended

Report No: 01973-001 1 2-2

177

236

Max~mum landlng gear operalrng speed - VLo

Maxrmurn landlng gear extended speed - V,


1

180

240



#

POWER PLANT LIMITATIONS

I ENGINE

REMARKS

Maximum operating limit (VMdMM,)

Normal operating range. Lower limit is maximum weight stall speed in the clean configuration (Vs). Upper limit is the maximum operating speed (V,,@h,).

Full flap operating range. Lower limit is maximum weight stall speed in landing configuration (V,). Upper limit is maximum speed with full flaps extended (VFE). -

MARKING

Red Line or

ReWhite Needle

Green Arc

White Arc

I Number of Engines

Engine Manufacturer

Engine Model Number

KIAS VALUE OR RANGE

236 or 0.48 M whichever is

lower

86 to 236

59 to 130


1

Pratt & Whitney Canada

PT6A-678

Report No: 01973-001 2-3

LlMl I A I IONS ?C XI1

ENGINE OPERATING LIMITS

The limits presented in each column shall be observed. The limits presented do not necessarily occur simultaneously. Refer to the Pratt 8 Whitney Engine Maintenance Manual for specific action i f limits are exceeded.

Report No: 01 973-001 2-4

OPERATING

CONDITION

TAKEOFF (9)

MAX. CONT.

MAX. CLIMB/ CRUISE

MIN IDLE

STARTING

TRANSIENT

MAX.

REVERSE L

Issued: June 10,1994 , Revision 5: May 10, 1996

SHP

1200

1000

900

TORQUE

PSI

( 1 )

44.34

36.95

61.00

( 4 )

34.25

MAX

ITT

" C

800

760

750

( 5 )

lo00

( 3

870

( 4 )

760

Ng Yo

( 8

104

104

50.7 (G.1.)

.64

( F. I. )

104

NP RPM

( 8 )

1700

1700

1P'O

( 4 )

1650

OIL

PRESS

PSI

( 2 )

90TO

135

90TO

135

60 MIN.

200MAX.

40 TO

200

( 4 )

90TO 135

OIL

TEMP

"C

( 6 ) ( 7 )

10TO

110

10TO

1 05

-40TO 110

-40 MIN.

-40 TO 110

10TO 1 05

SECTION 2

k 1 2 LIMITATIONS

( 1 ) Torque limit applies within a range of 1000 to 1700 propeller rpm. Torque is limited to 23.9 psi below 1000 propeller rpm.

( 2 ) Normal oil pressure Is 90 to 135 psi at gas generator speeds above 72%. With engine torque below 35.87 psi, minimum oil pressure is 85 psi at normal oil temperature (60 to 70" C). Oil pressures under 90 psi are undesirable. Under emergency condiiions, to complete a flight, a bwer oil pressure of 60 psi is permissable at reduced power level not exceeding 23.9 psi torque. Oil pressures below 60 psi are unsafe and require that either the engine be shut down or a landing be made as soon as possible using the minimum power required to sustain flight.

( 3 ) These values are time limited to 5 seconds maximum.

( 4 ) These values are time limited to 20 seconds maximum.

( 5 ) Applies over a speed range of 50.7% to 61.4% Ng rpm.

( 6 ) For increased service life of the engine oil, an oil temperature of between 74 to 80" C (60 to 70" MSN 273 AND UP and Post SB 79-003) is recommended. I

( 7 ) Oil temperature limits are -40" C to 105' C with limited periods of 10 minutes at 105 to 1 10°C.

( 8 ) 100% gas generator speed corresponds to 37468 rpm. 100% power turbine speed (N1) corresponds to 29894 rpm which also corresponds to 1700 rpm propeller speed.

( 9 ) Takeoff power is time limited to 5 minutes.


Report No: 01 973-001 2-5


OIL

OIL GRADE OR SPECIFICATION

Any oil specified by brand name in the latest revision of Pran 8 Whitney Service Bullelin14001 is approved.

Do not mix types or brands of oil.

OIL QUANTITY

Total Oil Capacity 3.6 US gal (13.6 liters)

Drain and Refill Quantity 2.0 US gal (7.6 liters)

Oil Quantity Operating Range 1.0 US gal (3.8 liters)

An oil quantity check is required for takeol. Takeoff is not approved with the OIL O N annunciator illuminated.

FUEL

APPROVED FUEL GRADES

JET A, JET-A-1. JET B, JP-4

Any other fuel which complies with the latest revision of Pratt 8 Whiney Service Bulletin 14004.

ANTI-ICING ADDITIVE

Anti-icing additive must be used for all flight operations in ambient temperatures below 0" C.

I Use anti-icing additive conforming lo MIL-DTL-27686 or MIL-DTL-85470.

Anti-icing additives should be in compliance to Prati 8 Whitney Service Bulletin 14004.

Additive concentration must be between a minimum of 0.06 % and a maximum of 0.15 % by volume.

1 CAUTION I THE CORRECT MIX OF ANTI-ICING ADDITIVE WITH THE FUEL IS IMPORTANT. CONCENTRATIONS OF MORE THAN THE MAXIMUM (0.15% BY VOLUME) WlLL CAUSE DAMAGE TO THE PROTECTIVE PRIMER AND SEALANTS OF THE FUEL TANKS. DAMAGE WlLL OCCUR IN THE FUEL SYSTEM AND ENGINE COMPONENTS.

Refer to Section 8, Handling, Servicing, and Maintenance for blending instructions.

Report No: 0 1973-00 1 2-6

Issued: June 10,1994 Revision 10: September 1,2000


PROPELLER

Propeller Manufacturer

Propeller Model Number

Number of Propellers

Number of Propeller Blades

Propeller Diameter

Minimum

Maximum

Propeller Operating Limits (Np)

Maximum Normal Operation 1,700 rpm

Maximum transient ( 20 sec.) 1,870 rpm

Maximum reverse 1,650 rpm

Stabilized operation on the ground between 350 and 950 rpm is not permitted.

Blade Angles at Station 42

Fine Pitch 19" +I- 0.1"

Maximum Reverse Pitch -17.5" +I- 0.5"

Feather 79.6" +I- 0.5"

Minimum pitch in flight 6"

STARTER

The engine automatic starting cycle shall be limited to the following intervals:

1. Sequence, 60 seconds OFF

2. Sequence. 60 seconds OFF

3. Sequence, 30 minutes OFF

Repeat as required.


Report No: 01973-001 2-7


GENERATOR

Maximum generator load limit as follows

POWER CONTROL LEVER OPERATION

GENERATOR

Starter / Generator

Second Generator

Power Control Lever operation aft of the idle detent is prohibited:

1. When engine is not running.

'Maximum load permitted for a 2 minute period per each one hour of operation.

MAX CONTINUOUS LOAD

300 AMP

115 AMP

2. During flight. Such operation may lead to loss of airplane control.

MAX LOAD FOR 2 MINUTES '

450 AMP

NI A

3. When engine is controlled by the Manual Override System. Such operation may lead to loss of airplane control or may result in an enginelpropeller overspeed condition and consequent loss of engine power.

CHIP DETECTOR

Takeoff is not approved with CHIP annunciator illuminated.

1 Report No: 01 973-001 2-8



POWER PLANT INSTRUMENT MARKINGS

MISCELLANEOUS INSTRUMENT MARKINGS

-!

RED. RADlDIA Max. Limit

104%

8OOn1 1.000' C

t10‘C

135/200

44 3.'6 1 .O

'Cabin Differential Pressure Caution Range is Amber.

Engine Speed (Ng)

ITT

Oil Temperature

Oil Pressure (psi)

Torque (psi)

Instrument

Oxygen Pressure

(Psi)

Cabin Differential

(Psi)

Issued: June 10. 1994 Revlsion 7. July 1. 1997

RED DINRAD Min. Limit

N/ A

NIA

N: A

40!60

N/ A

Report No: 01973 001 2-9

YELLOW ARC

Caution

N!A

760" lo 800" C

105" 10 110°C

N/ A

36.9 lo 44.3

YELLOW ARC

Caulion

N:A

NIA

10" to 55" C

60 to 90

N! A

RED RADIAL

Min. Limit

NIA

N!A

GREEN ARC

Norm Ops.

60 lo 104O/o

400" to 760" C

55' to 105" C

90 10 135

0 lo 36.9

YELLOW ARC

Caution

NIA

N/A

YELLOW ARC

Caulion

NIA

5.75 to 6 50'

GREEN ARC

Norm Ops

NiA

0 10 5.75

RED RAD.'DIA Max. L m t

1850 to 2000

6 50


WEIGHT LIMITS

Max~mum Ramp Welght

Max~mum Takeoff We~ght

Max~mum Landlng Welghl

Max~mum Zero Fuel Welght

Maxlmum Baggage Welght

Maxlmum Floor Loadlng - On Seat Ralls

On Cabln Floor

Repoil No: 01973-001 2-10

9,083 Ib (4.120 kg)

9.039 Ib (4.100 kg)

9.039 Ib (4.100 kg)

8,160 Ib (3.700 kg)

400 Ib (1 80 kg)

Issued: February 14, 1994 Revlsion 2: February 14. 1995


Forward Limit

I 1

NOTES

Straight line variation between points given.

The datum is I 18 in (3.0 m) forward of firewail.

It is the responsibility of the pilot to ensure that airplane is loaded properly.

See Section 6, Weight and Balance for proper loading instructions.


Report No: 01973-001 2-1 1


MANEUVER LIMITS

This airplane is certificated in the Normal Category. The normal category is applibk to aircraft intended for non-acrobatic operations. These include any maneuvers incidental to normal flying, stalls (except whip stalls), lazy eights, chandelles, and turns in which the bank angle does not exceed 60".

Acrobatic maneuvers, including spins, are not approved.

FLIGHT LOAD FACTOR LIMITS

Flight load limits with Naps up +3.4 g, -1.36 g

Flight bad limits with flaps down +2.0 g, -0.0 g

FLIGHT CREW LIMITS

Minimum required flight crew is one pilot in the left hand seat.

KINDS OF OPERATION

The Pilatus PC-12 is approved for the following types of operation when the required equipment is installed and operational as defined within the Kinds of Operation Equipment List:

1. VFR Day.

2. VFR Niiht.

3. IFR Day incl. CAT 1 approaches, single pilot.

4. IFR Niiht incl. CAT 1 approaches, single pilot.

5. Flight into Known Icing Conditions.

PNEUMATIC DEICING BOOT SYSTEM

I The wing and tail leading edge pneumatic deicing boot system must be activated at the first sign of ice formation anywhere on the aircraft, or upon annunciation from an ice detector system installed), whichever occurs first.

The wing and tail leading edge pneumatic deicing boot system may be deactivated only after leaving icing conditions and after the aircraft is determined to be clear of ice.

Issued: June 10.1994 Revision 9: September 1,1999


ICING LIMITATIONS

Icing conditions can exist when:

The outside air temperature (OAT) on the ground and for takeoff, or total air temperature (TAT) in flight, is 10°C or colder, and visible moisture in any form is present (such as clouds, fog or mist with visibility of one mile or less, rain snow, sleet and ice crystals).

The OAT on the ground and for take-off is 10°C or colder when operating on ramps, taxiways or runways, where surface snow, ice, standing water, or slush may be ingested by the engine, or freeze on the engine, or the engine nacelle.

There are visible signs of ice accretion on the aircraft.

Flight in icing conditions is only approved with all ice protection systems, generator 1 and generator 2 serviceable.

During flight in icing conditions, if there is a failure of any of the aircraft ice protection systems or generator 1 or generator 2, exit icing conditions. Contact ATC for priority assistance if required.

Prolonged flight in severe icing conditions should be avoided as this may exceed the capabilities of the aircraft ice protection systems.

During flight in icing conditions or flight with any visible ice accretion on the airframe. the following flap maximum extension limits apply:

- With operational airframe pneumatic deice boots 15" FLAP.

- After failure of the airframe pneumatic deice boots O" FLAP.

In the event of a balked landing go-around with residual ice on the airframe. the flaps should not be retracted from the 15" position.

Flight in freezing rain, freezing fog, freezing drizzle and mixed conditions causing ice accretion beyond the protected areas of the pneumatic boots is not approved.

The aircraft must be clear of all deposits of snow, ice and frost adhering to the lifting and control surfaces immediately prior to takeoff. I In the event of a balked landing (go around) with residual ice on the airframe, the landing gear and flaps may not fully retract after selection.

Operation of the pneumatic de-ice boot system in ambient temperatures below -40°C and above +4VC may cause permanent damage to the boots.

Issued: June 10, 1994 Revision 6: Dec 6. 1996

Report No: 01973-001 2-13


The lelt wing ~nspection light must be operative prior to flight into known or forecast icing condit~ons at night.

SEVERE ICING CONDITIONS

Severe icing may result from environmental conditions outside of those for which the airplane is certificated. Flight in freezing ram, freezing drizzle, or mixed icing conditions (supercooled l~quid water and ice crystals) may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming afl of the protected surfaces. This ice may not be shed using the ice protection systems, and may seriously degrade the perlormance and controllability 01 the airplane.

During Il~ght, severe lclng conditions that exceed those for which the airplane is certificated shall be determined by the following visual cues. If one or more of these visual cues exists, immed~ately request prior~ty handling from Air Traffic Control to facilitate a route or an altitude change to exit the icing conditions:

unusually extensive ice accumulation on the airframe and windshield areas not normally observed to collect ice

accumulation of ice beyond the active portions of the wing pneumatic boots

Care must be taken when using the autopliot that tactile cues, such as increased aileron forces, are not masked by the autoptilot function. Periodically disengage the autopilot to check lor abnormal forces.

Report No: 01 973-001 2-14



KINDS OF OPERATlON EOUIPMENT LIST

This airplane is approved for operations under day or night VFR, day or night IFR and flight into known icing conditions when the required equipment is installed and operating properly. The following systems and equipment list is predicated on a crew of one pilot and does not include specific flight and radiolnavigation equipment required by any particular country's operating regulations. The pilot in command is responsible for determining the airworthiness of the aircraft and assuring compliance with current operating regulations for each intended flight.

The zeros (0) used in the list below mean that the system andlor equipment was not required for type certification for that kind of operation. When (AR) appears for the number required it ind i tes As Required.

Deviations from this KOEL may be approved for the operation of a specilic aircraft if a proper MEL (Minimum Equipment List) has been authorized by the appropriate regulatory agency.

SYSTEM I EQUIPMENT VFR VFR IFR IFR ICING DAY NIGHT DAY WIGHT

Flight Instruments:

Airspeed lndicator Sensitive Altimeter Magnetic Compass Attitude Indicator (EADI) Attitude lndicator (Self contained) Rate of Turn lndicator (EADI) Slip-Skid lndicator (Inclinometer) Direclional lndicator (EHSI) AHRS Vertical Speed Indicator Clock

Englne Instruments (EIS):

Torquemeter 1 I 1 1 1 Propeller Tachometer (Np) 1 1 1 1 1 Inter Turbine Temperature Gauge (1T-r) 1 1 1 1 1 Gas Generator Tachometer (Ng) 1 1 1 1 1 Oil Pressure Gauge 1 1 1 1 1 Oil Temperature Gauge I 1 1 1 1

Miscellaneous Instruments (EIS):

Fuel Quantity Indicators L FUEL LOW Annunciator


Report No: 01973-001 2-15 1


SYSTEM I EQUIPMENT VFR DAY

VFR IFR IFR ICING NIGHT DAY NIGHT

I Miscellaneous lnstrs (EIS) (con'd):

R FUEL LOW Annuncialor DC Voltmeler DC Ammeter Outside Air Temperature Gauge

Engine:

Starter Generator Second Generator Inertial Separator INERT SEP Annunciator Engine Driven Low Pressure Fuel Pump FUEL PRESS Annunciator Electric Wing Tank Fuel Boost Pump L FUEL PUMP Annunciator R FUEL PUMP Annunciator Firewall Fuel Shuloll Valve FCU Manual Override System Torque Limiter PROP LOW P Annunciator lgnit~on System IGNITION Annunciator Fire Detect System ENG FlRE Annunciator FIRE DETECT Annunciator OIL QTY Annunciator CHIP Annunciator

Battery BAT OFF Annunciator BAT HOT Annunciator

I 2"' Battery (il installed) 2" BAT OFF Annunciator (if installed) 2" BAT HOT Annunciator (if Installed) Stall WarningISlick Pusher System AOA Probes PUSHER Annunciator AIR 1 GND Annunciator Vm0 Overspeed Aural Warning

CAWS

Report No: 01973-001 2-16

Issued: June 10.1994 Revision 8: September 1. 1998


SYSTEM I EQUIPMENT VFR VFR IFR IFR ICING DAY NIGHT DAY NIGHT '

Eleclrlcal (cont'd):

Longludinal (Stab) Trim System 1 Alternate Stab Trim System 1

' STAB TRIM Annunciator 1 Lateral Trim System 1 Directional Trim System 1 Triple Trim Indicator 1 Trim Interrupt System 1 Windshield Heat 2 WSHLD HT Annunciator 1 INVERTER Annunciator 0 GEN 1 OFF Annunciator 1 GEN 2 OFF Annunciator 0 ESNTL BUS Annunciator 1 AV BUS Annunciator 1 N ESNTL B Annunciator 1 BUS TIE Annunciator 1 Posl in Lights 0 Strobe Lights 0 Landing Lights 0 Taxi Light 0 Instrument and Panel Lighting 0 Audio System 1 Cockpit Speaker 1 Cabin Speaker 1 PUSHER ICE MODE Annunciator 0 DE ICE Annunciator 0 DE ICE BOOTS Annunciator 0 Deice Boot Timer 0 AOA Heater LH 1 AOA Heater RH 1 Probe Current Monitor 1 PROBES DE ICE Annunciator 1 AOA DE ICE Annunciator 1 Propeller Deice Timer 0 Propeller Deice Brush 0 Propeller Deice MOV 0 Propeller Deice Boots 0 Propeller Deice OAT Sensor 0 PROP DE ICE Annunciator 0 Left Wing Inspection Light 0

' Refer to Section 2 System and Equipment Limits

1 1 1 1 AR AR AR AR 1 1 1 1 1 1 1 1 1 1 1 1 0 0 0 1 0 0 0 1 0 0 0 1 0 0 0 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 0 0 0 1 0 0 0 1 0 0 0 1 0 0 0 4 0 0 0 2 0 0 0 0 0 0

Heated WindshieM for the actual limitation 1 Issued: June 10,1994 Revision 8: September 1, 1998

Report No: 01973-001 2-17


SYSJEM I EQUIPMENT

Mechanical Systems :

Landing Gear Actuating System HYDR Annunciator Gear Position IndicationsNVarning Emergency Gear Extension System Flap Control & Indication Flap Interrupt System

I FLAPS Annunciator Seat Restraints (each occupant) PASS DOOR Annunciator CAR DOOR Annunciator Firewall ECS Shutott Valve Emergency Ram Air Scoop Oxygen System PASS OXY Annunciator Deice Boot PRV Deice Boot EFCV's Deice Boot Pressure Switches Deice Boot, Inner Wing LH Deice Boot, Outer Wing LH Deice Boot, lnner Wing RH Deice Boot. Outer Wing RH Deice Boot, Tail LH Deice Boot, Tail RH

For Pressurized Flight:

Pressurization Control System (CPCS) Cabin Altimeter Cabin Vertical Speed Indicator Cabin Dinerentiat Pressure Indicator Maximum Cabin Altitude Warning CAB PRESS Annunciator ECS ECS Annunciator

Report No: 01 973-001 2-18

VFR VFR DAY NIGHT

IFR IFR ICING DAY NIGHT

Issued: June 10.1994 Revision 8: September 1. 1998 ,


FUEL LIMITATIONS

Total Fuel Capacity 406.8 US gal, 2,736.5 Ib (1.540 liters, 1,241.3 kg)

Total Usable Fuel MSN 101-140 400.4 US gal, 2,693.3 Ib (1,515.7 liters, 1,221.7 kg)

MSN 141 AND UP 402 US gal, 2,703.6 Ib (1,521.5 liters, 1,226.4 kg)

Total Unusable Fuel MSN 101-140 6.4 US gal. 43.2 Ib (24.3 liters, 19.6 kg) MSN 141 AND UP 4.8 US gal, 32.9 Ib (18.5 liters, 14.9 kg)

Maximum Fuel Imbalance 26.4 US gal, 178 Ib (100 liters, 80.6 kg) (Maximum 3 LCD segments on indicator)

NOTE

Usable fuel can be safely used during all Normal Category airplane maneuvers.

MAXIMUM OPERATING ALTITUDE LIMITS

Maximum Operating Altitude 30,000 ft (9,144 m)

OUTSIDE AIR TEMPERATURE LIMITS

Minimum Outside Air Temperature -55' C (-67" F)

Maximum Outside Air Temperature +50° C (122" F)

CABIN PRESSURIZATION LIMITS

Maximum cabin pressure differential is 5.75 psi (400 mbar).

Pressurized landing is not approved.


Report No: 01973-001 2-19 1


MAXIMUM PASSENGER SEATING LIMITS

Maximum number of occupants is 9 passengers plus pilot(s).

Refer to Section 6, Weight and Balance, for seat locations.

The PC-12 was ceftified with two basic cabin interior configurations, a Corporate Commuter and an Executive interior. Variations to the two basic interior configurations that have been approved together with general limitations are given below:

Corporate Commuter lnterior Code STD-9s nine standard seats. Corporate Commuter lntehor Code STO-6s-36 six standard seats and three seat bench. Executive lnterior Code EX-6s six executive seats. Executive lnterior Code EX-8s eight executive seats. Leave seats 5, 6, 7 and 8 vacant during takeoff and landing unless Seat in front is occupied. Executive lnteiir Code EX-4s-38 four executive seats and three seat bench. Executive Interior Code EX-6s-STD-2s six executive seats and two standard seats. Leave seats 5, 6. 7 and 8 vacant during takeoff and landing unless Seat in front is occupied. Executive Interior Code EX-4s-STD-4s four executive seats and four standard seats.

Piatus must be contacted to determine the modification work required to the aircraft. before any change to an interior configuration is made.

SYSTEMS AND EQUIPMENT LIMITS

STALL WARNlNOlSTICK PUSHER SYSTEM

Preflight function test required before takeoff.

System is required to function properly in normal mode for all flights and in ice mode for flight into known icing conditions.

TRIM SYSTEMS

Stabilizer normal and alternate, and rudder trim systems must function properly for all flights.

HEATED WINDSHIELD

Left Hand and Right Hand Heated Windshields must function properly for all flights. Exception, for IFR flights conducted into no known or forecast icing conditions at least

Report No: 0 1913-00 1 2-20

Issued: June 10, 1994 Revision 9: September 1, 1999

SECnON 2 LIMITATIONS

one heating zone of the windshield on the side of the pilot in command must funcHon properly.

FIRE DETECTION SYSTEM

PrefligM Function Test is required for takeoff.

System must function properly for all flights.

ENGINE ICE PROTECTION

Preflight Function Test is required for takeoff.

OXYGEN SYSTEM

A minimum oxygen supply of 10 minutes duration for each occupant is required for dispatch for pressurized flight above F1250.

NOTE

Some National Operating Requirements may require that a larger quantity of oxygen be carried on the aircraft.

Post SB 35-001. MSN 181 and UP - The oxygen system shut-off valve handle in the cockpit must be selected to on prior to engine start and throughout the duration of flight.

The oxygen masks for the crew must be connected for all flights.

For aircraft with the Corporate Commuter side wall paneling, oxygen masks must be connected and properly stowed for each passenger prior to takeoff when the aircraft is to be operated above 10,000 feet.

NOTE

In the executive interior configurations the oxygen masks are permanently connected.

EFlS

During EFS 40150 operation, the B e n d i n g EFS 40150 Pilot's Guide must be on board the airplane and immediately available to the pilot.

Standby artifiiial horizon must be operational for departure (IFR operation only).

No SG or DU flag may be visible prior to departure (IFR operation only).


Report NO: 01973-001 2-21 1

SECTION 2

PNEUMATIC DEICE SYSTEM

The pneumatic deice system boots are required to be instaUed for all flights.

Preflight function test required before takeoff and flight into known icing conditions.

The system is required to function properly for flight into known icing conditions.

PROBE HEAT

Preflight function test required before takeoff.

The system is required to function properly for IFR flight and flight into known king condition.

ENGINE INSTRUMENT SYSTEM (EIS)

EIS must function properly for all flights.

EIS system test required before takeoff.

Takeoff is not approved with the red EIS warning light illuminated after a system test.

FLAP SYSTEM CYCLE LIMITS

A flap cycle is defined as mQV@m@nt from 0" to 15' to 0" and from 0" to 15" to 40" to 0". Maximum number of cycles per hour -

Up to 25' C OAT 10

25" C to 50' C OAT 8

AUTOPILOT

During autopilot operation, the BendiiKing KFC 325 Pilot's Guide must be on board and immediately available to the pilot.

If the autopilot is to be used in flight, the entire preflight test must be successfully completed prior to each flight.

During autopilot operation, a pilot must be seated in a pilot position with seat belt fastened.

The autopilot (AP) and yaw damper (YD) must be OFF during takeoff and landing.

Altitude Select captures below 1000 feet AGL are prohibited.

The autopilot must be disengaged, when the airplane is below 1000 ft AGL, except in accordance with the conditions given below.



For airplanes equipped with a functioning Pilatus option radar altimeter installed, the autopilot must be disengaged below 200 ft AGL during ILS approach operations provided that the autopilot is coupled to glideslope vertical guidance of 6" or less. The system Is approved for Category 1 operation (Approach mode selected).

Do not override the autopilot to change pitch or roll attitude.

Continued autopilot operation Is prohibited following abnormal operation or malfunctioning prior to corrective maintenance.

I CAUTION 1

In accordance with FAA recommendation (AC 00-24A), use of "PITCH ATTITUDE HOLD" mode is recommended during operation in severe turbulence.

AHRS

Fly straight and level for 1 minute after each 15 minutes of continuous orbiting maneuvering (not applicable to LCR-93 AHRS).


Report No: 01 973-001 2-23


OTHER LIMITATIONS

PASSENGER SEAT LAP BELT EXTENSION

The lap belt extension PaR No. 959.30.01.588 can be used on all standard passenger seats Pad Nos. 959.30.01 501-520 and 525.22.12.01 11012. It's use is limited to those who need it and it shall be handed out by the pilot on a case by case basis before flight. The lap belt extension must not be used for strapping small children sitting on a persons lap.

CARGO LIMITATIONS

Maximum Freight Load 2200 Ibs (1 000 kg)

Cargo must be arranged to permit free access to the left hand cabin door and the right hand emergency overwing exit. No cargo must be placed on the seats.

All cargo must be secured by approved Cargo Restraints as described in Section 6. Tie Down Straps with a breaking strength of at least 1800 Ib per strap must be used. All CargoIContainers must be located against a Retaining Angle secured laterally to the seat rails.

Items up to a total weight of 66 Ib (30 kg) can be stowed in the cabin area without being strapped down providing a Cargo Net is installed in front of the items. Cargo Nets may only be installed on the attachments at Frames 24 and 27. No passengers must be seated rearward of a Cargo Net.

If an extendable baggage net is used the tie down fittings and the cargo strap fittings must have a minimum space of 5 inches between the fittings.

EXECUTIVE CABIN FLOODLIGHTS

The cabin floodlights that are Pre SB 33-008 in executive interior aircraft shall not be operated on the ground for more than 2 hours in any 4 hour period, if the outside air temperature is greater than 25" C.

Repon No: 01973-001 2-24



LUGGAGE LIMITATIONS

The luggage area maximum load is given in the following table. The load is dependent on the aircraft interior configuration and the Part No. of the luggage net installed (Post SB 25-01 9).

A Luggage Net must be installed at Frame 34 when luggage is stowed.

The luggage area maximum load is 500 Ib (225 kg) with an extendable luggage net installed (Post SB 25-010). The extendable luggage net andlor any luggage may not extend in front of frame 32. If the extendable luggage net is used without a three seat bench installed, there must be a clear area in front of the net as follows:

8 at least 280 mm forward of frame 32, when the net floor attachments are placed at frame 32 (the most forward position of the net)

at least 340 mm fonrvard of frame 34, when the net floor attachments are placed at frame 34

When an optional wardrobe is installed, the maximum weight limit for luggage items stowed in the bottom of the wardrobe is 35 Ib (15.9 kg). All stowed luggage in the wardrobe must be secured with the safety net.

STRUCTURAL LIMITATIONS

Refer to Chapter 4 of the PC-12 Aircraft Maintenance Manual, Pilatus Report Number 02049.

SMOKING

Smoking is not permitted in the cabin of aircraft equipped with a standard interior unless ashtrays are installed.


Report No: 01 973-001 2-25 1


PLACARDS - EXTERIOR

On exterior Cabin Door:

PULL HANDLE AND TURN TO OPEN

DO NOT OPEN DOOR WHEN ENGINE IS RUNNING

UNLESS IN EMERGENCY

OPEN

Report No: 01973-001 1 2-26

Issued: June 10,1994 Revision 1 1 : March 1.2003


On exterior Cargo Door:

PRESS HERE TO OPEN PULL HANDLE AND

PULL DOOR OUT DO NOT OPEN DOOR WHEN

ENGINE IS RUNNING UNLESS IN EMERGENCY

PULL TO OPEN

Issued: June 10,1994 Revision 11: March 1, 2003

Report No: 01973-001 2-27 1


Near Static Ports:

STATIC PRESSURE I KEEPCLEAR I

On left side Vertical Tail forward of Horizontal Stabilizer:

On Rudder (each side):

1 DO NOT PUSH I

Report No: 01973-001 1 2-28



[ On exterior Emergency Exit: (Not to Scale)

t

t 1 PUSH

1

PUSH IN AFTER RELEASE

r Inside left Engine Cowling:

, TURBINE OIL ACCEPTABLE OILS SEE P+W SB 14001

TOTAL SYSTEM CAPACITY 14,5 QRT 13,6 LTR

DO NOT MIX OIL TYPES

NOTE: The engine oil type used will be added to the placard prior to delivery of the aircraft. I

Issued: June 10.1994 Revision 1 1 : March 1,2003

Report No: 01 973-001 2-29 1


On Nose Landing Gear (each side):

On Nose Landing Gear:

I k DO NOT TURN BEYOND RED MARKS--+ I =

Near Fuel Filler:

MSN 101-140 MSN 141 AND UP

FUEL: ASTM-0-1655 JET A, JET A-1 AND JET B

(CPW 204 SPEC)

TOTAL CAPACITY 770 LTR. 203 US. GAL.

USABLE CAPACITY 758 LTR. 200 US. GAL.

1 REFUELING BONDING POINT 1

FUEL: ASTM-D-1655 JET A, JET A-1 AND JET B

(CPW 204 SPEC)

TOTAL CAPACITY 770 LTR. 203 US. GAL.

USABLE CAPACITY 761 LTR. 201 US. GAL.

On top surface of each Aileron and three places on top surface of each flap:

( DO NOT PUSH I On the main landing gear doors:

I TYRE PRESSURE 55 PSI^

On the nose landing gear doors:

I TYRE PRESSURE 60 PSI[

Report No: 01973-001 1 2-30


C r PILATUSI SECTION 2

i PC12 LIMITATIONS

PLACARDS - COCKPIT

/ On the Instrument Panel:

PRE SB 11 -002

VFE (15") 163 KlAS

1 vFE (4001 130 KlAS

POST SB 1 1-002 7

L I VFE ABOVE 15' 130 KlAS I

Vo (4100KG) 151 KlAS

Vo (2600KG) 120 KlAS

VMO 236 KlAS

MMO 0.48 M

L On the center panel below Display Unit:

STABILIZED PROPELLER OPERATlON ON GROUND (NOT FEATHERED) BETWEEN 350 AND 950 rpm IS PROHIBITED

k Near Fuel Quantity Indicator:

CAPACITY 1516 LTR

400 US. GAL 2693 LBS JET-A1

CAPACITY 1521 LTR

402 US. GAL 2704 LBS JET-A1

I MSN 101-140 MSN 141 AND UP

I

b Near Landing Gear Selector Handle:

Issued: June 10, 1994 C Revision 1 1 : March 1,2003

r j 236 KlAS

Report No: 01 973-001 2-31


Near Pressurization Controls:

BEFORE LANDING

On Center Console:

DIFF. PRESS. = 5.75 PSlD

M U MAX

M A 0' N U

15'

0 v E R R

N lo'

1 D 0 R

E 0 U E L N V

FEATHER 40'

OFF R

W S T S a5401 AND ~ 0 ~ 1 1 1 - 0 s ~

At aft end of Center Console:

CAUTION PCL OPERATION AFT OF IDLE DETENT IS PROHIBITED WHEN: - ENGINE NOT RUNNING - IN FLIGHT - WITH MOR OPERATION


Revision 11: March 1, 2003


On center console (FWD of trim indicators):

WARNING DURING FLIGHT IN IClNQ CONDITIONS OR FLIGHT WITH ANY VISIBLE ICE ACCRETION ON THE AIRFRAME THE FOLLOWING FLAP MAXIMUM EXTENSION LIMITS APPLY:- - WITH OPERATIONAL AIRFRAME PNEUMATIC DE-ICE BOOTS 15' FLAP. -AFTER FAILURE OF THE AIRFRAME PNEUMATIC DE-ICE BOOTS 0' FLAP.

C

At rear of Center Console:

FUEL EMER SOV HANDLE HYDRAULIC HAND PUMP HANDLE ECS EMER SOV HANDLE - - Near Fuel Shut off Valve Handle (Post SB 28-002 and MSN 181 and up):

On Fuel and ECS Firewall Shut off Valve Handles:

On Cockplt LH Side Panels near oxygen system controls:

I.,]


PASSENGER OXYGEN \ Report No: 01 973-001

2-33 1


Near Standby Magnetic Compass:

FOR CORRECT REMM SWITCH:

I NAV 1 UlSTRUMENl UWNG AS REOUAED WIWSHIEU) DEXE LH 1 R H OFF

AUXILIARY nunffi SYSTEMS OFF I I AUXYURY COWNO SYSTEMS OFF I MSN 101-320

I Near DV Window:

STANDBY COMPASS FW CORREC? REAOIM MTW.

MSN 322-400 AND AIRCRAFT WITH MODIFIED HEATING SYSTEM

DV WINDOW PRESS BUlTON

AND PULL INWARDS

On Left and Right Control Wheel:

Near Cockpit Oxygen Outlet:

I On Post SB 25-020 Narco 910 ELT Remote Control Panel, located on pilots lower left panel:

ACTIVATED WHEN FLASHlNG

ON

TO TEST OR RESET: PRESS ON, WAR 2 SECONDS PRESS ARM

- SEEOWNERSYLNUM ARY

EMERGENCY USE ONLY

Report No: 01973-001 2-34

Issued: June 10,1994 Revision 1 1 : March 1,2003 '


I I

On left Coc ~t Side Panel and right Cockpit Side Panel for a dual instrumented aircraft: (LH ShownXH Opposite)

OPERATIONAL LIMITATIONS

THlS AIRPLANE MUST BE OPERATED AS A NORMAL CATEGORY AIRPLANE IN COMPLIANCE WITH THE OPERATING LIMITATIONS

STATED IN THE FORM OF PLACARDS, MARKINGS AND MANUALS.

NO ACROBATIC MANEUVERS INCLUDING SPINS ARE APPROVED.

THIS AIRPLANE APPROVED FOR VFR, IFR, DAY 6 NIGHT OPERATION 6 ICING CONDITIONS.

WARNING DO NOT SMOKE WHILE OXYGEN IS IN USE

EMERGENCY GEAR EXTENSION

- AIRSPEED 110 KlAS - ENSURE LANDING GEAR HANDLE DOWN IF 3 GREENS NOT ILLUMINATED AFTER 30 SECONDS - EMERGENCY GEAR PUMP HANDLE (LOCATED AT PEDESTAL REARSIDE) AND PUMP SLOWLY - ACTIVATE UNTIL 3 GREENS OBTAINED IF 3 GREENS STILL NOT ILLUMINATED - YAW AIRCRAFT TO LOCK LH 6 RH GEAR

REDUCE AIRSPEED TO LOCK NOSE GEAR

L

L

1

I

C

b

b

3

L On the front side of the right cockpit bulkhead: 1

t

b FIRE EXTINGUISHER

!

b

b

1

k

I

b

Issued: June 10.1994

f Revision 1 1 : March 1.2003 ?

Report No: 01 973-001 2-35 1

SECTION 2 =PILATUSW LIMITATIONS PC12

PLACARDS - CABIN The following standard placards are installed In all aircraft.

On Interior Cabin Door:

EXIT \ DO NOT OPERATE IN FLIGHT /

DO NOT OPEN DOOR WHEN ENGINE IS RUNNINO

UNLESS IN EMERGENCY

AT ANY TIME

PUSH BUTTON FOR c=2

Report No: 01 973-001

1 2-36

(OPEN)

Issued: June 1 0 , 1 9 9 4

Revision 1 1 : March 1.2003 '


r On Interior Emergency Exit:

On lnterior Emergency Exit Handle:

On lnterior Cargo Door Handle Cover:

DO NOT REMOVE COVER IN FLIGHT

On lnterior Cargo Door Handle:

LIFT LOCKING LEVER AND PULL HANDLE PUSH DOOR OUT

On lnterior Cargo Door:

DO NOT OPEN DOOR WHEN ENGINE IS RUNNING

UNLESS IN EMERGENCY

On cabin to baggage area step (MSN 322-400 and aircraft with a modified heating system):

I 1

Issued: June 10,1994 b Revision 1 1 : March 1,2003 r

-- - -

Report No: 01 973-001 2-37


On forward and rear Cargo Doot Frame:

On lower Cargo Door Frame:

BEFORE LOADING CARGO

MAX FREIGHT LOAD = 1000 kg 12200 Ib

Above Baggage Area:

Max Load on Seat Rails

1000 kg/m2 205 Ib/tt2

1 MAX BAGGAGE LOAD = 70 kg 1155 Ib I

Max Load on Floor Panels

600 kg/m2 125 1blft2

MAX BAGGAGE LOAD = 100 kg 1220 Ib

CARGO MUST NOT OBSTRUCT ACCESS TO CABIN DOOR AND

EMERGENCY EXIT

MAX BAGGAGE LOAD = 120 kg 1265 Ib

or

MAX BAGGAGE LOAD = 180 kg 1400 ib

Post SB 25-010 the above placard is replaced by:

MAX BAGGAGE LOAD = 225 kg 1500 Ib (ONLY IF BAGGAGE NET 525.25.12.028 IS INSTALLED)

At interior fuselage cargo net attachment points:

Report No: 01 973-001 2-38



PLACARDS - 9 SEAT CORPORATE COMMUTER (Interior Code STD-9s). The cabin placards plus the following additional placards are those required for this interior.

On the rear of the left and right cockpit bulkheads, and on the rear of each seat:

FOR TAKEOFF AND LANDING - FASTEN SEAT BELT - SEAT BACK MUST BE FULLY UPRIGH

On the rear of each seat. except seat N0.5:

OXYOEN MASK LOCATED UNDER YOUR SEAT

On the rear of the seat No.5:

YGEN MASK LOCATED UNDER SEAT IN FRO

On the rear of the left cockit bulkhead:

Near each Passenger Oxygen Outlet and Cover:

1 OXYGEN I On the forward cargo door frame (Post SB 25-016 and MSN 261 and UP):


INTERIOR CODE: STD-9S

(SEE AFMIPOH SECTION 6)

Repotl No: 01973-001 2-39 1


PLACARDS - 6 SEAT CORPORATE COMMUTER AND A THREE SEAT 4

BENCH (Interior Code STD-6SQB). The cabin placards, the 9 seat commuter placards and the following replacement/additional placards are required for this interior. 1

On the rear of seats 5 and 6:

OXYGEN MASK LOCATED UNDER YOUR SEAT

On the left side of the bench seat:

LEFT PARTOF BENCH SEAT MUST BE FULLY AFT DURING THE ENTIRE FLIGHT

LEFT PART OF BENCH SEAT MUST BE I FULLY AFT DURING THE ENTIRE FLIGHT 1

When the large baggage net is installed:

MAX BAGGAGE LOAD = 225 k 1500 Ib I (ONLY IF BAGGAGE NET 525.15.12.028 18 INSTALLED) I

On the forward cargo door frame (Post SB 25-016 and MSN 261 and UP):

Report No: 01973-001 1 2-40

INTERIOR CODE: STD - 6 s - 3B


Issued: June 10.1994 Revision 11: March 1,2003 '


PLACARDS - 6 SEAT EXECUTIVE (Interior Code EX-6s-1 and -2). The cabin placards plus the following additional placards are those required for this interior.

On the rear of the left bulkhead:

FIRE EXTlNQUlSHER LOCAlED (ON COCKPIT SIDE RH BULK- I))

HEAD BEHIND CaPILOT SEAT

On the inside of the lavatory doors:

OXYGEN MASK INSIDE

COMPAA

On the inside of the optional wardrobe:

BAGGAGE ALLOWED ABOVE ED DURING TAKEOFF 6 LANDING NET MUST BE CLOSED FOR FUGHT

On the inside of the left and right cabinet drawers:

Near each executive seat: FOR TAKEOFF AND LANDING

- SEAT MUST BE FULLY UPRIGHT - SEAT MUST BE FULLY UPRIGHT. FULLY TO THE REAR OF CABIN

-TABLE UUST BE STOWED AND FULLY OUTBOARD

RMREST MUST BE LOWERE

POST SB ZSM3 MTrRKHI CODE EX-1


Report No: 01 973-001 2-4 1


Near the No. 5 seat (Post SB 25-003 and INTERIOR CODE EX-65-1):

LEAVE THIS SEAT VACANT DURING TAKEOFF ( LANDING UNLESS SEAT W FRONT IS OCCUPIED

On each Passenger Oxygen Mask Cover:

OXYGEN MASK INSIDE

On the armrest near each Passenger Oxygen Mask:

OXYGEN MASK OXYGENMASK

By the ashtray near each seat:

WHILE OXYGEN

Near the optional coat rail in the baggage compartment:

(MAX COAT RAIL S kg I l l I9

On the forward cargo door frame (Post SB 25-016 MSN 261 and UP): -

EX - 6s-2 (SEE AFMIPOH SECTION 6)

Report No: 01973-001 [ 2-42



PLACARDS - 4 SEAT EXECUTIVE AND A THREE SEAT BENCH (Interior Code EX-4s-3B).

The cabin placards, the 6 seat executive placards and the following replacemenVadditionaI placards are required for this interior.

On the rear of seats 3 and 4:

On the left side of the bench seat:

YGENMASKLOCATEDUNDERYOURSE

I LEW PART OF BENCH SEAT MUST BE FULLY AFT DURING THE ENTIRE FLIGHT 1

LEFT PART OF BENCH SEAT MUST BE FULLY AFT DURING THE ENTlRE FLIGHT

When the large baggage net is installed:

MAX BAGGAGE LOAD = 225 k 1500 1b (ONLY IF BAGGAGE NET 525.25.12.028 IS INSTALLED)

On the forward cargo door frame (Post SB25-016 and MSN 261 and UP):

INTERIOR CODE: EX-4s-36

(SEE AFWPOH SECTION 6)


Report No: 01 973-001 2-43 1


PLACARDS - 8 SEAT EXECUTIVE (Interior Code EX-IS) and a 6 SEAT !

EXECUTIVE AND 2 SEAT CORPORATE COMMUTER (Interior Code EX-6s-STD-2s)

I The cabin placards, the 6 seat executive placards and the following replacemenUadditionaI placards are required for this interior.

I Near seats 5,6, 7 and 8:

LEAVE MIS SEAT VACANT DURING TAKEOFF AND LANDING UNLESS SEAT IN FRONT IS OCCUPIED

On the forward cargo door frame (Post SB 25-016 and MSN 261 and UP):

INTERIOR CODE: EX - 8s


INTERIOR CODE: EX-6s-STD-2s


PLACARDS - 4 SEAT EXECUTIVE AND 4 SEAT CORPORATE COMMUTER (Interior Code EX-4s-STD-4s)

The cabin placards. the 6 seat executive placards and the following placard is required for this interior.

INTERIOR CODE: EX-4s-STD-4S



Revision 1 1 : March 1,2003 '

I I

PILOT'S OPERATING t HANDBOOK

1 Emergency Procedure 1 C

SECTION 3 EMERGENCY PROCEDURES

SECTION 3

EMERGENCY PROCEDURES

CONTENTS

Paragraph Subject Page

3-1 EMERGENCYPROCEDURES

GENERAL

AIRSPEED FOR EMERGENCY OPERATIONS

REJECTED TAKEOFF

ENGINE FAILURE

ENGINE POWER LOSS IN FLIGHT ENGINE FAILURE BEFORE ROTATION ENGINE FAILURE AFTER ROTATION - LANDING GEAR DOWN ENGINE FAILURE AFTER ROTATION - LANDING GEAR UP ENGINE FAILURE IN FLIGHT

AIR START

AIR START ENVELOPE AIR START - WITH STARTER

ENGINE EMERGENCIES

OIL PRESSURE OIL TEMPERATURE OIL CONTAMINATION OIL QUANTITY FUEL CONTROL UNIT FAILURE (MOR OPERATION) PROPELLER - LOW PITCH PROPELLER - OVERSPEED PROPELLER - UNDERSPEED

Issued: Februaly 14,1994 Revision 10: September 1,2000

Report No: 01 973-001 3-i



FIRE, SMOKE OR FUMES

ENGINE FIRE - ON GROUND ENGINE FIRE - IN FLIGHT COCKPITICABIN FIRE, SMOKE, FUMES SMOKE EVACUATION

EMERGENCY DESCENT

GENERAL MAXIMUM RANGE DESCENT - AFTER ENGINE FAlL -SINGLE BATTERY MAXIMUM RANGE DESCENT - AFTER ENGINE FAlL - DUAL BATTERY MAXIMUM RATE DESCENT

EMERGENCY LANDING

GLIDE DISTANCE AND SPEED FORCED LANDING (ENGINE CUT-OFFIFEATHER) LANDING WlTH MAlN LANDING GEAR UNLOCKED LANDING WlTH NOSE LANDING GEAR UNLOCKED LANDING WlTH GEAR UP LANDING WITHOUT ELEVATOR CONTROL LANDING WlTH IMMOBILIZED HORIZONTAL STABILIZER LANDING WITHOUT FLAPS DITCHING

LANDING GEAR I HYDRAULIC SYSTEM FAILURE

LANDING GEAR FAILS TO RETRACT HYDRAULIC SYSTEM FAILURE EMERGENCY EXTENSION

FLAPS FAILURE

STICK PUSHER FAILURE

INADVERTENT PUSHERISHAKER OPERATION

PUSHER SHAKER

ELECTRICAL TRlM

TRlM RUNAWAY STABILIZER TRlM NO MAlN STABILIZER TRlM NO STABILIZER TRIM, MAlN OR ALTERNATE

Report No: 01973-001 3-ii

Issued: February 14, 1994 Revision 1 1 : March 1.2003



ELECTRICAL SYSTEM FAILURES

ESNTL BUS BAT HOT GEN 1 OFF GEN 2 OFF BAT OFF INVERTER BUS TIE .AV BUS N ESNTL BUS BATTERY THERMAL RUNAWAY

ENGINE INSTRUMENT SYSTEM (EIS) FAILURE

MSN 101-111 MSN 1 12-999

FUEL SYSTEM

LOW FUEL PRESSURE FUEL PUMP FAILURE AUTO FUEL BALANCING FAILURE SUSPECTEDFUELLEAK LOSS OF FUEL QUANTITY INDICATION

CABIN CONDITIONING FAILURES

CABIN PRESSURIZATION FAILURE TEMPERATURE CONTROL SYSTEM FAILURE ECS FAILURE

DElCE SYSTEM

PROPELLER DElCE FAILURE IN ICING CONDITIONS BOOT DElCE FAILURE IN ICING CONDITIONS INERTIAL SEPARATOR FAILURE LH WINDSHIELD DElCE FAILURE IN ICING CONDITIONS AOA PROBE DElCE FAILURE IN ICING CONDITIONS PlTOT AND STATIC PROBE DElCE FAILURE IN ICING CONDITIONS PUSHER ICE MODE FAILURE IN ICING CONDITIONS

MISCELLANEOUS

AIR GND PASSENGER AND CARGO DOOR ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS) FAILURE EFlS FAILURE PITOTISTATIC SYSTEM FAILURE


Report No: 01973-001 3-iii


Paragraph Subject

3.20.6 MALFUNCTIONS 3.20.7 WHEEL BRAKE FAILURE 3.20.8 CAWS MALFUNCTION

1 3.21 AUTOPILOT

1 3.21.1 CHECKLIST 3.21.2 AMPLIFIED MALFUNCTION PROCEDURES

Page 1

Report No: 01 973-001 I 3-iv

Issued: February 14, 1994 Revision 1 1 : March 1,2003 '


3.0 EMERGENCY PROCEDURES

3.1 GENERAL

The recommended action to be taken in case of failure or in emergency situations are contained in this section. Some situations require rapid action, leaving little time to consult the emergency procedures. Prbr knowledge of these procedures and a good understanding of the aircraft system is a prerequisite for safe aircraft handling.

KNOW YOUR AIRCRAFT AND BE THOROUGHLY FAMILIAR WlTH IMPORTANT EMERGENCY PROCEDURES.

The emergency procedures use the terms ' Land as soon as possible" and "Land as soon as practical". For the purpose of these procedures the meanings are as follows:

Land as soon as possible - Land without delay at the nearest airport where a safe approach and landing is reasonably assured

Land as soon as practical - Landing airport and duration of flight are at the discretion of the pilot. Extended flight beyond the nearest suitable airport is not recommended

A. WARNING SYSTEMS: .

The PC-12 is fitted with a comprehensive warning system.

The Central Advisory and Warning System (CAWS) contains:

RED WARNINGS which require immediate action.

AMBER CAUTIONS which advise that a system is not functioning or is an alert to a precautionary situation.

GREENBLUE ADVISORIES which indicate a system is functioning.

Whenever a CAWS red or amber caption illuminates, an aural gong will sound in the audio system and the MASTER WARNING or CAUTION will illuminate.

The EIS contains red warnings and amber caution lamps to advise of out of limit parameters on the engine display. With a steady red or amber comes a blinking LCD display.

B. AURAL WARNINGS

Aural warnings are provided to warn of an aircraft or flap overspeed, stall conditions and possible landing gear-up approach to landing. When the autopilot is engaged, aural warnings also indicate the approach to, or deviation from, selected altitude and1 or autopilot disconnects or system failures.

CAWS and EIS warnings and cautions will remain illuminated as long as the initiating condiiion exists. The MASTER WARNING and CAUTION lamps should be pressed to reset them for furlher failures once the failure is identified.


Report No: 01 973-001 3- 1

SECTION 3 EMERGENCYPROCEDURES

3.2 AIRSPEEDS FOR EMERGENCY OPERATIONS

All airspeeds shown are with airplane in clean configuration under ISA conditions.

A. Operating Maneuvering Speed (V,):

9039 Ib (41 00 kg) 151 KlAS

83801b (3800 kg) 145 KlAS

7940 Ib (3600 kg) 141 KlAS

7500 Ib (3400 kg) 137 KlAS

7060 Ib (3200 kg) 133 KlAS

6610 Ib (3000 kg) 129 KlAS

6170 Ib (2800 kg)

5730 Ib (2600 kg)

124 KlAS

120 KlAS

B. Best Glide (Propeller feathered):

9039 Ib (4100 kg) 110 KlAS

8380 Ib (3800 kg) 106 KlAS

7720 Ib (3500 kg)

7060 Ib (3200 kg)

6400 Ib (2900 kg)

102 KlAS

97 KlAS

93 KlAS

I C. Landing Approach Speeds with ice accretion on the airframe:

After failure of: Minimum Approach Speed:

Pneumatic Deice Boots (flap positio~ limit O") 130 KlAS

AOA Probe Deice andor

105 KlAS

I Pitot and Static Probe Deice 105 KlAS andor

Pusher Ice Mode (flap position limit 15') 105 KlAS

D. Balked Landing (Go Around)

After failure of:

Pneumatic Deice Boots (flap position limit OD)

(TOIPwr, flaps 0°, Pusher Ice Mode) 130 KlAS

Report No: 01973-001 3-2

Issued: June 10,1994 Revision 5: May 10, 1996


3.3 REJECTED TAKEOFF (Not englne related)

1. PCL Idle

2. Reverse As required

3. Braking As required

IF THE AIRCRAFT CANNOT BE STOPPED ON THE REMAINING RUNWAY:

4. PCL Idle

5. Condition lever Cut-off1Feather

6. FUEL EMERG SHUT OFF Press latch down (if installed) and pull lever up

7. GANGBAR electrical system Operate

AFTER THE AIRCRAFT HAS STOPPED - EVACUATE.

I CAUTION 1

A REJECTED TAKEOFF MAY CAUSE OVERHEATING OF WHEEL AND BRAKE ASSEMBLY COMPONENTS. THE MAIN WHEELS AND BRAKES SHOULD BE INSPECTED FOR DAMAGE IN ACCORDANCE WITH THE RESPECTIVE COMPONENT MANUALS BEFORE THE NEXT FLIGHT.

Issued: June 10,1994 Revision 6: Dec 6, 1996

Report No: 01 973-001 3-3


3.4 ENGINE FAILURE

I 3.4.1 ENGINE POWER LOSS IN FLIGHT

Indications: Englne power reduction.

No response to PCL movement.

PCL movement does not give required englne power.

1. PCL Full forward

2. Manual Override Lever Pull upwards and move slowly forward, to achieve required engine power

If engine falls below 50% Ng or accelerates very slowly with ITT at limit (870°CI 20 seconds) or compressor stall:

3. Staner On

IF STARTER SELECTED ON, AHRS AND EFlS MAY GO OFF-LINE.

4. Accelerate Ng > 80%

Refer to para 3.6.5 for remainder of WARNINGS and CAUTIONS.

3.4.2 ENGINE FAILURE BEFORE ROTATION

1. PCL Idle


IF RUNWAY OVERRUN OR COLLISION IS LIKELY, THEN:

3. Condition lever Cut-oWFeather




I 3.4.3 ENGINE FAILURE AFTER ROTATION - LANDING GEAR DOWN

1. If altitude is not suiiiient to Land straight ahead, turning only to

I select a runway or field. avoid obswctions

2. Flaps 40"



3. Airspeed

4. PCL

KIAS > 78

ldle

5. Condition lever Cutoflffeather

6. FUEL EMERG SHUT OFF Press latch down (i installed) and pull lever up

AFTER TOUCH DOWN:



3.4.4 ENGINE FAILURE AFER ROTATION - LANDING GEAR UP

1. Landing Gear Down

2. N ESNTL bus ovrd sw Override

3. Flaps

4. Airspeed KIAS > 95 Flaps 15" KIAS '> 78 Flaps 40"

5. PCL Idle

6. Condition lever Cut-offIFeather


AFTER TOUCH DOWN:


3.4.5 ENGINE FAILURE IN FLIGHT

1. Autopilot

2. PCL

3. Condition lever

Disengage if in use

ldle

Cut-off1Feather

4. Remaining fuel Check

5. Air start (refer to Sect. 3.5)

6. If above 13500 fl , (refer to Sect. 3.8) make an emergency descent

7. If engine air start is not (refer to Sect. 3.9) successful, make a forced landing


Report No: 01 973-001 3-5


3.5 AIR START

3.5.1 AIR START ENVELOPE

ALTITUDE - AIRSPEED ENVELOPE

AIRSPEED KIAS

Figure 3-1. Relight Envelope

Issued: Febnrary 14, 1994 Revision 2: Feb~ary 14, 1995


3.5.2 AIR START - WITH STARTER

DO NOT AlTEMPT MORE THAN ONE AIR START. REPEATED AIR START ATTEMPTS COULD DISCHARGE THE BATTERY TO A LEVEL THAT WOULD NOT BE ABLE TO SUPPORT ESSENTIAL ELECTRICAL SERVICES.

1. PCL Idle

2. Condition lever Cut-offJFeather

3. FUEL EMERG SHUT OFF Full in

4. + Electrical loads Reduce

5. + ECS OFF

6. GEN 1 and 2 OFF

7. BAT switch ON BAT 1 and BAT 2 switches ON (if 2"6 battery installed)

8. Air start envelope Check

9. Starter ON

10. Ignition ON

1 1. Condition lever (NG >I 3%) GROUND IDLE

12. ITT and Ng Monitor

WHEN ENGINE HAS RELIT Ng 260%:

13. Ignition AUTO

14. Condition lever FLIGHT IDLE

15. GEN 1 and 2 ON

16. Electrical Equipment As required

17. ECS AUTO

NOTE

Items marked thus: + may be omitted if time is short.

AIR START GIVES A TRANSIENT DC VOLTAGE DROP. CAUSING THE AHRS TO LOSE ITS REFERENCE, FORCING THE SYSTEM TO REALIGN, WHEN THE VOLTAGE IS RESTORED. DURING THIS PERIOD THE AIRCRAFT SHOULD BE FLYING STRAIGHT AND LEVEL.


Report No: 01973-001 3-7


3.6 ENGINE EMERGENCIES

3.6.1 OIL PRESSURE

Indlcatlons: EIS CauUon oll bllnklng 40/mln. (after 20 secs, €IS warning oll bllnklng 80hnln). 011 Px 60 to 90 PSI:

1. Ng Check above 72 %

2. Torque Reduce to below 24 PSI

3. Aircrafl Land as soon as practical.

Indlcatlons: €IS cauUon oll bllnks 40/mln andlor EIS warnlng oil bllnks 80lmln. 011 Px below 60 PSI or above 135 PSI

Land as soon as possiMe using minimum torque.

II possible always retain glide capability to the selected landing area in case 01 total engine failure.

1 3.6.2 OIL TEMPERATURE

I Indlcatlons: EIS cautlon oll bllnklng 4Wmln. or EIS wamlng oll bllnWng 801mln oll temp above 105'C.

1. PCL Reduce power.

IF OIL TEMPERATURE DOES NOT RETURN TO NORMAL:

2. Landing gear Extend

3. Land as soon as practical.

Report No: 01 973-001 3-8

Issued: February 14,1994 Revision 2: February 14. 1995


3.6.3 OIL CONTAMINATION CHlP

Indication: CAWS CHIP captlon

A. ONGROUND:

Before engine start:

DO NOT START ENGINE

After engine start or after landing.

RETURN TO PARKING AREA. '

SHUT DOWN ENGINE.

INSPECT CHlP DETECTOR(S) AND ENGINE, IF REQUIRED.

8. IN FLIGHT

CHECK AND MONITOR ENGINE PARAMETERS

REDUCE POWER TO MINIMUM REQUIRED FOR SAFE FLIGHT

LAND AS SOON AS PRACTICAL

After landing:

INSPECT CHlP DETECTOR@) AND ENGINE, IF REQUIRED.

3.6.4 OIL QUANTITY

Indication: CAWS OIL QTY caption (Only on ground, before engine start and 60 seconds after shutdown).

DO NOT START ENGINE. I

Refill engine with an approved oil.


Report No: 01973-001 3-9


3.6.5. FUEL CONTROL UNIT FAILURE (MOR OPERATION) t

Indlcatlons: Englne falls to below normal Idle (value depends on alrcraft speed and altitude).

PCL movement does not change Ng.

1. PCL Full forward.

2. MOR lever Pull upwards and move lotward, slowly, to achieve required engine power.

IF ENGINE ACCELERATES VERY SLOWLY WlTH ITT AT LIMIT (870°C/20 SEC) OR COMPRESSOR STALL:

3. Starter ON

IF STARTER IS SELECTED ON, AHRS AND EFlS MAY GO OFF LINE.

4. Accelerate Ngr 80%

I CAUTION I ENGINE CAN TAKE SEVERAL SECONDS TO REACT TO MOR MOVEMENT.

MAX. FUEL FLOW, WITH MOR, IS 500 LBS. FULL FORWARD MOVEMENT AT LOW ALTITUDE MAY NOT GIVE FULL ENGINE POWER (MIN. 30 PSI TORQUE). FULL FORWARD MOVEMENT MAY EXCEED MAX PERMllTED ENGINE TORQUE AND ITT. THE TORQUE LIMITER AND NG GOVERNOR ARE INOPERATIVE. RESPECT ENGINE LIMITS BY RESTRICTING MOR LEVER MOVEMENT.

Report No: 01 973-001 3-1 0

Issued: F e b ~ a r y 14,1994 Revision 2: February 14, 1995


ENGINE RESPONSE MAY BE MORE RAPID THAN WHEN USING THE PCL. ADDITIONAL CARE IS REQUIRED TO AVOID EXCEEDING ENGINE LIMITATIONS.

WHEN MOR IS IN OPERATION, DO NOT PERMIT NG TO FALL BELOW 65%. FROM BELOW THIS SPEED, ENGINE IT1 LIMIT MAY BE EXCEEDED.

5. In descent and until touch down at least 75% Ng.

6 . After landing, Condition lever Cut-offFeather

DO NOT USE MOR ON GROUND FOR TAXIING.

DO NOT MOVE PCL AFT OF IDLE DETENT.

SELECTING REVERSE MAY CAUSE PROP TO GO TO FULL REVERSE PITCH AND/OR OVERSPEED.

BETA IS OPERATIVE WITH PCL MOVEMENT IN THE 2 CM FORWARD POSITION OF IDLE DETENT.

TOTAL LANDING DISTANCE IS INCREASED BY A FACTOR OF 2.

Issued: June 10,1994 Revision 5: May 10,1996

Report No: 01973-001 3-1 1


3.6.6 PROPELLER - LOW PITCH PROP LOW P

Indlcatlons: CAWS Prop low P warnlng captlon. Posslble excessive propeller speed or helght loss.

1. PCL ENSURE FORWARD OF IDLE DETENT.

IF IT IS NOT POSSIBLE TO MAINTAIN SPEED AND HEIGHT:

2. Condition Lever Cut-off1Feather

Carry out emergency descent (Sect. 3.8) and landing (Sect. 3.9)

3.6.7 PROPELLER - OVERSPEED

Indlcatlons: EIS caution prop dlglts bllnklng 40lmin or EIS warnlng prop dlglts bllnklng 8Olmln. Propeller speed above 1760 RPM.

1. PCL REDUCE POWER.

2. AIRCRAFT SPEED REDUCE

II NP remains between 1760 and 1870 RPM:

Continue flight, at low speed, using minimum possible power.

IF NP IS ABOVE 1870 RPM:

1. PCL Reduce power (to idle if necessary)

2. Aircraft speed Reduce to 120 KlAS or below

3. Aircraft Land as soon as possible.

If possible always retain elide capability, to the selected landing airfield, in case of total propeller failure.

In the event of heavy vibration or uncontrolled speed runaway, be prepared to shut down engine.

4. Condition lever Cut-ofWFeather

5. Aircraft

Report No: 01973-001 3-12

Descent (Sect. 3.8) and landing (Sect. 3.9) procedures.



3.6.8 PROPELLER - UNDERSPEED

A. IN FLIGHT

Indications: EIS caution, propeller digits blink 40/mln Propeller speed below 1640 RPM

1. PCL Increase power

2. Aircrafl speed Increase

8. ONGROUND

Indlcatlons: EIS caution, prop diglts blink 40lmin. Prop speed below 950 RPM. or EIS warning, prop digits blink 80lmln. Prop speed below 950 RPM more than 10 seconds.

1. PCL


Retard PCL aft of Idle Detent or lncrease power until NP is above 950 RPM.

Report No: 01 973-001 3-13


FIRE, SMOKE OR FUMES

3.7.1 ENGINE FIRE - ON GROUND

Indlcatlons: CAWS warnlng ENG FIRE. Possible smoke andlor fumes.

1. PCL Idle

2. Condition lever Cut-off/Feather

3. ECS EMER shut off Pull

4. Fuel EMER shut off Press latch down (if installed) and pull lever up

5. Radio

6. Electrical Power

7. Parking brake

8. Aircraft

9. Fire

Emergency call

OFF (use GANGBAR)

OFF (if possible)

Evacuate

Extinguish

3.7.2 ENGINE FIRE - IN FLIGHT

Indlcatlons: CAWS warning ENG FIRE. Posslble smoke andlor fumes.

1. Engine Power Reduce to minimum acceptable according to flight situation.


3. Oxygen masks ON (all aircraft occupants)

Procedure to put on the crew oxygen masks:

a. Remove the normal headset. b. Put the oxygen mask on. c. Put the normal headset back on. d. Disconnect the normal headset boom microphone connector from the

MIC connector on the sidewall. e. Connect the oxygen mask microphone connector to the MIC connector

on the sidewall.

4. Passenger Oxygen selector ON

5. Confirm that tire exists:

6. Fuel EMER shut off Press latch down (it installed) and pull lever up

7. Condition lever Cut-off1Feather

Carry out emergency descent (Sect. 3.8) andlor emergency landing (Sect. 3.9) procedures.

Report No: 01 973-001 3-14



3.7.3 COCKPITICABIN FIRE, SMOKE OR FUMES

1. Oxygen masks ON (all occupants)




on the sidewall.

2. Crew Oxygen 100%

3. Passenger Oxygen ON

4. Descend to below 12,500' INITIATE

5. Smoke Evacuation Procedure Carry out if required

6. Fire Extinguisher Use if required

7. Nearest Airfield Proceed to

As soon as time permits

If source is known:

1. Associated electrical equipment

2. ECS

If smoke/fumes persists:

1. ECS EMER shut off

2. CABIN PRESS switch

3. BUS TIE CB (overhead panel)

4. GEN 2 TIE CB (overhead panel)

5. GEN 1 and GEN 2

6. BAlT 2 (if installed)

Off (circuit breakers)

As required

Pull

DUMP

Pull

Pull

OFF

OFF

I CAUTION I SWITCH OFF ANY UNNECESSARY ELECTRICAL EQUIPMENT WHEN ON BATTERY POWER ONLY.

If smokelfumes persists:

7. BATT 2 (if installed) ON

8. GEN 1 and GEN 2 ON


Report No: 01 973-001 3-15


9. BAlT 1 OFF

If smokelfumes persists:

10. Avionics Bus 1 and 2 OFF

11. DV Window

If VMC:

Verify open

12. All electrical power OFF

3.7.4 SMOKE EVACUATION

If smoke or fumes present, carry out cockpitlcabin firelsmokelfumes procedure (Sect. 3.7.3) then:

1. Oxygen masks ON (all occupants)



MIC connector on the sidewall. a. Connect the oxygen mask microphone connector to the MIC connector

on the sidewall.

2. Crew Oxygen 100%

3. Passenger oxygen . ON


5. ECS pressurization switch DUMP

When cabin pressure differential is zero:

6. DV window open

7. Fans ON

8. Aircraft Descent to cabin altitude of 12,500' or below.

If satisfied that ECS Set ECS control switch to OFF and is not source of contamination then to AUTO. Set ECS pressurization

switch to AUTO.

Report No: 01 973-001 3-16



3.8 EMERGENCY DESCENT

3.8.1 GENERAL

The type of emergency descent will depend on the kind of failure and Ule aircraft situation.

Two types of descent are considered:

1. Engine failure, aircraft flown for maximum range.

2. Engine running, maximum descent rate.

The factors to be considered are:

I) Cabin altitude and oxygen duration.

ii) Electrical power endurance.

ili) Distance to suitable landing area.

iv) Flight conditions IMC, VMC, ICING.

v) Minimum safe altitude.

vi). Fuel reserves.

The pilot must consider the situation and priorities and adjust his actions accordingly.

Issued: February 14,1994 Revision 2: Feb~ary 14,1995

Report No: 01 973-001 3-1 7

SECTION 3 SPILATUSW EMERGENCY PROCEDURES PC12

I 3.8.2 MAXIMUM RANGE DESCENT - AFTER ENGINE FAlL - SINGLE BATTERY

1. PCL

2. Condition lever

Idle

Cut-off1Feather (to feather propeller)

3. Aircraft configuration Landing gear up and flaps to 0".

I CAUTION I IF LANDING GEAR AND OR FLAPS ARE EXTENDED GLIDE RANGE WlLL BE SEVERELY REDUCED. RETRACTING LANDING GEAR AND FLAPS WlLL REDUCE BATTERY ENDURANCE SIGNIFICANTLY AND MAY PREJUDICE SUBSEQUENT FLAPS LOWERING. AHRS AND EIS MAY FAlL DURING LANDING GEAWFLAP OPERATION.

4. Speed

I 5. Oxygen masks

110 KIAS for 9039 Ib (4100 kg) (for best glide speed see para 3.2.8) (In icing conditions 130 KIAS)

Prepare. Put on before cabin altitude exceeds 12,500 11.

I Procedure to put on the crew oxygen masks: l a. Remove the normal headset.

b. Put the oxygen mask on. c. Put the normal headset back on. d. Disconnect the normal headset boom microphone connector from the


on the sidewall. I

6. Passenger Oxygen Auto. Check contents. Calculate Oxygen duration and check flow to PAX.

I 7. Electrical load Monitor battery amps.

Electrical loads can be further reduced by selectively switching off the following:

a. External lights All off or as required b. De-ice systems In icing conditions, all off except

probes c. Internal lights All off, if flying at night - instrument

panel and cockpit flood lights as required

d. NavICom 2 Off e. ADF Off 1. Transponder OH if radar communication is not

required

Report No: 01973-001 3-18



g. Wx Radar Off, if in IMC - as required

If time permits, pulling the following CB's will assist to reduce electrical load:

RMl 1 Avionics 1 bus AHRS 2 (if installed) Avionics 1 and Avionics 2 bus EFlS 2 EHSl (if installed) Avionics 1 bus RMI 2 (if installed) Avionics 2 bus EFlS 2 EADl (if installed) Avionics 2 bus COPILOT PlTOT DE ICE Generator 1 bus

NOTE

During extended glide period engine low oil quantity warning may appear - disregard for air start.

Engine restart soon as possible (if applicable) refer to Sect 3.5 -perform only one start attempt.

If engine restart was not successful or not applicable:

8. Rate of descent

Below 10,000 ft:

9. Windshield heat

Adjust to achieve cabin altitude of 12,500 ff before Oxygen supply exhausted.

When required, to HEAVY (uses less current than LIGHT)

10. ECS EMER shut off Pull (cabin ventilation)

For forced landing (Sect. 3.9.2).

Issued: June 10,1994 I Revision 1 1 : March 1, 2003

Report No: 01973-001 3-19

SECTION 3 =PILATUSB EMERGENCY PROCEDURES PC12

3.8.3 MAXIMUM RANGE DESCENT - AFTER ENGINE FAlL - DUAL BATTERY

1. PCL Idle

2. Condition lever Cut-olllFeather (to feather propeller)

3. Aircraft configuration Landing gear up and flaps to 0".

[ CAUTION 1 IF LANDING GEAR AND OR FLAPS ARE EXTENDED GLIDE RANGE WlLL BE SEVERELY REDUCED. RETRACTING LANDING GEAR AND FLAPS WlLL REDUCE BAllERY ENDURANCE SIGNIFICANTLY AND MAY PREJUDICE SUBSEQUENT FLAPS LOWERING. AHRS AND EIS MAY FAlL DURING LANDING GEAWFLAP OPERATION.

4. Speed

5. Autopilot

6. Oxygen masks

1 10 KIAS for 9039 Ib (4100 kg) (for best glide speed see para 3.2.8) (In icing conditions 130 KIAS)

Engage NAV or HDG and IAS hold mode

Prepare. Put on before cabin altitude exceeds 12,500 11.




on the sidewall.


8. Electrical load Monitor battery amps and reduce load as follows:

a. External lights All off or as required b. De-ice systems In icing conditions, all oll- except

probes c. COPILOT PlTOT DE ICE Pull

circuit breaker (Generator 1 bus bar)

Report No: 01973-001 3-20



MONITOR BAT 1 AND BAT 2 AMPS. IF ONE INDICATION IS POSITIVE. SWITCH OFF AFFECTED BATTERY. AFTER 5 MlNS BATTERY CAN BE SWITCHED ON AGAIN. IF INDICATION STILL POSITIVE SWITCH BATTERY OFF.

NOTE

During extended glide period engine low oil quantity warning may appear - disregard for air start.

Engine restart as soon as possible (if applicable) refer to Sect 3.5 -perform only one start attempt

If engine restart was not successful or not applicable:

9. Rate of descent

Below 10,000 R:

10. Windshield heat

Adjust to achieve cabin altitude of 12,500 ft before Oxygen supply exhausted.

When required, to HEAVY (uses less current than LIGHT)

11. ECS EMER shut off Pull (cabin ventilation)

For forced landing (Sect. 3.9.2).

Issued: June 10,1994 Revision 11: March 1,2003

Report No: 01 973-001 3-21


I 3.8.4 MAXIMUM RATE DESCENT I

1. PCL IDLE

2. Landing gear Below 177 KIAS, down

3. Aircrafi speed M ~ d v ~ ~

4. Oxygen masks Prepare. Put on before cabin altitude exceeds 12,500 ft.



MIC connector on the sidewall. e. Connect the oxygen mask microphone connector to the MIC

connector on the sidewall.


1 CAUTION 1

IN TURBULENCE REDUCE SPEED TO 170 KIAS

6. Windshield heat As required

Report No: 01973-001 3-22



EMERGENCY DESCENT PROFILES

Issued: June 10.1994 h Revision 1 1 : March 1,2003

Report No: 01973-001 3-23 1

SECTION 3 .PPILATUSB EMERGENCYPROCEDURES PC12

3.9 EMERGENCY LANDING

3.9.1 GLIDE DISTANCE AND SPEED

Configuration:

1. Landing gear

2. Flaps

3. Condition lever

4. Best glide speed 1 10 KlAS for 9039 b (41 00 kg) (for best glide speed see para 3.2.8)

NOTE

A fully loaded battery will last as follows:

60 Amps for 20 minutes.


NOTE

Two fully loaded batteries (if installed) will last as follows:

- 60 Amps for 40 minutes.


Report No: 01 973-001 1 3-24 Issued: June 10,1994

Revision 1 1 : March 1,2003


3.9.2 FORCED LANDING (ENGINE CUT-OFFIFEATHER)

1. PCL

2. Condition lever

3. FUEL EMERG SHUT OFF

4. Electrical load

5. ECS Switch

6. Best glide speed

7. Seat backs

8. Seat belts

9. Passengers

IF LANDING SlTE ALLOWS:

Idle

Cut-offIFeather

Pull

Reduce

Dump

110 KlAS for 9039 Ib (4100 kg) (for best glide speed see para 3.2.8)

Upright

Fastened. Tighten lap straps

Brief. Instruct to sit upright

10. Landing gear DOWN

Use hand pump until 3 greens appear.

IF LANDING SlTE NOT SUITABLE FOR GEAR DOWN LANDING:

1 1. Keep landing gear UP

12. Flaps 40'

NOTE

To silence the flap aural warning pull the AUDIO circuit breaker (Battery busbar).

13. Final approach speed 78 KlAS

AFTER TOUCH DOWN:


15. Evacuate aircrafl after stop


Report No: 01 973-00 1 3-25 I


3.9.3 LANDING WlTH MAlN LANDING GEAR UNLOCKED

1. Confirm landing gear position by control tower or other aircraft.

I CAUTION 1

IF ONE MAIN LANDING GEAR IS NOT DOWN, IT IS RECOMMENDED TO LAND WlTH GEAR UP.

IF FAILED GEAR IS DOWN BUT NOT LOCKED:

2. Fuel weight Reduce

3. Passengers Brief

4. Flaps 40'


6. Touchdown Gently

7. Lower nose wheel immediately to maintain lateral control.

8. Use full aileron during roll-out. -- To lift the wing with the failed landing gear.

9. PCL Idle

10. Condition lever Cut-offJFeather


12. Evacuate aircraft after stop

Do not taxi the aircraft before deficiency is rectified.

Report No: 01973-001 1 3-26 Issued: June 10, 1994



3.9.4 LANDING WITH NOSE LANDING GEAR UNLOCKED

1. Passengers Brief

2. Flaps 40"


4. Land on main wheels, keep nose high.

5. Condition lever Cut-offlFeather


7. Lower nose wheel slowly

8. Avoid braking

9. Evacuate aircraft after stop.

3.9.5 LANDING WITH GEAR UP

1. Passengers

2. Approach

3. Flaps

4. Final approach speed

5. ECS Switch

WHEN RUNWAY IS ASSURED:

6. PCL

7. Condition lever

8. FUEL EMERG SHUT OFF

9. Flare out

AFTER TOUCH DOWN:

10. GANGBAR electrical system

11. Evacuate aircraft after stop

Issued: June 10,1994 I Revision 1 1 : March 1,2003

Brief

Standard

40'

78 KlAS

Dump

Idle

Cut-off1Feather

Pull

Operate

Report No: 01973-001 3-27 1

SECTION 3 =PILATUSW EMERGENCYPROCEDURES P(P

1 3.9.6 LANDING WITHOUT ELEVATOR CONTROL

I 1. Passengers Brief

2. Landing gear

3. Flaps

Down

40'


5. Set power as necessary to maintain speed and 300 to 500 ft rate of descent.

6. Use stab trim to adjust pitch.

WHEN CLOSING TO GROUND:

7. Reduce Rate of Descent by increasing pitch and/or power.

8. Reduce power progressively.

STALLS ARE NOT PROTECTED WlTH THE STICK PUSHER INOPERATIVE.

STALLS MUST BE AVOIDED WHEN THE STICK PUSHER IS INOPERATIVE. EXCESSIVE WING DROP AND ALTITUDE LOSS MAY RESULT DURING STALL WlTH FLAPS DOWN AND/OR WHEN POWER IS APPLIED.

3.9.7 LANDING WlTH IMMOBILIZED HORIZONTAL STABILIZER

I AS

FLAPS

LAND

I

Report No: 01973-001 1 3-28

Fly at IAS which reduces elevator forces to minimum.

At a safe altitude select flap required for landing.

Land using normal procedures holding elevator forces.

Issued: June 10,1994 Revision 11 : March 1,2003


3.9.8 LANDING WITHOUT FLAPS

1. Proceed as for normal approach

2. Landing gear Down


4. Landing Normal


6. Reverse As required

Landing distance will increase by 80%

3.9.9 DITCHING

1. Landing gear UP

HEAVY SWELL WITH LIGHT WIND, DITCH PARALLEL TO THE SWELL. STRONG WIND, DITCH INTO THE WIND.

2. Passengers

3. Flaps

Brief

40"


5. ECS switch Dump

6. Ditch with a low rate of descent.


8. Evacuate through the ovewing emergency exit only.


Report No: 01 973-001 3-29 1


3.10 LANDING GEARIHYDRAULIC SYSTEM FAILURE

3.10.1 LANDING GEAR FAILS TO RETRACT

Indications: All Landlng Gear indicator Llghts do not extingulsh.

1. If CAWS N ESNTL Caution is ON

Refer to electrical failure (Sect. 3.15)

2. If CAWS HYDR Caution is ON

Refer to emergency gear lowering (Sect. 3.10)

3.10.2 HYDRAULIC SYSTEM FAILURE

A. ON GROUND:

1. CAWS HYDR caution Continuously illuminated - requires maintenance.

6. IN FLIGHT:

CAWS HYDR caution continuously illuminated:

1. HYDR CTL circuit breaker Pull (Non Essential busbar)

2. Airspeed < 177 KIAS

After 200 minutes, landing gear may start to extend.

3. Before landing refer to emergency extension (Sect 3.10.3)




I 3.10.3 EMERGENCY EXTENSION

Indlcatlons: Incorrect lndicatlon on landing gear lndlcator lights. Red unlocked llghts on andlor green llghts not Illumlnated.

1. Airspeed 110 KlAS

2. Selector Down

IF 3 GREEN LIGHTS NOT ILLUMINATED WITHIN 30 SEC: 1 3. Hand pump Activate

Use full strokes, pump until 3 green lights are illuminated. t Complete lowering takes about 80 strokes.

IF 3 GREEN LIGHTS STILL NOT ILLUMINATED:

4. Yaw the aircraft left and right to lock the main landing gear.

5. Airspeed reduce to minimum safe to improve nose gear locking.

I IF 3 GREEN LIGHTS ILLUMINATED:

6. Land

If not successful refer to emergency landing (Sect. 3.9.5)

Issued: June 10, 1994 Revision 1 1 : March 1, 2003

Report No: 01 973-001 3-31 1


3-1 1 FLAPS FAILURE

A. FAILURE ON GROUND

Indlcatlon: CAWS FLAPS cautlon lllumlnated -means flap system falled

FLAP circuit breaker (Battery busbar)

Check. If tripped, wait 5 minutes, reset (ma. 2 attempts) and continue normal operation if caption goes off.

If unsuccessful: no flight permitted, maintenance action required.

B. FAILURE IN FLIGHT

Indlcatlon: CAWS FLAPS cautlon lllumlnated - means flap system falled

FLAP circuit breaker (Battery busbar)

Check. If tripped, wait 5 minutes, reset (rnax. 2 attempts) and continue normal operation.

If unsuccessful, land with flaps at the failed position.

C. FAILURE IN FLIGHT - WITH POTENTIAL FLAP ASYMMETRY

Indicatlon: CAWS FLAPS cautlon lllumlnated shortly followed by PUSHER cautlon lllumlnating - means flap system failed and stick pusher set to 'safe' mode

NOTE

In this mode the stick pusher remains operative but will push at a higher airspeed, (approximately 5 knots).

Airspeed Reduce to below 120 KlAS for indicated flaps position 30° or greater.

Land as soon as practical - with flaps at the failed position. Approach using EADI Angle of Attack display at the fast diamond. Landing distance will increase.

Report No: 01973-001 3-32



3.12 STICK PUSHER FAILURE

lndlcatlon: CAWS PUSHER cautlon lllumlnated

A. ONGROUND:

1. Pusher Test Carry out

If pusher caution persists:

NO FLIGHT PERMITTED, REQUIRES MAINTENANCE.

B. IN FLIGHT:

1. Pusher Test Carry out

If Shaker 1 and 2 active No further action required. and PUSHER caution extinguished

If Shaker 1 or 2 not active The aircraft is not stall protected and PUSHER caution persist indicates

2. Airspeed not below 1.3 Vs Flaps 0" 110 KlAS Flaps 15' 95 KlAS Flaps 40" 78 KlAS

STALLS MUST BE AVOIDED WHEN THE STICK PUSHER IS INOPERATIVE. EXCESSIVE WING DROP AND ALTITUDE LOSS MAY RESULT DURING STALL WITH FLAPS DOWN AND/OR WHEN POWER IS APPLIED.

STALL SPEEDS IN TURNS ARE HIGHER.

3.13 INADVERTENT PUSHERISHAKER OPERATION

3.13.1 PUSHER

Indlcatlon: non-commanded pusher operation, rapld nose pltch-down motlon.

NOTE

Control wheel force to stop pusher operation Is 60 to 65 pounds.

1. Control wheel HOLD against pusher action

2. PUSHER INTR switch PRESS and HOLD


Report No: 01 973-001 3-33 1

SECTION 3 =PILATUSW EMERGENCY PROCEDURES PC12

3. PUSHER SYS circuit breaker Pull (Battery busbar)

NATURAL STALLS ARE NOT PREVENTED WlTH THE STICK PUSHER INOPERATIVE.

STALLS MUST BE AVOIDED WHEN THE STICK PUSHER IS INOPERATIVE. EXCESSIVE WlNG DROP AND ALTITUDE LOSS MAY RESULT DURING STALL WlTH FLAPS DOWN AND/OR WHEN POWER IS APPLIED.

3.13.2 SHAKER

Indlcatlon: non-commanded shaker operation (automatic autopilot disconnect).

1. AOA Decrease

2. IAS Increase

If shaker continues to operate:

3. STALL WARN UNIT 1 circuit Pull breaker (Battery busbar)

4. STALL WARN UNIT 2 circuit Pull breaker (Generator 1 busbar)

APPROACHES TO STALLS ARE NOT WARNED AND NATURAL STALLS ARE NOT PREVENTED WlTH THE STICK SHAKER INOPERATIVE.

STALLS MUST BE AVOIDED WHEN THE STICK PUSHER IS INOPERATIVE. EXCESSIVE WlNG DROP AND ALTITUDE LOSS MAY RESULT DURING STALL WlTH FLAPS DOWN AND/OR WHEN POWER IS APPLIED.

5. Refer Stick Pusher failure (Sect. 3.12)

I Report No: 01973-001 1 3-34



3.14 ELECTRICAL TRlM

3.14.1. TRlM RUNAWAY

Indlcatlon: non-commanded trlm operation, rapidly lncreaslng out of trlm forces.

1. TRIM INTERRUPT switch INTR

2. CB of failed trim Pull

STAB TRlM and AIL TRlM CBs on BAT Bus STAB ALTERN and RUD TRlM CBs on GEN 1 Bus

3. TRIM INTERRUPT switch NORM

NOTE

Reduce speed if control forces are high.

If main stabilizer trim has failed:

4. Pitch trim use ALTERNATE STAB TRIM.

NOTE

The autopilot will disconnect when TRlM INTERRUPT is operated. If the ALTERNATE STAB TRlM has failed the N P is not operative.

3.14.2. STABILIZER TRlM

STAB TRlM CAWS warning 1 ON ground only.

Stabilizer trim Set for take-off.

3.14.3. NO MAIN STABILIZER TRlM

TRIM INTRRUPT Check at NORM

ALTERNATE STAB TRIM Operate as reqd

3.14.4. NO STABILIZER TRIM, MAIN OR ALTERNATE

Carry out

LANDING WITH IMMOBILIZED HORIZONTAL STABILIZER procedure (refer to Sect. 3.9.7)


Reporl No: 01 973-001 3-35 1


3.15 ELECTRICAL SYSTEM FAILURES

3.15.1 ESNTL BUS CAWS WARNING

Indicates that the BAT, GEN 1 or GEN 2 BUS volts is below 18 V.

1. OVERHEAD PANEL Check (Determine which BUS is off- line.)

Continue flight without service of failed BUS.

A. Battery Bus Fail: Battery 1 Bus Fail: Do not fly in icing conditions (If second banely installed)

1. INVERTER switch GEN

2. EFlS CMPST switch Operate if required

NOTE

Operation of the EFlS CMPST switch with the autopilot engaged, will cause the autopilot to disengage.

3. Pitch Trim Use alternate

4 Aircraft Descend below 12,500'

Land as soon as possible.

I CAUTION I THE FOLLOWING SYSTEMS ARE INOPERATIVE:

Engine StarVrelight Stick pusher Propeller feather Landing gear Indications Fire detection Low airspeed warning (MSN 101-180) Fuel quantity sensing Standby attitude (operative with EPS option) Fuel pump LH Airlground system

a Int + ext lights (partial) ECS Main pitch trim Aileron trim AOA plate heaters

Report No: 01 973-001 1 3-36 Issued: June 10, 1994



8. GEN 1 BUS FAlL

1. INVERTER switch BAT

2. EFlS CMPST switch Operate if required

3. Aircraft Do not fly in icing conditions

I CAUTION I

THE FOLLOWING SERVICES ARE INOPERATIVE:

Auto pilot (pitch modes) Auto fuel pump and fuel quantity measuring

system Alternate pitch trim RH fuel pump Rudder trim Oxygen contents Airlground system Int + Ext lights (partial) Inert sep AOA plate heaters

C. GEN 2 BUS FAlL

1. Aircraft Do not fly in icing conditions

I CAUTION

THE FOLLOWING SERVICES ARE INOPERATIVE:

Wing deice LH Windshield heat Taxi light Wing Inspection light Cabin electrical heater AOA vane heaters Pitotlstatic heating system


Report No: 01973-001 3-37 1

SECTION 3 I P I L A T U S S EMERGENCY PROCEDURES PC12

3.15.2 BAT HOT CAWS CAUTION

Indicates that the battery temperature is excessive.

1. Overhead Panel Identify the battery overtemp (If second battery installed)

2. Battery charge current Check

If charge current high:

3. Battery OFF (Do not select ON again)

If charge current normal:

4. Battery OFF

If Battery hot caption extinguishes, wait 5 minutes then:

5. Battery ON (Once only)

NOTE

The BAT HOT caution is inoperative on aircraft with a lead acid battery or batteries installed.

Issued: June 10,1Q94 Revision 1 1 : March 1, 2003

=PlLATUSI SECTION 3 PC12 EMERGENCY PROCEDURES

3.15.3 GEN 1 OFF CAWS CAUTION

Indicates that GEN 1 is OFF line.

NOTE

NON ESNTLB caution will also illuminate. GEN 2 TIE ON indicator will also illuminate on the overhead panel.

1. GEN 1 Master switch Confirm ON

2. GEN 1 reset Press

3. GEN 1 Master switch ON

If GEN 1 does not reset:

4. VIA Meter Select to GEN 2 (MSN 101-1 11)

5. Electrical load

6. EIS

Reduce until battery current is positive (not discharging) I

I

Monitor DC AMPS BAT and GEN 2

Services on the NON ESSENTIAL BUS are automatically switched OFF.

7. Before landing refer to landing gear emergency lowering (Sect 3.10)

THE FOLLOWING SERVICES ARE INOPERATIVE.

Hydraulic Control (Landing Gear) Vent Fans Flood Fans Cabin Reading Lights Cabin Flood Light 100% Landing Light (RH) Auxiliary Heating System Cooling System Cabin Services RH Windshield Heat AOA Heater Plates

Issued: June 10, 1994 Revision 11: March 1,2003

Report No: 01 973-001 3-39


3.15.4 GEN 2 OFF CAWS CAUTION

1. GEN 2 master switch Confirm ON

2. GEN 2 reset Press (Pre SB 24-010)

3. GEN 2 Master switch ON

If GEN 2 does not RESET:

4. Overhead Panel Confirm GEN 2 BUS failure indicator extinguished and GEN 2 TIE ON indicator illuminated. (GEN 1 supplies GEN 2 BUS).

5. EIS Check DC AMPS BAT positive/monitor

6. Continue flight - Report on landing.

I CAUTION 1 THE FOLLOWING SERVICES ARE INOPERATIVE.

AOA Heater Plates (MSN 101-260)

3.15.5 BAT OFF CAWS CAUTION

1. Overhead Panel Identify BAT 1 or Bat 2 loss (If second battery installed)

2. BAT master switch Confirm ON

3. Overhead Panel Confirm BAT BUS failure indicator extinguished. (GEN 1 supplies BAT BUS).

NO ENGINE RELIGHT POSSIBLE

4. EIS Monitor DC VOLTS, DC AMPS


3.15.6 INVERTER CAWS CAUTION and/or RMI flag

1. Overhead panel Select alternate inverter


Report No: 01973-001 1 3-40



3.15.7 BUS TIE CAWS CAUTION

NOTE

NON ESNTL bus indicator will also illuminate on the overhead panel ii AUTO Is selected.

1. VOLTS l AMPS

Overhead panel

Check BAT, GEN 1 (MSN 101-1 11). GEN 2 for normal chargeldischarge.

Check BAT1 and BAT 2 indicators (If second battery Installed)

NOTE

GEN 1 bus and AVIONIC 2 bus supplied from GEN 1.

IF NORMAL AMPS INDICATED AND NO OTHER ELECTRICAL CAUTIONS OR WARNING ILLUMINATED:

2. BUS TIE circuit breaker Reset (one attempt only) (overhead panel)

IF NO RESET:

Continue flight using remaining electrical services.

3. GEN2 Reduce load to below 130 amps/monitor

3.15.8 AV BUS CAWS WARNING

Indicates Avionics Bus I or 2 voltage below 18 V

1. Overhead Panel ldentify failed AVlONlC BUS

2. AVlONlCS MASTER Confirm ON

3. EFlS CMPST CMPST (if required)

NOTE


4. Continue flight without the services of the failed line.


Report No: 01973-001 3-41 1


3.15.9 N ESNTL BUS CAWS CAUTION

Normally indicates automatic load shedding after GEN 1 failure.

A. If GEN 1 off, warning is illuminated - refer to GEN 1 OFF (Sect. 3.15.3)

0. If GEN 1 (MSN 101-1 11) BAT (MSN 112-999) Charge current is positive:

1. AUTOIOVRD ON switch OVRD ON

2. N ESNTL B Caution OFF

3. Report failure after landing.

A FULLY CHARGED BAlTERY WlLL LAST FOR 20 MINUTES IF THE LOAD IS REDUCED BELOW 60 AMPS OR FOR 30 MINUTES IF THE LOAD IS REDUCED BELOW 50 AMPS.

TWO FULLY CHARGED BATTERIES (IF INSTALLED) WlLL LAST FOR 40 MINUTES IF THE LOAD IS REDUCED BELOW 60 AMPS OR FOR 60 MINUTES IF THE LOAD IS REDUCED BELOW 50 AMPS.

FROM MSN 112

3.15.10 BATTERY THERMAL RUNAWAY EIS WARNING BAT DC A

Indicates excessive increase of battery charging rate.

1. Overhead panel

2. Battery

Wait 5 minutes then:

3. Battery

Report No: 01 973-001 1 3-42

Identify the battery ovenemp (If second battery installed)

OFF

ON (once only)

Issued: June 10,1994 Revision 11 : March 1,2003 '


3.16 ENGINE INSTRUMENT SYSTEM (EIS) FAILURE

3.16.1 MSN 101-111

Indlcatlon: Torque, ITT, 011 Pressure or Propeller RPM suspect or frozen.

1. EIS ACQ 1 circuit breaker Pull (Battery busbar)

H problem still exists:

2. EIS ACQ 1 circuit breaker Reset (Battery busbar)

3. EIS ACQ 2 circuit breaker Pull (Generator 1 busbar)

If problem still exists:

4. EIS ACQ 2 circuit breaker Reset (Generator 1 busbar)

If ITT or NG indication inoperative, set toque as required - not to exceed value in the appropriate Torque Chart in the Performance Section.

NOTE

System failures within the EIS system are not indicated while the engine is running.

PrelPost flight inspection:

Indication: EIS caution light blinking.

5. EIS System Test (Test Panel)


Press Note failure code

Report No: 01973-001 3-43 1

SECTION 3 'SPILATUSV EMERGENCY PROCEDURES PC12 i

I

3.16.2 MSN 112 AND UP

Indication: Torque or ITT suspect, frozen or blank

1. €IS MENU button Press

Digital readouts of ITT and Torque will replace the normal OAT and ENDUR readouts on the EIS Display.

Indication: Any Indication other than Torque or ITT frozen or blank.

No action possible. Land as soon as practical.

If ITT or Ng indication inoperative, set torque as required - not to exceed value in the appropriate Torque Chart in the Performance Section.

Pre/Post flight inspection:

Indication: EIS caution light blinking or any indication blank.

2. EIS System Test (Test Panel)

Report No: 01973-001 1 3-44

Press Note failure code

Issued: June 10.1994 Revision 1 1 : March 1,2003 4

.(


PARAMETER OUT OF RANGE ACTION:

RED

Torque

ITT

Gas GEN RPM (NG)

Oil Temperature

Oil Pressure

Propeller RPM (NP)

OC Volts


1

ACTION:

Reduce power below 44 PSI

Reduce power below 800°C

Reduce power below 104% Ng

See Sect. 3.6 - Engine emergencies


See Sect. 3.6, Engine emergencies

Check electrical system status

AMBER

Torque

ITT

Gas GEN RPM (NG)

Oil Temperature

Oil Pressure

Propeller RPM (NP)

DC Volts

DC AMPS

OAT

Report No: 01 973-001 3-45 1

ACTION:

Reduce power below 44 PSI

Reduce power 800°C

Reduce power 104% Ng



See Sect. 3.6, Engine emergencies

Check electrical system status

Check 1. VIA meter switched to battery (MSN 101-111).

Check 2. Electrical system status. Reduce load below 60 A (20 min) Reduce bad below 50 A (30 min)

Switch ON "Probes Deice*


3.17 FUEL SYSTEM

3.17.1 LOW FUEL PRESSURE

Indication: FUEL PRESS CAWS CAUTION

1. Power Reduce to minimum to sustain flight

2. Fuel pumps ON

If FUEL PRESS caption remains ON:

3. Aircraft Land as soon as possible

Retain glide capability to landing area if possible.

NOTE

Fuel low pressure will normally cause the fuel pumps to come on automatically.

In this case the Indication are both FUEL PUMPS running continuously,cycling OFF/ON every 10-15 secs.

4. Fuel Pumps ON

L FUEL PUMP

5. Aircraft Descent to warmer air: (A possible cause is the fuel filter blocked with Ice crystals).

3.17.2 FUEL PUMP FAILURE

R FUEL PUMP

1. Fuel pump(s) AUTO

2. FUEL CONT circuit breaker Reset (Generator 1 busbar)

3. Fuel pump(s) ON

IF NOT SUCCESSFUL CONTINUE FLIGHT WITHOUT ELECTRIC FUEL PUMP@).

Report No: 01973-001 1 3-46



3.17.3 AUTO FUEL BALANCING FAILURE

Indlcatlon: EIS analogue fuel gauges Indicate 3 segments or more difference between left and right without automatic actlvatlon. Possibly aileron defleotion requlred for wlngs level flight, especially at low speed.

1. Fuel Pump (fuller side) ON

2. Fuel state Monitor

3. When fuel balanced Fuel Pump AUTO

3.17.4 SUSPECTED FUEL LEAK

1. Fuel pump (on leaking side) ON

2. FUEL PUMP CB (on good side) Pull (Battery or Generator 1 busbar)

3. Fuel state Monitor

Land as soon as practical.

3.17.5 LOSS OF FUEL QUANTITY INDICATION

lndicatlon: EIS display goes blank.

The automatic fuel balancing system will be inoperative.

Land as soon as practical.


Report No: 01 973-001 3-47 (


3.18 CABIN CONDITIONING FAILURES

3.18.1 CABIN PRESSURIZATION FAILURE

Indlcatlon: CAB PRESS CAWS WARNING

1. Pressure Indication: Check

A. If CAB P > 5.75 PSI:

1. Cabin Altitude Selector Check correct setting

2. Manual control valve Operate to reduce pressure diferential to required level

IF UNSUCCESSFUL:

3. ECS switch OFF

4. ECS EMERG shut off Pull

5. Oxygen masks ON


a. Remove the normal headset. b. Put the oxygen mask on. c. Put the normal headset back on. d. Disconnect the normal headset boom microphone connector from

the MIC connector on the sidewall. e. Connect the oxygen mask microphone connector to the MIC


6. PASSENGER OXYGEN valve AUTOION

7. Passengers INSTRUCT to don masks

8. Emergency descent Refer to Sect. 3.8.

B. If cab. alt. > 10700 H:

1. Oxygen masks ON


a. Remove the normal headset. b. Put the oxygen mask on. c. Put the normal headset back on. d. Disconnect the normal headset boom microphone connector from

the MIC connector on the sidewall. e. Connect the oxygen mask microphone connector to the MIC




4. ECS switch Check AUTO

Report No: 01973-001 3-48



5. ECS pressurization switch Check AUTO

6. Manual control valve Confirm fully closed.

IF UNSUCCESSFUL: I

7. Limit flight altitude to maintain cabin altitude < 12500 ft.

8. If necessary carry out emergency descent.

3.18.2 TEMPERATURE CONTROL SYSTEM FAILURE

1. ECS switch MAN

2. TEMPERATURE MAN switch HOT or COLD as required

3. TEMPERATURE MAN switch Control should be adjusted following each change in engine power, to prevent air temperature limits being exceeded.

3.18.3 ECS FAILURE I Indlcatlon: CAWS ECS cautlon

Reset the ECS:

1. ECS switch OFF

2. ECS swiich AUTO

IF UNSUCCESSFUL:

3. ECSSWITCH OFF

4. ECS EMERG shut off Pull

If aircraft alt. > 10700 ft:

5. Oxygen masks ON




on the sidewall.



8. Emergency descent Refer to Sect. 3.8.


Report No: 01 973-001 3-49


3.19 DEICE SYSTEM

3.19.1 PROPELLER DEICE FAILURE IN ICING CONDITIONS

Indication: DE ICE CAWS CAUTION with flashing PROP DE ICE CAWS ADVISORY.

11 WARNING ((

THE LOSS OF PROPELLER DEICE IN ICING CONDITIONS CAN CAUSE SEVERE DEGRADATION IN AIRCRAFT SPEED AND CLIMB PERFORMANCE.

1. PROP switch Set to OFF and wait 10 seconds

2. PROP switch Set to ON

3. PROP DE ICE circuit breakers Check. Do not reset unless tripped (Battery and Generator 1 busbar)

IF CAPTIONS RETURN TO NORMAL OPERATION AFTER 5 SECONDS:

4. Aircraft Continue flight and monitor system

IF CAPTIONS STAY IN FAILURE STATUS AFTER 5 SECONDS:

5. PROP switch

6. Aircraft

Maintain ON (together with INERT SEP OPEN) to maintain PUSHER ICE MODE

DEPART ICING CONDITIONS to positive IOAT atmosphere, if possible

IF PROPELLER VIBRATION OCCURS:

7. Power Set as required. If possible reduce to minimise vibration

8. Aircraft Avoid further icing conditions

IF PROPELLER VIBRATION CONTINUES OR ATTAINED PERFORMANCE DEGRADES:

9. Aircraft Land as soon as possible.



- - -


3.19.2 BOOT DEICE FAILURE IN ICING CONDITIONS

Indication: DE ICE CAWS CAUTION with flashlng DE ICE BOOTS CAWS ADVISORY.

A BOOT DEICE FAILURE IN ICING CONDITIONS CAN CAUSE DEGRADATION OF AIRCRAFT SPEED AND CLIMB PERFORMANCE AND A PREMATURE STALL. FLAP POSITION IS LIMITED TO 0" WITH THIS FAILURE.

1 .' BOOTS switch Set to OFF and wait 1 minute

2. BOOTS switch Set to 3MIN or 1 MIN

3. BOOTS DE ICE circuit breaker Check. Do not reset unless tripped. (Generator 2 busbar)

IF CAPTIONS RETURN TO NORMAL OPERATION:


IF CAPTIONS STAY IN FAILURE STATUS:

5. Aircraft DEPART ICING CONDITIONS to positive IOAT atmosphere, if possible

6. BOOTS switch Set to OFF

7. Aircraft

to

Avoid large or sudden changes in aircraft directional, longitudinal and lateral control until airframe is judged be free of residual ice.

8. Aircraft , Avoid further icing conditions

IF AIRFRAME IS FREE OF ICE ACCRETION:

9. Landing approach Flaps as reqd

IF AIRFRAME IS NOT FREE OF ICE ACCRETION:

10. Flap position Limited to 0"

1 1. Landing approach

Issued: June 10.1994 Revision 1 1 : March 1.2003

Keep minimum landing approach speed above 130 KIAS. The total landing distance will be longer by up to 90%.

Report No: 01 973-001 3-51 1


3.19.3 INERTIAL SEPARATOR FAILURE

Indlcatlon: INERT SEP CAWS ADVISORY does not come on with operation of inertial separator.

AN INERTIAL SEPARATOR FAILURE IN ICING CONDITIONS CAN CAUSE DEGRADATION OF THE AIRCRAFT ENGINE PERFORMANCE (AN INCREASE IN ITT).

AN INERTIAL SEPARATOR FAILURE DURING OPERATIONS IN FOREIGN OBJECT DAMAGE ENVIRONMENTS (FOD) MAY CAUSE LONG TERM ENGINE DETERIORATION AND SHOULD BE REPORTED FOR POST FLIGHT MAINTENANCE.

1. INERT SEP switch Set to CLOSED and wait 30 seconds

2. INERT SEP switch Set to OPEN

3. INERT SEP circuit breaker Check. Do not reset unless tripped. (Generator 1 busbar)

IF CAPTION RETURNS TO NORMAL OPERATION AFTER 30 SECONDS:


IF CAPTION STAYS IN FAILURE STATUS AFTER 30 SECONDS:

5. INERT SEP switch Maintain OPEN (together with PROP ON) to maintain PUSHER ICE MODE


7. Aircraft Avoid furlher icing conditions

IF ANY ATTAINED PERFORMANCE DEGRADATION CONTINUES:

8. Aircraft

Reporl No: 01973-001 1 3-52

Land as soon as possible.

Issued: June 10,1994 Revision 11 : March 1,2003 .


3.19.4 LH WINDSHIELD DEICE FAILURE IN ICING CONDITIONS

Indlcatlon: WSHLD HEAT CAWS ADVISORY does not come on with operation of windshield heating.

1. LH WSHLD circuit breaker Check. Do not reset unless tripped. (Generator 2 busbar)

2. LH WSHLD switch Cycle from OFF to LIGHT or HEAVY

IF CAPTION RETURNS TO NORMAL OPERATION:


IF CAPTION STAYS IN FAILURE STATUS AND FORWARD VISIBILITY THROUGH LH WINDSHIELD IS LOST:

4. Windshield Use RH windshield

IF TOTAL FORWARD VISIBILITY IS LOST:

5. Aircraft

hand.

DEPART ICING CONDITIONS to positive IOAT atmosphere, if possible. Interior fogging can be cleared by


IF WINDSHIELD HAS NOT CLEARED BY TIME OF LANDING:

7. Cabin pressure Make sure depressurized

8. DV window Use, if required.


Report No: 01973-001 3-53 1


3.19.5 AOA PROBE DEICE FAILURE IN ICING CONDITIONS

Indication: AOA DE ICE CAWS CAUTION with PROBES DE ICE CAWS ADVISORY.

AN AOA PROBE DEICE FAILURE IN ICING CONDITIONS CAN CAUSE A FALSE ACTIVATION OF THE STALL PROTECTION SYSTEM.

1. LH AOA SENS DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

2. RH AOA SENS DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

3. AOA PLATE HEAT Check. Do not reset unless tripped circuit breaker (Non Essential busbar)

4. PROBES switch Set to OFF and wait 3 minutes

5. PROBES switch Set to ON

IF CAPTIONS RETURN TO NORMAL OPERATION:

6. Aircraft Continue flight and monitor System

IF CAPTIONS STAY IN FAILURE STATUS:


STICK SHAKER MAY ACTIVATE AT HIGHER SPEEDS THAN NORMAL. IF THIS OCCURS, INCREASE SPEED UNTIL SHAKER STOPS.


9. Landing approach Keep minimum landing approach speed above 105 KlAS or shaker activation speed, whichever is highest. The total landing distance will be

longer by up to 55%.

I CAUTION 1 WHEN LANDING WITH AOA DEICE FAILURE THE EFlS AOA FAST SLOW POINTER MAY NOT BE CORRECT AND SHOULD NOT BE USED AS REFERENCE.

Report No: 01973-001 1 3-54

I


=PILATUS= SECTION 3 PC12 EMERGENCY PROCEDURES

3.19.6 PITOT AND STATIC PROBE DEICE FAILURE IN ICING CONDITIONS

Indication: PROBES DE ICE CAWS ADVISORY does not come on with operation of probes heating.

A PlTOT AND STATIC DEICE FAILURE IN ICING CONDITIONS CAN CAUSE AN INCORRECT INDICATION ON THE AS1 AND/OR ALTIMETER AND VSI.

1. PROBES switch Cycle from OFF to ON

2. PlTOT DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

3. STATIC DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

4. PROBES circuit breaker Check. Do not reset unless tripped (Battery busbar)



IF CAPTION STAYS IN FAILURE STATUS:

6. Autopilot Disconnect

7. Aircraft DEPART IClNG CONDITIONS to positive IOAT atmosphere, if possible



10. Landing approach Center EFlS AOA fast slow pointer with PUSHER ICE MODE and flaps 15". Maintain speed above shaker activation. The total landing distance will be longer by up to 55%.


Report No: 01973-001 3-55 1


3.19.7 PUSHER ICE MODE FAILURE IN ICING CONDITIONS

Indication: PUSHER CAWS CAUTION , PROP DE ICE and INERT SEP CAWS ADVISORIES are on. PUSHER ICE MODE CAWS CAUTION is off.

A FAILURE OF THE STALL WARNINGISTICK PUSHER SYSTEM TO RE-DATUM TO ICE MODE WHEN IN ICING CONDITIONS CAN LEAVE THE AIRCRAFT UNPROTECTED AGAINST THE NATURAL STALL WlTH RESIDUAL ICE ON THE AIRFRAME.

1. PUSHER test switch Press and hold for duration of Pusher test sequence (approx. 5 seconds)

(this identifies Pusher ice mode computer or selection failure).

IF FAILURE STAYS DURING TEST GO TO STEP 7

IF FAILURE DISAPPEARS DURING TEST BUT RETURNS AFTER COMPLETION OF TEST:

2. PROP switch Cycle from OFF to ON

3. INERT SEP switch Cycle from CLOSED to OPEN

IF CAPTIONS RETURN TO NORMAL OPERATION WITHIN 30 SECONDS:


5. PROPDE ICE Check. Do not reset unless tripped circuit breakers (Battery and Generator 1 busbar)

6. INERT SEP circuit breaker Check. Do not reset unless tripped (Generator 1 busbar)

IF CAPTION STAYS IN FAILURE STATUS:

7. Aircraft DEPART ICING CONDITIONS to positive IOAT atmosphere, if possible.


9. Landing approach Keep minimum landing approach speed above 105 KIAS. The total landing distance will be longer by up to 55%.

WHEN LANDING WlTH FAILED PUSHER ICE MODE, THE EFJS AOA FAST SLOW POINTER WILL NOT BE CORRECT AND SHOULD NOT BE USED AS REFERENCE.

Report No: 01973-001 1 3-56 Issued: June 10. 1994


=PILATUSW SECTION 3 PC12 EMERGENCY PROCEDURES

3.20 MISCELLANEOUS

3.20.1 AIR GND CAWS WARNING

INDICATES THE LEFT AND RIGHT WEIGHT ON WHEELS SWITCHES GIVE DIFFERENT INDICATIONS.

1. Landing gear

IF NO EFFECT:

Cycle down, up (once only)

A. IN THE AIR - the following systems could be affected:

Stick pusher Inoperative

Landing gear handle solenoid Goes to locked. Possible to lower gear but not to raise it

Transponder Not operative

Hydraulic Pump Not operative with engine OFF

ECS If ground idle selected, cabin will depressurize

Prop deice Not operative with engine off

The following CAWS captions may give Incorrect Indications:

The EIS propeller caution and warnings may be incorrect.

B. ON THE GROUND - the following systems could be affected:

Stick shaker -Aural warning Active

CAB PRESS

CHIP

Pressurization Cabin pressurizes to altitude on pressurization controller

PROP LOW P HYDR

Landing gear handle Up selection possible, gear will retract

STAB TRIM 1

Transponder Can transmit on ground


Report No: 01 973-001 3-57 1

SECTION 3 =PILATUSW EMERGENCY PROCEDURES PC12

3.20.2 PASSENGER AND CARGO DOOR

Indication: PASS DOOR or CAR DOOR CAWS WARNING Illuminated: Indicates that passenger or cargo door Is not correctly locked.

On ground: Visually check for the correct locking of the door latches (green indicators visible)

Passenger door check the handle lock pin for freedom of movement

During flight:

1. Reduce IAS

2. Start a slow descent

3. Decrease cabin press by selecting a higher cab. alt. and max cabin rate.

3.20.3 ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS) FAILURE

Indication: Attitude fall andlor heading fail ON EFIS.

A. Aircraft with single AHRS installation

1. Attitude and heading reference Use standby attitude and compass

2. Overhead panel Check at least one avionic bus is powered

3. Left and right windshield de icing switches OFF

4. Auxiliary heating system switch OFF MSN 321 and UP: CABIN HEATING switch OFF WF HEATER circuit breaker Pull (Generator 2 busbar)

5. Foot heater switch (if installed) OFF

6. Auxiliary cooling system (if installed) OFF

7. Aircraft Fly wings level, constant speed

8. Avionic 1 and 2 bus switches Both OFF, wait 5 secs then both ON

9. Aircraft Maintain wings level, constant speed, pitch attitude 10" or less from level. Wait 1 min to see if system realigns.

If no realignment continue flight using standby instruments.

Report No: 01 973-001 I 3-58



NOTE

If a yaw rate sensor is installed the yaw rate reference is independent of AHRS function.

B. Aircraft with second AHRS installation

1. AHRS 1IAHRS 2 switch Select AHRS 2

3.20.4 EFlS FAILURE

Indication: DU goes blank.

EFlS CMPST switch CMPST

NOTE


Indlcatlon: EFlS goes blank.

Use alternate instruments as required.

3.20.5 PlTOT I STATIC SYSTEM FAILURE

Indlcatlons: Incorrect Indication ON AS1 andlor altlmeter and VSI.

Probes switch Check ON. (Probes caption illuminated ON CAWS)

A If AS1 malfunctions:

Cruise and descent Only using known power settings and aircraft attitudes.

Approach Use AOA (AN) display in EADI. (Needle center gives 1.3 V stall)

B. If Altimeter malfunctions:

Below 10,000 ft Depressurize aircraft.

Cabin altitude selector Select actual aircraft altitude on outer scale.

When cabin pressure differential approaches 0:

ECS switch Dump

Use cabin altimeter to give approximate aircraft altitude.


Report No: 01 973-001 3-59 1


3.20.7 WHEEL BRAKE FAILURE

Indications: 1) Wheel brakes ineffective

2) Pedal excessively sofl when pressed.

Landing + Taxi Use reverse power, BETA and Nose wheel steering.

3.20.8 CAWS MALFUNCTION

indication: MASTER CAUTION ilght flashing wlth no CAWS captions on.

MASTER CAUTION light Press to reset CAWS.



lP lLATUSB SECTION 3 PC12 EMERGENCY PROCEDURES

3.21 AUTOPILOT

3.21.1 CHECKLIST

The four step procedure listed under paragraph A should be among the basic airplane emergency procedures that are committed to memory. It is important that the pilot be proficient in accomplishing all four steps without reference to this manuat.

A. Autopilot Malfunction (abrupt control and/or airplane motion, or AP TRIM CAWS warning annunciation with a warning tone).

Accomplish Items 1 and 2 simultaneously.

1. Airplane Control Wheel - GRASP FIRMLY and regain aircraft control.

2 Autopilot Disengage Switch - PRESS to disengage the autopilot (pilot or co-pilot wheel).

3. Aircraft - RETRIM manually as needed.

4. AUTOPllOT circuit breaker - PULL. (Avionic 2 busbar)

DO NOT ATTEMPT TO RE-ENGAGE THE AUTOPILOT FOLLOWING AN AUTOPILOT OR AUTOTRIM MALFUNCTION.

Maximum Altitude losses due to autopilot malfunction:

Confiauration .@!uBS Cruise, Climb, Descent 300 ft

Maneuvering 20 R

APR 3" ILS 70 11

APR 6" ILS 40 ft

Issued: June 10,1994 Revision 1 1 ; March 1,2003

Report No: 01 973-001 3-61 1

SECTION 3 SPILATUSW EMERGENCYPROCEDURES PC12

B. Autopilot Disengagement. The autopilot may be manually disengaged by any of the following methods:

1. Press the Autopilot Disengage switch on the pilot's or copilot's control wheel.

2. Actuate the airplane Electric Stabilizer Manual Trim Engage switch (momentary).

3. Press the autopilot (AP) mode selector switch (disengage).

4. Pull the AUTOPILOT circuit breaker (Avionic 2 busbar).

5. AVIONICS MASTER 2 switch to the off position.

C. Airplane Stall - Automatic Autopilot disengagement will occur by the stall warning system if stick shaker speed is approached.

D. Aileron (amber ROLL) mistrim:

NOTE

In the case of an aileron out of trim force for longer than 6 seconds, a steady, yellow, ROLL annunciation will appear on the EADI. The annunciation will flash after 10 seconds if not attended to. Failure to retrim at this point may result in either roll forces exceeding the authority of the autopilot or a sharp rolling motion upon autopilot disconnect. Autopilot operation in the presence of a continuing roll annunciation should not be continued.

Reporl No: 01 973-001 1 3-62 Issued: June 10,1994


WPILATUSW SECTION 3 PC12 EMERGENCY PROCEDURES

3.21.2 AMPLIFIED MALFUNCTION PROCEDURES

The following paragraphs are presented to supply additional information for the purpose of providing the pilot with a more complete understanding of the recommended course of action for an emergency situation.

1. An autopilot or autopilot trim disengage occurs when there is an uncommanded deviation in the airplane flight path due to a system malfunction. In some cases, and especially for autopilot trim, there may be little to no airplane motion, yet the red AP TRIM annunciator may illuminate and an alert tone sound. The KFC 325 autopilot incorporates monitors that detect abnormal airplane motion, therefore, i f the airplane for any reason is moved rapidly in pitch or roll the autopilot may be disconnected automatically.

The main concern in reacting to an autopilot or autopilot trim malfunction, or to an automatic disconnect of the autopilot, Is in maintaining control of the airp\ane. \mmediate\y grasp the control wheel and press the Autopilot Disengage Switch. Manipulate the controls as required to safely maintain operation of the airplane within all of its operaling limitations. Manual trim should be used as needed to relieve control forces.

With the AP and YD inode OFF or disengaged, the servo motors are no longer connected to the airplane's flight controls; autopilot trim will also be isolated from the alternate pitch trim and rudder trim system. Finally, the AUTOPILOT circuit breakers must be pulled to completely disable these systems.

DO NOT ATTEMPT TO RE-ENGAGE THE AUTOPILOT FOLLOWING AN AUTOPILOT1 AUTOTRIM MALFUNCTION UNTIL CORRECTIVE SERVICE ACTION HAS BEEN PERFORMED ON THE SYSTEM.

2. It is important that all portions of the autopilot system are preflight tested prior to each flight In accordance with the procedures published herein in order to assure their integrity and continued safe operation during flight. As a safety feature, autopilot functions cannot be engaged prior to successfull completion of the preflight test (mode controller TEST), due to an integrated system lock-out.


Report No: 01973-001 3-63 1


Repart No: 01973-001 1 3-64



PILOT'S OPERATING HANDBOOK

Normal Procedures

SECTION 4 NORMAL PROCEDURES

SECTION 4

NORMAL OPERATING PROCEDURES

CONTENTS

Paragraph Subject

NORMAL PROCEDURES

4.1 GENERAL

4.2 AIRSPEEDS FOR NORMAL OPERATIONS

PREFLIGHT INSPECTION EMPENNAGE RIGHT WlNG TRAILING EDGE RIGHT WlNG LEADING EDGE NOSE SECTION LEFT WING LEADING EDGE LEFT WING TRAILING EDGE CABIN COCKPIT

4.4 BEFORE STARTING ENGINE 4.4.1 PROCEDURE

4.5 ENGINE STARTING 4.5.1 NORMAL (USING AIRPLANE BATTERY) 4.5.2 WITH EXTERNAL POWER 4.5.3 DRY MOTORING RUN

4.6 BEFORE TAXIING

4.7 TAXIING

4.8 BEFORE TAKEOFF

4.9 TAKEOFF

4.10 FLIGHT INTO KNOWN ICING CONDITIONS


Page

4-1 I

Report No: 01973-001 4-1



CLIMB

CRUISE

DESCENT

BEFORE LANDING APPROACH CHECK FINAL CHECK

BALKED LANDJNG (GO-AROUND)

LANDING NORMAL SHORT FIELD

AFTER LANDING

SHUTDOWN

OXYGEN SYSTEM

AUTOPILOT AUTOPILOT OPERATION SUMMARY FLIGHT DIRECTOR OPERATION

1 4.21 NOISE LEVEL 4-42

AMPLIFIED PROCEDURES 4-43

4.22 CROSSWIND OPERATION 4-43

4.23 FLIGHT IN ICING CONDITIONS 4-44

4.24 SEVERE ICING CONDITIONS 4-45

1 4.25 PREVENTING OF FROZEN /LOCKED BRAKES 4-49 j

Issued: Februaiy 14,1994 Revision 10: September 1,2000



NORMAL OPERATING PROCEDURES

4.1 GENERAL

This section provides the normal operating procedures. All of the procedures required by regulation as well as those procedures which have been determined as necessary for the operation of thi i airplane are provided.

Normal operating procedures associated with optional systems or equipment which require supplements are contained in Section 9, Supplements.

Pilots must familiarize themselves with these procedures to became proficient in the normal operation of the airplane.

It is recommended that these procedures be followed for the normal operation of the aircraft, however a Short Checklist has been produced on an easy to use card. When the aircraft has been in extended storage, had recent major maintenance or been operated from prepared unpaved surfaces the full preflight inspection procedure given in thi i section is recommended.

PILOTS WHO FLY AT HlGH ALTITUDE MUST BE AWARE OF THE PHYSIOLOGICAL PROBLEMS ASSOCIATED WITH PROLONGED FLIGHTS AT SUCH ALTITUDES. DEHYDRATION AND THE SLOW ONSET OF HYPOXIA MAY BE NOTICED IN THE PASSENGERS.

PASSENGER COMFORT MAY BE INCREASED BY AN OCCASIONAL INTAKE OF FLUIDS. PROLONGED HlGH ALTITUDE FLIGHTS REQUIRE WARM CLOTHING AND MONITORING OF THE CABIN TEMPERATURE AND THE PHYSICAL STATE OF THE CREW AND PASSENGERS.


Report No: O f 973-001 4- 1

SECTION 4 NORMALPROCEDURES

4.2 AIRSPEEDS FOR NORMAL OPERANONS

Airspeeds for normal operations are listed bebw. Unless otherwise noted, all airspeeds are based on a maximum takeofl weight of 9,039 Ib (4,100 kg) at sea level under ISA standard day conditions. I

Takeofl (VR): Flaps 15" 78 KlAS

Maximum Climb: Best Angle (VX) 110 KlAS

Best Rate (VY) Flaps 0": Up to 10,000 A 120 KlAS 15,000 fl 115 KlAS 20,000 I and above 110 KlAS

Recommended Climb Speed with Flaps retracted and Pusher Ice Mode 125 KlAS

Recommended Holding Pattern Speed Range for Flight into Known Icing Conditions with Flaps retracted 140 to 170 KlAS

Maximum Operating Maneuvering Speed (Vo) (9,039 IW 4,100 kg) 151 KlAS

Maximum Flaps Extended (VFE): Flaps 15" 163 KlAS Flaps 40" 130 KlAS

Maximum Landing Gear: Extension (VLO) 177 KlAS Retraction (VLO) 177 KlAS Extended (VLE) 236 KlAS

Landing Approach Speed (based on Maximum Landing Weight of 9,039 IW 4,100 kg):

Flaps 0" 110 KlAS Flaps 15" 95 KlAS Flaps 40" 78 KlAS with residual ice on the airframe Flaps 15", Pusher Ice Mode 105 KlAS

Balked Landing (Go-Around): TOIPwr, Flaps 15" 95 KlAS TOIPwr, Flaps 40" 80 KlAS TOIPwr, Flaps 15". Pusher Ice Mode 105 KlAS

Maximum Demonstrated Crosswind for Takeon and Landing (not a limitation): Flaps 0° 30 kts Flaps 15" 25 kts

I Flaps 30" 20 kts I Flaps 40" (landing only) 15 kts

RepoR NO: 01973-001 4-2

Issued: June 10,1994 Revision 6: Dec 6,1996


4.3 PREFLIGHT INSPECTION

4.3.1 EMPENNAGE

1. Luggage

2. Cargo (Combi Interior)

3. Tie Down Straps (Combi interior)

4. Cargo Door

CHECKED and SECURED

CHECK that cargo is located against retainer angles installed on seat rails.

CHECK fimings properly inserted into seat rails and that the straps are tight.

After cargo loading I unloading: CHECK lower attachment lugs for condition.

5. Hydraulic system Make sure nitrogen pressure is in the (MSN 230 and UP) green sector and the level indicator

shows lull.

6. Cargo Door CLOSED and LOCKED (check for green flags)

7. Static ports CHECK CLEAR of OBSTRUCTIONS

8. Tail tie-down DISCONNECTED

9. External Power Door CLOSED

If a larger capacity oxygen system is installed in the rear fuselage:

10. Oxygen rupture disc INTACT


Report No: 01973-001 4-3


1 1. Rudder and trim tab

12. Vertical stabilizer

13. Elevator assembly

14. Horizontal stabilizer

15. Deicing Boots

16. Static discharge wicks

17. Dorsal and ventral fairings

18. General condition

19. Banery Compartment

4.3.2 RIGHT WlNG TRAILING EDGE

1. Flaps

2. Aileron



4.3.3. RIGHT WlNG LEADlNG EDGE

1. NavIStrobe light

2. Fuel tank vent

3. Fuel quantity and filler cap

4. Pitot probe

5. AOA probe

Report No: 0 t 973-001 4-4

CHECK VISUALLY

CHECK VISUALLY

CHECK VISUALLY

CHECK VISUALLY, Stabilizer Trim Mark within green range.

CHECK VISUALLY

CHECK

CHECK

CHECK

CHECK CLOSED

CHECK CONDITION

CHECK CONDITION

CHECK

CHECK

CHECK CONDITION

CLEAR of OBSTRUCTIONS

CHECK and SECURE

COVER REMOVED and CHECKED

COVER REMOVED CHECK FREE MOVEMENT



6. Wing tie-downlwheel chocks DISCONNECTED and REMOVED

7. De-Icing boot CHECK GENERAL CONDITION

8. Right main landing gear CHECK

9. Right brake assembly CHECK

10. Two fuel drains SAMPLE and SECURE

1 1. General condition CHECK

4.3.4 NOSE SECTION r

A. SERVICE BAY (RIGHT) (If a standard oxygen system Is installed):

1. Oxygen Lever CHECK OPENIGUARDED (MSN 101-180 Pre SB 35-001)

2. Oxygen Press CHECK

3. Oxygen and ECS Doors CLOSED

4. Oxygen rupture disc INTACT

DO NOT TOUCH OUTPUT CONNECTORS OR COUPLING NUTS OF IGNITION EXCITER WITH BARE HANDS.

B. Engine Area:

1. Cowling RH CHECK and SECURE

2. Propeller - a. Blade Anchor RE MOVED and STOWED

b. Blade CHECK

c. De-Icing Boots CHECK GENERAL CONDlT ION

d. Spinner CHECK

3. Air Inlet Covers REMOVED and STOWED

4. Air Inlets CHECK OIL COOLER ECS and GENERATOR for OBSTRUCTIONS

5. Exhaust System CHECK

6. Nose Gear and Doors CHECK

7. Wheel Chocks REMOVED


Report No: 01 973-001 4-5


8. Engine drain mast (LH) CHECK. No leaks permitted

9. MSN 106 & UP Engine drain (LH) SAMPLE and SECURE

10. Oil Quantity CHECK SIGHT GLASS AND

DIPSTICK FOR SECURITY

Check oil level in green range of sight glass within 10 to 20 minutes alter engine shut down. If engine has been shut down for more than 30 minutes, check dipstick indication and # it indicates that oil is needed, start the engine and run at ground idle for 5 minutes. Recheck oil level using dipstick and refill if necessary.

1 1. General Condition CHECK

12. Cowling LH CHECK and SECURE

13. Windshield CHECK CLEAN

C. SERVICE BAY (LEFT)

1. Fuel Filter SECURE INDICATOR FLUSH

2. Fuel Filter drain SAMPLE AND SECURE

3. Fuel Compartment Doors CLOSED

4. Air Separator drain SAMPLE AND SECURE

4.3.5 LEFT WING LEADING EDGE

I 1. Two fuel drains SAMPLE and SECURE

2. Left main landing gear

3. Left brake assembly

4. De Icing boot

5. AOA Probe FREE

6. Wing tie-downlwheel chocks

7. Fuel quantity and tiller cap

8. Fuel tank vent

9. Nav/Strobe light


CHECK

CHECK

CHECK GENERAL CONDITION

COVER REMOVED and CHECK

MOVEMENT

DISCONNECTED and REMOVED

CHECK and SECURE

CLEAR of OBSTRUCTIONS

CHECK CONDITION

CHECK

Report No: 01 973-001 4-6

Issued: June 10,1994 Revision 10: September 1.2000


4.3.6 LEFT WING TRAILING EDGE


2. Aileron and trim tab

3. Flaps


4.3.7 CABIN

1. Main Entry Door

2. Hand luggage

3. Passenger Seat

4. Passenger Seat Belts

5. Overwing emergency exit

6. Fire Extinguisher

For flights above 10,000 11 altitude:

7. Passenger oxygen masks

4.3.8 COCKPIT

1. Flight Control Lock

2. External Power switch

3. Battery Master switch Battery 1 and 2 switches

4. Generator 1 8 2 switches

5. AV Bus 1 and 2 switches

6. Non-essential Bus

7. Circuit breakers

8. Parking Brake Handle

9. Ignition switch

10. Fuel Pumps switch


CHECK SECURITY and CONDITION

CHECK CONDITION

CHECK CONDITION

CHECK

CLOSED and LOCKED (check for 6 green flags) If required, fit thermal blanket. Make sure door handle and latch are accessible.

CHECK Backrests in upright position (for takeoff and landing)

FASTENED

LOCK PIN REMOVED, EXIT CHECKED and LOCKED

CHECK ATTACHMENT and PRESSURE

CONNECTED AND STOWED (for each passenger)

REMOVED and placed in STOWAGE POINT

OFF

OFF OFF (if 2* battery installed)

OFF

OFF

AUTO

CHECK IN. (Except FLAP MAlNT if installed)

SETIPUSH BRAKE PEDALS

AUTO

AUTO

RepoR No: 01 973-001 4-7


1 1. AHRS SLAVE AHRS IlAHRS 2 (if installed) SLAVE

12. Cooling switch OFF

13. Fan switches (2) LOW

14. Heating System OFF

15. EFlS CMPST switch NORM

16. Audio switch

17. External Lights switches (5)

18. De-Ice switches (5)

19. Landing Gear Handle

20. ECS switch

21. ECS pressurization switch

22. Trim lnterrupt switch

23. Flap Interrupt switch

NORM

OFF

OFF

DN

OFF

AUTOIGUARDED

NORMIGUARDED

NORWGUARDED

TO PREVENT DAMAGE TO ENGINE CONTROLS, DO NOT MOVE THE POWER CONTROL LEVER AFT OF THE IDLE DETENT WITH ENGINE NOT RUNNING.

24. Manual Override Lever

25. Power Control Lever

26. Condition Lever

27. Flap Lever

28. Cockpit/ lnstrumentl Cabin Light switches

29. Fuel Firewall Shut-on lever

30. Hydraulic hand pump handle

31. ECS Firewall Shut-off lever

Report No: 01973-001 1 4-8

OFF

IDLE DETENT

CUT-OFFIFEATHER

0"

OFF

FULLY IN

STOWED

FULLY IN



4.4 BEFORE STARTING ENGINE

4.4.1 PROCEDURE

1. Preflight inspection COMPLETE

2. Seats ADJUSTED and LOCKED

3. Seat belts FASTENED

4. Battery Master switch ON Battety 1 and 2 switches ON (if 2"d battery installed)

5. External power ON (if available)

6. Landing Gear 3 greens

7. Fuel Contents

8. FUEL PUMP LH switch ON and AUTO FUEL PUMP RH switch ON and AUTO

9. OXYGEN lever

10. Oxygen pressure gage

11. PASSENGER OXYGEN switch

CHECK

CHECK equal L & R levels

CHECK for audible operation

CHECK for audible operation

ON (if installed)

CHECK 1.850 psi MAX

ON. CHECK CAWS PASS OXY caption on. SET switch to AUTO

12. Oxygen masks AVAILABLE for all occupants, as required

I

13. Pilot ventilation window CLOSED and LOCKED

14. External lights AS REQUIRED

NOTE

Avoid prolonged use of the beacon and logo lights (if installed), as this can cause a decrease in battety power and affect the engine starting.

Issued: June 10, 1994 Revision 1 1 : March 1.2003

Report No: 01973-001 4-9


4.5 ENGINE STARTING

4.5.1 NORMAL (USING AIRPLANE BATTERY)

I 1. Voltmeter CHECK (24 VDC min). 4

Both indicators (if 2"4 battery installed)

2. Fuel quantity CHECK I

3. ElS System TEST

4. Test Fire Wrn switch TEST (ENG FIRE and F DETECT lights on)

5. Test Lamp LAMP (Warning, caution, advisory lights on)

6. Propeller area CLEAR, Confirm CLEAR of 1 obstructions

7. Starter switch ON

a. Oil pressure CHECK RISING

b. Fuel Press Caution It OFF

c. Ng

8. Condition Lever

NOTE

STABLE between 13% and 20%

GROUND IDLE

Set FLIGHT IDLE for a cold engine. (oil temperature below + 5") until NG above 50%, then set GROUND IDLE.

Apply maximum possible brake pressure prior to engine Start.

a. ITT MAXIMUM 1 ,OOO°C LIMITED TO 5 Sec. 800" - 870" C limited to 20 sec

b. Ng STABLE above 50%

If Ng stays below 50% then:

c. CONDITION LEVER FLIGHT IDLE

d. ITT MONITOR

9. If combustion is not initiated within 10 sec of moving Condition Lever to GROUND OR FLIGHT IDLE, then:

a. Condition Lever CUT-OFFIFEATHER

b. Starter switch RESET

c. Allow min 30 sec draining period, then refer to DRY MOTORING RUN.

Report No: 01 973-001 4-10



10. Starter sequence

11. Engine instruments

12. Fuel Totalizer

13. Generator 1 switch

14. Generator 2 switch

15. INVERTER BAT or GEN

16. AV Bus 1 and 2 switches

17. RadioslAvionics

18. NavlStrobe Lights switch

19. ECS switch

20. Temperature setting

21. Cooling or Heating System

COMPLETED

CHECK

RESET

ON-Check CAWS GEN 1 OFF

ON-Check CAWS GEN 2 OFF

CHECK FUNCTION CHANGE SELECTION

ON

AS REQUIRED

AS REQUIRED

AUTO

AS REQUIRED

AS REQUIRED

22. Inertial Separator OPEN, if operating on unprepared surface

4.5.2 WITH EXTERNAL POWER

1. Battery Master switch ON Battery 1 and 2 switches ON (if 2"4 battery installed)

2. External Power Unit Connected, then ON

3. External Power switch ON

4. Voltmeter CHECK 28 VDC

NOTE

Aircraft Post SB 24-008 and MSN 231 & UP. The external power control unit on the aircraft will disconnect the EPU if the output voltage is above 29.5 or below 23 VDC.

5. Cooling and Heating System OFF

6. Fuel quantity CHECK

7. ElS System TEST

8. Test Fire switch

9. Test Lamp

10. Propeller area

TEST (ENG FIRE and FIRE DETECT lights on)

LAMP (Warning, caution, advisory lights on)

CLEAR, Confirm CLEAR of obstructions


Report No: 01 973-001 4-11


I 11. Starter switch ON

a. Oil pressure CHECK rising

b. Fuel Press caution It OFF I

c. Ng STABLE between 13% and 20%

I 12. Condition Lever GROUND IDLE

NOTE

Set FLIGHT IDLE for a cold engine (oil temperature below + 5") until NG above 50% then set GROUND IDLE.

I Apply maximum possible brake pressure prior to engine Start.

a. ITT Maximum 1 ,0OO0C LIMITED TO 5 Sec , 800" - 870' C limited to 20 sec

b. Ng Stable above 50%

If Ng stays below 50% then:

C. CONDITION LEVER FLIGHT IDLE

d. ITT CHECK I

13. If combustion is not initiated within 10 sec of moving Condition Lever to GROUND OR FLIGHT IDLE, then:

a. Condition Lever CUT-OFFIFEATHER

b. Starter switch RESET

c. Allow min 30 sec draining period, then refer to DRY MOTORING RUN.

14. Starter sequence COMPLETED

15. Engine Instruments CHECK

16. Generator 1 switch ON - Check CAWS GEN 1 OFF

17. Generator 2 switch ON - Check CAWS GEN 2 OFF

18. External Power switch OFF

19. External Power Unit OFF then disconnect

20. AV Bus 1 and 2 switches ON

21. INVERTER BAT or GEN CHECK FUNCTION CHANGE SELECTION

22. Fuel Totalizer RESET

23. RadioslAvionics AS REQUIRED

24. NavIStrobe Lights switch AS REQUIRED

I Report No: 01973-001 1 4-12


EPILATUSW SECTION 4 ?c %I1 NORMAL PROCEDURES

25. ECS switch AUTO

26. Temperature setting AS REQUIRED

27. Cooling or Heating System AS REQUIRED

28. Inertial Separator OPEN, if operating on unprepared surface

4.5.3 DRY MOTORING RUN

Allow min 30 sec draining period, then:

1. Condition Lever CUT-OFFIFEATHER

2. Power Control Lever IDLE DETENT

3. Ignition CB PULL

4. Battery MasterlExt pwr switch ONION

5. Fuel Pumps switches ON

6. Starter switch ON

After 15 seconds: Should a fire persist, as indicated by sustained ITT, CLOSE the Fuel Emerg Shut-off valve at this point and continue motoring for another 15 sec.

7. Starter switch RESET

8. Fuel Pumps switches AUTO

9. Ignition CB RESET

10. Battely MasterlExt pwr switch OFFIOFF

Observe starter cooling OH limits, then initiate applicable engine start procedure.

Issued: June 10,1994 Revision 8: September 1. 1998

Report No: 01 973-001 4-13


1.6 BEFORE TAXIING

1. AHRS CHECK - NO FLAGS AHRS llAHRS 2 (if installed) CHECK - NO FLAGS

2. Flaps Lever 15"

3. Autopilot Test Button PRESS momentarily and NOTE:

a. All annunciator lights on (ROLL, PTRM and AP TRlM FAlL annunciators flashing).

b. After approximately 5 seconds, all annunciator lights off except AP which will flash 12 times prior to extinguishing. Note the aural disconnect and the AP TRlM FAlL tones.

IF THE AP TRlM FAlL WARNING ON THE CAWS STAYS ON, THEN THE AUTOTRIM DID NOT PASS PREFLIGHT TEST. THE AUTOPILOT CIRCUIT BREAKER MUST BE PULLED.

c. Control Wheel HOLD to keep from moving.

d. AP Button PRESS to engage autopilot.

e. Control Wheel that

MOVE fore, aft, left & right to verify the autopilot can be overpowered.

f. Autopilot Disengage switch PRESS. Verify that the autopilot disconnects and all flight director modes are cancelled.

4. PCL

I 5. PUSHER test switch (Test Panel)

SET 5 - 10 psi

PRESS AND HOLD

l a. PCL Set to idle.

I b. Elevator Control PULL

1 c. CAWS PUSHER ICE MODE CHECK ON caution

I d. Shaker and Audio Warning for 2 sec Break for 1 sec Shaker and Audio Warning for 2 sec Break for 1 sec Pusher, Shaker and Audio Warning CHECK correct operation

When pusher operates:

I e. CAWS PUSHER caution CHECK OFF

Report No: 01 973-001 4-1 4



f. Pusher test switch

g. CAWS PUSHER ICE MODE caution

h. PUSHER INTR switch (Control Wheel)

i. CAWS PUSHER caution

j. PUSHER INTR switch

k. CAWS PUSHER caution

6. CAWS panel

7. LH WSHSLD switch

8. RH WSHSLD switch

9. PROBES switch

10. INERT SEP switch

1 1. EXTERNAL LIGHTS switches

12. Parking Brake Handle

IF ICING CONDITIONS EXPECTED

13. PROP switch

14. PUSHER ICE MODE caption

15. BOOTS switch

16. All DE ICING switches


RELEASE

CHECK OFF

PRESS and HOLD

CHECK ON after 3 sec

RELEASE

CHECK OFF

CHECK (No red warning lights)

CHECK green light WSHLD HEAT on when switched to LIGHT and HEAVY

CHECK green light WSHLD HEAT on when switched to LIGHT and HEAVY

Set to ON and check green light PROBES DE ICE on amber light AOA DE ICE off

Set to OPEN and check green light INERT SEP comes on within 30 seconds. Set switch to CLOSED

AS REQUIRED

RELEASE

Set to ON and check green light PROP DE ICE comes on within 5 seconds

Check amber light comes on when PROP switch set to ON with INERT SEP switch at OPEN

Set to 1 MIN and CHECK for a minimum of one minute green light DE ICE BOOTS is on

Set to OFF

Report No: 01 973-001 4-15

SECTION 4 I IPILATUSF NORMAL PROCEDURES ?c %I,

4.7 TAXIING

1. Passenger(s) BRIEF

2. Brakes CHECK

3. Flight instruments CHECK

4. LH and RH WSHLD switches Set to LIGHT or HEAVY as required

I CAUTION 1

TO AVOID POSSIBLE PROPELLER DAMAGE, DO NOT ALLOW STABILIZED PROPELLER OPERATION BETWEEN 350 AND 950 RPM (PROPELLER NOT FEATHERED).

NOTE

Beta range (aft of idle detent) may be used during taxi to control taxi speed and reduce wear on brakes.

4.8 BEFORE TAKEOFF

Parking Brake

Takeoff power setting

Fuel quantity

Friction lock

Engine instruments

Flight instruments

Trim If CG is 236 inches (6 meters) or turlher aft of datum

Flaps

Flight controls

CAWS panel

Radios/Avionics

Pilot ventilation window

1 Reporl No: 01 973-001 ( 4-16

SET

CALCULATED

CHECK

ADJUST

CHECK

CHECK and SET

SET GREEN LINES SET GREEN DIAMOND

15"

FREE and CORRECT

CHECK (No warning RED lights)

AS REQUIRED

CLOSED and LOCKED



13. Cabin Pressure Controller SET to cruise altitude + 500 tt SET rate knob to mid position

14. Condition Lever FLIGHT IDLE

15. De ice Probes switch ON

16. Windshield Heat AS REQUIRED

17. Nav/Strobe switch ON

18. DC Amps Battery CHECK (15 amps maximum. If greater Batteries (if 2"6 battery installed) than 15 amps, delay takeoff until

indication at or below 15 amps) I

19. Transponder ON

20. Parking Brake RELEASE

IF ICING CONDITIONS EXPECTED SET THE DE ICING SWITCHES AS FOLLOWS:

21. PROP ON

22. INERTSEP OPEN

23. BOOTS 3 MIN or 1 MIN as required

NOTE

To switch between 3 MIN and 1 MIN cycles during operation: set system to OFF, wait until BOOTS DE ICE caption is off then set system as required.

24. LH and RH WSHLD switches LIGHT or HEAVY as required


Report No: 01 973-001 4-17 1


4.9 TAKEOFF

1. EHSl CHECK HDG

2. ECS switch OFF (If torque as per Static Takeoff Torque chart in Section 5 is below flat rating)

3. Power Control Lever SET (Under certain hot andlor high aitfield altitude the engine power is below the torque limiter setting and manual power setting is required according to Static Takeoff Torque chart in Section 5)

p Z i E i E r 1

THE TORQUE LIMITER ASSISTS THE PILOT IN SETTING THE ENGINE POWER. THE PILOT IS RESPONSIBLE TO RESPECT ALL ENGINE OPERATING LIMITS.

NOTE

Increasing airspeed might cause torque and In to increase.

4. Engine instruments:

a. Torque MONITOR

b. ITT MONITOR

c. Ng MONITOR

d. Oil Temp/Pressure MONITOR

5. Rotate at VR. initial climb at VX or VY, as required

6. Brakes PRESS to stop wheel rotation

AHer lift-off and positive rate of climb:

7. Landing Gear Handle UP

8. Flaps 0" above 100 KlAS

9. Taxi Light OFF

I 10. External Lights AS REWIRED

1 1. Yaw Damper AS REQUIRED

12. WX Radar AS REQUIRED (if installed)

RepoR NO: 01973-001 4-18



4.10 FLIGHT INTO KNOWN ICING CONDITIONS

Icing conditions are defined in Section 1

NOTE

Flight in icing conditions is only permitted with full operational status of all aircraft deicing systems. The deicing systems may be activated before takeoff.

FLIGHT IN ICING CONDITIONS IS PROHIBITED IF THERE IS A KNOWN FAILURE OF ANY OF THE ICE PROTECTION SYSTEMS OR A FAILURE OF GENERATOR 1 OR GENERATOR 2.

DURING FLIGHT IN ICING CONDITIONS OR FLIGHT WITH ANY VISIBLE ICE ACCRETION ON THE AIRFRAME. THE FOLLOWING FLAP EXTENSION LIMITS APPLY:

- WITH OPERATIONAL AIRFRAME PNEUMATIC DEICE BOOTS = 15" FLAP.

- AFTER FAILURE OF THE AIRFRAME PNEUMATIC DEICE BOOTS = On FLAP.

BEFORE ENTERING ICING CONDITIONS SET THE DE ICING SWITCHES AS FOLLOWS

1. PROP ON

2. INERT SEP OPEN

3. BOOTS 3 MlN or 1 MIN as required

4. LH and RH WSHLD switches LIGHT or HEAVY as required

NOTE

When DE ICING switch PROP is set to ON and INERT SEP is set lo OPEN. the stick shakerlpusher system is automatically reset to provide stall protection at lower angles of attack. The CAWS amber caption PUSHER ICE MODE comes on to inform the aircrew of this mode change. In this mode the shaker and pusher are activated at higher airspeeds.

Issued: June 10. 1994 Revision 6: Dec 6, 1996

Report No: 01973-001 4-19 (


DURING ICING CONDITIONS:

5. Wing leading edge

6. CAWS

MONITOR for continual shedding ol ice

MONITOR lor correct function of ice protection systems

IF ANY OF THE AIRCRAFT ICE PROTECTION SYSTEMS FAIL DURING FLIGHT IN ICING CONDITIONS. EXIT ICING CONDITIONS. CONTACT ATC FOR PRIORITY ASSISTANCE IF REQUIRED.

AFTER DEPARTURE OF ICING CONDITIONS WITH RESIDUAL AIRFRAME ICE

7. PROP Maintain ON

8. INERT SEP Maintain OPEN

This ensures that the stick shakerlpusher system is maintained in PUSHER ICE MODE.

9. BOOTS As required

10. LH and RH WSHLD LIGHT or HEAVY as required

11. Flaps Do not extend beyond 15" or if extended do not retract to 0"

AFTER REMOVAL OF RESIDUAL AIRFRAME ICE

12. PROP OFF

13. INERT SEP CLOSED

14. BOOTS

15. LH or RH WSHLD

16. Flaps

Report No: 01 973-001 1 4-20

OFF

LIGHT or HEAVY as required

AS REQUIRED

Issued: June 10. 1994 Revision 6: Dec 6. 1996


rr 4.11 CLIMB

1. Ice Protection system

1 2. Autopilot

I 3. Power Control Lever

I

4. ECS switch

5. Engine instruments:

a. Toque

b. ITT

c. Ng

6. Temperature setting

4.12 CRUISE

1. Altimeters

2. AHRS

3. Cabin Pressurization

AS REQUIRED

AS REQUIRED

SET (According to Climb Torque chart for best performance or 720°C ITT recommended)

AUTO

MONITOR

MONITOR

MONITOR

AS REQUIRED

SET 1013.2t29.92

CHECK

Monitor* 'Confirm cabin pressure differential is < 5.75 psi (i.e. gauge green arc) -

If cabin pressure differential > 5.75 psi check cabin altitude selection correct. If cabin pressure differential still > 5.75 psi there is a malfunction of the cabin pressure control system. Refer to section 3.18 to determine appropriate action.

4. Power Control Lever SET (According to Cruise Toque table)

5. Engine Instruments MONITOR

Aircraft Post SB 79-003 and MSN 273 - 400. Oil temperature excursions into the yellow Caution range during cruise and descent are acceptable.

6. Ice Protection system AS REQUIRED


Report No: 01973-001 4-21


4.13 DESCENT


2. Power Control Lever SET to desired toque

3. Cabin Pressure Controller SET to field elevation + 500 R

4. Windshield Heat AS REQUIRED

4.14 BEFORE LANDING

4.14.1 APPROACH CHECK


2. Altimeter SET

3. Fuel Quantity CHECK

4. Landing Gear DOWN (below 177 KIAS)

5. Landing Lights AS REQUIRED

6. External Lights AS REQUIRED

7. Flaps - Normal SET 15'( max 163 KIAS) - With residual airframe ice SET 15" (max 163 KIAS) - Boot failure Maintain at O0

8. Speed AOA Centered - Pusher Normal Mode 95 KlAS at 9039 Ib (4100 kg) - Pusher Ice Mode 105 KlAS at 9039 Ib (4100 kg)

Boot or AOA Deice or 130 KlAS PUSHER ICE MODE failure

I CAUTION 1

ON LANDING APPROACH AFTER BOOT FAILURE (FLAPS 0") OR AFTER PUSHER ICE MODE FAILURE OR AFTER AOA DEICE FAILURE THE EFlS AOA FAST SLOW POINTER WILL NOT BE CORRECT AND SHOULD NOT BE USED AS REFERENCE.

9. Passengers

10. Inertial Separator

Repon No: 01973-001 4-22

Brief

OPEN, H operating on unprepared surface



4.14.2 FINAL CHECK

1. Landing Gear 3 Green Lights

2. Flaps - Normal SET 40" (below 130 KIAS) - With residual airframe ice Maintain at 15"

3. Speed AOA Centered - Pusher Normal Mode 78 KlAS at 9039 Ib (4100 kg) - Pusher Ice Mode 105 KlAS at 9039 Ib (4 100 kg)

Boot or AOA Deice or 130 KlAS PUSHER ICE MODE failure

4. Pressurization 0 Diff Pressure

5. Autopilot DISENGAGED

6. Yaw Damper (prior landing) DISENGAGED

NOTE

For crosswind information, refer to para 4.2 and Section 5

For minimum autopilot use height, refer to Section 2 (Autopilot).

Issued: June 10, 1994 Revision 6: Dec 6, 1996

Report No: 01973-001 4-23 (


4.15 BALKED LANDING (GO-AROUND)

1. Go Around switch (il autopilot engaged)


3. Cl~mb airspeed

4. Flaps - Normal - W~th residual airframe c e - Boot failure

5. Climb airspeed - Pusher Normal Mode - Pusher Ice Mode - Bool failure

6. Landing Gear Handle

PRESS

SET (According to the Balked Landing Torque chart in Section 5)

80 KlAS

SET 15" (max 163 KIAS) Mainlain al 15" Maintain at 0"

95 KIAS 105 KIAS 130 KIAS

Up with positive rate-01-climb

7. Flaps - Normal AS REQUIRED - With residual airframe ice Maintain at 15" - Boot failure Maintain O0


IN THE EVENT OF A BALKED LANDING (GO-AROUND) WITH RESIDUAL ICE ON THE AIRFRAME, THE FLAPS SHOULD NOT BE RETRACTED. THE LANDING GEAR MAY NOT FULLY RETRACT AFTER SELECTION (REMAINING RED INDICATION).

I Report No: 01973-001 1 4-24



4.16 LANDING

4.16.1 NORMAL

1. TOUCH DOWN MAlN WHEELS FIRST.

2. DO NOT FLARE WITH HlGH PITCH ANGLE.

3. Power Control Lever IDLE

4. Condition Lever GROUND IDLE

5. Braking AS REQUIRED

4.16.2 SHORT FIELD

1. TOUCH DOWN MAIN WHEELS FIRST.

2. DO NOT FLARE WlTH HlGH PITCH ANGLE.

3. Power Control Lever IDLE

4. Reverse SELECT MAX (if desired)

5. Brake FIRM

6. Power Control Lever IDLE (before airplane stops)

4.17 AFTER LANDING

When runway vacated:


2. Condition Lever

3. Flaps

4. External Lights

5. De-Ice switches (5)

6. Transponder

7. WX Radar


AS REQUIRED

GROUND IDLE

UP

AS REQUIRED

OFF

STBY

STBY (if installed)

Report No: 01973-001 4-25


4.18 SHUTDOWN

FOR ANY INDICATION OF ENGINE FIRE AFTER SHUTDOWN, IMMEDIATELY DO DRY MOTORING RUN PROCEDURE.

NOTE

Allow ITT to stabilize at least two minutes at ground idle. Monitor compressor deceleration after shutdown for possible engine damage.

1. Power Control Lever IDLE DETENT

2. Parking Brake SETIPEDALS PUSH

3. Cooling and Heating system OFF

4. External Lights switches OFF

5. ECS switch OFF

6. Inertial Separator CLOSED

7. AV Bus 1 and 2 switches OFF

8. Generator 1 and 2 switches OFF

9. Condition Lever CUT-OFFIFEATHER

10. Oxygen shut-off lever OFF (if installed)

11. Lighting switches OFF 1 12. CAWS OIL QTY warning CHECK. Refill engine with an (60 secs minimum atter shutdown) approved oil

13. Battery Master switch OFF Battery 1 and 2 switches OFF (if 2* battery installed)

( 14. Flight Control Lock INSTALLED

( 15. Main entry door REMOVE thermal blanket (if installed)

1 16. Wheel chocks AS REQUIRED 1 17. Tie downs AS REQUIRED

MAKE SURE PROPELLER ANCHOR IS PROPERLY INSTALLED TO PREVENT POSSIBLE ENGINE DAMAGE DUE TO WINDMILLING WITH ZERO OIL PRESSURE.

1 18. Propeller anchor INSTALLED

1 19. External covers INSTALLED

Report No: 01973-001 4-26



4.19 OXYGEN SYSTEM

1. Oxygen Pressure Gauge NOTE READING

2. Outside Air Temperature NOTE READING

3. Percentage of Full Bottle DETERMINE from the ' Oxygen Available with Partially Full Bottle " graph, Figure 4-1.

4. COMPUTE Oxygen Duration in minutes:

a. Determine the Oxygen Duration in minutes for a full bottle for the number of connected passenger oxygen masks and pilots from the " Oxygen Duration with Full Bottle " graph. Figure 4-2.

I

b. Multiply the Full Bottle Duration by the percent of Usable Capacity to obtain the available oxygen duration in minutes.

5. Turn the Oxygen shut-off lever (if installed) and Passenger Oxygen control valve to ON. Insert the connector of each mask into an outlet and verify proper oxygen flow to the mask. For flights above 10,000 feet leave the masks connected to the outlets and turn the Oxygen Control Valve to AUTO.

OXYGEN AVAILABLE WITH PARTIALLY FULL BOllLE

PERCENT OF USEABLE CAPACITY (%)

Figure 4-1. Oxygen Available with Partially Full Bottle


Report No: 01973-001 4-27


Figure 4-2. Oxygen Duration with Full Bottle (Standard Oxygen System) (Sheet 1 of 2)

No. of Pax Oxygen

Masks Connected

0

1

2

Report No: 01973-001 4-28


Oxygen Duration

Pax plus 1 Crew Mask on

9

Oxygen Duration

Pax plus 2 Crew Masks on

Diluter1 Demand

(min)

141

70

47

Diluter1 Demand

(min)

71

47

35

100 %

(min)

59

42

32

14

100 %

(min)

29

24

21

12 13 10


Figure 4-2. Oxygen Duration with a Full Bottle (Larger Capacity Oxygen System) . (Sheet 2 of 2)

No. of Pax Oxygen

Masks Connected

0

1

2

Issued: June 10,1994 Revision 1 1: March 1,2003

Report No: 01 973-001 4-29

Oxygen Duration

Pax plus 1 Crew Mask on

9

Diluter1 Demand

(min)

477

237

159

Oxygen Duration

Pax plus 2 Crew Masks on

100 %

(min)

200

142

108

Diluter1 Demand

(min)

240

159

118

I 47 41 I

100 %

(min)

98

81

71

44 34

SECTlON 4 RPILATUSS NORMAL PROCEDURES PC XI1 1

4.20 AUTOPILOT .i

I 4.20.1. AUTOPILOT OPERATION SUMMARY

THE PILOT IN COMMAND MUST CONTINUOUSLY MONITOR THE AUTOPILOT WHEN IT IS ENGAGED, AND BE PREPARED TO DISCONNECT THE AUTOPILOT AND TAKE IMMEDIATE CORRECTIVE ACTION - INCLUDING MANUAL CONTROL OF THE AIRPLANE AND/OR PERFORMANCE OF EMERGENCY PROCEDURES - IF AUTOPILOT OPERATION IS NOT AS EXPECTED OR IF AIRPLANE CONTROL IS NOT MAINTAINED.

DURING ALL AUTOPILOT COUPLED OPERATIONS THE PILOT IN COMMAND MUST USE PROPER AUTOPILOT COMMANDS AND USE THE APPROPRIATE COMBINATION OF ENGINE POWER, WING FLAPS, AND LANDING GEAR TO ENSURE THAT THE AIRPLANE DOES NOT EXCEED AIRPLANE OPERATING LIMITATIONS.

NOTE

The autopilot is inoperative with two faiied inverters.

1. Before takeoff

Autopilot and Yaw Damper - DISENGAGE.

IF THE AIRPLANE IS NOT LEVEL AND THE YD IS INADVERTENTLY ENGAGED, THERE IS A POSSIBILITY OF RUDDER AUTOTRIM MOVEMENT. PRIOR TO TAKEOFF, CONFIRM THAT THE RUDDER TRIM IS SET TO THE GREEN MARK.

Report NO: 01 973-001 4-30



2. lnflight Autopilot Engagement

(Verify the airplane is in a trimmed condition prior to autopilot engagement.)

a. AP Button - PRESS. Note AP and FD annunciators ON. If no other flight director modes are selected at the time of autopilot engagement the mode of operation will be flight director wings level and pitch attitude hold.

DO NOT HELP THE AUTOPILOT OR HAND-FLY THE AIRPLANE WlTH THE AUTOPILOT ENGAGED AS THE AUTOPILOT WILL RUN THE TRlM TO OPPOSE YOUR CONTROL MOVEMENT. A MISTRIM OF THE AIRPLANE. WITH ACCOMPANYING LARGE ELEVATOR AND/OR RUDDER CONTROL FORCES. MAY RESULT IF THE PILOT MANIPULATES THE CONTROLS MANUALLY WHILE THE AUTOPILOT IS ENGAGED.

3. Climb or Descent in Pitch Attitude Hold.

a. Using CWS

1) CWS Button - PRESS and MOVE aircrafl nose to the desired attitude.

2) CWS Button - RELEASE. Autopilot will maintain aircratl pitch attitude up to the pitch limits.

b. Using Vertical Trim

1) VERTICAL TRlM Control - PRESS either up or down to modify aircraft attitude at a rate of 0.75 deglsec up to the pitch limits.

2) VERTICAL TRlM Control - RELEASE when desired aircrafl attitude is reached. The autopilot will maintain the desired pitch attitude.

Issued: June 10, 1994 Revision 6: Dec 6, 1996

Report No: 01 973-001 4-31 1


4. Altitude Hokl

a. ALT Mode Selector Button - PRESS. Note ALT mode annunciator ON Autopilot will maintain the pressure altitude.

NOTE

In accordance with FAA recommendation(ACO0-24B), use 01 basic "PITCH ATTITUDE HOLD' mode is recommended during operation in severe turbulence.

b. Change selected altitudes.

1) Using CWS (Recommended lor altitude changes greater than 100 H.)

a) CWS Button - PRESS and fly aircraH to desired altitude

b) CWS Button - RELEASE when desired altitude is reached. The autopilot will maintain the pressure altitude.

2) Using Vertical Trim (Recommended lor altitude changes less than 100 ft.)

a) VERTICAL TRlM Control - PRESS either up or down. Vertical Trim will seek an altitude rate ol change of about 500 tpm.

b) VERTICAL TRlM Control - RELEASE when desired pressure alMude is reached. The autopilot will maintain Ihe desired pressure all~tude.

Report No: 01973-001 1 4-32

Issued: June 10.1994 Revision 6: Dec 6, 1996


5. Vertical speed and Altitude Preselect Operation

NOTE

The altitude alert annunciator is illuminated 1000 2 50 feet prior to the selected altitude, goes out 200 f 50 feet prior to the selected altitude and illuminates momentarily when the selected altitude is reached. Once the selected altitude is reached the reillumination of the light signifies that the 200 t 50 feet "safe band" has been exceeded and will remain on until 1000 t 50 leet from the selected altitude. An aural tone accompanies the alert annunciator.

a. Altitude Preselect

1) ALTITUDE SELECT (SET) knobs - PUSH small knob to the "IN' position.

VERIFY UNIT IS DISPLAYING ALTITUDE SELECT WINDOW PRIOR TO INITIATING ANY CHANGE IN THE SELECTED ALTITUDE VALUE.

2) ALTITUDE SELECT (SET) knobs - ROTATE lo select and automatically ARM the desired altitude. Note the ARM mode annunciation ON. (The pilot must terminate an existing ALT HOLD mode or glideslope coupling to acquire ALT ARM.)'

3) ALTITUDE SELECT (ARM) button - PUSH to engage or cancel altitude ARM as necessary (Altitude ARM is inhibited in glideslope.). May be used to bring up the flight director i f not already engaged.


Report No: 01 973-00 1 4-33 1


ENGAGING GO AROUND WILL CANCEL ALT ARM. PILOT EFFORT SUCH AS RESELECTING THE DESIRED ALTITUDE, OR PUSHING THE ARM BUTTON, IS REQUIRED TO ARM ALTITUDE HOLD IN GO AROUND.

4) Airplane - ESTABLISH ATTITUDE to intercept the selected altitude.

b. Vertical Speed Select

1) VERTICAL SPEED SELECT (SET) knobs - PULL small knob to the "OUT" position.

VERIFY UNIT IS DISPLAYING VERTICAL SPEED SELECT WINDOW PRIOR TO INITIATING ANY CHANGE IN THE SELECTED VERTICAL SPEED VALUE.

2) VERTICAL SPEED SELECT (SET) knobs - SELECT the desired vertical speed.

3) VERTICAL SPEED (ENG) button - PUSH to engage ~erticrll speed hold mode.

NOTE

The vertical speed display will flash after tive seconds. Push the small knob "IN" to cancel the flashing and restore the altitude display.

Report No: 01 973-001 1 4-34



c. Changing Vertical Speed

1) Using CWS

a) CWS button - PRESS and HOLD.

b) Airplane - Establish desired vertical speed.

C) CWS bullon - RELEASE.

2) Using Vertical Trim Control

VERTICAL TRIM CONTROL - PRESS either up or down lo increase or decrease the vertical speed. Displayed verfical speed changes 100 fpm for every second the control is held down.

WHEN OPERATING AT OR NEAR THE BEST RATE OF CLIMB AIRSPEED AND USING VERTICAL SPEED HOLD, IT IS EASY TO DECELERATE TO AN AIRSPEED ON THE BACK SlDE OF THE POWER CURVE (A DECREASE IN AIRSPEED RESULTS IN A REDUCED RATE OF CLIMB). CONTINUED OPERATION ON THE BACK SlDE OF THE POWER CURVE IN VERTICAL SPEED HOLD MODE WlLL RESULT IN APPROACHING THE STICK SHAKER SPEED AND AN AUTOMATIC AUTOPILOT DISENGAGE.

WHEN OPERATING AT OR NEAR THE MAXIMUM AUTOPILOT SPEED. IT WlLL BE NECESSARY TO REDUCE POWER IN ORDER TO MAINTAIN THE DESIRED RATE OF DESCENT AND NOT EXCEED V&Mm.


Report No: 0 1973-00 t 4-35 1


6. lndicated Airspeed Hold

a. IAS Mode Selector button - PRESS. Note the IAS mode annunciator ON. The autopilot w~ll maintain the current indicated airspeed.

b. Change Selected lndicated Airspeed.

1) Using CWS (recommended for airspeed changes 01 10 KIAS or greater).

a) CWS button - PRESS and fly aircraft to desired airspeed

b) CWS button - RELEASE when desired airspeed is reached. The autopilot will maintain the desired airspeed.

2) Using Vertical Trim (recommended for airspeed changes less than 10 KIAS)

a) VERTICAL TRIM control - PRESS either up or down. Vertical Trim will seek a new airspeed at a rate of about .75 knot per second.

b) VERTICAL TRIM Control - Due to the lag in response, RELEASE approximately 3 to 6 knots prior to reaching the desired airspeed as read on the airplane airspeed indicator. The autopilot will maintain the new stabilized airspeed.

7. Heading Hold

a. Heading Hold

1) Heading Selector Knob - SET BUG to desired heading.

2) HDG Mode Selector Button -.PRESS. Nole HDG mode annunciator ON. Autopilot will automatically lurn the aircraft to the selected heading.

b. Command Turns (Heading Hold mode ON)

1) Heading Selector Knob - MOVE BUG to the desired heading. Autopilot will automatically turn the aircraft to the new selected heading.

1 Report No: 01 973-001 1 4-36

Issued: June 10.1994 Revision 6: Dec 6, 1996 ,


8. Roll Attitude Hold [HOG Mode Off)

a. CWS button - PRESS and hold while manually banking the airplane to any desired angle between 6' and the maximum autopilot limit of 25" t 4".

b. CWS button - RELEASE. Roll attitude hold will maintain the desired bank angle. HOG. NAV and APR mode selections will cancel roll altitude hold.

c. Manual Heading Changes

1) CWS Button - PRESS and MANEUVER aircraft to the desired heading.

2) CWS Button - RELEASE when wings level. Autopilot will maintain aircraft in wings level mode.

NOTE

Aircraft heading may change in the wings level mode due to an aircraft out of trim condition.

9. Reduced Bank Angle

HALF BANK (HE) Mode Button - PRESS. The commanded bank angle will be reduced to 12" +- 5". This mode is functional during HOG and NAV mode operations but will be automatically deselected and inhibited during APR coupled operations.

10. Soft Ride

SOFT RIDE (SR) Mode Button - PRESS. This mode softens the autopilots commands to provide a smoother ride during operations in turbulence.

The normal autopilot performance [maintaining heading, wings level, attitude, airspeed andlor altitude) may be degraded by use of the Soft Ride mode. Soft Ride is automatically deselected and inhibited during APR coupled operations.

~ssued: June 10.1994 Revision 6: Dec 6, 1996

Report No: 01973-001 4-37 /

SECTION 4 =PILATUSE NORMAL PROCEDURES ?c %)I

11. NAV Coupling 1

a. Course Bear~ng Pointer - SET to desired course.

b. Heading Selector Knob - SET BUG to provide desired intercept angle. 1

c. NAV Mode Selector Button - PRESS.

1) If the Course Deviation Bar is greater than 40% to 60% of full scale from center: the aircraft will continue in HOG mode (or wings level il ; HDG not selected) with the NAV ARM annunciated; when the computed capture point is reached HDG will disengage, ARM will ext~nguish and the selected course will be automatically captured and , tracked.

2) If the D-Bar is less than 40% to 60% of full scale from center, the HOG mode will disengage upon selecting NAV mode; the NAV annunciator will illuminate and the captureltrack sequence will automatically begin.

12. Approach (APR) Coupling

a. Course Bearing Pointer - SET lo desired course.

b. Heading Selector Knob - SET BUG to provide desired intercept angle.

c. APR Mode Selector Button - PRESS.

1) I1 the Course Deviation Bar is greater than 40°/0 to 60% of full scale from center: the aircraft will cont~nue in HDG mode (or wings level if HDG not selected) with the APR ARM annunciated; when the computed capture point is reached the HDG will disengage, ARM will extinguish and the selected course will be automatically captured and tracked.

2) If the D-Bar is less than 40% to 60% of full scale from center, the HDG mode will disengage upon selecting APR mode; the APR annunciator will illuminate steady and the captureltrack sequence will automatically begin.

I CAUTION I

RECOMMEND SETTING FLAPS TO 15" PRIOR TO GLIDESLOPE CAPTURE. POOR GLIDESLOPE CAPTURE AND INITIAL TRACKING MAY OCCUR IF FLAPS ARE EXTENDED GREATER THAN 15" PRIOR TO COMPLETION OF GLIDESLOPE CAPTURE.

Reporl No: 01 973-001 I 4-38

Issued: June 10, 1994 Revision 6: Dec 6. 1996 ,


THE AUTOPILOT MUST BE DISENGAGED WHEN THE AIRPLANE IS BELOW 1000 FT AGL IN COUPLED ILS APPROACH MODE, WHEN THE AIRPLANE IS NOT' EQUIPPED WITH A FUNCTIONING PILATUS OPTION RADAR ALTIMETER.

13. BC Approach Coupling

a. Course Bearing Pointer - SET to the ILS front course inbound heading.

b. Heading Selector Knob - SET BUG to provide desired intercept angle.

c. APR Mode Selector Button - PRESS. APR to aulomatically activate BC when the airplane heading differs from the front course bearing set in 13.a. above by more than 105" (refer to EFlS BC annunciation).

1) If the Course Deviation Bar is greater than 40% to 60% of full scale from center, the aircraft will continue in HOG mode (or wings level i f HOG not selected) with APR ARM BC annunciated; when the computed capture point is reached HOG will disengage. ARM will extinguish and the selected course will be automatically captured and tracked.

2) I f the 0-Bar is less than 40% to 60% of full scale from center, the HDG mode will disengage upon selecting APR mode; the APR BC annunciators will illuminate and the captureltrack sequence will automatically begin.


Report No: 01973-001 4-39 (


14. Glideslope Coupling

NOTE

Glideslope couplrng is rnhibited when operating in NAV or APR BC modes. Glideslope coupling occurs automatically in the APR mode.

a. APR Mode - ENGAGED

NOTE

Autopilot can capture glideslope from below or slightly above the beam wh~le operating in either PITCH ATTITUDE HOLD. IAS HOLD. VS HOLD, or ALT HOLD modes. Capture is better assured lrom below the beam.

b. Near glideslope centering - NOTE GS annunciator ON.

NOTE

It after glideslope coupling the glideslope signal becomes inadequate (GS llag in view), the GS annunciator will flash at least six times before extinguishing and the syslem will transfer to PITCH ATTITUDE HOLD. If a valid glideslope signal returns within six seconds, the system will automatically recouple. II a valrd glideslope signal does not return within six seconds, the autopilot must once again capture the glideslope beam to achieve glideslope coupling.

Report No: 01973-001 ( 4-40

Issued: June 10. 1994 Revision 6: DeC 6. 1996


15. Missed Approach

a. Power Lever GA Switch - PRESS to engage the go-around mode. This will provide a wings level, pitch up command. Note GA mode annunciator OW. Forward PCL to the required setting.

b. Activate AlP modes as desired.

r CAUTION I

ENGAGING GO AROUND WILL CANCEL ALL MODES. PILOT EFFORT IS REQUIRED TO ACTIVATE AJP MODES AS DESIRED IN GO AROUND.

C. MISSED APPROACH - EXECUTE.

d. Lateral Guidance (Select one)

1) Heading Mode - SET BUG and PRESS HOG.

2) NAV Mode - PRESS NAV.

3) APR Mode - PRESS APR. Glideslope coupling will be inhibited so that the LOC can be tracked outbound and the autopilot will not couple to false glideslope signals.

16. Before Landing

Autopilot Disengage Switch - PRESS to disengage AP and YO.

Issued; June 10,1994 Revision 6: Dec 6, 1996

Report No: 01973-001 4-41 (

- -


CPLATUS W PC XI)

I 4.20.2 FLIGHT DIRECTOR OPERATION

I

NOTE

The Flight Director modes of operation are the same as those used for autopilot operations except the autopilot is not engaged and the pilot must maneuver the aircraft to satisfy the Flight Director commands.

No aulotrim lunctions are provided by the autopilot.

4.21 NOISE LEVEL

The noise levels slated below have been verified and approved by FOCA in noise level test llights conducted in accordance wtth ICAO Annex 16. Chapter 10 and Swiss VEL. This airplane model is in compliance with all ICAO Annex 16 and Swiss VEL noise standards applicable to this type.

ICAO Annex 16, Chapter 10 76.7 dB(A)

Swiss VEL 73.2 dB(A)

The noise level stated below has been verified and approved by the FAA in noise level test tlrghts conducted in accordance with FAR Part 36, Appendix G. This airplane model is in compliance with FAR 36 noise standards applicable to this type.

FAR Part 36, Appendix G 73.2 dB(A)

No determinalion has been made by the Certifying Authority that the noise levels of this airplane are or should be acceptable or unacceptable for operation at, into, or out of, any airport.

Report NO: 01 973.001 4-42



AMPLIFIED PROCEDURES

C

4.22 CROSSWIND OPERATION

The maximum demonstrated crosswind for takeoff and landing for all flap configurations is shown in para 4.2.

ON RUNWAYS WITH POOLS OF STANDING WATER AND/OR POOR BRAKING ACTION IT MAY NOT BE POSSIBLE TO MAINTAIN THE CENTERLINE AND/OR THE CORRECT ALIGNMENT OF THE AIRCRAFT ON THE RUNWAY IN CONDITIONS OF STRONG CROSSWIND.

For further information on crosswind operation refer to Section 10.


Report No: 01973-001 4-43


4.23 FLIGHT IN ICING CONDITIONS

Icing conditions can exist when:

The Outside Aa Temperature (OAT) on the ground and for takeoff, or Total Air Temperature (TAT) in flight, is 10°C or colder, and visible moisture in any form is present (such as clouds, fog or mist with visibility of one mile or less, rain snow, sleet and ice crystals).

The OAT on the ground and for takeofl is 10°C or colder when operating on ramps, taxiways or runways, where surface snow, ice, standing water, or slush may be ingested by the engine, or freeze on the engine, or the engine nacelle.

There are visible signs of ice accretion on the aircraft.

Severe icing may result from environmental conditions during flight in freezing rain, freezing drizzle, or mixed icing conditions (supercooled liquid water and ice crystals) which may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming all of the protected surfaces.

Information on the removal of snow, ice and frost from the aircrafl is provided in Section 10.

Freezing rain, freezing fog. freezing drizzle and mixed conditions and descent into icing clouds from above freezing temperatures can result in excessive accretion of ice on the protected surfaces. They may also result in runback ice forming beyond the protected surfaces over a large percentage of the chordwise extent of the lifting surfaces. This ice cannot be shed and it may seriously degrade performance and control of the aircraft.

Flight In severe icing conditions should be avoided, as this may exceed the capabilities ot the a~rcraft ice protection systems. Severe icing conditions can be identified by excessive Ice accretion on Ihe visible pans of the airframe including the prote~ted surfaces. This might affect the aircraft performance and handling qualities, and cause significant loss in powerplant performance. If this occurs request priirity assistance from ATC to facilitate a route or an altitude change to exit the icing conditions.

Operation on deep slush or snow covered runways greater than 1 inch (2.5 cm) may result in contamination of the flap drive mechanism resulting in failure to retract. I f possible operation on deep slush and snow compacted runways should be avoided.

For takeon and landing on runways covered with surface snow, ice, standing water, or slush, the inertial separator must be open.

For flight in heavy precipitation the inertial separator must be open.

Detection of icing conditions and ice accretion on the aircraft is by pilot visual identification on the left hand wing leading edge. A wing inspection light is provided for night time operations.

Prior lo entering icing conditions, activate all ice protection systems as required.

If not already activated, select all systems as required, immediately icing conditions are identified.

The procedures lor selection ol the ice protection systems are provided in Section 4.

Report No: 01973-001 4-44

~ssued: June 10,1994 Revision 8: September 1, 1998


During all icing encounters or times with visible ice accretion on any part of the airframe fhe flaps must not be extended beyond certain limits. These limits eliminate the possibility of tailplane stall which results in an uncontrolled aircraft pitch down moment.

- With operational airframe pneumatic deice boots 15" flap - After failure of the airframe pneumatic deice boots 0' flap.

The minimum recommended speeds for icing encounters and with residual ice on the airframe are :-

- Climb. Flaps 0". Pusher Ice Mode = 125 KlAS - Holding Pattern, Flaps 0' = 140 KlAS to 170 KlAS - Landing Approach. Flaps 15", Pusher Ice Mode = 105 KlAS - Landing Approach, Flaps 0°, Boot Failure

Pusher Ice Mode = 130 KlAS - Balked Landing (Go-Around) Flaps 15" Pusher Ice Mode 5 105 KlAS - Balked Landing (Go-Around) Boot failure Flaps 0° Pusher Ice Mode = 130 KlAS

Flight in icing conditions is only permitted with full operational status of all aircrafl deicing systems. This includes :-

- Propeller Deice - Wing and Horizontal Tail Deice Boots - Inertial Separator - Windshield Deice - Probes Deice - Stick Pusher Ice Mode

The propeller de-ice is aclivaled from the DE-ICING switch panel by the switch labeled PROP being moved 10 ON. When activated and operating correctly, the green CAWS caption labeled PROP DE ICE will be continuously illuminated. In this mode the propeller de-ice system will be automatically selected to the correct cycle with reference to outside air temperature. No further aircrew input is required. If a system failure occurs when activated, the green CAWS caption will blink and the amber CAWS caption labeled DE ICE will be cor~tinuously illuminated. An aural gong will sound.

The wing and horizontal tail de-ice boots are activated from the DE-ICING switch panel by the switch labeled BOOTS being moved to either 3MlN or IMIN. 3MIN is to be selected in icing conditions with moderate ice accretion rates as judged by the aircrew. 1MlN is to be selected in icing conditions with high ice accretion rates. When activated in either 1MIN or 3MIN mode and operating correctly, the green CAWS caption labeled DE ICE BOOTS will be continuously illuminated. H a system failure occurs when activated, the green CAWS caption will blink and the amber CAWS caption labeled DE ICE will be continuously illuminated. An aural gong will sound.

The engine inertial separator is activated to its open (icing encounter) position lrom the DE ICING switch panel by the switch labeled INERT SEP being moved to OPEN. Once activated the inertial separator door will reach its lully open position in approximately 30 seconds. When fully open the green CAWS caption labeled INERT SEP will be continuously illuminated. If the door does not reach its lully open position


Report No: 01973-001 4-45 1


or moves away lrom its fully open position when still selected, the green CAWS caption will be extinguished.

The LH slde and RH side windshield deice is activated from the DE ICING swlch panel by two switches labeled LH WSHLD and RH WSHLD respectively, being moved to either LIGHT or HEAVY depending on the severity of the icing encounter. When activated in either mode and operating correctly, the green CAWS caption labeled WSHLD HEAT will be continuously illuminated. If a system failure occurs when activated, the green CAWS caption will extinguish.

Deicing ol all probes. AOA (vane and mounting plate), pitot and static, is activated from Ihe DE ICING switch panel by a switch labeled PROBES being moved to ON. When aclivated and operating correctly, the green CAWS caption labeled PROBES DE ICE will be continuously illuminated. If deicing of the pitot andlor static probes fails when selected, then the green CAWS caption will be extinguished. If deicing of the AOA probes fails when selected, then the amber CAWS caption labeled AOA DE ICE will be continuously illuminated. An aural gong will sound.

When the propeller de-ice is selected to ON and the inertial separator selected to OPEN, the stall protection system, slick pusherlshaker system is re-datumed to provide both shake and push functions at lower angles of attack and higher speeds. This is to protecl against the natural stall through the affects of residual ice on the protected surfaces ol the airfoil leading edges. When the system is in the re-datum mode, lhe aircrew are alerted by illurninalion of the amber CAWS caption labeled PUSHER ICE MODE. An aural gong will sound. Failure of the system in ice mode will result in the caption being extinguished and the AMBER CAWS caption labeled PUSHER w~ll be continuously illuminated. An aural gong will sound.

Night time flight in icing conditions is only authorized with full operational StatUS of all the aircraft de-icing systems above, plus the wing inspection light.

The wing inspection light is activated from the EXTERNAL LIGHTS switch panel by the switch labeled WING being moved to ON. No functional or failure indications are provided.

A full description of all of the de-ice systems, their switch terminology and caution and warning logic is provided in Section 7.

The probes de-ice should be selected to on, prior to, and during all flights.

During the icing encounter the pneumatic de-ice boots will operate continuously in elther 3min or lmin cycle mode as selected by the aircrew. During this time the aircrew should frequently monitor the continual shedding of ice lrom the wing leading edge and the airframe for ice accretion on all visible surfaces that could affect aircraft controllab~lity. It should be noted that some residual ice will be maintained on the wing leading edge during cycling of the boots.

During the icing encounter continue lo monitor the CAWS for correct function of the ice protection systems.

During flight in icing conditions the aircraft may be subject to a slight degradation in aircratl performance and engine performance. This may be recognized by a required increase in engine power to maintain a constant indicated airspeed and an increased engine IT1 to maintain a constant power respectively. If failure of any of the ice protection systems occurs this degradation may become more severe. ARer such

Report No: 01973-001 1 1 4-46

Issued: June 10,1994 Revision 8: September t , 1998


failure the pilot should make immediate arrangements for departure of king condiiions as soon as practicable. If required ATC priority assistance should be requested.

The emergency procedures, concerning failure of the ke protection systems during flight in king condiiions, are provided in Section 3.

On departure from king condiiions the inertial separator (INERT SEP) and the propeller deke system (PROP DE ICE) should be kept OPEN and ON respectively until all visible and unprotected areas of the aircraft are observed as being free of ice. This protects the engine from possiMe k e ingestion and maintains the stick shakerlpusher computer in PUSHER ICE MODE therefore protecting the aircraft agalnst the onset of natural stab. The flaps are not to be extended beyond 15" or in the case of deice boot failure, left at 0'. H the flaps are in an extended position, do not retract them until the airframe is clear ot ke.

On departure of king conditions the deice boots are to be selected OFF and the windshield heat is to be selected as required for good visibility, irrespective of the presence of residual ke.

Once all visible protected and unprotected areas are observed as being free of ice then the inertial separator and the propeller deice system can be selected CLOSED and OFF respectively. This will return the stick shakerlpusher computer to its normal mode. The flaps can be extended or retracted to any required position.

When performing a landing apprbach after an king encounter and with residual ice on I the airframe the minimum landing speedS defined above should be observed. This will prevent stick shaker activation AIM PUSHER ICE MODE.

When performing a landing approach after an king encounter and with residual ice on the airframe the flap limitations defined above must be observed. I In case of a balked landing go around after an king encounter, the climb speeds defined above should be maintained. This will prevent stick shaker activation in PUSHER ICE MODE.

In case of a balked landing go around after an king encounter, the flap position should not be changed and should be maintained at the approach position.

The landing gear can be retracted but a locked indication may not be achieved due to ice contamination of the up position switch striker.

Use of ICE X (B.F. Goodrich Brand Name) improves the shedding capabili of the pneumatic de-ice boots. Its use (see Aircraft Maintenance Manual) is recommended but not mandatory.


Report No: 01 973-001 4-47


System I Setting

I

BOOTS 1 lMlN

LIGHT

PROP

ON with MODE INERT

ON

SEP to I OPEN

An aural gong sounc

CAWS STATUS

System 011

System Functional

DE ICE BOOTS GREEN continuous

4

PROP DE ICE GREEN continuous

INERT SEP GREEN continuous

WSHLD HEAT GREEN continuous

PROBES DE ICE GREEN continuous with AOA DE ICE AMBER off

I

PUSHER ICE I MODE AMBER* continuous

D when the amber caption

DE ICE BOOTS GREEN off

System Failed

DE ICE BOOTS GREEN blinking with DE ICE AMBER' continuous

PROP DE ICE GREEN blinking with DE ICE AMBER' continuous

INERT SEP GREEN off

WSHLD HEAT GREEN off

(Pitot-Static Heating System Failure) PROBES DE ICE GREEN off

(AOA Heating System Failure) PROBES DE ICE GREEN continuous with AOA DE ICE AMBER' continuous (3 minutes delay)

PUSHER ICE MODE AMBER ofi with PUSHER AMBER' continuous

comes on

PROP DE ICE GREEN off

INERT SEP GREEN off

WSHLD HEAT GREEN off

PROBES DE ICE GREEN off with AOA M ICE AMBER' continuous (3 minutes atter PROBES switch set to OFF)

PUSHER ICE MODE AMBER off

Figure 4-3. Ice Protection Systems CAWS Advisoly

Report No: 01973-001 1 4-48



4.24 SEVERE ICING CONDITIONS

Severe king may result from environmental condiions outside of those for whkh the airplane is certlicated. Flight in freezing rain, freezing drizzle, or mixed king condiiions (supercooled liquid water and k e crystals) may result in ice buikl-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming aft of the protected surfaces. This k e may not be shed using the ke protection systems, and may seriously degrade the performance and controllability of the airplane.

The following weather conditions may be conductive to severe in-flight king:

visible rain at temperatures below 0 degrees Celsius ambient air temperature

droplets that splash or splatter on impact at temperatures below 0 degrees Celsius ambient air temperature

The following procedures are for exiting the severe king environment and are applicable to all flight phases from takeoff to landing. Monitor the ambient air temperature. While severe king may form at temperatures as cold as -18 degrees Celsius, increased vigilance is warranted at temperatures around freezing with visible moisture present. If the visual cues speclied in Section 2 for identifying severe icing conditions are obsewed, accomplish the following:

report the weather conditions to Air Traffic Control

immediately request priority handling from Air Traffic Control to facilitate a route or an altitude chanae to exit the severe kina conditions in order to avoid extended exposure to fliiht condiiions more sev&e than those for which the airplane has been certificated

4.25 PREVENTING OF FROZENILOCKED BRAKES

Frozen brakes could occur when fully saturated with water and operated at or below +41° F (+5" C) OAT. The following procedures will help to prevent frozen brakes:

PRE FLIGHT INSPECTION

Apply approximately two ounces (59 cc) of undiluted isopropyl alcohol to the outer diameter of the brake assembly with a spray bottle

TAXI BEFORE TAKEOFF

If operating conditions permit use the brakes to slow down the aircraft and minimize the use of propeller reverse in order to drive off excessive moisture

TAKEOFF

Do not operate the brakes after takaon Avoid prolonged gear down flight in freezing conditions


Reporl No: 01973-001 4-49


LANMNG

Do not operate the brakes before landing If operating conditions permit avoid prolonged gear down flight in freezing conditions If operating conditions permit use the brakes to slow down the aircraft and minimize the use of propeller reverse in order to drive off excessive moisture

TAXI AFTER LANDING

If operating conditions permit use the brakes to slow down the aircraft and minimize the use of propeller reverse in order to drive off excessive moisture

AIRCRAFT WASHING

Install wheeVtire covers to minimize water ingress

URCRAFT PARKING

When the aircraft is parked, apply the parking brake to help keep moisture out of the brake assembly

If the aircraft is to be parked for a long p e r i i of time with temperatures below 32" F (0 " C), apply approximately two ounces (59 cc) of undiluted isopropyl alcohol to the outer diameter of the brake assembly with a spray bottle. Apply the alcohol after the brake assembly has cooled enough to touch but before it is cold

r CAUTION . I

DO NOT USE DEFROST AGENT ETHYLENE GLYCOL ON CARBON BRAKES, THIS CAN HAVE AN ADVERSE EFFECT ON VIBRATION, OXIDATION AND WEAR OF THE CARBON.

Report No: 01973-001 4-50


b

I PILOT'S OPERATING I HANDBOOK

Performance

SECTION 5 PERFORMANCE

SECTION 5

PERFORMANCE

TABLE OF CONTENTS

Subject

GENERAL

STANDARD TABLES

FIG. 5-1, FAHRENHEIT TO CELSIUS TEMPERATURE CONVERSION FIG. 5-2, INDICATED OAT CORRECTION FIG. 5-3. ISA CONVERSION FIG. 5-4. U.S. GALLONS TO LITERS CONVERSION FIG. 5-5. FEET TO METERS CONVERSION FIG. 5-6, POUNDS TO KILOGRAMS CONVERSION FIG. 5-7, INCHES TO MILLIMETERS CONVERSION FIG. 5-8. WIND COMPONENTS

AIRSPEED CALIBRATION

FIG. 5-9, AIRSPEED CALIBRATION - FLAPS RETRACTED FIG. 5-10, AIRSPEED CALIBRATION - FLAPS EXTENDED

ALTIMETER CORRECTION

FIG. 5-1 1, ALTIMETER CORRECTION

STALLSPEED

FIG. 5-12. STALL SPEEDS KlAS - FLIGHT IDLE POWER (STANDARD UNITS) FIG. 5-13, STALL SPEEDS KlAS - FLIGHT IDLE POWER (METRIC UNITS) FIG. 5-14, STALL SPEEDS KCAS - FLIGHT IDLE POWER (STANDARD UNITS) FIG. 5-15, STALL SPEEDS KCAS - FLIGHT IDLE POWER (METRIC UNITS)

Page

5-1


Report No: 01973-001 5-1


Subject Page

TAKEOFF PERFORMANCE

FIG. 5-16, STATIC TAKEOFF TORQUE 5-17 FIG. 5-17, ACCELERATE-STOP DISTANCE - FLAPS 15" (STANDARD UNITS) 5-18 FIG. 5-18, ACCELERATE-STOP DISTANCE - FLAPS 15" (METRIC UNITS) 5-19 FIG. 5-19, TAKEOFF GROUND ROLL - FLAPS 15" (STANDARD UNITS) 5-20 FIG. 5-20, TAKEOFF GROUND ROLL - FLAPS 15" (METRIC UNITS) 5-21 FIG. 5-21, TAKEOFF TOTAL DISTANCE - FLAPS 15" (STANDARD UNITS) 5-22 FIG. 5-22. TAKEOFF TOTAL DISTANCE - FLAPS 15" (METRIC UNITS) 5-23

CLlMB PERFORMANCE

FIG. 5-23, MAXIMUM CLlMB TORQUE FIG. 5-24, MAXIMUM RATE OF CLlMB - FLAPS 15" (STANDARD UNITS) FIG. 5-25, MAXIMUM RATE OF CLlMB - FLAPS 15" (METRIC UNITS) FIG. 5-26. MAXIMUM RATE OF CLlMB - FLAPS 0" (STANDARD UNITS) FIG. 5-27, MAXIMUM RATE OF CLlMB - FLAPS 0' (METRIC UNITS) FIG. 5-28, CRUISE CLlMB AIRSPEED SCHEDULE FIG. 5-29, RATE OF CLlMB - CRUISE CLlMB (STANDARD UNITS) FIG. 5-30, RATE OF CLlMB - CRUISE CLlMB (METRIC UNITS) FIG. 5-31, TIME TO CLlMB - CRUISE CLlMB (STANDARD UNITS) FIG. 5-32, TlME TO CLlMB - CRUISE CLlMB (METRIC UNITS) FIG. 5-33. FUEL USED TO CLIMB - CRUISE CLlMB (STANDARD UNITS) FIG. 5-34, FUEL USED TO CLlMB - CRUISE CLlMB (METRIC UNITS) FIG. 5-35, DISTANCE TO CLlMB - CRUISE CLIMB (STANDARD UNITS) FIG. 5-36, DISTANCE TO CLlMB - CRUISE CLlMB (METRIC UNITS)

CRUISE PERFORMANCE

FIG. 5-37, MAXIMUM CRUISE POWER FIG. 5-38, LONG RANGE CRUISE FIG. 5-39. MAX ENDURANCE CRUISE FIG. 5-40, SPECIFIC AIR RANGE (7000 LB) FIG. 5-41, SPECIFIC AIR RANGE (8000 LB) FIG. 5-42, SPECIFIC AIR RANGE (9000 LB) FIG. 5-43, HOLDING TlME AND FUEL

DESCENT PERFORMANCE

FIG. 5-44, TlME TO DESCEND FIG 5-45, FUEL USED TO DESCEND (STANDARD UNITS) FIG. 5-46. FUEL USED TO DESCEND (METRIC UNITS) FIG. 5-47, DISTANCE TO DESCEND FIG. 5-48, POWER OFF GLIDE TlME (STANDARD UNITS) FIG. 5-49, POWER OFF GLIDE TlME (METRIC UNITS) FIG. 5-50. POWER OFF GLIDE DISTANCE

Report No: 01 973-001 5-ii



Subject Page

BALKED LANDING CLIMB PERFORMANCE

FIG. 5-51. BALKED LANDING TORQUE 5-67 FIG. 5-52. RATE OF CLIMB - BALKED LANDING (STANDARD UNITS) 5-68 FIG. 5-53. RATE OF CLIMB - BALKED LANDING (METRIC UNITS) 5-69

LANDING PERFORMANCE

FIG. 5-54, LANDING TOTAL DISTANCE - FLAPS 40" (STANDARD UNITS) 5-70 FIG. 5-55, LANDING TOTAL DISTANCE - FLAPS 40" (METRIC UNITS) 5-71 FIG. 5-56. LANDING GROUND ROLL - FLAPS 40" (STANDARD UNITS) 5-72 FIG. 5-57. LANDING GROUND ROLL - FLAPS 40" (METRIC UNITS) 5-73 FIG. 5-58. LANDING TOTAL DISTANCEWITH THE USE OF REVERSE 5-74 THRUST - FLAPS 40" (STANDARD UNITS) FIG. 5-59. LANDING TOTAL DISTANCE WITH THE USE OF REVERSE 5-75 THRUST - FLAPS 40' (METRIC UNITS) FIG. 5-60. LANDING GROUND ROLL WlTH THE USE OF REVERSE THRUST 5-76 - FLAPS 40" (STANDARD UNITS) FIG. 5-61, LANDING GROUND ROLL WlTH THE USE OF REVERSE THRUST 5-77 - FLAPS 40" (METRIC UNITS)

FLIGHT IN ICING CONDITIONS 5-78

STALL SPEEDS ENGINE TORQUE TAKEOFF PERFORMANCE CLIMB PERFORMANCE HOLDING ENDURANCE LANDING PERFORMANCE BALKED LANDING PERFORMANCE

FLIGHT PLANNING EXAMPLE 5-80 I

Issued: February 14, 1994 Revision 6: Dec 6, 1996

Report No: 01973-001 5-iii

SECTION 5 =PILATUS PERFORMANCE PC XI1 I


Report NO: 01 973-001 5-iv

Issued: February 14,1994 Revision 2: Februaly 14, 1995


GENERAL

This section contains all of the required and complimentary performance data lor airplane operation. Airplane performance associated with optional equipment and systems which require supplemenls is provided in Section 9. Supplements.

The performance information presented in this section is derived from actual flighl test data corrected to standard day conditions and analytically expanded for the different parameters such as weight, altitude. and temperature, etc. This information does not account lor many factors that the pilot must evaluate before each takeon such as pilot proficiency, aircraft condition, runway surface and slope other than that spec~fied, or the effect of winds aloft. When necessary, a performance chart (table) will specify the aircraft configuration and the procedure to achieve the published performance.

A Flight Planning Example is provided to assist the pilot in the preflight performance calculations as required by the operating regulations. Each performance chart (table) has an example plotted to indicate the proper sequence in which to use the chart and determ~ne accurate performance data.

All performance data is limited to between the -55" C (-67" F) and +50° C (122" F) outside air lemperature limits. Some tables presented in !his section show data for lemperatures below -55" C (-67" F) which is purely lor ease of interpolation between data points. These temperature areas in the tables are shaded.

Performance data regarding takeoff, landing and accelerate-stop distances is presented up to 10.000 It. This does not, however, imply an operational limitat~on ol the aircraft. F~eld performance data at higher altitudes can be supplied under special request.

The slall speeds shown in the performance charts are achieved at an entry rate of t knotlsecond. Maximum altitude loss observed during Ihe stall was 300 feet. Durtng an accelerated stall, a rapid pitch-down in excess of 30" may result w~lh an allitude loss of up to 500 feet.

By setting the climb torque as defined in this section, the recommended ITT of 720°C (see Section 4 Climb) could be exceeded.

When landing with llaps set to less than 40". the total landing distances will be increased by the following factors:

I Issued: February 14. 1994 Revision 7: July t . I 997

b

FLAP SETTING FACTOR

Report No: 0 1973-00 1 5- 1

SECTION 5 PP lLATUSl PERFORMANCE ?c %I) 1

1

FAHRENHEIT TO CELSIUS CONVERSION 1

I Figure 5-1. Fahrenheit to Celsius Conversion

Report No: 01 973-00 1

I 1 5.2 Issued: February 14, 1994

Revision 2: February 14, 1995


INDICATED OAT CORRECTION SUBTRACT EMPERATURE CORRECTION FROM INDICATED OAT

S T ~ A U D DAY (ISA) ASSUMES A PROBE RECOVERY FACTOR OF 60% ASSUMES ZERO INSTRUMENT ERRORS

URE CORRECTION 5 'C

IAS - INDICATED AIRSPEED - KNOTS

Figure 5-2. Indicated OAT Correction I Issued: Feb~ary 14,1994 Re~iSion 2: Feb~ary 14, 1995

Report No. 01973-001 5-3 I


ISA TEMPERATURE CONVERSION I

80 1

Report No: 01 973-001 / 5-1

OUTSIDE AIR TEMPERATURE - 'C

. . . , . . . . . . . . I . , . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . a

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

. , ,

Figure 5-3. ISA Conversion

. . . . . I , . , . . . . . . . . . . . .

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

Issued: February 14,1994 Revision 2: Feb~ary 14,1995

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

EXAMPLE. .

OUTSIDE AIR TEMPERATURE -12 'C PRESSURE ALTITUDE - 25.000 FT DIFFERENCE FROM ISA 23 'c -

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

. . . . . . . . . . . . . . . . . . . . . . . . . I , . , . , , a

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

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

. . . ( . . . . , . . . . . . . . . . . . . . . . . . . . , .

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

I , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I . . . . . . . . . . . . . . . . . . . . . . . . . .

. . ,


US GALLONS TO LITERS CONVERSION

LKXllD MEASURE - US (UUONS

Figure 54. U.S. Gallons to Liters Conversion

Issued. Feb~ary 14, 1994 Revislon 2: February 14,1995

Report No: 01973-001 5-5 I


I m , j

FEET TO METERS CONVERSION

DISTANCE - FEET

I Report No: 01973-001 1 5-6

Figure 5-5. Feet to Meten Converjon



POUNDS TO KILOGRAMS CONVERSION

Figure 56. Pounds lo Kilograms Conversion I Issued: February 14,1994 Revidon 2: February 14,1995


INCHES TO MILLIMETERS CONVERSION

DISTANCE - INCHES

I Figure 5-7. Inches to Millimeters Conversion

Report No: 01 973-00 1 15-11

Issued: FebNary 14,1994 Revision 2: February 14,1995


TAKEOFF AND LANDING CROSSWIND COMPONENT

WIND SPEED 31 KNOTS ANOLE BFTWEEN WIND AM) FLIGHT PATH 34'

R M PATH HEADWIND COMPOMNT 25.7 KNOTS N A Y CENTRELINE WMKNENT -- 17.3 KNOTS

0 10 20 30 40 50

CRO6WW COMPONENT - KNOTS

Figure 5-8. Wind Components I I

Issued: February 14, 1994 Revision 2: February 14,1995

Report No: 01973-001 5-9 I


AIRSPEED CALIBRATION - FLAPS RETRACTED EXAMPLE:

INDICATED AIRSPEEDS ASSUME IAS - INDICATED AIRSPEED 190 KNOTS ZERO INSTRUMENT ERROR CAS - CALIBRATED AIRSPEED 194 KNOTS

g, & 1, 1, 1, 160 180 200 no 240 260

IAS - INDICATED AIRSPEED - KNOTS

Figure 5-9. Airspeed Calibration - Flaps Retracted

Report No: 01 973-001 5-10


SECTION 5

AIRSPEED CALIBRATION - FLAPS EXTENDED EXAMPLE: IAS - INDICATED AIRSPEED 91 KNOTS

INDICATED AIRSPEEDS ASSUME FLAPS 15' ZERO INSTRUMENT ERROR CAS - CALIBRATED AIRSPEED 93 KNOTS

60 80 100 120 140 160 180

IAS - INDICATED AIRSPEED - KNOfS

Figure 5-10. Airspeed Calibration - Flaps Extended


Report No: 01973-001 5-1 1


mPILATUSJ PC XII

ALTIMETER CORRECTION ADD ALTIMETER CORRECTION TO INDICATED ALTITUDE TO OBTAIN CORRECTED ALTITUDE ALTIMETER ERROR SHOWN IS FOR FLAPS RETRACTED WITH FLAPS EXTENDED THE ERROR NEVER EXCEEDS 20 FT

60 80 100 120 140 160 180 200 220 240 260 t

US - INDKXTED ARSPEED - KNOTS 4

Figure 5-1 1. Altimeter C0f~t3ction

Repofl No: 01973-001 5-12

Issued: Febmary 14,1994 Revision 2: February 14,1995

STALL SPEEDS - FLIGHT IDLE POWER (STANDARD UNITS)

EXAMPLE: MIGHT 7200 LB

NOTES: RAPS - 40' STW IS WINED BY PUSHER ACTIVATION ANGLE OF BANK 45 DEG LANDING GEAR POSITION HAS NO EFFECT STALL SPEED 63 KlAS

W E l W - LB ANGLE OF BANK - DEG

STALL SPEEDS - FLIGHT IDLE POWER (METRIC UNITS)

EXAMPLE: WEIGHT 3250 KG

NOTES: FLAPS - 40' STAU IS DEFINED BY PUSHER ACTIVATION ANGLE OF BANK 45 DEG LANDING GEAR POSITION HAS NO EFFECT STALL SPEED 63 KlAS

WEIGKT - KG ANGLE OF BAN( - DEG

STALL SPEEDS - FLIGHT IDLE POWER (STANDARD UNITS)

U(AMPLE:

WEIGHT 7200 LB NOTES: RAPS - 40' STALL IS DEFINED BY PUSHER ACTIVATION ANGLE OF BANK 45 DEG LANDING GEAR POSITION HAS NO EFFECT STAU SPEED 65 KCAS

loo00 9000 8000 7000 6000 0 20 40 60

WIGHT - LB ANGLE OF BAN( - DEG


EXAMPLE: W E l M 3250 KG

NOTES: RAPS - 40' S T W IS DEFINED BY PUSHER ACTlVATlON ANGLE OF BANK 45 DEG LANDING GEAR POSITION HAS NO EFFECT STAU. SPEED 65 KCAS

4500 4000 3500 3000 2500 0 20 40 60

W E l W - KG ANGLE OF BANK - DEG


L STATIC TAKEOFF TORQUE

TORQUE WILL INCREASE WlTH INCRfiAWG AIRSPEED

PROPELLER SPEED 1700 RPM I N E W SEPARATOR CLOSE0 MAXIMUM TOAWE REDUCTION WITH EXAMPLE:

MERTUL SEPARATOR WEN : ALTlTUM m R - 1.2 PSI IN NON ICINQ CONbmONS OAT 28 'C - 2.1 PSI IN IClNQ CONDmONS ENGINE MRQUE 41 PSI

-40 -20 0 20 40 60

OVlSlM AIR TEMPERATURE - 'C

Figure 5-16. Static Takeoff Torque

Report No: 01 973-001 5-17

Issued: February 14, 1994 Revision 7: July 1, 1997

SECTIO

N 5

PERFO

RM

AN

CE

See FLIGH

T IN IC

ING

CO

ND

ITION

S para for info on effect of icing

Figure 5-17. Accelerate - Stop D

istance - Flaps 15" (standard units) Report N

o: 01 973-001 5-18

Issued: February 14, 1994 R

evision 10: September 1,2000

ACCELERATE-STOP DISTANCE - FLAPS 15' (METRIC UNITS)

AssamW C(MmONS: PWER4+W AT 1.1 Vsq m L N E R A T O R O U H ) t O L E FmD SURFACE: TAW42

SECTIO

N 5

PERFO

RM

AN

CE

See FLIGH

T IN IC

lNO

CO

NO

lTlON

S para for Info on effect of lclng

Figure 5-20. Takeoff Ground R

oll - Flaps 15" (metric units)

Issued: February 14,1994 R

evision 10: September 1,2000

Report N

o: 01 973-001 5-21

SECTIO

N 5

PERFO

RM

AN

CE

I See FLIG

HT IN

lClN

G C

ON

DITIO

NS

para for info on effect of icing

Figure 5-21. Takeoff Total Distance - Flaps 15" (standard units)

Repon No: 01973-001

5-22 Issued: February 14, 1994

Revision 10: Septem

ber 1,2000

TAKEOFF TOTAL DISTANCE - FLAPS 15' OVER 15 M OBSTACLE; (MEIRIC UNTTS)

O(AMSE:

M T ~ ~ D E Boaom ASWXTED COMmoN8: 'w - KG Vm - KIAS V= - KUS OAT 18 'C UFT OR AT 1.1 Vs, 2800 63 78 WWWT 35W KG OBSTACLE AT 1.3 Vs, 3300 67 03 HEADWPOCa)rPOF(MT 8 KT REFER TO THE SPECD SCHEWLE TABLE 3700 71 88 UPHUCQWNEM 1 % FBD SURACE: TARMkC 4100 75 93 TMEWFTOTALMSTANCE 7 W M


MAXIMUM CLIMB TORQUE PROPELLER SPEED 1700 RPM INERTIAL SEPARATOR CLOSED MAXIMUM TOAQUE REDUCTK)N WITH EXAMPLE:

INERTIA SEPARATm WEN : ALTITUDE 5000 f=T - 1.2 PSI IN NON ICING ~ ~ O N S OAT 29 'c - 2.1 PSI IN ICING IONS ENGINE TORQUE 33 PSI

Figure 5-23. Maximum Climb Torque

Report No: 01 973-001 5-24

Issued: February 14,1994 Revision 5: May 10,1996

SE

CTIO

N 5

PE

RFO

RM

AN

CE

I Figure 5-24. M

aximum

Rate of C

limb - F

laps 15" (standard units) I

Issued: February 14. 1994

I R

evision 7: July 1. 1997 R

eport No: 0 1973-00 1

5-25

SE

CTIO

N 5

SP

ILA

TU

Sg

P

ER

FOR

MA

NC

E

-?c %I1

Figure 5-25. Maxim

um R

ate of Cllm

b - Flaps 15" (metric units)



Revision 7: July 1. 1997

=PILATUS* S

EC

TION

5 1

?C %

I1 P

ER

FOR

MA

NC

E

I

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for Info. on effect of lclng

I Figure 5-26. M

aximum

Rate of C

limb - Flaps 0'

(standard unils) I


evision 7: July 1. 1997 t


5-27

SE

CTIO

N 5

SP

LA

TU

SS

P

ER

FOR

MA

NC

E

'PC %I1

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for info on etfect of icing

I Figure 5-27. M

aximum

Rate ot C

limb - Flaps 0" (m

etric units)

Report No: 0 1973-001


Revision 7

: July 1, 1997

L =PLATUSE SECTION 5 I ?c XI1 PERFORMANCE

CRUISE CLIMB AIRSPEED SCHEDULE

ASSOCIATED CONDITIONS: PROPELLER SPEED 1700 RPM INERTIAL SEPARATOR CLOSED

1 Figure 5-28. Cruise Cl~mb Airspeed Schedule

Issued: February 14. 1994 Revision 5: May 10, 1996

1

Report No 01973-001 5-29

RATE OF CLIMB - CRUISE CLIMB (STANDARD UNITS)

ASSOCIATED CONDITIONS: MAXIMUM CLIMB POWER LANDING GEAR RETRACTED FLAPS UP AIRSPEED SCHEDULE: SEE RELEVANT TABLE

EXAMPLE: ALTITUDE BOO0 FT OAT 15 'C AIRCRAFT WEIGHT 7500 LB RATE OF CLIMB 1725 FPM

=PILATUSs S

EC

TION

5 -P

C X

II' P

ER

FOR

MA

NC

E

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

. .\. . . . . . - . -

. . - . . . . . . . I

I:: ....:. p3.j$gky'$-;i$ I

..

.

I

(;;::i;;ii

! Figure 5-30. R

ate of Clim

b - Cruise C

limb (m

etric units)

Issued. February 14. 1994 R

evision 7: July 1, 1997 R

eport No: 01973-00 1

5-31

SE

CTIO

N 5

cPILATUSrC P

ER

FOR

MA

NC

E

-PC XII-

I F

igure 5-31. Time to C

l~m

b - Cruise C

limb (standard units)

I

Repon N

o: 01973-001 5-32


evision 7:

July 1. 1997

SP

ILA

TU

Sg

S

EC

TION

5 ?c X

I1 P

ER

FOR

MA

NC

E

k Figure 5-32. Tim

e to Clim

b - Cruise C

limb (m

etric units)

Issued: February 14. 1994 R

evision 7: July 1. 1997 !

Reporl N

o. 01973.00 1 5- 33

SE

CTIO

N 5

PE

RFO

RM

AN

CE

Figure 5-33. Fuel Used to C

limb - C

ruise Clim

b (standard unils) d

Report N

o: 01 973-001 5-34


evis~on 7: July 1. 1997

=PlLATUS= S

EC

TION

5 -PC

XII-

PE

RFO

RM

AN

CE

I Figure 5-34. Fuel U

sed to Clim

b - Cruise C

limb (m

etric units) '1

Issued. February 14. 1994 R

evision 7: July 1. 1997 1

Report N

o: 01973-00 1 5-35

DISTANCE TO CLIMB - CRUISE CLIMB (STANDARD UNITS)


EXAMPLE. ALTlTUDE 25000 FT OAT -30 'C AIRCRAFT Wl&fl 7950 LB DISTANCE TO CClMB 55 NM

-60 -40 -20 0 20 40 60 loo00 9000 Boo0 7000 6000 so00

OUTSIDE AIR TEMPERATURE - 'C AlRCRAn WEIGHT - LB

SE

CTIO

N 5

PERFO

RM

AN

CE

t F

igure 5-36. Distance to C

limb - C

ruise Clim

b (metric units)

I Issued: F

ebruary 14. 1994 R

evision 7: July 1. 1997

I

Report N

o: 01973-00 1 5-37

SE

CTIO

N 5

=PILAW

S= PER

FOR

MA

NC

E -PC

XII-

1 I I

I F

gu

re 5-37. Marlm

um C

ruise Power (S

heet 1 01 4) 1

Report No. 01973-001

5-38 Issued: February 14. 1994

Rev~slo

n 7: July I. 1997

I

SE

CTIO

N 5

PE

RFO

RM

AN

CE

I Figure 5-37. M

aximum

Cruise Pow

er (Sheet 2 of 4)

I Issued: February 14. 1994 R


eport No: 01973-001

5-39

SECTION 5 =PILATUSS: PERFORMANCE 'PC %I1 i

MAXIMUM CRUISE POWER NOTE: IOAT. TOROUE AND FUEL FLOW 6ASED ON I000 Ib (3828 kg)

@ t W l b @7000lO @WWlh @8000b (2721 kg) (3176 kg) (3621 kg) (4082 hg)

ISA Anludc IOAT OAT Torqw Fwlllow IAS TAS IAS TAS IAS TAS IAS TAS ( t ) (fl) ( t ) ( t ) (pn) (Iblh) (kglh) (kt) (M) (kl) (kl) (M) (M) (M) (M) 0 0 19 !? 31) 613 278 229 ??? 2 8 . . -232 2?! .22! 2%. _2?!

Z% ! < . ! I - 3 6 9 594 ?9 2% -137 -2%. .L32 .31. .?E 224'..235 bOOT, I ? 7 369 576 E! 224 _% _ gQ_ _ 2 1 2 3 2% -222 239

Figure 5-37. Maximum Cruise Power (Sheet 3 of 4)

Report No: 01 973-001 5-40


mPkATUS= SECTION 5 PC XII- PERFORMANCE

MAXIMUM CRUISE POWER NOTE: WAT, TORQUE AND FUEL FLOW EASED ON woo ~b (m rg,


Issued: February 14,1994 Revision 5: May 10, 1996 - - - - - -

Report No: 01 973-001 5-41

SE

CTIO

N 5

%PLATUS=

PE

RFO

RM

AN

CE

?c XII-

Figure 5-38. Long Range C

ruise (Sheet 1 of 4)

Report N

o: 01 973-001 5-42

I


evision 5: May 10,1996

SE

CTIO

N 5

PE

RFO

RM

AN

CE

: Febr n 5: I



uary 14, 1994 R

eport No: Ol!

day lo, 1996

SE

CTIO

N 5

PE

RFO

RM

AN

CE



Report N

o: 01 973-001 5-44


evision 5: May 10, 1996

LONG RANGE CRUISE NOTr:101TrU8eDON1000L~kgj

MAXIMUM ENDURANCE CRUISE rn m A l m mSPEE0 * 110 Kis CONSTANT

SE

CTIO

N 5

=PILATUS= P

ER

FOR

MA

NC

E

-PC x

B

1

Figure 5-39. Maxim

um E

ndurance Cruise (S

heet 3 of 4)

4

Report N

o: 01 973-001 5

48


evision 5: May 10, 1996

!

MAXIMUM ENDURANCE CRUISE NOTE: INOICATED AIRSPEED IS 110 KTS CONSTANT

SPECIFIC AIR RANGE WEIGHT 7000 Ib (3175 kg) - ISA-20"

EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI

FLAPS UP TRUE AIRSPEED 234 KT

INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.58 NMllb

- - 80 100 120 140 160 180 200 220 240 260 280 300

TRUE AIRSPEED - KT


SPECIFIC AIR RANGE MIGHT 7000 Ib (3175 kg) - ISA

EXAMPLE: ASSOCIATED ~~NDITIONS: PRESSURE ALTITUDE ~SOOO FEET LANDING GEAR RETRACTrD TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 239 KT INERTIAL SEPARATa CLOSED SPECIFIC AIR RANGE 0.595 NMIlb

80 1100 120 140 160 180 260 220 240 260 280 300

TRUE AlRSPEED - KT

Figure 5-40. Specific Air Range - 7000 Ib (Sheet 2 of 3)


-

Report No: 01 973-001 5-51


SPECIFIC AIR RANGE WEIGHT 7000 Ib (3175 kg) - ISA+20°

EXAMPLE: ASSOCIATED CONDITIONS: PRESSWE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 243 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.6 NWlb

TRUE AIRSPEED - KT

I Figure 540. Specific Air Range - 7000 Ib (Sheet 3 of 3)

Report No: 01973-001 5-52


SPECIFIC AIR RANGE WEIGHT 8000 Ib (3629 kg) - ISA-20'

U(AMPLE:

ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 232 KT - -

INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.575 ~h;Vlb I I

TFIUE AIRSPEED - KT

SECTION 5 GPILATUSg PERFORMANCE ?C XI1 j

SPECIFIC AIR RANGE WEIGHT. 8000 Ib (3629 kg) - ISA

EXAMPLE: ASSOCIATED CONDITIONS. PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 237 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.585 NMllb

80 100 120 140 160 180 200 220 240 260 280 3W

TRUE AIRSPEED - KT


1 Report No 01973-001 5-54




EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 241 KT INERTIAL SEPARATCM CLOSED SPECIFIC AIR RANGE 0.59 NMilb

TRUE AIRSPEED - KI


Issued: February 14. 1994 Revision 5: May 10. 1996

Report No: 01 973-001 5-55


EXAMPLE:

ASSOCIATED CONDITIONS: PRESSURE AlTrrUDE 15000 FEET W D l N G GEAR RETRACTED TORQUE 27 PSI

RAPS UP TRUE AIRSPEED 230 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.57 NMb

TRUE AIRSPEED - KT


SPECIFIC AIR RANGE WEIGHT 9000 Ib (4082 kg) - ISA

EXAMPLE: ASSOCIATED C0M)ITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEM RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 234 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.58 NMllb

TRUE AIRSPEED - KT

Figure 542. Specific Air Range - 9000 Ib (Sheet 2 of 3)


Report No: 01973-001 5-57



EXAMPLE: ASSOCIATED CONDITIONS. PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 239 KT INERTIAL SEPARATOA CLOSED SPECIFIC AIR RANGE 0.585 NMllb


Report No: 01 973-001 5-58

Issued: February 14.1994 Revision 5: May 10, 1996


HOLDING TlME AND FUEL LANDING GEAR RETRACTED - FLAPS UP ISA STANDARD DAY EXAMPLE: AIRSPEED 150 KlAS HOLDING TIME 2.6 HR POWER FOR LEML FLIGHT ALTITUDE 10000 FT INERTIAL SEPARATOR CLOSED FUEL REWIRED 900 LB

2500

See FLIGHT IN ICING CONDITIONS para for Info on effect of icing

Figure 5-43. Holding Time and Fuel


Report No: 01 973-001 5-59


TlME TO DESCEND ASSOCIATED CONDITIONS: W I N G GEAR RETRACTED - FLAPS UP POWER AS REQUIRED TO DESCEND AT 2000 FPM EXAMPLE: AIRSPEED: MACH 0.48 OR 236 KIAS. KTITUOE 22000 FEET WHICHEVER IS LOWER TIME 11 MIN

Figure 5-44. Time to Descend

Report No: 01973-001 5-60


m 99 am mj ~13~01 SI MA~H~~HM 'mn scz t10 WO m 81 Oa6L ~3M~V wddOOOZlV~3aol01~~tl3MOd 3. OC- IVO d~~-~~Z)NI~ uoowz 3am1l~ :SNOIU~~ aun-

~~

FUEL USED TO DESCEND (METRIC UNITS)

ASSOCIATED C O M I T W : LANDING GEAR RETRACTED - F I N S UP PWER AS REOUIRED TO DESCENO AT 2000 FPM MACH 0.48 OR 238 K W , WHICHEVER IS LOWER

EXAMPLE: ALrrmDE 20000 FT OAT -30 'C AIRCRAFT WEIGHT 3600 KG FUEL USED 30 KG

40 40 -20 0 20 40 6 0 ~ 1 O O O 3 5 0 0 3 0 0 0 2 5 0 0

OUT- AIR M R A T U R E - C AIRCRAFT WEIGHT - KG


DISTANCE TO DESCEND ASSOClAlEO CONOITIOFIS: LANDING GEAR RETRACTED - FLAPS UP POWER AS REWNIED TO U(AMPLE:

DESCEND AT 2000 FPM ALmUDE 25000 n AIRSPEED: MACH 0.48 OR 236 KIAS, OAT -32 'C WHICHEVER IS LOWER DISTANCE !% NM

Figure 5-47. Distance to Descend

Issued: ~ebiuar~ 14,1994 Revision 5: May 10, 1996

Report No: 01 973-001 5-63

POWER-OFF GLIDE TIME (STANDARD UNITS) - - ---

Weight-LB KlAS - -- - -- ASSOCIATED CONDITIONS: 9040 110 EXAMPLE: POWER OFF 8380 108 ALTITUDE 2001X) FT PROPELLER FEATHERED 7720 102 OAT -30 'C IANDING GEM RETRACTED 7060 97 AIRCRAFT WEIGHT 6290 LB W S UP 6400 92 GLIDE TIME 29 MIN

50

- 6 0 4 - 2 0 0 20 40 6 0 1 0 0 0 0 g O O O 8 0 0 0 7 0 0 0 6 0 0 0

UlT!jIDE AIR TEMPERATURE - 'C AIRCRAFT WEIGHT - LB


POWER-OFF GLIDE DISTANCE t

(VALID FOR ALL AIRCRAFT WEIGHTS)

ASSOCIATED CONDITIONS: POWER OFF PROPELLER FEATHERED LANDING GEAR RETRACTED FLAPS UP

EXAMPLE: I

lrLTrruw 25000 R OAT -30 'C 1 GLIDE DISTANCE 64 NM t

Figure 5-50. Power-off Glide Distance

I Report No: 01973-001 5-66



BALKED LANDING TORQUE PROPELLER SPEED 1700 RPM AIRSPEED 80 KlAS INERTIAL SEPARATOR CLOSED MAXIMUM TORQUE REDUCTION WITH EXAMPLE:

INERTIAL SEPARATOR OPEN : ALTITUDE 8000 R 1.2 PSI IN NON ICING CONDITK3lS OAT 13 'C

- 2.1 PSI IN ICING CWDITIONS ENGINE TORQUE 39.9 PSI

-60 -40 -20 0 20 40 60


Figure 5-51. Balked Landing Torque

Issued: February 14. 1994 Rev~sion 7: July I , 1997

Report No. 01973-001 5.67

RATE OF CLIMB - BALKED LANDING (STANDARD UNITS)

ASSOCIATED CONDmlO(r(S: TAKEOFF POWER LANDING GEAR EXTENDED FLAPS 40' AIRSPEED 80 KlAS

EXAMPLE: A L T ~ E 7000 n OAT 22 'C AIRCRAFT WEIGHT 7500 LB RATE OF CLIMB 1000 FPM

-60 -40 -20 0 20 40 60 1 O O O O ~ 8000 7000 6000

m I D E AIR TEMPERATURE - 'C AIRCRAFT WEIGHT - LB

RATE OF CLIMB - BALKED LANDING (METRIC UNITS)

ASSOCIATED CONDITIONS: TAKEOFF POWER LANDING GEAA EXENDED W S 40' AIRSPEED 80 KlAS

EXAMPLE: AlTlR.DE 7000 Fi OAT 22 'C AIRCRAFT WEIGHT 3400 KG RATE OF am 1000 FPM

OUTSIDE AIR TEMPERATURE - 'C AIRCRAFT WEIGHT - KG

LANDING TOTAL DISTANCE - FLAPS 40' FROM 50 FT; (STANDARD UNITS)

ASSOCIATED COM)nm Ex.AMRE

APPROACH AT 1.3 VS; UTmRE 60Mn

REFER TO THE SPED S H E M E 1-E WEIGHT -- LB V ~ p s - KlAS OAT 18 'c AERAGE BRAKING TECMNlm 64@3 67 WEKW? 7 1 6 L8 GRo~JNO !OLE AFTER TOUCH DOWN 7 W 72 HEADWYO COMPONEM 8 KT

RUWAY SURFACE. TARMAC 8200 76 UPHU MMPONENT 1 h

LANDING TOTM DISTANCE 1970 n

LANDING TOTAL DISTANCE - FLAPS 40' FROM 15 M; (METRIC UNITS)

~ A E O C O M m ExMelE APPROICH AT 1.3 Vs, M W BMY) FT RERR TO 1W WEED S M W L E TABLE W E M - KG VH - KlAS OAT 18 % AVERAGE BRAKING ECHNKXlE 2900 87 W E I W 3500 KG OAOUND IDLE KTER TOUCH 3300 72 HE4WlND COkPOHHl 8 KT W A Y SURFACE TARUU: 3100 78 UPHUCMPONNl 1 % SEE SECTION 2 . UMITATIONS 4100 60 WING TOTK USlANCE 6CQ M

- w -J OUTSlM AIR W E R A N R E - 'C , Y

WEIGHT - KG

LANDING GROUND ROLL - FLAPS 40" (STANDARD UNITS)

ASSOCIATED MMIIIONS: AMRAOE W I N G E C H N I W GROUN) IDLE AFTER TOUCH RUNWAY SWIFACE. TARMAC SEE SECTION 2 . LlMrrATmJS

U(MRCE:

A L T r m M 60W OAT 18 H l E l W n l 6 HEADWIND m N I 8 UPHILL CDhR0NEN-r 1 LANDING GCIOUK) ROU. 1110

LANDING GROUND ROLL - FLAPS 40' (METRIC UNITS)

AssmAw C O H ) m :

A V E W WM TEWYXlE GAOUNI lME PFTER TOUCH WW RUNWAY SURFACE: TARMAC SEE S E m 2 - LIMITATICS

KTrmDE 8000 f7 OAT 18 'C WEIGHT 5500 KG NAMMH)CCMPOEENT 8 K T LlPHLC cOkp0Hh-r 1 Y. IJNDIffi G R O W RQI 340 M

OUISE AIR m T L R E - 'C WKm - KG WHOcOWU€NT-KT 9 0 P E - Y

LANDING TOTAL DISTANCE WITH REVERSE THRUST - FLAPS 40" FROM 50 FT; (STANDARD UNITS)

ASSCCIATED CONDITIONS. EWIMPLE APPROACH AT 13 VIq NTT(UIE 6000 R REFER TO THE SPEED SCHEME TABLE WEIGHT - LB Vur - K- OAT

18 'C

AVERAGE BRMlNG TECHWlOUE 6400 67 W E I M 7716 LB FUU R M R S E THRUST WfER TOUCH OOWN 7300 72 HEADWIND COMPONENT 8 KT

R W A Y SURFACE TAAMAC 8200 76 VPH(XL COMWNENT 1 9.

SEE SECTKXJ 2 - LlMlTATKWSS 9100 80 W I N G TOTN DISTANCE 1675 Fl

..

OUTSIDE UR T E ~ A T U R E - %

L 3P

ILA

TU

SS

S

EC

TION

5 PC

%!I

PE

RFO

RM

AN

CE

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for info on

effect of icing

Figure 5-59. Landing Total Distance w

ith the use ol Reverse Thrust - Flaps 40" (m

etric units) I


evision 6: Dec 6. 1996

Report N

o: 01 973-00 1 5-75

SE

CTIO

N 5

PE

RFO

RM

AN

CE

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS


I Figure 5-60. Landing G

round Roll w

ith the use ol Reverse Thrust - Flaps 40' (standard units)

Report N

o: 01973-001 5-76


evision 6: Dec 6, 1996

SE

CTIO

N 5

PE

RFO

RM

AN

CE

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for info

on effect of icing

Figure 5-61. Landing Ground R

oll with the use o

l Reverse Thrust - Flaps 40" (m

etric units) I


evision 6: Dec 6, 1996

Report N

o: 01973-001 5-77


FLIGHT IN ICING CONDITIONS

STALL SPEEDS

I When operat~ng in PUSHER ICE MODE, the stick pusher computer automatically reduces the shaker and pusher settings by 8" AOA. W~th operational deice boots this results in an increase of the stall speed at the maximum takeofl weight of 10 kts with 0" flaps and 7 kts with 15" flaps.

NOTE

Flap position is limited to maximum 15" in icing conditions with operational deice boots.

With failed deice boots and ice accretion on the lilting surfaces an increase of the stall speed at the maximum takeoff weight ol 14 kts with 0" flaps is to be expected.

NOTE

Flap position is limited to maximum 0" in icing conditions with lailed deice boots.

The wings level stall speeds at the maximum takeoff weight of 9039 Ib (4100 kg) and with power at fl~ght idle are summarized in the lollowing table:

ENGINE TORQUE

" FLAPS

0"

15"

When the engine inlet inertial separator is open and during flight at altitudes above 5000 A, the maximum torque available can be reduced by up to 1.2 psi in non icing conditions, and up to 2.1 psi in icing conditions.

Report No: 01 973-001 5-78

Non icing

Icing conditions

Pneumatic boots failure

Non icing

Icing conditions


,

STALL SPEED

KlAS

86

96

100

74

8 1

KCAS

88

98

102

76

83

PLATUS= SECTION 5 'PC x F PERFORMANCE

TAKEOFF PERFORMANCE

When taking-off in or into known icing conditions, the flaps must be set to 15" and the rotation speed increased by 9 KIAS. The speed at 50 ft (15 m) height will be correspondingly increased by 12 KIAS. As a result, the takeoff ground roll distance wiH be increased by 29%. The takeoff distance will be increased by 42% and the accelerate-stop distance by 30%.

The takeoff ground roll and takeoff total distances given in the performance charts will be longer with operations on slush or snow covered runways.

CLIMB PERFORMANCE

During flight in icing conditions, the maximum rate of climb can be reduced by up to 1000 FPM. After failure of the pneumatic deice boots, the maximum rate of climb can be reduced by up to 1200 FPM below 15000 ft and by up to 1400 FPM at higher altitudes.

HOLDING ENDURANCE

During holding flight in icing conditions, a higher engine torque is required to maintain level flight. Increases in engine fuel flow between 25% and 50010 are to be expected with respect to non icing conditions.

LANDING PERFORMANCE

After icing encounters and with visible ice accretion on the airframe, the landing is performed with 15" flaps and an approach speed of 105 KIAS. As a result, the landing ground roll distance can be increased by up to 90% without reverse thrust applied, and by up to 8W0 with reverse thrust applied. The landing total distances will correspondingly be bnger by up to 55% without reverse thrust applied, and by up to 45% with reverse thrust applied.

After failure of the airframe pneumatic deice boots in king condiiions, the landing is performed with 0° flaps and an approach speed of 130 KIAS. As a result, the ground roll distance can be increased by up to 160% without reverse thrust applied, and up to 140% with reverse thrust appiied. The landing total distances will correspondingly be bnger by up to 90% without reverse thrust appiied, and by up to 75% with reverse thrust applied.

BALKED LANDING PERFORMANCE

During flight in icing conditions, the balked landing rate of climb with 15' flaps can be up to 100 FPM bwer than in non icing conditions with 40° flaps. After failure of the pneumatic deice boots the rate of climb can be up to 400 FPM lower.


Report No: 01973-001 5-79


Report No: 01 973-001 6-80

FLIGHT PLANNING EXAMPLE

GENERAL

Before performance calculations can begin, it will be necessary to determine the aircrafl loading. Refer to Section 6, Weight and Balance to calculate the actual aircraft loading. FOI this flight planning example, the sample aircraft loading in Section 6, Fig. 6-8, will be used.


Aircraft Configuration:

Takeoff Weight 8798 1b

Usable Fuel 1650 Ib

Departure Airport Conditions:

Field Pressure Altitude 4000 ft

OAT +17"C (ISA +lO°C)

Wind Component 9 kt (headwind)

Runway Slope 1% (uphill)

Field Length 3690 A

Total Trip Distance 765 nm

Total Trip Distance 765 nm

Destination Alrport Conditions:

Field Pressure Altitude 2000 ft

OAT + 16°C (ISA + 5°C)

Wind Component 6 M (headwind)

Runway Slope 1.5% (downhill)

Field Length 2550 ft

Cruise Conditions:

Pressure Altitude FL 280

Forecast Temperature -31°C (ISA +lO°C)

Forecast Wind Component 10 kt (headwind)

TAKEOFF

Apply the departure airport conditions and the aircraft weight to the appropriate takeofl performance charts and check that the corresponding distances are less than the available field length at the departure airport.

Apply the departure airport conditions to the Takeoff Power Chart to determine maximum torque to be applied before brake release.


CLIMB

NOTE

The dimb performance chart assumes a no wind condition. The pilot must consider the effect of the winds alofl when computing time, fuel, and distance to climb. The fuel to climb includes the fuel consumed during the takeoff run.

Apply the cruise conditions of pressure altitude and temperature (respectively 28000 It and ISA + 10°C in this case) to the appropriate chart to determine the time, fuel, and distance to climb from sea level to the cruise altitude at the specified takeoff weight (8798 Ib in this case). Next. apply the departure airport conditions (respectively 4000 f l and ISA + 10°C in this case) to the same chart to determine those same values to climb from sea level to the departure airport. Subtract the values for the departure airport from those for the cruise altitude. The remaining values are the time, fuel, and distance to climb from the departure airport to the cruise altitude.

DESCENT

NOTE

Cllmb

From S. L. to 28000 ft

From S. L. to departure airport

Departure airport to 28000 ft.

The descent performance chart assumes a no wind condition. The pilot must consider the effect of the winds aloft when computing time, fuel, and distance to descend.

Fuel

180 1b

25 Ib

155 1b

Time

26 min

3 min

23 min

Apply the cruise conditions of pressure altitude and temperature (respectively 28000 fl and ISA + 10°C in this case) to the appropriate chart to determine the time, fuel, and distance to descend from cruise altitude to sea level. The weight at the beginning of the descent is not known exactly at this stage, but it can be estimated in practice as shown in the following table:

Dlstance

78 nm

7 nm

71 nm


Report NO: 01973-001 5-8 1


descent beginning of descent

8798 1b - 1650Ib +3001b + 100 1b = 7548 1b

'As required by operating regulations; here a reserve corresponding to 45 min hold at 5000 11 is assumed.

Next, apply the destination alrport conditions (respectively 2000 A and ISA + 5°C in this case) to the same chart to determine those same values to descend from the destination airport to sea level. Subtract the values for the destination airport from those for the cruise altitude. The remaining values are the time, fuel, and distance to descend from the cruise altitude to the destination airport.

CRUISE

Descent

From 28000 ft. to S. L.

From destination airport to S. L.

From 28000 A. to destination airport

Calculate the cruise distance by subtracting the climb and descent distances from the total trip distance. Select a cruise power setting and refer to the appropriate chart to determine the true airspeed and fuel flow for the forecast cruise conditions. Adjust the true airspeed for the winds aloft headwind component to determine the ground speed. Divide the cruise distance by the ground speed to determine the cruise time. Calculate the cruise fuel required by multiplying the fuel flow by the cruise time.

Time

14 min

1 min

13 min

By assuming an average cruise weight of 8500 Ib. Maximum Cruise Power setting for 28000 H. at ISA +1VC yields 250 KTAS at 325 Iblhr.

Report No: 01973-001 5-82

Fuel

86 lb

8 Ib

78Ib ,

= Cruise distance

= 632 nm

- Total trip distance

765 nm

Issued: February 14.1994 Revision 6: Dec 6, 1996

Distance

66 nm

4 nm

62nm

- Climb distance

-71 nm

- Descent distance

- 62 nm

EPILATUSm SECTION 5 ?c XI1 PERFORMANCE

Cruise Speed

250 KTAS

Cruise Distance

632 nm

LANDING

+I- Headwind Component

-10 kt

Calculate the estimated landing weight by the subtracting the weight of the fuel for climb. descent, and cruise from the takeoff weight.

1

= Ground Speed

240 KTAS

I Ground Speed

1 240 kt

= Cruise Fuel

= 855 1b

Cruise Time

2.63 hr

1

= Cruise Time

= 2.63 hr

(2hr 38 min)

x Fuel Flow

x 325 Iblhr

Apply the destination airport conditions and the calculated aircraft weight to the appropriate landing performance charts and check that the corresponding distances are less than the available field length at the destination airport.

Takeoff weight

8798 1b


Report No: 01973-001 5-83

- Climb Fuel

- 155 Ib

- Descent Fuel

- 78 1b

- Cruise Fuel

- 855 Ib

1

= Landing Weight

= n 1 0 l b .


TOTAL FLIGHT TIME

The total flight t~me IS the sum of the t~me to climb, descent, and cruise.

TOTAL FUEL REQUIRED

Cl~mb Time

23 min

The total luel required is the sum ol the fuel consumed during engine start and ground operation, takeoff and climb, descent, and cruise.

+ Descent Time

+ 13min

Report No: 01973-001 5-84

Ground Ops

40 1b


+ Cruise Time

+ 2 hr 38 min

= Total Time

3 hr 14 min

t TO & Climb

+ 1551b

+ Descent

+ 78 1b

+ Cruise

+ 855 1b

+ Reserve

+ 300 1b

= Total

=I428 1b


Weight and Balance

SECTION 6 WEIGHT AND BALANCE

SECTION 6

WEIGHT AND BALANCE

TABLE OF CONTENTS

Subject

GENERAL

AIRPLANE WEIGHING

PREPARATION WEIGHING PROCEDURE WlTH LOAD PLATES WEIGHING PROCEDURE WlTH JACKS AND LOAD CELLS

WEIGHT AND BALANCE RECORD

GENERAL LOADING RECOMMENDATIONS

CARGO HAZARDOUS MATERIALS

WEIGHT AND BALANCE DETERMINATION FOR FLIGHT

COMPLETION OF THE LOADING FORM COMB1 CONVERSION

EQUIPMENT LIST

INTERIOR CONFIGURATIONS

CORPORATE COMMUTER INTERIOR CODE STD-9S CORPORATE COMMUTER INTERIOR CODE STD-6SdB EXECUTIVE INTERIOR CODE EX-6s-1 (MSN 101-171) EXECUTIVE INTERIOR CODE EX-6s-2 (MSN 172-999) EXECUTIVE INTERIOR CODE EX8S EXECUTIVE INTERIOR CODE EX-4s-38 EXECUTIVE INTERIOR CODE EX-6s-STD-2s EXECUTIVE INTERIOR CODE EXJS-STD-4s


Page

Report No: 01 973-001 6-i






GENERAL

This section contains the information required to determine the Basic Empty Weight and Moment of the aircraft, adjust the B.E.W. 81 M as equipment is added or removed, and calculate aircraft loading for various flight operations. Sample loading forms are provided.

To achieve the performance designed for the aircraft it must be flown with approved weight and center of gravity limits.

It Is the responsiblllty of the pilot in command to make sure that the aircrafl does not exceed the rnaxlmum weight limits and is loaded within the center of gravity range before takeoff.

Weight in excess of the maximum takeoff weight may be a contributing factor to an accident. especially with other factors of temperature. airfield elevation and runway conditions. The aircraft's climb, cruise and landing performance will also be affected. Loads that the aircraft was not designed for may be put on the structure, particularly during landing.

The pilot should routinely determine the balance of the aircraft since it is possible to be within the maximum weight limit and still exceed the center of gravity limits. Information regarding the Basic Empty Weight can be found on the Weight and Balance Record in this section. Installed equipment information can be found in the Equipment List at the back of this manual. Using the basic empty weight and moment together with the Loading Form the pilot can determine the weight and moment for the loaded aircraft by computing the total weight and moment and then determine whether they are within the Center of Gravity Envelope.

Issued: June 10,1994 i Revision 9: September 1, 1999

Report No: 01 973-001 6-1


AIRPLANE WEIGHING

PREPARATION

1. Make sure that all applicable items listed on the airplane equipment list are Installed in their proper locations.

2. Clean airplane. Remove dirt, excessive grease, water, and foreign items.

3. Completely defuel the fuel tanks. Use the wing fuel drain ports for the completion ot the task.

4. Fill oil, hydraulic fluid, and all other operating fluids to full capacity.

5. Make sure that the flaps are fully retracted and that the flight controls are in the neutral position.

6. Place crew seats in the center position and make sure the cabin passenger seats are in the correct positions. Refer to the relevant Interior Configuration Code Seat Location Charl in this Section.

7. Close access panels and passenger cabin door.

8. Make sure that all tires are inflated to normal operating pressure.

9. Place airplane in a closed hangar to prevent scale reading errors due to wind.

Report No: 01973-001 6-2



WEIGHING PROCEDURE WITH LOAD PLATES

LEVELING

Open the cargo door and place a level across the seat tracks. Adjust the main gear tire pressure until the airplane is laterally level. Place the level along the top of the inboard seat track and adjust the nose tire pressure until the airplane is longitudinally level. Refer to Section 8 of this Handbook for more information concerning airplane leveling. Remove the level and carefully close the cargo door.

WEIGHING

1. Refer to the manufacturers instructions and position the aircraft on the load plates. I 2. With the airplane level and brakes released, record the weight shown on each scale in

the appropriate section on Figure 6-1 A, Airplane Weighing Form. I 3. Record the tare weight for each applicable scaie on Figure 6-lA, Airplane Weighing I

Form.

4. Record nose gear (a) and the main gear (b) dimensions on Figure 6-IA, Airplane 1 Weighing Form.

5. Subtract tare weight from applicable scale reading for net weight and record in appropriate section on Figure 6-l A, Airplane Weighing Form. I

6. Calculate Arm of the nose gear (A) and the main gear (B) and record in appropriate section on Figure 6-lA, Airplane Weighing Form. I

7. Calculate airplane C.G. Arm and record on Figure 6-2, Airplane Basic Empty Weight. I 8. Adjust weight and moment for unusable fuel and optional equipment installed after 1

airplane weighing to determine airplane Total Basic Empty Weight and Moment.

9. Calculate Basic Empty Weight C.G. I 10. Update Figure 6-3, Weight and Balance Record, as required. I 11. After weighing return tire pressures to operational values. Refer to Section 8 for I

instructions.

Issued: June 10,1994 L Revision 1 1 : March 1,2003

Report No: 01 973-001 6-3

-

SECTION 6

I

F P L A T U S F WEIGtiT AND BALANCE PC12 i

TOTAL AIRCRAFT WEIGHT

C G. ARM DATi

Report No: 01 973-001 6-4

Figure 6-1. Airplane Weighing Form (Sheet 1 of 3)



Landing Gear

Nose

Len Main

Right Main

Figure 6-1A. Airplane Weighing Form (Sheet 2 of 3)


Report No: 01 973-001 6-5

Symbol

(a)

(b)

(C)

Arm (A)

(mm)

2842

2840

2838

2835

2833

283 1

2828

2826

2824

282 1

281 9

2817

2814

Dimension (a)

(mm)

0

20

40

60

80

100

120

140

160

180

200

220

240

Arm (A) (in)

111.90

111.82

11 1.74

111.62

111.54

111.46

111.34

111.27

111.19

11 1.07

110.99

11 0.91

110.79

Dimension

mm

Average (b)

mm

- -

(L+R) / 2

Arm

in (mm)

(A)

(6)


MAIN GEAR ARM r'-l

I Dimension (b) is strut extension

Calculate the airplane C.G. arm as weighed:

Dimension (b)

(mm)

110

130

150

170

190

210

230

250

270

290

31 0

C.G. Arm (In or m) = N x A + (L + R) x B T

Where: A = Nose Landing gear arm B = Main Landing gear arm N = Nose Landing gear weight L = Left main landing gear weight R = Right main landing gear weight T = Total weight of N + L + R

Arm (8) (in)

254.78

254.46

254.07

253.60

253.04

252.41

251.71

250.88

249.97

248.91

247.73

Figure 6-1A. Airplane Weighing Form (Sheet 3 of 3)

Report No: 01 973-001 Issued: June 10,1994 6-6 Revision 11 : March 1,2003

Arm (B)

(mm)

6471

6463

6453

6441

6427

641 1

6393

6372

6349

6322

6292


WEIGHING PROCEDURE WITH JACKS AND LOAD CELLS

LEVELING

Put the jacks in position below the wing and tail jacking points. The fuselage jacking points must not be used. Refer to the manufacturers instructions for the use of the load cell equipment. Position the load cells and adapters and slowly raise the aircraft clear of the ground.

Open the cargo door and place a level across the seat tracks. Place the level along the top of the inboard seat track and adjust the tail jack until the airplane Is longitudinally level. Refer to Section 8 of this Handbook for more information concerning airplane leveling. Remove the level and carefully close the cargo door.

WEIGHING

1. With the airplane level, record the weight shown on each load cell in the appropriate section on Figure 6-1 B, Airplane Weighing Form.

2. Calculate airplane C.G. Arm and record on Figure 6-2, Airplane Basic Empty Weight. The C. G. Arm calculation formula is:

3. Adjust weight and moment for unusable fuel and optional equipment installed after airplane weighing to determine airplane Total Basic Empty Weight and Moment.

4. Calculate Basic Empty Weight C.G.

5. Update Figure 6-3, Weight and Balance Record, as required.


Report No: 01 973-001 6-6A

SECTION 6 =PILATUSW WEIGHT AND BALANCE PC12

r 8uk

scab Poriuon "'""" Pb9tpp';" Tail Jacklnp Polnl T

Left Ualn Jacking Polnl L

Right k i n ~ s c ~ n g Point R

TOTAL AIRCRAFT WEIGHT

Report No: 01973-001 6-6B

Figure 6-1 B. Airplane Weighing Form


€PLATUS9 SECTION 6 P(l2 WEIGHT AND BALANCE

AIRPLANE USEFUL LOAD - NORMAL CATEGORY OPERATION

t

The Basic Empty Weight. C.G., and Useful Load are for the airplane as licensed at the factory. These ligures are only applicable to the specific airplane serial number and registration number shown. Refer to Figure 6-3. Weight and Balance Record when modifications to the airplane have been made.

Model:

Figure 6-2. Airplane Basic Empty Weight

Issued: June 10, 1994 Report No: 01 973-001 Revision 9: September 1,1999 6-7

= Useful Load Ib (kg)

I

Ramp Weight Ib (kg)

Serial No.:

- Basic Empty Weight Ib (kg)

Moment Ib-in (mkg)

9746 (1 12.31)

7422 (87.38)

Registration No.:

Item

1. Airplane Weight, C.G. arm, and moment. (As weighed in FQure 6-1)

2. Unusable Fuel

3. Optional equipment, if applicable



6. TOTbL BASIC EMPTY WEIGHT AND MOMENT (Sum of 1 t h ~ 5)

Date:

Weight Ib (kg)

MSN 101-140 43.2 (19.6) MSN 141 AND UP 32.9 (14.9)

C.G. Arm in (m)

225.6 (5.73)

225.6 (5.73)


WEIGHT AND BALANCE RECORD

Figure 6-3. Weight and Balance Record is a log of the modifications that occurred after the airplane was licensed at the factory. Any change to the permanently installed equipment or airplane modifications which effect the airplane Basic Empty Weight or Total Moment must be entered in Figure 6-3. Weight and Balance Record. The last entry on the Weight and Balance Record will be the current airplane Basic Empty Weight and Total Moment.


SE

CTIO

N 6

WE

IGH

T AN

D B

ALA

NC

E

Figure 6-3. Weight and B

alance Record

(Sheet 1 of 2) Issued: June 10,1994 R

evision 9: September 1, 1999

Repon No: 01973-001

6-9

SECTIO

N 6

WEIG

HT A

ND

BA

LAN

CE

Figure 6-3. Weight and Balance R

ecord (Sheet 2 of 2)


Issued: June 10, 1994 6-10

Revision 9: Septem

ber 1, 1999


-

GENERAL LOADING RECOMMENDATIONS

The following general loading recommendation is Intended as only a guide. Refer to Section 2 for Seating and Cargo Limitations. The pibt in command must refer to the appropriate moment charts, loading form, and the C.Q. Envelope to determine that the airplane is properly loaded.

Pilot Only or Pilot and Copilot (Right Front Passenger)

Load rear cabin area as desired to remain within the afl C.G. limit. Cargo may have to be placed forward of the cargo door if the rear baggage area is loaded near its maximum limit. Fuel load may be limited by maximum weight.

Pilot, Copibt (Right Front Passenger), Passengers

Load passengers afl from passenger seat 1 as desired. With up to 6 occupants and no rear cargo, ballast may be needed in the rear baggage area to bring the C.G. afl of the forward limit at low fuel loads. Load rear cabin area as desired to remain within the aft C.G. limit. Fuel bad may be limited by maximum weight.

Load fuel equally between the left and right wing fuel tanks.

CARGO

Before loading airplane, attach the tail support stand to prevent the tail from contacting the ramp surface while ground personnel are in the aft cabin during the loading process.

Observe the maximum floor and seat rail load limits given on Ule placard on the forward and rear cargo door frame. Figure 6-4 gives the cabin dimensions and loading areas.

Secure heavy cargo in the cabin area with tie-down straps attached to seat rail anchor points. Refer to Fiure 6-5 for cargo mass calculation and Figure 6-6 lor tie-down strap and cargo net installation.

Refer to the Illustrated Parts Catalog (IPC) Chapter 25 for the part numbers of the approved cargo restraint nets, tie down straps, load carriers and retaining angles.


Report No: 01973-001 6-11


-- --

HAZARDOUS MATERIALS

Protection against the damaging effects of leakage of hazardous materials has not been provided in the cargo area. Provisions should be made for protection if carriage of these materials is planned.

4

1

In addition to the pilot in command, other personnel used for loading and unloading should be properly trained concerning the handling, storage, loading and unloading of harardous materials it they are to be carried.

Information and regulations pertaining to the air transponation of hazardous materials is I

outlined in the Code of Federal Regulations (CFR) Title 49 and in the International Civil Aviation Organization (ICAO) Technical Instructions for the Safe Transport of Dangerous Goods by Air.



FORWARD CABIN DOOR

CARGO DOOR

COMPARTMENT VOLUME (WlWlu OIMENSICU(ED AREAS)

CABIN tm.rs na (s.4 IIP) BAGGAGE 96.0 na (0.45 ma)


\ I CARGO MAY BE LOCATED ANYWHERE IN SHADED AREA

ALL CARGO MUST I)E SEWRED BY APPROVED CARGO RESTRAINTS 8

Figure 6-4. Loading Areas

Report No: 01973-001 6-13


MAXIMUM ALLOWABLE MASS PER SINGLE CONTAINER ( WITHOUT SPECIAL EQUIPMENT )

The maximum allowable mass is based upon the package dimensions, vertical c.g., and the number of seat rails used to secure the ti8 down straps.

1. Measure height and length (along aircraft longitudinal axis).

2. Determine vertical c.g. (c).

3. Calculate values for Kh and Kc as follows:

4. Refer to the appropriate configuration chart to determine whether proposed method of securing cargo is adequate for the size and mass of package.

Fiure 6-5. Maximum Cargo Mass (Sheet 1 of 4)

Report No: 01973-001 Issued: June 10,1994 6-14 Revision 9: September 1, 1999


MAXIMUM ALLOWABLE MASS PER SlNOLE COUTAINER ( WITHOUT SPECIAL EQUIPMENT )

EXAMPLE

1. Package:

Length (along 40" (1.016 m) bng. axis)

Mass 200 Ib (90.7 kg)

2. Determine vertical c.g.

Balance package on edge and draw a vertical line up from the corner of the P=wJe.

Repeat for opposite side.

Measure height of line intersection for vertical c.g. (c).

3. Calculate values for Kh and Kc

F i r e 6-5. Maximum Cargo Mass (Sheet 2 of 4)


Report NO: 01973-001 6-15


MAXIMUM ALLOWABLE MASS PER SINGLE CONTAINER ( WITHOUT SPECIAL EQUIPMENT )

EXAMPLE (CONT'D) I

4 Refer to the Cargo Tie Down Configuration Chart for four seat rails. Determine I

chart area applicable to the package mass of 200 lb ( 90.7 kg ). A package , mass of200 Ib is more than 110 Ib and less Ulan 220 Ib, so only areas A, 8, C, , and D apply.

5. Plot Kh and Kc. Determine that the plotted point falls within the 220 Ib limit. Proposed cargo loading and tie down is satisfactory. I

If the plotted point falls outside the limit for package mass, change package '

orientation to lower vertical c.g. to bring Kh/Kc point within mass limit.

Figure 6-5. Maximum Cargo Mass (Sheet 3 of 4)



CARGO TIE DOWN CONFIGURATION CHART (APPROVED LOADING OF STRAPS = 1800 LBS OR 816 KGS)

CARGO SECURED LONGITUDINALLV TO TWO SEAT RAILS

MAX MISS NOT

EXCEEOING LO (KG) AREA

m(lSOl A

nol1wl AB

rlo~so~ A ~ C

CARGO SECURED LONGITUDINALLY TO FOUR SEAT RAILS

MAX MISS NOT

EXCEEOING LBIKGI AREA

GfflIMOl A

44Ol2Wl AB

330llMI ABC

~ ( l w l ABCD

I lo 1501 ABCDE

Figure 6-5. Maximum Cargo Mass (Sheet 4 of 4)


Report No: 0 1973-001 6-17

SECTION 6 EPLATUSW WEIGHT AND BALANCE K12

- CARGO TIE WWN

Figure 6-6. Cargo Restraint Installation (Sheet 1 of 2)

Report No: 01973-00 1 Issued: June 10.1994 6-18 Revisin 9: September 1,1999


ANCHOR PIATE

N S NET FLOOR

ATTACHMENT Fm- & EXTENDED L U W E FORWAR0 NET

ATTACHMENT FlTIlNOS

AlTACWHENT SEAT RAIL FITTINGS R,

Figure 6-6. Cargo and Luggage Restraint Installation (Sheet 2 of 2)

Issued: June 10.1994 ~eport No: 01973-001 Revision 9: September 1,1999 6-19


WEIGHT AND BALANCE DETERMINATION FOR FLIGHT , 1

This section contains the crew seats, baggage, fuel load moments charts and C of G envelopes in LB-IN and MKG. I

1

Refer to the lnterior Configurations section for the passenger seat moments. Find the correct Moment Chart tor the interior configuration of the aircraft (Interior Code No. on Post SB 25-016 aircraft).

A Sample Loading Form and a blank Loading Form tor ownersloperators use are given in Figure 6-8. Instructions on how to use the charts, complete the loading form and to determine if center of gravity is within approved limits are given below.

COMPLETION OF THE LOADING FORM

Enter the current Running Basic Empty Weight and Total Moment from Figure 6-3 in the appropriate space on the Loading Form, Figure 6-8.

1 Enter the weights of all of the crew, passengers, items stowed in cabinets and baggage to be loaded, in the appropriate space on the Loading Form, Figure 6-8. 1

Use the Moment Charts in Figure 6-7, to determine the moment for the crew and baggage.

Use the correct Interior Code No. Moment Chart in the Interior Configurations section, to determine the moment for the passengers.

Enter the moment of each item in the appropriate space on the Loading Form, Figure 6-8.

Add the weight and moment of all of the items to the Basic Empty Weight and Moment of the airplane to determine the Zero Fuel Weight and Moment. Divide the moment by the w e e l to determine the C.G. arm,

Locate this point in the C.G. Envelope, Figure 6-9. If the point falls within the envelope, the loading meets the weight and balance requirements.

Use the Moment Chart in Figure 6-7, to determine the moment of the fuel load.

Enter the weight and moment of the fuel in the appropriate space on the Loading Form, Figure 6-8.

Add the fuel weight and moment to the calculated Zero Fuel Weight and Moment to determine the Ramp Weight and Moment. Divide the moment by the weight to determine the C.G. arm.

Locate this point in the C.G. Envelope, Figure 6-9. If the point falls within the envelope, the loading meets the weight and balance requirements.

RepoR No: 01973-001 6-20



Subtract the weight and moment of the fuel allowance for engine start and ground operations and add the moment due to the retraction of the landing gear to the Ramp Weight and Moment to determine Takeoff Weight and Moment. Divide the moment by the weight to determine the C.G. arm.

Locate this point in the C.G. Envelope, Figure 6-9. If the point falls within the envelope, the loading meets the weight and balance requirements for takeoff.

COMB1 CONVERSION

A Combi Conversion can be made from the removal of cabin seats from a Corporate Commuter and the removal of cabin seats and furnishings from an Executive Interior aircraft. The Combi Interior consists of 2 crew seats and payload or a combination of seats and payload. Cargo nets can be installed to attachment points at frames 24 and 27. Refer to Section 2 for the Cargo Limitations.

The airplane is weighed at the factory before the time of delivery. When other interior configurations are required, adjust the Basic Empty Weight and Moment and complete the landing form as follows:

Make a temporary mark on the seat rail at the forward edge of the Corporate Commuter Seat@) or mark position of the Executive Seat anachment fittings of the seat@) to be removed with masking tape or similar material to expedite re-installation. Remove the passenger seats and furnishings as required.

Use the passenger seats and furnishings weight and moment data in the relevant Interior Code section and determine the total weight and moment difference of the interior items removed from the aircraft.

Example: Three Seat Bench removed. Frame 27 Cargo Net installed.

NOTE: The moment figure is obtained by multiplying the weight of the item by the fuselage station given in the Seat Location Chart.

ITEM

Three Seat Bench

Frame 27 Cargo Net

Total Value

Enter the Total Value on line 2 of the Loading Form, Figure 6-8.


WEIGHT LB (KG)

- 125.2 (56.8)

+ 3.6 (1.65)

- 121.6 (55.15)

Report No: 01973-001 6-21

MOMENT LB IN (M KG)

- 40586 (467.61)

+ 1049 (12.21)

- 39537 (455.4)


Calculate the cargo moment as follows:

Locate one of the luggage net floor attachment points at frame 34. Measure distance from the attachment point to the center of the cargo i.e.35 in (0,889 m). The fuselage station dimension at the luggage net attachment point is 361.15 in (9,170 m) The arm of the cargo is the fuselage station dimension of the net attachment point minus the distance to the center of the cargo.

Example: Distance to cargo center - 35 in (0,889 m) Net Fuselage Station a 361.15 in (9,170 m) Cargo Arm = 361.15 in -35 in = 326.15 in (9,170 m-0,889 m = 8,281 m)

Enter the cargo arm and the weight of the cargo plus tie down straps and cargo arm on the Loading Form.

Complete the remainder of the Loading Form as given above.

When re-installing the passenger seats, return the seats to their original positions and verify the dimensions as shown in the Seat Location Chart for the aircraft configuration. Secure the arresting pin on the Corporate Commuter Seat(s) or install the locking needles on the Executive Seat(s). Remove the temporary seat rail marks.

Report NO: 01973-001 6-22

Issued: June 10.1994 Revision 9: September 1, 1999

1 IPLATUSSr L

SECTION 6

I K12 WEIGHT AND BALANCE

CREW OCCUPANT MOMENTS LB - IN ) ARM 160.27 IN'

Figure 6-7. Moment Chart (Sheet 1 of 4)

CREW OCCUPANT MOMENTS ( Kg - m ) AAM 4.071 m*

Report No: 01973-001 6-23

Issued: June 10,1994

1 Revision 9: September 1, 1999

Arm for center position only. Adjust arm 0.69 inch for each hole from center position. Maximum seat travel is +I- 4 holes or +I- 2.76 inches from center positbn.

WEIGHT

Ib

WEIGHT

Ib

MOMEM

kg - m

WEIGHT

Ib

MOMENT

Ib - in'

MOMENT

Ib - In'

200 32054

210 33657

220 35259

230 36862

240 38465

WEIGHT

kg

MOMENT

kg - m

150 24040

160 25643

170 27246

180 28849

190 30451

MOMENT

Ib - in'

50 8014

60 9616

70 11219

80 12822

90 14424

MOMENT

Ib - in'

WEIGHT

kg

MOMENT

kg - m

WEIGHT

kg

25 101.78

30 122.13

35 142.49

40 162.84

45 183.20

100 16027

110 17630

120 19232

130 20835

140 22438

WEIGHT

Ib

' Arm lor center positiin only. Adjust arm 0.018 meter for each hole fr0m center position. Maximum seat travel is +I- 4 holes or +I- 0.070 meters from center posilion.

75 305.33

80 325.68

85 346.04

90 366.39

95 388.75

50 203.55

55 223.91

60 244.26

65 264.62

70 284.97

100 407.10

105 427.46

110 447.81

115 468.17

120 488.52

MOMENT

K g - m

WEIGHT

kg


REAR BAGGAGE AREA MOMENTS ( LB - IN ) STANDARD NET AT FRAME 34 - ARM 371.0 IN


Report No: 01973-001 Issued: June 10.1994 6-24 Revision 9: September 1, 1999

REAR BAGGAGE AREA MOMENTS ( Kg - m ) STANDARD NET AT FRAME 34 - ARM 9.420 M

MOMENT Ib - in

WEIGHT b

MOMENT kg - m

MOMENT Ib - in

310 115,010 320 118,720 330 122,430 340 126,140 350 129,850 360 133,560 370 137.270 380 140.980 390 144,690 397 147,287

10 3,710 20 7,420 30 11,130 40 14,840 50 18.550

60 22,260 70 25,970 80 29,680 90 33,390 100 37,100

MOMENT kg - m

WEIGHT kg

WEIGHT ks

WEIGHT Ib

MOMENT Ib - in

WEIGHT Ib

155 1460.10 160 1507.20

165 1554.30

170 1601.40 175 1648.50 180 1695.60

WEIGHT ks

WEIGHT Ib

110 40,810 120 44,520 130 48,230 140 51,940 150 55,650 160 59360 170 63,070 180 66,780 190 70,490 200 74,200

MOMENT Ib - in

105 989.10 110 1036.20

115 1083.30

120 1130.40 125 1177.50 130 1224.60 135 1271.70

140 1318.80

145 1365.90

150 1413.00

5 47.10 10 94.20 15 141.30

20 188.40 25 235.50 30 282.60 35 329.70

40 376.80

45 423.90

50 471.00

MOMENT kg-m

210 77,910 220 81,620 230 85,330 240 89,040 250 92,750 260 96,460

270 100,170 280 103.880 290 107,590 300 111,300

55 518.10 60 565.20 65 612.30

70 659.40 75 706.50

80 753.60 85 800.70

90 847.80

95 894.90

100 942.00

WEIGHT kg

MOMENT kg - m



Issued: June 10,1994 Report No: 01973-001 Revision 9: September 1, 1999 6-25

-

REAR BAGGAGE AREA MOMENTS ( LB - IN ) EXTENDABLE NET AT FRAME 32 - ARM 361.0 IN

REAR BAGGAGE AREA MOMENTS ( Kg - m ) EXTENDABLE NET AT FRAME 32 - ARM 9.17 M

WEIGHT Ib

WEIGHT kg

10 3610 20 7220 30 10831 40 14441 50 18051 60 21661 70 25272 80 28882 90 32492 100 36102 110 39713 120 43323 130 46933

MOMENT Ib - in

5 45.85 10 91.70 15 137.55 20 183.40 25 229.25 30 275.10 35 320.95 40 366.80

45 . 412.65 50 458.50 55 504.35 60 550.20 65 596.05

MOMENT kg - m

WEIGHT Ib

MOMENT Ib - in

WEIGHT kg

140 50543 1% 54154 160 57764 170 61374 180 64984 190 68594 200 72205 210 75815 220 79425 230 83035 240 86646 250 90256 260 93866

MOMENT kg - m

MOMENT kg - m

70 641.90 75 687.75 80 733.60 85 779.45 90 825.30

95 871.15 100 917.00 105 962.85 110 1008.70

115 1054.55 120 1100.40 125 1146.25 130 1192.10

WEIGHT Ib

270 97476 280 101087 290 104697 300 108307 310 111917 320 115528 330 119138 340 122748 350 126358 360 129969 370 133579 380 137189 390 140799

WEIGHT kg

135 1237.95 140 1283.80 145 1329.65 150 1375.50 155 1421.35 160 1467.20 165 1513.05 170 1558.90 175 1604.75 180 1650.60 185 1696.45 190 1742.30 195 1788.15

400 144409 410 148020 420 151630 430 155240 440 158850 450 162461 460 166071 470 169681 480 173291 490 176902 500 180512

MOMENT Ib - in

200 1834.00 205 1879.85 210 1925.70 215 1971.55 220 2017.40 225 2063.25

MOMENT kg - m

WEIGHT Ib

WEIGHT kg

MOMENT Ib - in


FUEL LOAD MOMENTS (LB - IN)

NOTE

Unusable fuel is considered in empty weight. The chan shows only additional fuel.

FUEL LOAD MOMENTS ( Kg - m )

Fi~ure 6-7. Moment Chart (Sheet 4 of 4)


MOMENT

Ib - in

WEIGHT

kg

MOMENT

Ib - in

2200 51 1463

2300 534839

2400 558130

2500 581450

2600 604724

2700 628029

WEIGHT

Ib

MOMENT

Ib - in

WEIGHT

Ib

WEIGHT

Ib

MOMENT

kg-m

MOMENT

kg - m

1500 347656

1600 371079

1700 394500

1800 417912

1900 441347

2000 464746

2100 488120

WEIGHT

Ib

100 22572

200 45161

300 67776

400 90443

500 113351

600 136538

700 159955

MOMENT

Ib - in

1100 6497.53

1150 6793.90

1200 7090.37

1250 7385.69

MOMENT

kg-m

WEIGHT

kg

50 286.64

100 573.59

150 860.84

200 1149.27

250 1441.88

300 1738.40

350 2037.52

WEIGHT

kg

800 183555

900 2071 11

1000 230572

1100 253974

1200 277441

1300 300811

1400 324221

WEIGHT

kg

400 2337.14

450 2635.13

500 2932.34

550 3230.45

600 3526.99

650 3824.03

700 4122.29

MOMENT

kg - m

750 4419.61

800 4717.33

850 5014.59

900 5312.14

950 5608.06

1000 5905.10

1050 6201.26


ITEM I MOMENT 1, ib-In

1

1. Baslc Empty Weight

2. Combi Interior Conversion

3. Pilot

PC-12 EXAMPLE LOADING FORM INTERIOR CODE: STD-9s I

1 4. Copilot (Right Seat Passenaer) 1 170

1 5. Passenger 1

6. Passenger 2

7. Passenger 3

8. Passenger 4 170

9. Passenger 5 170

10. Passenaer 6 170

I 1 1. Passenger 7

17. a. Rear Baggage (net at frame 32) 1 215 ( 3 6 L 0 = [ 7 9 7 6 5 \ I b. Rear Baggage (net at frame 34) 370.87 (9.420)

12. Passenger 8

13. Passenger 9

14. Optional Wardrobe

15. LH Cabinet

16. RH Cabinet

170

170

r L

I L

31 1.03 (7.900)

344.03 (8.738)

191 .OO (4.851)

212.10(5.387)

21 1.19 (5.3641

18. Cargo

19. Zero Fuel Welght MZFW 8160 Ib (3700 kg)

I L

t

b

52875

58485

(Sumof 1 thw18)

20. Fuel

7358

Figure 6-8. Example Loading Form (Sheet 1 of 2)

Issued: June 10,1994 Reporl No: 01 973-001 , Revision 11 : March 1,2003 6-27

21. Ramp Welght MRW 9083 lb (4120 kg) (Sumof l 9 + 2 0 )

22. Less Fuel for Ground Operations

23. Add Moment Due to Gear Retraction

24. Takeoff Welght MTOW 9039 Ib (4100 kg) (Sum of 21 -22+23)

1650

229.49

382790 1 9008

40

8968

1 81 4268

-

229.92

229.97

21 97058

9369

538(6.2)

21 88227

SECTION 6 €PILATUS= WEIGHT AND BALANCE PC12

Sum of 1 thru 18

Figure 6-8. Loading Form (Sheet 2 of 2) Report No: 01 973-001 Issued: June 10,1994 6-28 Revision 9: September 1, 1999

SECTIO

N 6

WE

IGH

T AN

D B

ALA

NC

E

Figure 6-9. C. G

. Envelope (S

heet 1 of 2) Issued: June 10,1994


Revision 9: Septem

ber 1, 1999 6-29

SECTIO

N 6

PILAW

S= W

EIGH

T AN

D B

ALA

NC

E

K12

Report N

o: 01973-001 6-30

Figure 6-9. C. G

. Envelope (S

heet 2 of 2) Issued: June 10,1994

Revision 9: Septem

ber 1, 1999


EQUIPMENT LIST

Refer to Pilatus Report No. 02647, Airplane equipment List, attached to the back of this report. The equipment list itemizes the installed equipment included in the Basic Empty Weight indicated in Figure 6-2 of this Airplane Flight Manual.

Issued: June 10.1994 Revision 9: September I, 1999

Report No: 01 973-001 6-31


INTERIOR CONFIGURATIONS 1

The PC-12 was designed and certified initially with two basic cabin interior configurations, a Corporate Commuter (Code STD-9s) and an Executive interior (Code EX-6s). The Corporate Commuter interior consists of two crew seats and 9 standard passenger seats. The Executive interior consists of two crew seats and 6 executive seats with forward storage cabinets and a toilet.

1

Variations to the two basic interior configurations are continuously being developed. The ' various configurations that have been approved are given below. Before using them it is the operators responsibility to check whether they require authorization by their regulatory authority. Some of the configurations require structural and system modifications, check with the Service Bulletin Index for the applicable SB's.

A Code Number is given to each interior configuration. The code will be shown on a placard 1

which will be installed on the cargo door frame (Post SB 25-016). The placard code gives the type and number of seats that are installed in the aircraft. Before making any changes to the interior configuration, contact Pilatus to make sure that any modification work or SB's are identified for embodiment. The placard must then be changed to show the correct code for the new configuration.

It is possible for aircraft with the executive interior to have more than one placard installed on the cargo door frame. The removal or installation of the rear seats must be done in accordance with an approved configuration. The correct weight and moment charts for the configuration must then be used for weight and balance determination for flight.

An optional three seat bench can be installed at the rear of the cabin in a Corporate Commuter and an Executive (Post SB 25-014) aircraft. A large baggage net (Post SB 25-010) can be installed in these configurations.

The following code numbers have been allocated and the seat locations are given in the following sub-sections:

8 CORPORATE COMMUTER lnterior Layout CODE STD-9s (nine standard seats)

8 CORPORATE COMMUTER lnterior Layout CODE STD-6s-38 (six standard seats and three seat bench)

8 EXECUTIVE lnterior Layout CODE EX-6S (six executive seats)

8 EXECUTIVE lnterior Layout CODE EX-BS (eight executive seats)

8 EXECUTIVE lnterior Layout CODE EX-4s-3B (four executive seats and three seat bench)

EXECUTIVE lnterior Layout CODE EX-6s-STD-2S (six executive seats and two standard seats)

EXECUTIVE lnterior Layout CODE EX-4s-STD-4s (four executive seats and four standard seats)

Report NO: 01973-001 6-32


SECTION 6 WEIGHTSAND BALANCE

CORPORATE COMMUTER INTERIOR CODE STD-9S

GENERAL

The basic Corporate Commuter lnterior consisting of 9 standard passenger seats. The section contains the following information:

passenger seat location chart

permitted passenger seat Part Nos. that can be installed

passenger seat and furnishings weight and moment chart (standard and metric units)

passenger seat occupant moment charts (standard and metric units)


Report No: 01 973-001 Interior Code STD-9s. Page 6-01-1

SECTION 6 =PILATUSW WEIGHT AND BALANCE PC12

CORPORATE COMMUTER INTERIOR CODE STD-9S

SEAT LOCATIONS

178.88' 211.17 244.17 277.17 310.17 343.17 FUSELAGE STATION 4.560 m 5.364 m 6.202 m 7.040 m 7.878 m 8.716 rn

DISTANCE FROM 0' 31.28' 64.28' 87.28' 130.28' 163.28' DIVIDER AFT SURFACE 0 m 0.705 m 1.633 m 2.471 m 3.308 m 4.147 m

I

I DISTANCE FROM 0' 34 2 5 67.25' 100.25' 133.25 DIVIDER AFT SURFACE 0 m 0.870 m 1.708 m 2.546 m 3.385 m

FUSELAGE 7 7 % 214.13' 247.13 280.13' 313.13' 5.436 m 6.277 m 7.1 15 rn 7.954 m

NOTE: PAX 9 SEAT INSTALLATION IS NOT PERMITTED ON THE LEFT HAND SIDE OF THE AIRPLANE CABIN.

CENTER ARRESTING PIN

NOTE: CABIN SEAT LOCATION IS DEFINED AS THE DISTANCE FROM THE AFT SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE CENTER ARRESTING PIN ON EACH SEAT.

Report No: 01973-001 Interior Code STD-9s. Page 6-01-2



CORPORATE COMMUTER INTERIOR CODE STD-SS

PERMITTED PASSENGER SEAT PART Nos. THAT CAN BE INSTALLED

SEAT NO.

NOTE:



PC-1 2

The lap belt extension Part No. 959.30.01.588 can be installed on all of the above seats.

flXED CUSHIONS

959.30.01.501 959.30.01.503 959.30.01.505 959.30.01.507 959.30.01.509 959.30.01.51 1 959.30.01.513 959.30.01.515 959.30.01.517 959.30.01.519

959.30.01.502 959.30.01.504 959.30.01.506 959.30.01.508 959.30.01.510 959.30.01.512 959.30.01 514 959.30.01.516 959.30.01.518 959.30.01.520

PC-1 2145

REMOVABLE CUSHIONS

525.22.12.01 1

525.22.12.012

FIXED CUSHIONS

959.30.01.51 1 959.30.01.513 959.30.01.515 959.30.01.517 959.30.01.51 9

959.30.01.512 959.30.01.514 959.30.01.516 959.30.01.518 959.30.01.520

REMOVABLE CUSHIONS

525.22.1 2.01 1

525.22.12.012


CORPORATE COMMUTER INTERIOR CODE STD-QS

PASSENGER SEATS AND FURNISHINGS WEIGHT AND MOMENT CHART 1 i

Adjust the aircraft Basic Empty Weight on the Loading Form for items removedladded when convening to or from a Combi Interior Conversion. 4

1

When installing the extendable baggage net refer to Section 2 tor the Luggage Limitations.

FR 24 CARGO NET

FR 27 CARGO NET

FR 32 EXTENDABLE BAGGAGE NET

FR 34 BAGGAGE NET

RepoR No: 01973-001 Interior Code STD-9s. Page 6-01-4


3.6 (1.65)

3.6 (1.65)

6.44 (2.92)

3.6 (1.65)

941 (10.96)

1049 (12.21)

2325 (26.78)

1335 (15.13)


CORPORATE COMMUTER INTERIOR CODE STD-SS

PASSENGER SEAT OCCUPANT MOMENT CHART

Issued: June 10.1994 Revision 9: September I , 1999

WEIGHT

Ib

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240


PAX 1

215.00 in

10750

12900

15050

17200

19350

21500

23650

25800

27950

30100

32250

34400

36550

38700

40850

43000

45150

47300

49450

51600

PASSENGER

PAX2

212.03 in

10602

12722

14842

16963

19083

21203

23323

25444

27564

29684

31805

33925

36045

38166

40286

42406

44527

46647

48767

50888

SEAT

PAX3

248.00 in

12400

14880

17360

19840

22320

24800

27280

29760

32240

34720

37200

39680

42160

44640

47120

49600

52080

54560

57040

59520

OCCUPANT

PAX4

245.03 in

12252

14702

17152

19603

22053

'24503

26953

29404

31854

34304

36755

39205

41655

44106

46556

49006

51457

53907

56357

58808

PAX5

281.00 in

14050

16860

19670

22480

25290

28100

30910

33720

36530

39340

42150

44960

47770

50580

53390

56200

59010

61820

64630

67440

LB - IN )

PAX?

314.00 in

15700

18840

21980

25120

28260

31400

34540

37680

40820

43960

47100

50240

53380

56520

59660

62800

65940

69080

72220

75360

MOMENTS (

PAX6

278.03 in

13902

16682

19462

22243

25023

27803

30583

33364

36144

38924

41705

44485

47265

50046

52826

55606

58387

61167

63947

66728

PAX8

311.03 in

15552

18662

21772

24883

27993

31103

34213

37324

40434

43544

46655

49765

52875

55986

59096

62206

65317

68427

71537

74648

PAX9

344.03 in

17202

20642

24082

27523

30963

34403

37843

41284

44724

48164

51605

55045

58485

61926

65366

68806

72247

75687

79127

82568


CORPORATE COMMUTER INTERIOR CODE STD-OS


Report No: 01973-001 Interior Code STD-9s. Page 6-01 -6


PAXQ

8.738 m

218.5

262.2

305.8

349.5

393.2

436.9

400.6

524.3

568.0

611.7

655.4

699.1

742.8

786.5

830.1

873.8

917.5

961.2

1004.9

1048.6

WEIGHT

kg

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

PAX 1

5.461 m

136.5

163.8

191.1

218.4

245.7

273.1

300.4

327.7

355.0

382.3

409.6

436.9

464.2

491.5

518.8

546.1

573.4

600.7

628.0

655.3

PASSENGER

PAX2

5.386 m

134.6

161.6

188.5

215.4

242.4

269.3

296.2

323.1

350.1

377.0

403.9

430.8

457.8

484.7

511.6

538.6

565.5

592.4

619.3

646.3

SEAT

PAX3

6.299 m

157.5

189.0

220.5

252.0

283.5

315.0

346.5

378.0

409.4

440.9

472.4

503.9

535.4

566.9

598.4

629.9

661.4

692.9

724.4

755.9

OCCUPANT

PAX5

7.137 m

178.4

214.1

249.8

285.5

321.2

356.9

392.6

428.2

463.9

499.6

535.3

571.0

606.7

642.4

678.1

713.7

749.4

785.1

820.8

856.5

PAX4

6.224 m

155.6

186.7

217.8

249.0

280.1

311.2

342.3

373.4

404.5

435.7

466.8

497.9

529.0

560.1

591.3

622.4

653.5

684.6

715.7

746.9

MOMENTS

PAX6

7.062 m

176.6

211.9

247.2

282.5

317.8

353.1

388.4

423.7

459.0

494.3

529.7

565.0

600.3

635.6

670.9

706.2

741.5

776.8

812.1

847.4 - --

( KG - M PAX7

7.976 m

199.4

239.3

279.1

319.0

358.9

398.8

438.7

478.5

518.4

558.3

598.2

638.0

677.9

717.8

757.7

797.6

837.4

877.3

917.2

957.1 -

)

PAX8

7.900 m

197.5

237.0

276.5

316.0

355.5

395.0

434.5

474.0

513.5

553.0

592.5

632.0

671.5

711.0

750.5

790.0

829.5

869.0

908.5

948.0 - - --


CORPORATE COMMUTER INTERIOR CODE STD-6s-3B

GENERAL

This configuration is a variation of the basic Corporate Commuter interior and consists of 6 standard passenger seats and a 3 seat bench. It is the operators responsibility to check before using this configuration whether they require authorization by their regulatory authority. The following information is given:



passenger seats and furnishings weight and moment chart (standard and metric units)



Report No: 01 973-001 Interior Code STD-6s-36. Page 6-02-1



SEAT LOCATIONS

178.88- 2 1 1 . 1 ~ 244.17- 277.17- FUSELAGE STATION 4.569 m 5.364 m 6.202 m 7 . W m

DISTANCE FROM 0' 31.28" M.28' 87.28' DIVIDER AFT SURFACE 0 m 0.795 m 1.633 m 2.471 m

I

DISTANCE FROM i- 32.25 65.25 98.25' 134.25' DIVIDER AFT SURFACE 0 m 0.819 m 1.657 m 2.496 m 3.410 m

FWD FACING SEAT CENTER AWESTING PIN C

NOTE:

Report No: 01 973-001 Interior Code STD-6s-3B. Page 6-02-2

I

CENTER OF ARRESTING PIN

Issued: June 10,1994 Revision 1 1 : March I, 2003



I

b PERMlTlED PASSENGER SEAT PART Nos. THAT CAN BE INSTALLED

I NOTES: The extendable baggage net Post SB 25-010 can be used with this interior.

SEAT NO.

1,3,5

2,4,6

7,8,9

The lap belt extension Part No. 959.30.01.588 can be installed on seats 1 thru 6 only.

A bulkhead plus curtain Part No. 525.24.12.023 can be installed at frame 32.

Issued: June 10,1994 Revision 1 1 : March 1, 2003 ,

FIXED CUSHIONS

501.30.01.501 501.30.01.503 501.30.01.505 501.30.01.507 501.30.01.509 501.30.01.51 1 501 .30.01.513 501.30.01.515 501.30.01.51 7 501.30.01.519

501 .30.01.502 501.30.01.504 501.30.01.506 501.30.01.508 501.30.01.51 0 501 .30.01.512 501.30.01.514 501.30.01.516 501.30.01.518 501.30.01.520

959.30.01.801 959.30.01.802 959.30.01.803 959.30.01.804 959.30.01.805 (Bench Seat)

Report No: 01973-001 Interior Code STD-6s-38. Page 6-02-3

PC-12

REMOVABLE CUSHIONS

525.22.12.01 1

525.22.12.01 2

959.30.01.801 959.30.01.802 959.30.01.803 959.30.01.804 959.30.01.805 (Bench Seat)

PC-1

FIXED CUSHIONS

501.30.01.51 1 501.30.01.513 501.30.01.515 501.30.01.517 501.30.01.519

501 .30.01.512 501.30.01.514 501.30.01.516 501.30.01.518 501 .30.01.520

959.30.01.801 959.30.01.802 959.30.01.803 959.30.01.804 959.30.01 805 (Bench Seat)

2/45

REMOVABLE CUSHIONS

525.22.12.01 1

525.22.12.012

959.30.01.801 959.30.01 802 959.30.01.803 959.30.01.804 959.30.01.805 (Bench Seat)


- -- --

CORPORATE COMMUTER INTERIOR CODE STD-6S9B

PASSENGER SEATS AND FURNISHINGS WEIGHT AND MOMENT CHART

Adjust the aircraft Basic Empty Weight on the Loading Form for items removedladded when converting to or from a Combi Interior Conversion.

I Report No: 01 973-001 ' Interior Code STD-6S-36. Page 6-02-4 Issued: June 10, 1994



CORPORATE COMMUTER INTERIOR CODE STDbS3B



WElG HT

Ib

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

Report No: 01973-001 Interior Code STD-6s-3B. Page 6-02-5

PAX 1

213.0 0 in

10650

12780

14910

17040

19170

21300

23430

25560

27689

29819

31949

34079

36209

38339

40469

42599

44729

46859

48989

51119

PASSENGER

PAX 2

212.0 3 in

10602

12722

14842

16963

19083

21203

23323

25444

27564

29684

31805

33925

36045

38166

40286

42406

44527

46647

48767

50888

SEAT

PAX 3

246.0 0 in

12300

14760

17220

19680

22140

24600

27060

29520

31979

34439

36899

39359

41819

44279

46739

49199

51659

54119

56579

59039

PAX 5

279.0 0 in

13950

16740

19530

22320

25110

27900

30690

33480

36269

39059

41849

44639

47429

50219

53009

55799

58589

61379

64169

66959

OCCUPANT

PAX 4

245.0 3 in

12252

14702

17152

19603

22053

24503

26953

29404

31854

34304

36755

39205

41655

44106

46556

49006

51457

53907

56357

58808

MOMENTS (

PAX 6

278.0 3 in

13902

16682

19462

22243

25023

27803

30583

33364

36144

38924

41705

44485

47265

50046

52826

55606

58387

61167

63947

66728

LB - IN )

PAX 7

321.1 2 in

16056

19267

22478

25690

28901

32112

35323

38534

41746

44957

48168

51379

54590

57802

61013

64224

67435

70646

72958

77069

PAX 8

321.1 2 in

16056

19267

22478

25690

28901

32112

35323

38534

41746

44957

48168

51379

54590

57802

61013

64224

67435

70646

73858

77069

PAX 9

321.1 2 in

16056

19267

22478

25690

28901

32112

35323

38534

41746

44957

48168

51379

54590

57802

61013

64224

67435

70646

73858

77069

SECTION 6 WEIGHT AND BALANCE - -

CORPORATE COMMUTER INTERIOR CODE ST0-6S-38


R w r t NO: 01973-001 Interior Code STD-6s-38. Page 6-02-6


)

PAX 8

8.156 m

203.91

244.69

285.48

326.26

367.04

407.82

448.60

489.39

530.17

570.95

611.73

652.52

693.30

734.08

774.86

815.64

856.43

897.21

937.99

978.77

PAX 9

8.156 m

203.91

244.69

285.48

326.26

367.04

407.82

448.60

489.39

530.17

570.95

61 1.73

652.52

693.30

734.08

774.86

815.64

856.43

897.21

937.99

978.77

MOMENTS

PAX 6

7.061 m

176.55

211.86

247.17

282.48

317.79

353.10

388.41

423.72

459.03

494.34

529.65

564.96

600.27

635.58

670.89

706.20

741.51

776.82

812.13

847.44

OCCUPANT

PAX 5

7.087 m

177.16

212.60

248.03

283.46

318.89

354.33

389.76

425.19

460.62

496.06

531.49

566.92

602.35

637.79

673.22

708.65

744.08

779.52

814.95

850.38

( KG - M

PAX 7

8.156 m

203.91

244.69

285.48

326.26

367.04

407.82

448.60

489.39

530.17

570.95

61 1.73

652.52

693.30

734.08

774.86

815.64

856.43

897.21

937.99

978.77

WEIGHT

ko

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

SEAT

PAX 3

6.248 m

156.21

187.45

218.69

249.93

281.17

312.42

343.66

374.90

406.14

437.38

468.62

499.86

531.11

562.35

593.59

624.83

656.07

687.31

718.55

749.80

PAX 4

6.224 m

155.60

186.71

217.83

248.95

280.07

311.19

342.31

373.43

404.55

435.67

466.79

497.90

529.02

560.14

591.26

622.38

653.50

684.62

715.74

746.86

PAX 1

5.410 m

135.25

162.30

189.35

216.40

243.45

270.51

297.56

324.61

351.66

378.71

405.76

432.81

459.86

486.91

513.96

541.01

568.06

595.11

622.16

649.21

PASSENGER

PAX 2

5.386 m

134.64

161.57

188.50

215.42

242.35

269.28

296.21

323.14

350.06

376.99

403.92

430.85

457.78

484.70

511.63

538.56

565.49

592.42

619.34

646.27


EXECUTIVE INTERIOR CODE EX6S-1

GENERAL

The basic Executive Interior consisting of 6 executive passenger seats. There are two seat location configurations for the basic executive interior, they have been given the codes EX-6s- 1 (MSN 101-171) and EX-6s-2 (MSN 172 and UP). The section contains the following information:

passenger seat location charts for EX-6s-1

permitted passenger seat Part Nos. that can be installed for EX-6s-1

passenger seats and furnishings weight and moment charts for EX-6s-1 (standard and metric units)

passenger seat occupant moment charts for EX-6s-1 (standard and metric units)


Report No: 01 973-001 lnterior Code EX-6s-1. Page 6-03-1


EXECUTIVE INTERIOR CODE EX-GS-1

SEAT LOCATIONS CD PLAYER

c A & T (OPTIONAL)

DISTANCE FROM 0 10.05- 31.88' 31.30- DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.795 m

179.85 190.00' 211.85' 211.18' 4.571 m 4.826 m 5.381 rn 5.364 m

PAX 6

I I I ARRESTINO PIN LOCATION

CA&ET PA XI^ PAX w I PAX 5

DISTANCE FROM C 31.88' 56.W 89.18- 137.18' DIVIDER A F I SURFACE 0 m 0.809 rn 1.427 m 2.265 m 3.484 rn

179.85' 21 1.83' 236.13' 269.13' 317.13' FUSELAGE 4.571 m 5.380 m 5.998 m 6.836 m 8.055 m

AFT FACINQ SEAT FWD FACINQ SEAT

f -l

CENTER OF

NOTE: CABIN SEAT LOCATION IS DEFINED AS THE DISTANCE FROM THE AFT SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE ARRESTING PIN ON EACH SEAT.

RepoR No: 01 973-001 ' Interior Code EX-6s-1. Page 6-03-2 Issued: June 10,1994

Revision 11 : March 1.2003


EXECUTIVE INTERIOR CODE EX-6s-1



SEAT NO.

1

2

3,5

4,6

Report No: 01973-001 Interior Code EX-6s-1. Page 6-03-3

PC-1 2 AND PC-1 2145

MSN

PACIFIC SCIENTIFIC RESTRAINT

959.30.01.601

959.30.01.602

959.30.01.603 959.30.01.609

959.30.01.604 959.30.01.610

101-171

SCHROTH RESTRAINT

959.30.01.61 3 or 959.30.01.625 959.30.01.617 959.30.01.627 959.30.01.619 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.614 or 959.30.01.626 959.30.01.618 959.30.01.628 959.30.01.620 959.30.01.630 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01.661 959.30.01.645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01.654 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648




Adjust the aircraft Basic Empty Weight on the Loading Form for items removed/added when converting to or from a Combi Interior Conversion.

CD PLAYER (optional)

When installing the extendable baggage net refer to Section 2 for the Luggage Limitations.

FR 24 CARGO NET

FR 27 CARGO NET


FR 34 BAGGAGE NET

Report No: 01 973-001 Interior Code EX-6s-1. Page 6-03-4

~ssued: June 10,1994 Revision 9: September 1, 1999 ,

3.6 (1.65)

3.6 (1.65)

6.44 (2.92)

3.6 (1.65)

941 (10.96)

1049 (12.21)

2325 (26.78)

1335 (15.1 3)

L EPILATUSF SECTION 6 b P(12 WEIGHT AND BALANCE

EXECUTIVE INTERIOR CODE EX6S-1



1

WEIGHT

Ib

50

60

70

80

90

100

110

1 20

130

140

150

160

170

180

190

200

210

220

230

240

Report No: 01973-001 Interior Code EX-6S-1. Page 6-03-5

PASSENGER SEAT

PAX 112

234.09 in

11705

14046

16387

18728

21068

23409

25750

28091

30432

32773

351 14

37455

39796

42137

44478

46819

49160

51501

53842

56183

OCCUPANT

PAX 3 4

276.12 in

13806

16567

19329

22090

. 24851

27612

30374

33135

35896

38657

41419

44180

46941

49702

52464

55225

57986

60747

63508

66270

MOMENTS ( LB - PAX

5

324.12 in

16206

19447

22688

25929

291 70

32412

35653

38894

42135

45376

4861 7

51859

55100

58341

61582

64823

68064

71306

74547

77788

IN )

PAX 6

334.12 in

16706

20047

23388

26729

30070

33412

36753

40094

43435

46776

501 17

53459

56800

60141

63482

66823

70 164

73506

76847

80188

SECTiON 6 WEIGHT AND BALANCE

EXECUTIVE INTERIOR CODE EX-1 4



>


- M )

PAX 6

8.487 rn

212.16

254.60

297.03

339.46

381.89

424.33

466.76

509.19

551.63

594.06

636.49

678.92

721.36

763.79

806.22

848.65

891.09

933.52

975.95

1018.39

WEIGHT

ka

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

OCCUPANT

PAX 3/14

7.014 rn

175.34

210.41

245.47

280.54

315.61

350.68

385.75

420.81

455.88

490.95

526.02

561.08

596.15

631.22

666.29

701.35

736.42

771.49

806.56

841.63

PASSENGER SEAT

PAX li2

5.946 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

654.06

683.79

71 3.52

MOMENTS ( KG

PAX 5

8.233 rn

205.81

246.98

288.14

329.30

370.46

41 1.63

452.79

493.95

535.12

576.28

617.44

658.60

699.77

740.93

782.09

823.25

864.42

905.58

946.74

987.91



GENERAL

The basic Executive lnterior consisting of 6 executive passenger seats. There are two seal location configurations for the basic executive interior, they have been given the codes EX-6s- 1 (MSN 101-171) and EX-6s-2 (MSN 172 and UP). The section contains the following information:

passenger seat location charts for EX-6s-2

permitted passenger seat Part Nos. that can be installed for EX-6s-2

passenger seats and furnishings weight and moment charts for EX-6s-2 (standard and metric units)

passenger seat bccupant moment charts for EX-6s-2 (standard and metric units)

Issued: June 10, 1994

I Revision 9: September 1, 1999 Report No: 01 973-001

lnterior Code EX-6s-2. Page 6-04-1



SEAT LOCATIONS

RH CDPUYER CABlNfT (OPTIONAL) PAX 8

DISTANCE FROM 0' 10.05 31.89' 31.30' 147.18' DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.795 m 3.738 m

179.95 190.00' 211.85 21 1.19' FUsEUGEsTATIoN 4.571 m 4.826 m 5.381 m 5.364 m 327.13'

8.309 m

ARMSTINQ PIN LOCATION

ET PAX 1/2 PAX 314

DISTANCE FROM 0' 31.88' 56.1C 89.lC 135.18' DIVIDER AFT SURFACE 0 m 0.809 m 1.427 m 2.265 m 3.435 m

F U S E ~ ~ G ~ STATION :,77!% ::I2% 236.13' 269:13' 315.13' 5.998 m 8.836 m &OM m

AFT FACING SEAT FWD F A C W SEAT

I

NOTE: CABIN SEAT LOCATION IS DEFWED AS THE DISTANCE FROM THE A R SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE ARRESTING PIN ON EACH SEAT.

Repon NO: 01973-001 Interior Code EX-6s-2. Page 6-04-2

Issued: June 10,1994 Revision 1 1 : March 1. 2003


EXECUTIVE INTERIOR CODE EXbS-2

PERMlllED PASSENGER PART Nos. THAT CAN BE INSTALLED -

Issued: June 10,1994 I Revision 1 1 : March 1,2003

SEAT NO.

1

2

3,5

4.6


PO12 AND PC-12/45

MSN 172


959.30.01.601

959.30.01.602

959.30.01.609

959.30.01.610

and UP

SCHROTH RESTRAINT

959.30.01.613 or 959.30.01.625 959.30.01.617 959.30.01.627 959.30.01.619 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.614 or 959.30.01.626 959.30.01.61 8 959.30.01.628 959.30.01.620 959.30.01.630 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01.661 959.30.01.645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01.654 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648


a PILATUS W PC12



Adjust the aircraft Basic Empty Weight on the Loading Form for items removedladded when converting to or from a Combi lnterior Conversion.


PC-12 AND PC-12/45


Report No: 01973-001 lnterior Code EX-6s-2. Page 6-04-4

FR 24 CARGO NET

FR 27 CARGO NET


FR 34 BAGGAGE NET


3.6 (1.65)

3.6 (1.65)

6.44 (2.92)

3.6 (1 65)

941 (10.96)

1049 (12.21)

2325 (26.78)

1335 (15.13)


- -

EXECUTIVE INTERIOR CODE EXdS-2



I

WEIGHT

Ib

50

60

70

80

90

100

110

120

1 30

140

1 50

160

170

1 80

190

200

210

220

230

240


PASSENGER SEAT

PAX 1l2

234.09 in

11705

14046

16387

18728

21068

23409

25750

28091

30432

32773

351 14

37455

39796

421 37

44478

46819

49160

51501

53842

56183

OCCUPANT

PAX 314

276.12 in

13806

16567

19329

22090

24851

27612

30374

33135

35896

38657

41419

44 180

46941

49702

52464

55225

57986

60747

63508

66270

MOMENTS ( LB - PAX

5

322.13 in

1 61 07

19328

22549

25771

28992

322 13

35434

38656

41877

45098

48320

51541

54762

57984

6 1205

64426

67648

70869

74090

773 12

IN )

PAX 6

334.12 in

16706

20047

23388

26729

30070

33412

36753

40094

43435

46776

501 17

53459

56800

601 41

63482

66823

70 1 64

73506

76847

801 88


- - -



Repon NO: 01973-WI lnteriir Code EX-6s-2. Page 6-04-6

EXECUTIVE INTERIOR CODE EX 6s-2


- M)

PAX 6

8.487 m

212.16

254.60

297.03

339.46

38 1.89

424.33

466.76

509.19

55 1.63

594.06

636.49

678.92

721.36

763.79

806.22

848.65

891.09

933.52

975.95

1018.39 -

WEIGHT

kQ

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

PASSENGER SEAT

PAX 112

5.946 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

69.06

683.79

713.52

OCCUPANT

PAX 314

7.014 m

175.34

210.41

245.47

280.54

315.61

350.68

385.75

420.81

455.88

490.95

526.02

56 1.08

596.15

631.22

666.29

701.35

736.42

ni .49

806.56

841.63

MOMENTS ( KG

PAX 5

8.182 m

204.55

245.46

286.38

327.29

368.20

409.1 1

450.02

490.93

531.84

572.75

613.66

654.57

695.48

736.39

77730

818.21

859.13

900.04

940.95

981.86


EXECUTIVE INTERIOR CODE EX8S

GENERAL

This configuration is a variation of the basic executive interior and consists of 8 executive passenger seats. It is the operators responsibility to check before using this configuration whether they require authorization by their regulatory authority. The following information is given:






Report No: 01 973-00 1 Interior Code EX-8s. Page 6-05- 1


EXECUTIVE INTERIOR CODE EX-8S

SEAT LOCATIONS

RH CDPUYER CABINET (OPTIONAL)

DISTANCE FROM 0' 10.05. 31.89' 31.30' DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.795 m

179.85 1SO.M)' 211.85' 211.19' FUSELAQEsTAnON 4.571 m 4.826 m 5.361 m 5.364 m

PAX 112 PAX w PAX sm PAXM

DISTANCE FROM 0' 31.88' 58.1C 8S.lB' 121.16' 153.1(r DIVIDER AFT SURFACE 0 m 0.809 m 1.427 m 2.265 m 3.078 m 3.891 m

178.95' 211.83' 236.14' 269.13. 301.13' 333.13' FUSELAGE STAnoN 4.571 m 5.380 m 5.998 m 8.836 m 7.648 m 8.462 m

AFT FACING SEAT FWD FACING SEAT

CENTER OF

NOTE: CABIN SEAT LOCAT(ON IS DEFINE0 AS THE OSTAhCE FROM THE A F l SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE ARRESTING PIN ON EACH SEAT.

Report No: 01 973-001 Interior Code EX-8s. Page 6-05-2

Issued: June 10,1994 Revision 11 : March 1, 2003


EXECUTIVE INTERIOR CODE EX4S



SEAT NO.

1

2

3,5, 7

4,6,8

Report No: 01 973-001 lnlerior Code EX-8s. Page 6-05-3

PC-1 2 AND PC-1 2/45

959.30.01.613 or 959.30.01.625 959.30.01.61 7 959.30.01.627 959.30.01.619 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.61 4 or 959.30.01.626 959.30.01.618 959.30.01.628 959.30.01.620 959.30.01 $30 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01 661 959.30.01.645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01 $54 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648

SECTION 6 FSPILATUS= WEIGHT AND BALANCE PC12 'I

4

EXECUTIVE INTERIOR CODE EX-IS

PASSENGER SEATS AND FURNISHINGS WEIGHT AND MOMENT CHART I

1 1

I

BAGGAGE NET

Adjust the aircraft Basic Empty Weight on the Loading Form for items removedfadded when converting to or from a Combi Interior Conversion.

When installing the extendable baggage net refer to Section 2 for the Luggage Limitations. ,

Report No: 01973-001 Interior Code EX-8s. Page 6-05-4

Issued: June 10,1994 Revision 9: September 1, 1999 ,

I

I GPILATUSm SECTION 6 K12 WEIGHT AND BALANCE

EXECUTIVE INTERIOR CODE EX-BS


Issued: June 10.1994 Revision 9: Seplember 1, 1999

L

WEIGHT

Ib

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

Repon No: 01973-001 Interior Code EX-8s. Page 6-05-5

PASSENGER SEAT

PAX 112

234.09 in

11705

14046

16387

18728

21068

23409

25750

28091

30432

32773

351 14

37455

39796

42137

44478

46819

49160

51501

53842

56183

OCCUPANT

PAX 314

276.12 in

13806

16567

19329

22090

24851

27612

30374

33135

35896

38657

41419

44180

46941

49702

52464

55225

57986

60747

63508

66270

MOMENTS ( LB

PAX 516

308.12 in

15406

18487

21569

24650

27731

308 12

33894

36975

40056

43137

46219

49300

52381

55462

58544

61625

64706

67787

70868

73950

IN )

PAX 7ls

340.12 in

17006

20407

23809

27210

3061 1

34012

37414

40815

44216

476 17

51019

54420

57821

61222

64624

68025

7 1426

74827

78228

81630


EXECUTIVE INTERIOR CODE EX-8s

PASSENGER SEAT OCCUPANT MOMENT CHART i

Report No: 01973-001 Interior Code EX-8s. Page 6-05-6

-

1


( KG - M )

PAX 718

8.639 m

215.98

259.17

302.37

345.57

388.76

431.96

475.15

518.35

561.54

604.74

647.94

691.13

734.33

777.52

820.72

863.91

907.1 1

950.31

993.50

1036.70

MOMENTS

PAX 516

7.826 m

195.66

234.79

273.92

313.05

352.19

391.32

430.45

469.58

508.71

547.84

586.98

626.1 1

665.24

704.37

743.50

782.63

821.77

860.90

900.03

939.16

CREW AND

WEIGHT

k~

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

1 20

PASSENGER

PAX 112

5.946 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

654.06

683.79

713.52

SEAT OCCUPANT

PAX 314

7.014 m

175.34

210.41

245.47

280.84

315.61

350.68

385.75

420.81

455.88

490.95

526.02

561.08

596.15

631.22

666.29

701.35

736.42

771.49

806.56

841.63


EXECUTIVE INTERIOR CODE EX-4s-3B

GENERAL

This configuration is a variation of the basic Executive interior and consists of 4 executive passenger seats and a 3 seat bench. It is the operators responsibility to check before using this configuration whether they require authorization by their regulatory authority. The following information is given:


permitted.passenger seat Part Nos. that can be installed


passenger seat occupant moments (standard and metric units)


1 Revision 9: September I, 1999 Report No: 01973-001

Interior Code EX-4s-3B. Page 6-06-1

SECTION 6 XPILATUSW WEIGHT AND BALANCE PC12


SEAT LOCATIONS

RH CDPLAYER CABINET (OPIIONAL)

DISTANCE FROM 0' 10 05' 31 88' 31.30 DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.785 m

178.85 180.00' 211.86 211.18' FUSELAGE 4.571 m 4.826 m 5.381 m 5.364 m

ARRESTIN5 PIN LOCATION

DISTANCE FROM 0' 31.88' 56.18' 89.1B 133.18' DIVIDER AFT SURFACE 0 m 0.808 m 1.427 m 2.265 m 3.383 m

178.95' 211.83' 236.13' 269.13' . 313.15 FUSELAGE 4.571 m 5.380 rn 5.888 m 6.836 m 7.954 m

AFT FACINQ SEAT FWD FACINQ SEAT

CENTER OF ARRESTING PIN ---------

NOTE: . CABIN SEAT LOCATION IS DEFINED AS THE DISTANCE FROM THE AFT SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE ARRESTING PIN ON EACH SEAT.

Report No: 01973-001 Interior Code EX-4s-3B. Page 6-06-2


~- ~

%PILATUSP SECTION 6 PC12 WEIGHT AND BALANCE

EXECUTIVE INTERIOR CODE EX4S-3B


NOTES: The extendable baggage net Post SB 25-010 can be used with this interior.

A bulkhead plus curtain Part No. 525.24.12.023 can be installed at frame 32.

SEAT NO.

1

2

3

4

5,6,7



PC-12 AND


959.30.01.601

959.30.01.602

959.30.01.603 959.30.01.609

959.30.01.604 959.30.01.610

PC-12/45

SCHROTH RESTRAINT

959.30.01.613 or 959.30.01.625 959.30.01.617 959.30.01.627 959.30.01.61 9 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.614 or 959.30.01.626 959.30.01.618 959.30.01.628 959.30.01.620 959.30.01.630 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01.661 959.30.01 645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01.654 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648

959.30.01.801 or 959.30.01.804 959.30.01 302 959.30.01 305 959.30.01 303 (Bench Seat)



PASSENGER SEATS AND FURNISHINGS WEIGHT AND MOMENT CHART 4

1

Adjust the aircraft Basic Empty Weight on the Loading Form for items removedladded when 4

convening to or from a Combi Interior Conversion.


1


Issued: June 10, 1994 Revision 9: September 1, 1999 ,



PASSENGER SEAT OCCUPANT MOMENTS


I

PASSENGER

WEIGHT

Ib

50

60

70

60

90

100

110

120

130

140

150

160

1 70

180

190

200

210

220

230

240

Report No: 01973-001 lnteriir Code EX-4s-3B. Page 6-06-5

SEAT OCCUPANT

PAX 112

234.09 In

11705

14046

16387

18728

21068

23409

25750

28091

30432

32773

351 14

37455

39796

42137

44478

46819

49160

51501

53842

56183 ~

MOMENTS (

PAX 314

276.12 in

13806

16567

19329

22090

24851

27612

30374

33135

35896

38657

41419

441 80

4694 1

49702

52464

55225

57986

60747

63508

66270

LB - IN )

PAX 51617

320.12 in

16006

19207

22409

25610

2881 1

32012

35214

38415

41616

44817

48019

51 220

54421

57622

60824

64025

67226

70427

73628

76830

SECTION 6 WEIGHT AND BALANCE --


PASSENGER SEAT OCCUPANT MOMENTS



( KGM )

PAX 51W

8.131 m

203.28

243.93

284.59

325.25

365.90

406.56

447.21

487.87

528.52

569.18

609.84

650.49

691.15

731.80

772.46

813.1 1

853.77

894.43

935.08

975.74

WEIGHT

ka

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

PASSENGER SEAT

PAX 112

5.946 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

654.06

683.79

713.52

OCCUPANT MOMENTS

PAX 314

7.014 m

175.34

210.41

245.47

280.54

315.61

350.68

385.75

420.81

455.88

490.95

526.02

561.08

596.15

631.22

666.29

701.35

736.42

771.49

806.56

841.63


EXECUTIVE INTERIOR CODE EXbS-STD-2s

GENERAL

This configuration is a variation of the basic Executive interior and consists of 6 executive passenger seats and 2 standard seats. It is the operators responsibility to check before using this configuration whether they require authorization by their regulatory authorily. The following information is given:



passenger seats and furnishings weigM and moment chart

passenger seat occupant moments (standard and metric units)


Reporl No: 01973-001 Interior Code EX-6s-STD-2s. Page 6-07-1

SECTION 6 WPILATUSV WEIGHT AND BALANCE PC12

SIX EXECUTIVE AND TWO STANDARD INTERIOR CODE EX-6s-STD-2s

SEAT LOCATIONS

LAVATORY1 RH CD PLAYER WARDROBE CABINET (OPTIONAL)

DISTANCE FROM 0' 10.05. 31.89' 31.30" DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.785 m

178.85 19O.W* 211.85' 211.19' FUSELAGE 4.571 m 4.826 m 5.381 m 5.364 m

DISTANCE FROM 0' 31.88' 56.18' 89.18' 121.11T 160.18" DIVIDER AFT SURFACE 0 m 0.809 m 1.427 m 2.265 m 3.078 m 4.067 m

179.85' 211.83" 236.14' 269.13' 301.13' 940.13' FUSELAGE 4.571 rn 5.380 m 5.998 m 6.836 m 7.649 m 8.638 m

AFT FACJNO SEAT FWD FAUNQ SEAT

CENTER OF ARRESTING PIN ----------

NOTE: C4BIN SEAT LOCATION IS DEFINED AS THE DISTANCE FROM THE SURFACE ON THE FORWARD DIVIDER PANEL TO THE CENTER OF THE ARRESTING PIN ON EACH SEAT. CENTER ARRESTING PIN

CABIN SEAT 7 AND 8 LOCATION IS DEFINED FROM THE CENTER OF THE CENTER ARRESTM PIN ON EACH SEAT.

Report No: 01 973-001 lnteiir Code EX-6s-STD-2s. Page 6-07-2



EXECUTIVE INTERIOR CODE EX6S-STD-2s

PERMlllED PASSENGER SEAT PART Nos. THAT CAN BE INSTALLED


SEAT NO.

1

2

3,5

4,6

7

8

Report NO: 01973-001 Interior Code EX-6s-STD-2s. Page 6-07-3

PC-12 AND PC-12/45

' 959.30.01.613 or 959.30.01.625 959.30.01.617 959.30.01.627 959.30.01.619 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.614 or 959.30.01.626 959.30.01.61 8 959.30.01.628 959.30.01.620 959.30.01.630 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01.661 959.30.01.645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01.654 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648

525.22.12.01 1

525.22.12.012


EXECUTIVE INTERIOR CODE EX6S-STD-2s

PASSENGER SEATS AND FURNISHINGS WEIGHT AND MOMENT CHART 4

Adjust the aircraft Basic Empty Weight on the Loading Form for items removedladded when : converting to or from a Combi Interior Conversion.

FA 24 CARGO NET

FR 27 CARGO NET


FR 34 BAGGAGE NET

When installing the extendable baggage net refer to Section 2 for the Luggage Limitations. !

Report No: 01 973-001 lnterior Code EX-6s-STDQS. Page 6-07-4

3.6 (1.65)

3.6 (1.65)

6.44 (2.92)

3.6 (1.65)


941 (10.96)

1049 (12.21)

2325 (26.78)

1335 (15.13)


EXECUTIVE INTERIOR CODE EX-6SSTD-2s



WEIGHT

Ib

50

60

70

80

90

100

110

120

130

140

150

160

170

180

190

200

210

220

230

240

Report No: 01 973-001 Interior Code EX-6s-STD-2s. Page 6-07-5

PASSENGER SEAT

PAX 112

234.09 in

11705

14046

16387

18728

21068

23409

25750

28091

30432

32773

351 14

37455

39796

421 37

44478

46819

49160

51501

53842

56183

OCCUPANT

PAX 314

276.12 in

13806

16567

19329

22090

24851

27612

30374

33 135

35896

38657

41419

44180

46941

49702

52464

55225

57986

60747

63508

66270

MOMENTS ( LB - PAX 516

308.12 in

15406

18487

21569

24650

27731

30812

33894

36975

40056

43 137

46219

49300

52381

55462

58544

6 1625

64706

67787

70868

73950 ,

IN )

PAX 718

341 .OO in

17050

20460

23870

27280

30690

34100

37510

40920

44330

47740

51 150

54560

57970

61380

64790

68200

71610

75020

78430

81840


EXECUTIVE INTERIOR CODE EXdESTDZS


Repon NO: 01973-001 Interior Code EX-6s-STD-2s. Page 6-07-6


- M )

PAX 718

8.661 m

216.54

259.84

303.15

346.46

389.76

433.07

476.38

51 9.68

562.99

606.30

649.61

692.91

736.22

779.53

822.83

866.14

909.45

952.75

996.06

1039.37

WEIGHT

ka

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120

PASSENGER SEAT

PAX 112

5.946 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

654.06

683.79

713.52

OCCUPANT

PAX 314

7.014 m

175.34

210.41

245.47

280.84

315.61

350.68

385.75

420.81

455.88

490.95

526.02

561.08

596.15

631.22

666.29

701.35

736.42

771.49

806.56

841.63

MOMENTS ( KG

PAX 516

7.826 m

195.66

234.79

273.92

313.05

352.19

391.32

430.45

469.58

508.71

547.84

586.98

626.1 1

665.24

704.37

743.50

782.63

821.77

860.90

900.03

939.16


EXECUTIVE INTERIOR CODE EX-4s-STD-4s

GENERAL

This configuration is a variation of the basic Executive interior and consists of 4 executive passenger seats and 4 standard seats. It is the operators responsibility to check before using this configuration whether they require authorization by their regulatory authority. The following information is given:



passenger seats and furnishings weight and moment chart

w passenger seat occupant moments (standard and metric units)




FOUR EXECUTIVE AND FOUR STANDARD INTERIOR CODE EX-4s-STD-4S

SEAT LOCATIONS LAVATORY1 RH CDPUYER WARDROBE CABINET (OPTIONAL)

DISTANCE FROU 0' 10.05 31.88' 31.30' DIVIDER AFT SURFACE 0 m 0.255 m 0.810 m 0.785 m

178.95* 180.00' 211 85' 211.18' FUSEUGE STAnON 4.571 m 4.826 m 5.381 m 5.364 m

DISTANCE FROM 0' 31.88' 56.18' 68.18' 128.18 162.18' DIVIDER AFT SURFACE 0 m 0.808 m 1.427 m 2.265 m 3.281 m 4.1 19 m

178.85' 211.83" 236.14' 26B.lY 309.13' 342.13' 4.571 m 5.380 m 5.898 m 6.838 m 7.852 m 8.690 m

AFT FACING SEAT

CENTER OF ARRESTING PIN .-----

NOTE: CABIN SEAT LOCATION IS DEFINED AS THE DISTANCE FROM THE AFT SURFACE ON THE FORWARD DIVIDER PAhEL TO THE CENTER OF THE ARRESTING P H ON EACH SEAT.

FWD FACING SEAT

I

CENTER ARRESTING PIN

NOTE: CABIN SEAT 5 THRU 6 LOCATION IS DEFINED FROM THE CENTER OF THE CENTER ARRESTING PIN ON EACH SEAT.

I

Report No: 01 973-001 In te r io r Code EX-4s-STD-4s. Page 6-08-2

Issued: June 10.1994 Revision 1 1 : March 1,2003 ,


EXECUTIVE lNTERlOR CODE EX-4S-STD-4S

PERMlTfED PASSENGER SEAT PART Nos. THAT CAN BE INSTALLED


SEAT NO.

1

2

3

4

5,7

6,8


PC-1 2 AND PC-1 2145

959.30.01.613 or 959.30.01.625 959.30.01.617 959.30.01.627 959.30.01.619 959.30.01.629 959.30.01.621 959.30.01.631 959.30.01.623

959.30.01.614 or 959.30.01.626 959.30.01.618 959.30.01.628 959.30.01.620 959.30.01.630 959.30.01.622 959.30.01.632 959.30.01.624

959.30.01.615 or 959.30.01.649 959.30.01.633 959.30.01.651 959.30.01.635 959.30.01.653 959.30.01.637 959.30.01.655 959.30.01.639 959.30.01.657 959.30.01.641 959.30.01.659 959.30.01.643 959.30.01.661 959.30.01.645 959.30.01.663 959.30.01.647

959.30.01.616 or 959.30.01.650 959.30.01.634 959.30.01.652 959.30.01.636 959.30.01.654 959.30.01.638 959.30.01.656 959.30.01.640 959.30.01.658 959.30.01.642 959.30.01.660 959.30.01.644 959.30.01.662 959.30.01.646 959.30.01.664 959.30.01.648

525.22.12.01 1

525.22.12.012


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

EXECUTIVE INTERIOR CODE EX-4s-STD-4S


Adjust the aircralt Basic Empty Weight on the Loading Form for items removed/added when converting to or from a Combi lnterior Conversion.


FR 24 CARGO NET

FR 27 CARGO NET


FR 34 BAGGAGE NET



5 (2.5)

3.6 (1.65)

3.6 (1.65)

6.44 (2.92) .

3.6 (1.65)


1162 (13.41)

941 (10.96)

1049 (12.21)

2325 (26.78)

1335 (15.13)


EXECUTIVE INTERIOR CODE EX4S-STD-4S



1

Report No: 01973-001 Interior Code EX-4s-STD-4s. Page 6-08-5

WEIGHT

Ib

50

60

70

80

90

100

110

120

130

140

150

160

1 70

180

190

200

210

220

230

240

PASSENGER SEAT

PAX 112

234.09 in

1 1705

14046

16387

18728

21 068

23409

25750

28091

30432

32773

351 14

37455

39796

42137

44478

4681 9

49160

51501

53842

561 83

OCCUPANT

PAX 314

276.12 in

13806

16567

19329

22090

24851

27612

30374

33135

35896

38657

41419

44180

46941

49702

52464

55225

57986

60747

63508

66270

MOMENTS ( LB - PAX 5/6

310.00 in

15500

18600

21700

24800

27900

31000

34100

37200

40300

43400

46500

49600

52700

55800

58900

62000

65 100

68200

71 300

74400

IN )

PAX 718

343.00 in

17150

20580

24010

27440

30870

34300

37730

41 160

44590

48020

5 1450

54880

58310

6 1740

65170

68600

72030

75460

78890

82320


EXECUTIVE INTERIOR CODE EXSS-STD4S


Report No: 01973-001 Interior Code EX-4s-STD-4s. Page 6-08-6


M )

PAX 718

8.7122 m

217.81

261.37

304.93

348.49

392.05

435.61

479.17

522.73

566.29

609.85

653.42

696.98

740.54

784.10

827.66

871.22

914.78

958.34

1001.90

1045.46

MOMENTS ( KG -

PAX 516

7.8740 m

196.85

236.22

275.59

314.96

354.33

393.70

433.07

472.44

51 1.81

551.18

590.55

629.92

669.29

708.66

748.03

787.40

826.77

866.1 4

905.51

944.88

OCCUPANT

PAX 314

7.01 35 m

175.34

210.41

245.47

280.54

315.61

350.68

385.75

420.81

455.88

490.95

526.02

56 1.08

596.1 5

631.22

666.29

701.35

736.42

771.49

806.56

841.63

WEIGHT

kg

25

30

35

40

45

50

55

60

65

70

75

80

85

90

95

100

105

110

115

120 ..

PASSENGER SEAT

PAX 1 I2

5.9460 m

148.65

178.38

208.1 1

237.84

267.57

297.30

327.03

356.76

386.49

416.22

445.95

475.68

505.41

535.14

564.87

594.60

624.33

654.06

683.79

713.52

=PILATUSS SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

SECTION 7


TABLE OF CONTENTS

Subject

GENERAL

AIRFRAME

GENERAL FUSELAGE EMPENNAGE WINGS

FLIGHT CONTROLS

GENERAL AILERON ELEVATOR RUDDER TRIM FLAPS INDlCATlONMlARNlNG SYSTEM

LANDING GEAR

GENERAL LANDING GEAR DESCRIPTION HYDRAULIC SYSTEM DESCRIPTION MAIN HYDRAULIC SYSTEM OPERATION INDICATION I WARNING SYSTEM EMERGENCY EXTENSION SYSTEM AIR I GROUND SYSTEM BRAKES WHEELS AND TIRES

BAGGAGE COMPARTMENT

Issued: February 14, 1994 Revision 6 Dec 6, 1996

Page

Report No: 0 1973-001 7-1

SECTION 7 .CPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

Subject Page 4

SEATSlRESTRAlNT SYSTEMS

SEATS SEAT BELTS AND SHOULDER HARNESSES

CARGO TIE-DOWNS

DOORS, WINDOWS AND EXITS

FORWARD CABIN DOOR CARGO DOOR WINDOWS lNDlCATlONMARNlNG EMERGENCY EXIT I AIRCRAFT SECURITY

CONTROL LOCKS

ENGINE

DESCRIPTION AND OPERATION AIR INDUCTION

CONTROLS Power Lever Control Manual Override Lever Condition Lever

ENGINE FUEL OIL STARTING IGNITION ACCESSORIES FIRE DETECTION ENGINE INSTRUMENT SYSTEM (EIS) TORQUE LIMITER INDICATION I WARNING

PROPELLER

GENERAL DESCRIPTION OPERATION PROPELLER DE ICE INDICATION / WARNING


t PILATUSg SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Subject Page

FUEL

GENERAL DESCRIPTION OPERATION INDICATION I WARNING

ELECTRICAL

GENERAL

DESCRIPTION Power Supplies Bus Bars Bus Tie Interlock Non Essential Bus Circuit Breakers AC Power Controls and Indicators

OPERATION MALFUNCTIONS, CAUTIONS AND WARNINGS

LIGHTING

INTERIOR EXTERIOR

ENVIRONMENTAL CONTROL SYSTEM

GENERAL DESCRIPTION OPERATION INDlCATIONMlARNlNG

HEATING SYSTEM

GENERAL DESCRlPTlON OPERATION INDICATION /WARNING

COOLING SYSTEM


Issued: February 14,1994 Revision 8 September 1, 1998

Report No: 01 973-001 7-iii

SECTION 7 AIRPLANE AND SYSTEMS DESCRIPTION

Subject Page

CABIN PRESSURIZATION CONTROL SYSTEM


OXYGEN

GENERAL DESCRIPTION

I K,"i:I,"," I wARNING OPTIONAL LARGER CAPACITY OXYGEN SYSTEM

COCKPIT ARRANGEMENT

GENERAL DESCRIPTION

CENTRAL ADVISORY AND WARNING SYSTEM (CAWS) 7-108

GENERAL DESCRIPTION

PlTOT STATIC SYSTEM

GENERAL DESCRIPTION INDICATION I WARNING

STALL WARNINGISTICK PUSHER SYSTEM

GENERAL DESCRIPTION OPERATION INDICATION 1 WARNING

PNEUMATIC WING DEICE SYSTEM


COMFORT FEATURES

GENERAL

Report NO: 01 973-001 7-iv

Issued: February 14,1994 Revision 11: March 1,2003

=PILATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

Subject Page

CABIN FEATURES

GENERAL CORPORATE COMMUTER INTERIOR EXECUTIVE INTERIOR

GENERAL AVIONICS INSTALLATlON

AUDIO PANEL General

AVIONICS General

ATTITUDE AND HEADING REFERENCE SYSTEM

GENERAL DESCRIPTION OPERATION BUILT-IN TEST EQUIPMENT (BITE) INDICATION I WARNING

STANDBY ATTITUDE INDICATOR


EMERGENCY LOCATOR TRANSMITTER

DESCRIPTION OPERATION

ELECTRONIC FLIGHT INSTRUMENTATION SYSTEM EFS 40150

GENERAL

DESCRIPTION Electronic Attitude Director lndicator Electronic Horizontal Situation Indicator EFlS Control Panel Multi Function Display (if installed) Multi Function Display Control Panel (if installed) Weather Radar Display (if installed) Weather Radar Control Panel (if installed)

OPERATION INDICATION I WARNING

Issued: February 14, 1994 Revision 11: March 1,2003

Report No: 01 973-001 7-v

SECTION 7 =PLATUSW AIRPLANE AND SYSTEMS DESCRIPTION 'PC %I1 ,

4

Subject Page

AUTOPILOT 7-1 67

GENERAL 7-167

DESCRIPTION Mode Controller Altitude / Vertical Speed Preselect

OPERATION Autopilot Pitch Limits

! Altitude / Vertical Speed Preselect I Control Wheel Steering

Autopilot Disconnect Manual Trim Engage

INDICATION /WARNING Autopilot Altitude 1 Vertical Speed Preselect

I

Report No: 01973-001 7-vi


9PfLATUSW SECTION 7 ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

GENERAL

Section 7 of this Pilot's Operating Handbook contains infomation related to h e detailed description and operation of the airplane and its systems.

AIRFRAME

GENERAL

The airplane is a low wing, T-tail, single engine, retractable landing gear type designed to transport passengers, cargo, or various combinations of both passengers and cargo. Constmction is conventional semimonocoque, primarily incorporating aluminum alloy, but composite structures are used in certain areas.

Rush keting is used where appropriate to minimize drag. Access panels are inslalled to facilitate inspection and maintenance. The complete airframe is electrically bonded to eliminate electro-magnetic interference and static discharge wicks are used to reduce static charges while inflight.

FUSELAGE

The fuselage consists of the engine area, nose gear assembly, cockpit, cabin, and ail fuselage. The engine area contains the powerplant, and associated accessories. The engine cowling is constructed from a carbonlnomex honeycomb material while the engine mount is welded steel tubing and bolted to the firewall in four places. The firewall is titanium and protected by insulation material.

A two piece windshield, two side windows, and a direct vision (DV) window provide cockpit visibility. The two piece windshield is glass while the two side windows and the DV window are stretched acrylic. All windows are of two ply laminated design.

The cabin area is from the cockpit to the aft pressure bulkhead and contains the forward cabin door, the cargo door, and an emergency overwing exit. The nine cabin windows are two ply laminated monolthic stretched acrylic and incorporate dry neoprene seals. Airplane avionics are mounted under the cabin floor, running the length of the center cabin, and are accessible through quick release panels. The cabin cany-through spar attachment fillings are one piece machined aluminum. Fuselage fairings are constructed from either carbonlnomex or aramidlnomex honeycomb material.

A Safety net is installed ail of the rear pressure bulkhead to protect the bulkhead from damage during maintenance.

issued: February 14, 1994 Revision 2: February 14,1995

Report No: 01973-001 7-1 1


EMPENNAGE

The empennage is a T-tail design with the horizontal stabilizer mounted on top of the vertical stabilizer. The aft fuselage is attached to the cabin at the aft pressure bulkhead. The vertical and hor~zontal stabilizer assemblies are conventional aluminum construction. The horizontal stablllzer 1s a trimmable structure. The dorsal and ventral fin fairings are kevlar honeycomb matenal.

WINGS

The wings are of conventional construction incorporating lront and rear spars, ribs, and skin. The lront and rear spars are mainly from machined aluminum alloy plate. Both spars include fuselage and integral landing gear attachment points, while the rear spar also integrates flap actuator attachment points. Main load carrying ribs are machined from aluminum alloy plate All other ribs are formed sheet metal. The ribs incorporate lightening holes to reduce weight and integral beads for stiffening. The wing skin is stiflened clad aluminum alloy sheet riveted to the spars and nbs. Access panels are in the wing bonom only.

Each wing is attached to the fuselage using three titanium shear pins and, at the aft upper l~tting, one steel tension bolt.

Each wing contains an integral fuel tank, aileron, flaps, deice boot, and main landing gear. The luel tanks are located between ribs 3 and 16, forward of the main spar to lhe nose rib and between ribs 6 and 16 behind the main spar to the rear spar.

The ailerons are conventional construction with a single spar and ribs. The aileron access panels are a carbonfnomex honeycomb consttuction. The ailerons are mass balanced and the aileronlwing gap is sealed.

Each wlng incorporates a single piece Fowler flap of conventional construction, with three support arms and associated linkages. The wing trailing edges above the flaps are foam core covered with carbon laminate while the flap fairings are a carbon laminate wilh nomex honeycomb reinforcement strips.

A surlace mounted deice boot is attached to lhe nose skin of each wing. Each wing has a main landing gear attached to the front and rear spar, with a carbon fiberhornex honeycomb gear door attached to the leg. The wing tips are constructed of carbon fibermoneycomb and metal strips for lightning protection.

Report No: 01973-001 7-2

Issued: February 14,1994 Revision 2: February 14, 1995 ,

*PILATUS-- SECTION 7 -PC x F AIRPLANE AND SYSTEMS DESCRIPTION

FLIGHT CONTROLS

GENERAL

The flight control system is conventional using push-pull rods and carbon steel cables. Electric trim systems are provided for the aileron, rudder, and elevator. All trim systems can be disconnected in the event of a runaway condition.

When the flaps start to extend an aileronlrudder interconnect system operates to move the rudder at the same time as the ailerons are operated. The more the flaps are extended, the larger the movement of the rudder when the ailerons operate. This decreases the quantity of input necessary from the rudder pedals at low airspeeds.

AILERON

The ailerons are connected to the cockpit control wheels by control cables in the fuselage and push-pull rods in the wings. Each aileron is attached to the wing at two hinge points. Each aileron is equipped with a minimum of two static wicks to dissipate static charges to the atmosphere.

The left aileron incorporates a trim tab which is electrically operated from the cockpit. Refer to 1 Trim system, this section, for more information.

ELEVATOR

The elevator is a two piece unit attached to the horizontal stabilizer at a total of five hinge points and is connected to the cockpit control wheel by carbon steel control cables. A down spring is installed in the control circuit to improve longitudinal stability. The elevator is equipped with static wicks to dissipate static charges to the atmosphere.

I Pitch trim is provided by positioning the horizontal stabilizer. Refer to Trim system, this section, for more information.

RUDDER

The rudder is a single piece unit attached to the vertical stabilizer at two hinge points and is connected to the cockpit rudder pedals by carbon steel control cables. Both pilot and copilot rudder pedals are adjustable by use of a crank located between each set of rudder pedals. Clockwise rotation of the crank moves the pedals aft. The rudder is equipped with static wicks to dissipate static charges to the atmosphere.

The rudder incorporates a trim tab that is electrically operated from the cockpit. Refer to Trim system, this section, for more information.


Report No: 01973-001 7-3

SECTION 7 r P I L ATUSE AIRPLANE AND SYSTEMS DESCRIPTION ?C XI1

TRlM

I The aileron, horizontal stabilizer and rudder trim are electrically operated. Aileron and horizontal stabilizer trim operation are controlled by a switch on the outboard horn of each control wheel while rudder trim operation is controlled by a switch located on the Engine Power Control Lever. Prior to selecting pitch and aileron trim press and hold the pitch trim engage switch located on the forward side of each outboard control wheel horn.

I Pitch trim is accomplished by an electrically controlled actuator connected to the moveable horizontal stabilizer. The secondary trim motor, installed in the same actuator, is controlled by the autopilot and can also be used as a backup system (alternate stabilizer trim) by the pilot. Alternate pitch trim can be accomplished by pressing the ALTERNATE STAB TRlM switch in the desired direction.

The leading edge of the stabilizer moves down for nose up trim and up for nose down trim. At the root of the leading edge of the leH horizontal stabilizer there is a trim range indicator which has markings to show full travel in either direction and a takeoff trim range. During the preflight inspection this external trim indicator should be used to verify cockpit trim position indication.

In the event of uncommanded trim operation, all trim operation can be stopped by pressing the TRlM INTR switch located ahead of the Engine Control Quadrant on the center console.

Report No: 01 973-001 7-4

Issued: June 10,1994 Revision 5: Mav 10. 1996 ,

=PlLATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

FLAPS I Each wing trailing edge has a single piece Fowler type flap supported by three flap arms. The flaps are controlled by a selector handle located to the right of the power controls on the center console. The flaps may be set to one of the four preset positions 0°, 15O, 30' and 40" by moving the handle to the appropriate position. If the flap lever is not at one of the four preset positions, the Flap Control and Warning Unit (FCWU) will drive the flaps to the nearest preset position. A flap position indicator is located near the top of the left instrument panel.

The flaps are electrically actuated. There is a single flap Power Drive Unit (PDU) installed below the cabin floor at the rear main frame. It drives screw actuators at the inboard and middle stations through flexible shafts. The screw actuators are connected to the flap actuating arms.

The flap control system incorporates a failure detection system. The system can detect a failure of a flexible shaft by disconnection or jamming, potentially resulting in flap asymmetry or failure of the system to achieve the selected flap position. The system can detect a failure of a single actuator, potentially resulting in single flap panel twisting. If a failure is detected, the FCWU disconnects the power to the PDU and the CAWS FLAPS caution will come on. This condition cannot be reset by pilot action, a landing should be made IAW the EMERGENCY PROCEDURES as maintenance action is required.

A rotation sensor is Installed on each of the outer flap screw actuators. These sense the rotation of the flexible shafts and give signals to the FCWU. The FCWU monitors these signals for asymmetrical flexible shaft rotation of more than 20 rotations (caused by a broken inner flap drive shaft). If failure is detected the FCWU disconnects the power to the PDU and the CAWS FLAPS caution will come on. This condition cannot be reset by pilot action. To detect satisfactory system operation, the FCWU monitors the left sensor for 10 rotations of the flexible shaft in the first 5 seconds (7 seconds with a modified FCWU) of a flap up or down I selection. If the selected flap position is not achieved the FCWU disconnects the power to the PDU and the CAWS FLAPS caution will come on.

There are five position sensors in the flap system, one at each center flap actuating arm, one at each inner flap actuating arm and one on the flap position lever, which give signals to the FCWU. The FCWU monitors the signals from the left and right flap sensors for flap asymmetry (caused by a broken inner flap drive shaft). If an asymmetry of 5" is sensed, power to the PDU is disconnected and the CAWS FLAPS caution will come on. The FCWU also monitors the signals from the left and right flap sensors for twisting of the left or right flap (caused by a broken outer flap drive shaft or unequal movement of the flap screw actuators). If a failure is detected, the FCWU disconnects the power to the PDU and the CAWS FLAPS caution will come on.

Additionally if flap asymmetry or twist is detected and the flap angle is greater than ZO, the stick pusher will default to a 'safe' mode and the CAWS PUSHER caution will come on 10 seconds later. In the 'safe' mode the stick pusher will operate at approximately 5 kts higher airspeed for the failed tlap position.


Report No: 01973-001 7-5

SECTION 7 =PILATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

Pre SB 27-013. If the PDU motor draws excess current, the excess current will open a remote circuit breaker, disconnects the power to the PDU and the CAWS FLAPS caution will come on. The remote circuit breaker will also open a FLAP circuit breaker on the Battery Bus circuit breaker panel. After waiting a period for cooling the FLAP circuit breaker can be reset and normal flap operation resumes. This is the only pilot re-senable failure and cycling the flap circuit breaker if it has not opened will not reset any other failure mode detected.

Post SB 27-013. If the Power Drive Unit (PDU) motor overheats or a stalled motor condition is detected, a signal from the PDU will open the FLAP circuit breaker on the Battery Bus circuit breaker panel. The FCWU then removes the up or down command to the PDU and the CAWS FLAPS caution will come on. After waiting for a period of 5 minutes the FLAP circuit breaker can be reset and normal flap operation resumes. This is the only pilot re-settable failure and cycling the flap circuit breaker if it has not opened will not reset any other failure mode detected.

To avoid an inadvertent flap down command at high speed, the flap down relay is disabled when the flap selector handle is in the 0' position.

Flap system operation may be stopped at any time by lifting the switch guard and pressing the INTERRUPT FLAP switch on the center console to INTR. The CAWS FLAPS caution will then come. If the switch is moved back to the NORM position, normal operation will not resume, even if the FCWU does not detect any failures.

A FLAP GROUND RESET switch (Post SB 27-006 and MSN 322 and UP) is installed on the maintenance test panel (right sidewall behind the co-pilot seat). The FLAP GROUND RESET switch is only operational on the ground for maintenance purposes.

Report NO: 01 973-001 1 7-6 Issued: June 10,1994


=PILATUSF SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

INDICATION I WARNING SYSTEM

A three-axis trim position indicator is located on the center console. The triple trim indicator shows a pictorial presentation of the trim position of the aileron trim tab (roll), horizontal stabilizer (pitch) and rudder trim tab (yaw). Some versions of the triple trim indicator include three white indicator lights, one for each trim axis. When the autopilot activates the autotrim system, the white light for the applicable axis illuminates.

A warble tone at 850 and 854 Hz will sound when a stabilizer trim runaway of the main system is sensed.

The Central Advisory and Warning System (CAWS) annunciator panel includes a STAB TRIM warning light. The illumination of this warning caption will illuminate after 5 seconds (MSN 101-180) 1 60 seconds (181 & UP) when weight is on the wheels and the trim position is unsafe for takeoff.

The Central Advisory and Warning System (CAWS) annunciator panel includes a FLAPS caution light. This caution will come on when the FCWU shuts down the system because it sensed an asymmetric flap condition, a twist in the left or right flap, flexible shaft asymmetry, or flap FCWU malfunction. The caution will also come on if the FLAP circuit breaker on the Battery Bus circuit breaker panel opens.

A flap position indicator is located on the pilots left instrument panel. The indicator face is marked with the positions 0°, 15", 30° and 40" and has a red warning caption. Flap position is shown by a pointer which moves in relation to flap movement. The red warning caption and a 1600 Hz aural warning tone interrupted at 5 Hz are activated anytime the airspeed is above the maximum limit for the current flap setting.

Issued: June 10,1994 , Revision 1 1 : March 1, 2003 Report No: 01 973-001

7-7

SECTION 7 =PILATUS~ AIRPLANE AND SYSTEMS DESCRIPTION PC12

AIRSPEED INDICATOR

165 KTS 130 KTS P

F W POSITION INDICATOR AND WAF!NW

FLAP CONTROL AND

WARNING UNIT

I Figure 7-1. Flap S y s t e m ( P r e SB 27-013)

Report No: 01 973-001 I 7-7A

Issued: June 10,1994 Revision 1 I : March 1,2003

b TCPILATUSF SECTION 7 P(12 AIRPLANE AND SYSTEMS DESCRIPTION

AIRSPEED FLAP POSITION INDICATOR INDICATOR AND WARN!NG

OVEASPEED EATERY

CAWS F6 ' 0..

WARN 1

t I FLAP

QEN 1 BUS

1

FLAP CONTROL

- -

I Figure 7-1. Flap System (Post SB 27-013)

Issued: June 10,1994 , Revision 1 1 : March 1,2003 Report No: 01973-001

7-78 1

SECTION 7 =PILATUSW AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

LANDING GEAR

GENERAL

Refer to Figure 7-2, Landing Gear System, for system operation.

The landing gear is a conventional tricycle configuration that is extended and retracted using hydraulic pressure produced by an electrically powered hydraulic pump. Landing gear extension and retraction is the only function of the hydraulic system and landing gear operation is completely automatic upon pilot gear selection.

A nitrogen charged accumulator is used to maintain hydraulic pressure and hold the landing gear in the retracted position following a hydraulic system failure. If required, the landing gear can be lowered manually through a combination of tree-falling and the emergency landing gear hand pump.

Landing gear position and waming indications consist of three indicator lights (greenlred) and an aural tone (gear waming tone < 10 % torque) with a silencer button.

Nosewheel steering is accomplished by mechanical nosewheel steering and by differential braking.

Aircraft braking is controlled by toe pedals that operate brake assemblies attached to the left and right landing gear. Propeller reverse also contributes to aircraft braking. Refer to Propeller system, this section, for more information.

Report No: 01973-001 ( 7-8 Issued: February 14, 1994


=PILATUSE SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

LANDING GEAR DESCRlPf ION

The nose gear is a fluid and nitrogen filled shock strut. The shock strut consists of a piston and fork assembly that slides inside a cylinder. A toque link connects the pistonlfork assembly to the cylinder. The cylinder is mounted inside the nosewheel well. The nose gear is locked in the extended position by pufflng the folding strut in an over-center position. A spring is attached to the nose gear to assist in free fall during emergency extension. The nose gear doors are spring loaded to the open position and are mechanically closed during nose gear retraction. The nose gear retracts rearward into the nosewheel well and is completely enclosed by the gear doors when the landing gear is retracted.

Both main landing gear are trailing link types. A fluid and nitrogen filled shock strut connects the trailing link to the main leg hinge point. Each main gear actuator incorporates a mechanical down-lock and a gear down and locked indicator switch. The main landing gear doors consist of a single door that is attached to the main gear leg and the outside edge of the main gear wheel well. Each main gear retracts inward into the main gear wheel well. With the landing gear retracted the main landing gear wheel and tire assemblies are not enclosed and protrude out of the main gear wheel well approximately one inch (25.4 mm).

All landing gear are held in the fully retracted position by hydraulic pressure. No mechanical up-locks are required.

Nose wheel steering is accomplished using the rudder pedals which are mechanically connected to the nosewheel. Additional nosewheel steering is done through differential braking. Use of rudder pedal only will turn the nosewheel i 12 degrees from center while differential braking will turn the nosewheel i 60 degrees from center. A shimmy damper is installed on the nose landing gear strut to eliminate nosewheel oscillations.

The tires are a low pressure type that allow operations from soft and unimproved fields.

Issued: June 10.1994 Revision 5: May 10, 1996

Report No: 01973-001 7-9

SECTION 7 EPILATUSE AIRPLANE AND SYSTEMS DESCRIPTION K12

HYDRAULIC SYSTEM DESCRIPTION

The main hydraulic system comprises a power pack located in the hydraulic service bay in the left wing root behind the main spar, a nitrogen charged accumulator in the same location, a landing gear selector valve mounted in the cockpit mechanically linked to the pilots control panel selector handle by a pusWpull rod and three actuators, one for each landing gear leg.

The power pack features an electrically driven variable displacement hydraulic pump abng with associated filtration, pressure regulation and failure protection systems, and an integral reservoir with a visual sight gauge. It provides the main source of hydraulic pressure and flow to facilitate landing gear extension and retraction. It is pressurized with nitrogen from a pressure cylinder to prevent pump cavitation during low temperature operations. Filtration consists of an integral filter with automatic by-pass and visual by-pass indication. (pop out button). Thermal protection is provided.

MSN 231 and UP have a larger nitrogen pressure cylinder which also contains hydraulic fluid thus increasing the system fluid volume. A visual level indicator is installed at the rear of the cylinder and can be seen at the wing root when the cargo door is open. This enables the hydraulic fluid level status to be seen during a pre flight inspection.

Electrical power supply for the system is provided from the BATTERY POWER LINE and is applied to the power pack, causing the pump to operate when low hydraulic pressure is sensed by the system pressure switch. The hydraulic control circuit is powered from the NON- ESSENTIAL BUS.

The selector valve is a two position, four way rotary type located in the cockpit and mechanically actuated by the landing gear selector handle.

Report No: 01 973-001 7-10

~ssued: June 10,1994 Revision 9: September 1,1999

EPILATUSE SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

The nitrogen charged accumulator is present for fluid thermal expansion and to ensure system pressure is maintained after power pack failure. Its size accounts for normal system pressure leakage rates to ensure that after such a failure, the landing gear will be maintained in the retracted position for 200 minutes. In this time however the gust load limit is reduced to 3.39.

NOTE

The landing gear may partially extend during a gust load that is greater than 3.39 and then return to the retracted position. Afler 200 minutes, the landing gear may start extending due to loss of system pressure.

The accumulator is charged via a charging valve located in the service bay with pressure indicated on a gauge at the same location. A service selector valve, located in the service bay, allows the system to be operated from a ground hydraulic service unit.

The actuators are of the linear type with the main landing gear actuators also incorporating the down locking mechanism.

Cockpit controls consist of the following - - A landing gear selector handle is located on the pilot's lower right panel and facilitates

extension or retraction of the landing gear. It acts directly, via a rod, on the landing gear selector valve. The handle is equipped with an electrical spring loaded solenoid which prevents it from moving to the retracted position when the airplane is on the ground. The airplane on ground status is sensed by the airlground system which comprises two proximity switches and associated targets, one on each main landing gear leg and a relay.

- An emergency landing gear hand pump and operating handle, located at the rear of the center console, is used to assist in free fall emergency landing gear deployment after failure of the main system.

Issued: February 14, 1994

, Revision 2: February 14, 1995 Report NO: o 1973-00 1

7-1 1 I


MAIN HYDRAULIC SYSTEM OPERATION

When the airplane IS on the ground and the electrical system is energized with the engine running and oil pressure is above 60 psi the system pressure limit switch, which senses accumulator nitrogen pressure, will maintain the hydraulic system pressure between the set limils of 2,450 and 2,800 psi

Movement of the landing gear selector valve to pass pressure to the appropriate side of the main landing gear and nose landlng gear actuator retracts the landing gear, which is then held there by continued application of hydraulic pressure. When the airplane is on the ground or in flight and the engine oil pressure is below 60 psi, the power pack is isolated to prevent pump cavitation in cold conditions. In normal operation, the landing gear is extended by moving the landing gear selector valve to apply pressure to the other side of all three actuators. Locking of the main landing gear actuators is accomplished by internal locks within the actuator housings which are actualed and released by the application of hydraulic pressure in the appropriate sense. The nose landing gear IS held in its extended position by an over-center two piece drag link.

The hydraulic system is isolated in the following conditions .

- After the loss of the GEN 1 which will also isolate the NON-ESSENTIAL BUS (auto load shed) due to a GEN 1 or engine failure.

- After power pack overheat via thermal protection integral to the power pack itself

Report No: 01973-001 7-12



Figure 7-2. Landing Gear System (Sheet 1 of 4)


1

Report No: 01 973-001 7-13

SECTION 7 3PILATUS W AIRPLANE AND SYSTEMS DESCRIPTION ?C XI1

BAT BUS

HYDRAULIC POWER PACKAGE

- - - -

NOTE: HYDRAULIC CONFROL SYSTEM

P I


BAT BUS

I Report No: 01973-001 7-14


GEN 1

1% PRESSURE LOW CAWS 1

CAWS SWITCH

I %PILATUSW SECTION 7

i ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

b I

GEAR HANDLE

GEN 1 BUS I

RH AIR J WEIGHT ON LEG RELAY

NOTE INDICATION SYSTEM A - - Figure 7-2. Landing Gear System

(Sheet 3 of 4)

GND

AIRGND SWlTCH


I Revision 2: February 14, 1995

Report No: 01973-001 7-15

SECTION 7 =PILATUS= AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

OVERHEAD PANEL

--\ /-

I CAWS

Report No: 01973-001 7-16

w LG CONTROL PANEL




lNDlCATlOWARNlNC3 SYSTEM

Extended position indication Is provided by micro switches internal to the main landlng gear actuators and a proximity switch on the nose landing gear drag Ink. Retraction position indication Is provided by proximity switches on the landlng gear doors.

Landing gear position Is indicated in the cockpit by three dual-indication llghts (one for each leg), located to the left of the landlng gear selector handle. Each light is split into twa sections with the top section being a red light while the bottom sectlon is a green Ilght. When a landing gear is down and locked, the bottom green light will be illuminated. Anytime a landing gear Is in transit, or In an unsafe condition, the top red light will be illuminated. Both lights will be off with a landing gear fully retracted.

The Central Advisory and Warning System (CAWS) annunciator panel includes a caution light labeled HYDR. This caution provides two different advisories - I - In flight a continuously illuminated caption means that main system pressure has fallen

below operatbnal limits (nominal 1,800 psi) and can not be relied upon for proper landing gear system operation. The annunciator is illuminated 30 seconds after low pressure Is sensed by a low pressure switch mounted in the power pack pressure supply line upstream of the landing gear selector valve.

- On ground after landing a continuously illuminated caption means that the power pack has been automatically Initiated in flight more that six times in an hour by the system 1 pressure switch in order to maintain system pressure. This indicates that the pressure leak rate from the accumulator is too high or a low fluid level. Pilot initiated landing gear cycle will reset this counter to zero.

All three landing gear indicator lights will come on red and an audible warning tone will be heard through the overhead speaker and through the headphone audio when the following conditions are met with electrical power applied to the airplane:

- Anytime the landing gear control handle Is UP, while in the air, with -

- flaps set to more than 15"; or - airspeed is less then 130 KIAS; and power setting is less than 10 psi.

The audible warning tone can be turned off, i f flaps are not set to more than 15", by pressing the SILENCER button on the landing gear control panel.


Report No: 01 973-001 7-17

SECTION 7 'FEPLATUSB AIRPLANE AND SYSTEMS DESCRIPTION K12

EMERGENCY EXTENSION SYSTEM

The emergency extension system is required to ensure that the landing gear can be extended I and locked after a failure of the main system. This is accomplished by pulling the HYDR CTL

CB and releasing the hydraulic pressure in the actuators by selecting landing gear DOWN via the selector handle. This returns the hydraulic fluid back to the reservoir, thus allowing the landing gear to free-fall. Emergency extension of the nose landing gear is assisted by a spring strut to overcome aerodynamic loads. If necessary, yawing the airplane to use the aerodynamic load may assist the emergency extension.

I To manually extend the landing gear, pull the HYDR CTL circuit breaker and set the landing gear selector handle to DN with airspeed less than 110 KIAS. This will allow the landing gear to free fall. If the landing gear do not completely extend and show three green indicator lights, pull the emergency landing gear hand pump handle out of the aft center console and begin pumping. Stop pumping when all three landing gear indicator lights are green. Experience has shown that approximately 60 strokes of the hand pump are required in adverse conditions. Stow the emergency landing gear hand pump handle back into the aft center console before landing.

If the landing gear has been extended manually for training purposes, the landing gear can be I retracted normally HYDR CTL CB in. If the HYDR caution light on the CAWS annunciator

panel was on due to low hydraulic system pressure, wait until the HYDR caution light goes out. 1 If lhere is no apparent landing gear malfunction, make sure that airspeed is less then 177

KIAS, then move the landing gear selector handle to UP.

NOTE

The landing gear hydraulic control circuit is on the NON ESSENTIAL BUS. In the event that generator 1 goes off-line with a subsequent auto load shed of the NON ESSENTIAL BUS, normal landing gear operation is not available. Gear extension must be accomplished by the Emergency Gear Extension procedure.


IdBCPLATUSE SECTION 7 P(l2 AIRPLANE AND SYSTEMS DESCRIPTION

AIR I GROUND SYSTEM

The airplane is equipped with independent weight on wheel detectors, one on each main landing gear. Each weight on wheel signal is generated by a proximity switch. Each proximity switch controls a relay that provides the AIWGND signal to the different systems.

The LH weight on wheel signal is sent to the following systems:

LH Stick Pusher Computer Cabin Pressurization Control System Central Advisory and Warning System Engine Instrument System Hydraulic System Control

The RH weight on wheel signal is sent to the following systems:

RH Stick Pusher Computer Landing Gear Indication Attitude and Heading Reference System Transponder

If a disparity occurs between the detector signals (proximity switch or relay), a failure signal is muted to the Central Advisory and Warning System to activate the red AIWGND warning annunciator after a delay of a maximum of 10 seconds. If both weight on wheel detectors become inoperative at the same time (within a short period of time, maximum 10 seconds), the AIWGND warning will not be activated.

I Failure of a weight on wheel signal, while the alrplane Is on the ground (AIR signal provided), will have the following effects on the systems:

LH and RH Stick Pusher The Stick Shaker and aural stall warning may activate. Computers

Cabin Pressurization The system will pressurize the cabin to the altiiude set on the controller.

Landing Gear Indication Selector Handle Solenoid will be retracted and it will be possible to move the Selector Handle to the UP position and retract the landing gear.

Transponder Transponder may be active.


Report No: 01 973-001 7-19


Failure of a weight on wheel signal, while the alrplane Is In the air (GND signal provided), will have the following effects on the systems:

LH and RH Stick Pusher Stick Pusher is inoperative. Stick Shaker and aural stall Computers warning still operative from the other computer

Cabin Pressurization No effect on system operation as long as condition lever is not set to GROUND IDLE.

Hydraulic System Control No effect on system operation as l0ng as engine is running. If the engine stops, the hydraulic pump will be inoperative due to NON ESSENTIAL BUS shedding and the loss of the AIRIGROUND engine out sensor bypass. The landing gear must be extended in accordance with

I the Emergency Extension Procedures. I

Landing Gear Indication Selector Handle Solenoid will go to the locked position preventing gear retraction. Landing gear may still be extended by normal operation.

Transponder Transponder will not be activated.

Failure of a weight on wheel signal will have the following effects on the systems at any time:

Central Advisory and The following warnings and cautions will blink giving Warning System erroneous annunciation: CAB PRESS, PROP LOW P,

STAB TRIM, CHIP, and HYDR.

Engine Instrument The propeller rpm indication will be correct but the System associated warnings and cautions will be erroneous

Repoll No: 01973-001 7-20


3PILATUSW SECTION 7 P(12 AIRPLANE AND SYSTEMS DESCRIPTION

BRAKES

Refer to Figure 7-3, Brake System, for system operation.

Aircraft braking is provided by two brake assemblies, one bolted to each main landing gear axle. The brakes are controlled by toe pedals attached to each rudder pedal assembly. The pilot and copilot left toe brakes operate the left brake while the pilot and copilot right toe brakes operate the right brake.

The brake system consists of a brake fluid reservoir, four brake master cylinders, a left and right shuttle valve, a parking brake valve, and two brake assemblies. If the pilot and copilot simultaneously apply pressure to the same side brake pedal, the one applying the greatest pressure will control the braking.

The brake system is separate and independent from the airplane hydraulic system. The brake fluid reservoir is located on the right hand side of the cabin sidewall and incorporates a fluid level indicator.

A separate brake master cylinder, located in the cockpit footwell, is mechanically connected to each toe pedal. There is no mechanical connection between the pilot and copilot brake pedals. Two shuttle valves, a left and a right, are used to combine inputs from their respective pilot and copilot brake pedals. Pressing a brake pedal causes the applicable brake master cylinder to force brake fluid through the respective shuttle valve and parking brake valve to the brake assembly.

Pre SB 32-013 brake assemblies have carbon friction surfaces and Post SB 32-013 have steel friction surfaces. The performance of the two brake assemblies is similar. Each brake assembly incorporates two brake lining wear indicators. As the brake linings wear, the pins will be pulled into the piston housing. When the system is pressurized and the pins are flush with

I the piston housing, the brake linings must be overhauled. I The parking brake valve has two OH-center cams that hold open poppet valves whenever the parking brake is released. This allows hydraulic fluid flow through the brake system. When the parking brake is set, the OH-center cams are rotated to allow the poppet valves to close. This traps brake fluid under pressure between the parking brake valve and the brake assemblies.

To set the parking brake, pull the PARKING BRK T-handle fully out and rotate to lock, then evenly press both brake pedals. Release pedal pressure and the brakes will remain set. To release the brakes, rotate and push the PARKING BRK T-handle fully in.

WHEELS AND TIRES

The wheels are split-hub type, the main wheels have three fusible plugs which melt when there is too much heat from the brakes. Tubeless tires are installed on the wheels and each wheel has a tire inflation valve and an overinflation safety plug. The main wheels have fairings on the outer hubs which make the wheels aerodynamically smooth when the .landing gear is retracted. The main wheels are modified when brake assemblies with steel friction surfaces (Post SB 32-013) are installed.


Report No: 01 973-001 7-2 1


PARKINO BRAKE VALVE (SHOWN WITH P M N Q BRWE NOT SET)

TO LEFT BRAKE TO RIGHT BRAKE ASSEMBLY ASSEMBLY

KEY

-3 . RESERVOLR BRAKE FLUM

-LEFT BRME SYSTEM

- RlOHT B-E SYSTEM

Figure 7-3. Brake System

Report No: 01 973-001 1 7-22 Issued: February 14, 1994


EPLATUS J SECTION 7 P(12 AIRPLANE AND SYSTEMS DESCRIPTION

BAGGAGE COMPARTMENT

A baggage compartment is provided at the rear of the cabin and is accessible during flight. A krggage net is secured at twelve anachment points to secure the baggage. An extendible baggage net (Post SB 25-010) can be installed to secure baggage in front of and in the baggage compartment. The floor attachments at the front of the net can be moved between frames 32 and 34.

1

SEATSlRESTRAlNT SYSTEMS

SEATS

The crew seats are adjustable both fore and aft and vertically. These seats also have lumbar support adjustment. The controls for seat adjustment are located centrally under the front of the seat pan (Pre SB 25-001). Post SB 25-001 and MSN 161 and UP the up and down adjustment handle is under the front and the fore and an adjustment handle is at the rear of the seat. All armrests can be moved upwards. The inner arms can also be turned through 90" before being moved upwards, to provide free access to get in and out of the seat.

The standard passenger seats have a reclining backrest, sliding headrest, and a folding inner armrest. Post SB 25-003 seats have a luggage restraint bar installed on the bottom front part of the seat structure. This allows small luggage to be put below the seat.

The executive seats are leather upholstered. with 90" swivel and 3.4 in (86.4 mm) of forwardlrear travel. A reclining backrest, sliding headrest, sliding armrest, magazine pocket and a restraint system are Wed. The seat position control is located on the forward edge of the arm. Puling up on control handle will allow the seat to be moved to the desired position. Releasing the control handle will lock the seat in position. The control for the back recline is a round push button located in the inner surface of the arm. Depressing the button will allow the seat back angle to be adjusted.

The optional three seat bench comprises three seats installed on a pedestal which goes across the width of the bench. Each seat has a reclining backrest, sliding headrests and a three point restraint system. The left seat has a handle which when moved upwards allows the seat to be moved forwards to make baggage loading easier.

SEAT BELTS AND SHOULDER HARNESSES

Each crew seat is equipped with a four-point restraint system consisting of an adjustable lap belt and a dual-strap inertia reel-type shoulder harness. Each passenger seat is equipped with a three-point restraint system consisting of an adjustable reel-type lap belt and an inertia reel- type shoulder harness.

CARGO TIE-DOWNS

Tie-down anchor points fit into the seat rails and lock into place by an over-center lever. Tie- down straps can be secured to these anchor points.


Report No: 01973-001 7-23

SECTION 7 '6?PLATUSW AIRPLANE AND SYSTEMS DESCRIPTION ?C XI1 1

4

DOORS, WINDOWS, AND EXITS

FORWARD CABIN DOOR

The forward door is located in the front left fuselage, immediately aft of the cockpit, and is 4 ft 5 ( in (1.35 m) high by 2 ft 0 in (0.61 m) wide. The door can be opened or closed from either side ,

and is secured by six locking pins, These can be checked visually from inside the cabin to verify engagement The door is hinged at the bottom and has an integral stepshandrail assembly which automatically extends and retracts as the door is opened or closed. A non- inflatable rubber gasket attached to the door seals the gap to allow the cabin to pressurize ' when the door is closed.

To open the door from the outside, pull outward on the aft end of the handle. Rotate the handle clockwise to the vertical position then pull outward on the door. As the door opens, the steps . and the handrail will be puled from the stowed position. Close the door by lifting the door into position, allowing the steps and handrail to fall into the stowed position, and rotating the handle counterclockwise. To open the door from the inside, lift the latch and rotate the handle clockwise to the open position and push the door open. To close, pull the door closed and allow the steps and handrail to fall into the stowed position before rotating the handle counterclockwise.

The cabin door is an emergency exit and it must be accessiMe at all times.

The CAWS annunciator PASS DOOR will illuminate when the door is not properly closed and locked. 4

CARGO DOOR 1

The cargo door is located in the aft left fuselage and Is 4 ft 4 in (1.32 m) high by 4 ft 5 in (1.35 m) wide. It is secured by locking pins which can be checked visually fmm outside the airplane to verify engagement. The door is hinged at the top and swings up out of the way to facilitate loadiig and unloading. A gas cylinder assists in door operation and hdds the door in the open position. A non-inflatable rubber gasket attached to the door seals the gap to allow the cabin to pressurize when the door is dosed.

To open the door from the outside, push the button and pull the handle outward and upward. The gas cylinder will assist in raising the door to the open position. To close, pull down on the strap to bring the door almost closed and stow the Strap. Push the door closed and push handle in until flush and the button pops back to the lock position. To open the door from the inside, lift the lever and pull handle to unlock and then push open the door. To close, pull down on the strap to bring the door almost closed and stow the strap. Pull the door closed and push handle down to the lock position.

An optional motor may be installed to assist the closure of the cargo door. To operate, press and hold the switch located aft of the cargo door until the door has lowered to the near closed position. The power supply to the motor is from the BAT DIR BUS and is disconnected by a

Report No: 01 973-001 7-24



microswitch which is operated by the drive mechanism when the door is nearly closed. The door must be manually pushed and locked to the closed position.

The CAWS annunciator CAR DOOR will illuminate when the door is not properly closed and locked.

WINDOWS

A two-piece windshield and two side windows provide cockpit visibility. Both pilot and copilot windshields are laminated twin-layer mineral glass with an embedded polyvinyl butyrol (PVB) layer. The windshield incorporates twin-zone dual electric heating elements for defogging and anti-icing capability. Both side windows are stretched acrylic with inner 2 mm thick double- glazed acrylic windows. A separate direct vision (DV) window, also stretched acrylic, is installed in the left side window. This can be opened to provide pilot visibility/smoke evacuation during emergencies and can be used to provide additional airflow during ground operations.

Windshield heat is controlled by two switches. LH WSHLD and RH WSHLD, on the pilot's lower right switch panel. These switches offer two heat levels to be used as required for defog and anti-ice. The windshield is protected from an overheat condition by a temperature sensor. This sensor w l remove current from the windshield heat circuit when the windshield surface temperature is above 60" C.

The cabin has four windows on the left side and five on the right side. All of the windows are stretched acrylic with integral sliding shades.

Windshield heat operation is controlled by two switches on the DE-ICING control panel (pilots lower right switch panel). The system is selected by moving the switches RH WSHLD and LH WSHLD to LIGHT or HEAVY. A steady green CAWS caption WSHLD is illuminated. in the event of a failure the green CAWS caption will be extinguished.

EMERGENCY EXIT

The overwing emergency exit is located over the right wing and is 2 ft 2 in (0.68 m) high by 1 ft 6 in (0.49 m) wide. This exit contains a window and can be quickly opened from either inside or outside when required. A non-inflatable rubber gasket attached to the exit seals the gap to allow the cabin to pressurize when the exit is in place. To open the exit from inside, remove cover and pull handle to release exit locking mechanism and pull inward. To open from the outside, push on the release lever and push exit inward. Refer to Figure 7-4, Emergency Exit.

AIRCRAFT SECURITY

To secure the aircraft when parked, install the lock pin in the emergency exit (Refer to Fig. 4) and lock the cargo and passenger door locks. Lock the sewice door under the rear fuselage, if a lock is installed.


Report No: 01973-001 7-25

SECTION 7 =PlLATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

OUTSIDE CABIN LOOKING INBNBOARD

INSIDE CABIN LOWING WTBOARO

Figure 7-4. Emergency Exit

Report No: 01973-001 7-26


EPRATUSE SECTION 7 ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

CONTROL LOCKS

The elevator and ailerons can be secured by placing a control lock through the hole in the collar

l and control column when the elevator is full down and the ailerons are neutral. For flight the control lock is stowed in a stowage point located on the cockpit left sidewall to the rear of the pilots seat. The rudder is held in position by the mechanical connection with the nose wheel steering.

THE CONTROL LOCK MUST BE REMOVED BEFORE TAKEOFF.

Issued: June 10,1994 t Revision 5: May 10, 1996

Report No: 01973-001 7-27

SECTION 7 ePILATUS= AIRPLANE AND SYSTEMS DESCRIPTION PC XC

DESCRIPTION AND OPERATION

Refer to Figure 7-5, PT6A-678 Engine, for engine configuration.

This a~rplane is powered by the Pratl 8 Whitney PT6A-678, which is a light weight, reverse flow, free turbine engine. In addition to the gas generator section, the PT6A-67B incorporates a power section with the power turbine and propeller reduction gearbox, an integral oil system, and an accessory gearbox for mountings for various accessories.

Air enters the compressor through an annular plenum chamber. The compressor consists of four axial stages and a single centrifugal stage. Stator vanes between each stage of compression diffuse the air, raise its static pressure, and direct it to the next stage of compression. From the centrilugal compressor, air flows through a diffuser tube, then changes direction 180 degrees as it flows into the combustion chamber. A compressor bleed valve is installed on the gas generator case at the 3 o'clock position. It automatically opens to spill interstage compressor air to prevent compressor stall.

The combustion chamber consists of two perforated annular sections bolted together with a large exit duct. Compressed air enters the combustion chamber through the perforations, where it is mixed with fuel and ignited. The rapidly expanding gas is directed through another 180 degree direction change into the turbine.

The turbine consists of a single stage compressor turbine and a two-stage power turbine. As the gas exits the combustion chamber, it is directed onto the compressor turbine, which powers the compressor. From the compressor turbine, the gas is directed to the two-stage power turbine which drives the propeller via the propeller reduction gearbox. Engine inter turbine temperature (ITT) is measured between the compressor and power turbines.

Gas flow is directed into the exhaust duct from the turbine. The exhaust duct has an annular inlet which leads exhaust gas to a bifurcated duct connected to two opposed exhaust ports. The exhaust duct is made from heat resistant nickel alloy metal and incorporates mounting flanges for the exhaust nozzles.

AIR INDUCTION

The air induction system is integrated into the front and rear lower cowlings and comprises of an air inlet and inlet duct, a plenum, and an inertial separator.

Report No: 01973-001 1 7-28

Issued: June 10, 1994 Revision 4: November 24, 1995

FUEL MANIFOLD ADAPTER AND COMBUSTION NOZZLE ASSEMBLY CHAMBER LINER

PROPELLER GOVERNOR EXHAUST TURBINE ACCESSORY

REDUCTION FUEL FLOW DIVIDER CENTRIFUGAL COMPRESSOR GEARBOX AND PURGE VALVE IMPELLER ROTOR ASSEMBLY

SECTION 7 SPLATUS= AIRPLANE AND SYSTEMS DESCRIPTION SO x d 1

The air inlet consists of a crescent shaped metal leading edge through which hot exhaust is 7 passed to prevent ice accumulation. The exhaust gas is extracted from the left hand side exhaust stub by the means of a 1.5 inch diameter pitot probe inserted into the stub itself. It then passes through the lip, consisting of a sealed chamber, before exiting into the right hand , stub through a 1.5 inch discharge tube. The probes are connected to the exhaust lip by 1.5 , inch diameter metal ducts complete with integral connectors. The inlet duct, which connects the inlet lip to the plenum, consists of a diverging nozzle following the same general shape as the inlet lip. 4

4

The plenum consists of a sealed circular metal canister surrounding the engine compressor inlet screen. It is here that the engine draws air to be compressed for combustion and services

1

supply.

The inertial separator is of the 'fixed geometry' design and provides engine induction system protection when operating in icing or FOD conditions. It can be used for takeoff when operating in a FOD environment. It comprises of a fixed No. 2 mesh screen attached to the rear wall of the plenum covering a percentage of the inlet area, a moveable outlet door and electrical actuator situated directly above the oil cooler outlet exit, and a converging by-pass duct.

In normal operations (non icing, non FOD) the outlet door is closed which seals the by-pass and provides the induction air with a single flow path to the plenum and engine through the porous No. 2 screen.

In icing or FOD conditions the actuator is retracted to open the outlet door. This allows a flow path past the plenum to ambient and increases the pressure ratio across the inlet system. The increased pressure ratio has the effect of accelerating heavy particles present in the inlet air. which then go straight past the plenum and into the by-pass duct before exiting through the outlet door. In icing conditions the porous No. 2 screen ices to restrict the flow path of sdid particles which can not turn into the plenum and thus further assist in engine protection. However the pressure of the air to the engine, with the inertial separator open, is also reduced with consequent reduction in available engine performance.

The inertial separator outlet door operation is controlled by a switch on the DE-ICING control panel (pilots lower right switch panel). The door is selected open by moving the switch labelled INERT SEP to OPEN. A steady green CAWS caption INERT SEP is illuminated when the door reaches its fully open position. When the door is selected OPEN but does not reach its selected position, the INERT SEP caption will not illuminate indicating a failure condition.

I After failure of the inertial separator, the aircrew should prepare for departure of icing conditions I as soon as possible.

Report No: 01973-001 7-30

Issued: June 10, 1994 Revision 5: May 10, 1996 :

'FSPILATUSF SECTION 7 ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

ED MESH SCREEN

PASS DOOR SHOWN IN THE N-ICINO (CLOSED) POSITION

ACTUATOR

t Figure 7-6. Inertial Separator

Issued: February 14, 1994 Revision 2: February 14,1995 ---

Report No: 01973401 7-3 1 I

SECTION 7 =PLATUS AIRPLANE AND SYSTEMS DESCRIPTION PC XI1

CONTROLS

Refer to Figure 7-7, Engine and Propeller Controls I Indication.

Eng~ne power is controlled by POWER CONTROL, CONDITION, and MANUAL OVERRIDE levers located on the center console.

POWER CONTROL LEVER

The POWER CONTROL lever (PCL) selects the required engine power (Ng) and in certain conditions it dlrectly controls the propeller pitch. The PCL has a flight and a ground operating range separated by an idle detent. The flight operating range is forward of the detent. As the PCL is moved foward of the idle detent the minimum propeller pitch (6" to 12") is &recUy controlled by the PCL while the propeller is in an underspeed condition during low engine power at a low airplane speed. When the PCL is moved further forward, engine power and airplane speed increase until each are high enough for the propeller to operate in a constant speed mode. In this mode, the Conslant Speed Unit (CSU) selects the propeller pitch to maintain a propeller speed of 1700 rpm.

When the PCL is at the idle detent, the gas generator is at idle and the propeller is at minimum pitch. A lilting action to raise the PCL over the detent is required to move the PCL into the ground operating range.

P WARNING 1

DO NOT MOVE THE PCL BELOW THE IDLE DETENT WHEN THE ENGINE IS NOT OPERATING TO PREVENT DAMAGE TO THE CONTROL LINKAGE.

PCL OPERATION AFT OF THE IDLE DETENT IS NOT PERMllTED IN FLIGHT OR WHEN ENGINE OPERATION IS CONTROLLED BY THE MANUAL OVERRIDE LEVER.

Aft of the Idle detent is the ground operating range or beta mode. The Nf governor limits the propeller speed to an underspeed concbtion to give the beta valve fun authority in controlling the propeller pitch. The engine power and propeller pitch are directly controlled by the PCL. Initial PCL movement aft of the idle detent adjusts the propeller pitch while the gas generator remains at idle and can be used to control taxi speed. Further aft movement causes the propeller to move into the reverse range followed by an increase in engine power.

Report No: 01 973-001 7-32


EPtLATUS= SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

I I PILOT'S LOWER LEFT PANEL

\ STARTERAND IGNITION SWITCHES

\ ENGINE INSTRUMENT DISPLAY UNIT /

PILOT'S LOWER RIGHT PANEL

( PROPELLER DE ICE SWITCH

INERTIAL SEPARATOR SWITCH

- FUEL FIREWALL SHUTOFF VALVE

(POST SB 28 - 002 AND MSN 161 AND UP)

Figure 7-7. Engine and P r o p e l l e r Controls I Indication

Issued: June 1 0 , 1994 Revision 6: Dec 6. 1996

Report No: 0 1 9 7 3 - 0 0 1

7-33

SECTION 7 'f PILATUS W AIRPLANE AND SYSTEMS DESCRIPTION ?c %)I

MANUAL OVERRIDE LEVER

The MANUAL OVERRIDE lever (MOR) is located on the center console to the left of PCL. The MOR controls the engine power in case of a pneumatic fa~lure of the engine fuel control or in case of a PCL-system failure.

The MOR directly operates the fuel metering valve by mechanically compressing the pneumatic bellows (Py pressure) in the FCU. In case of a Py leak, the engine spools down to minimum fuel flow with no response lo PCL inputs. Minimum fuel flow varies from 35% N on the ground to 7O0lO N~ at 30.000 11. With forward speed the engine would spool down very L w ~ y (from iae to minimum fuel flow it lakes approximately 1 to 2 minutes) and gives the pilot enough time to recover engine power with the MOR before Ng drops below 50%. Maximum fuel flow available from a minimum fuel flow condition is 512 Ibhr wh~ch gives approximately 750 SHP.

During normal engine operation the MOR lever is in the full aft or OFF position with the indicator thread attached. In case of a PCL or pneumatic failure, the PCL should be positioned at the maximum position to allow full MOR authority and in order to limit the propeller pitch to 12" which allows for a more stable Constant Speed Unit operation. I f Ng is above 50%, unrestricted use of the MOR is permitted. If the engine is allowed to drop below 50% Ng, starter assistance may be required to recover Ng above 50%. Movement of the MOR lever should be done very carefully so as not to cause compressor surges and ITT overtemperature. However on the ground with no forward speed it is not possible to recover engine speed.

To operate the MOR, lift up on the lever and move it forward as desired towards the MAX position. There is a small deadband forward of the OFF position before the engine responds to movement of the MOR lever.

Report No: 01 973-001 7-34

Issued: February 14, 1994 Revision 2: Feb~ary 14, 1995

WPILATUSE SECTION 7 ?a %I) AIRPLANE AND SYSTEMS DESCRIPTION

I WARNING I PCL OPERATION AFT OF THE IDLE DETENT IS NOT PERMITTED WHEN ENGINE OPERATION IS CONTROLLED BY THE MANUAL OVERRIDE LEVER.

THE MOR LEVER MUST BE IN THE OFF POSITION PRIOR TO ENGINE START TO PREVENT A HOT START.

NOTE

During MOR operalion the Py pressure has no authority which causes the loss of the toque limiting function and Nf governor operation (no reverse power to be used). Maximum Ng is also not limited.

Issued: Feb~ary 14,1994 I Revision 2: Feb~arv 14. 1995

Report No: 01973-001 7-35 I

SECTION 7 WPILATUS AIRPLANE AND SYSTEMS DESCRIPTION PC XI1

CONDITION LEVER

The Condition Lever has three positions and is used to select the gas generator idle speed, shut down the engine, and feather the propeller. The GROUND IDLE position (G.I.) is tor ground operation only. This selling insures that the propeller speed remains above the prohlbiled range (350 rpm to 950 rpm) for ambient conditions up to a temperalure of approximately 45" C. The FLlGHT IDLE position (F.1.) is selected for flight operation which provldes sufficient bleed air flow to maintain cabin pressurization al minimum Ng and to give smooth engine response to PCL movement during approach and landing. The CUT- OFFIFEATHER position mechanically stops the fuel low to shut down the engine and electrically feathers the propeller at the same time. The low fuel pressure switch is inoperalive when the Conation Lever is in the CUT-OFFIFEATHER position.

To move the Condition Lever from the CUT-OFF position the entire lever needs to be tifled. To move the lever to CUT-OFFIFEATHER it has to be lifted to get over the hard stop. Movements from GROUND IDLE to FLIGHT IDLE and vice versa need only pressure on the lever towards the new positions to overcome the soft stop.

The engine STARTER and IGNITION switches are located on the pilot's lower left switch panel.

Report NO: 01 973-001 7-36

Iswed: Febnrary 14,1994 Revision 2: February 14, 1995

3PILATUSW SECTION 7 PC XI1 AIRPLANE AND SYSTEMS DESCRIPTION

ENGINE FUEL

Refer to Figure 7-8, Engine Fuel System, for system configuration. For airplane fuel storage and distribution, refer to Fuel system, this section.

The engine fuel system consists of an oil to fuel heat exchanger, a high pressure engine driven fuei pump, a fuel control unit, a fuei flow transducer, a fuel flow divider and dump valve, and the fuel nozdes. .

Fuel is delivered to the fuelloil heat exchanger from the low pressure engine driven pump. The oil to fuel heat exchanger pre-heats the fuel, to eliminate the chance of ice formation in the fuel, and reduces the oil temperature. The high pressure engine driven fuel pump delivers fuel to the fuel control unit after It passes through the oil to fuel heat exchanger. See Fuel System for more information.

The fuel control unit is controlled by the POWER CONTROL and CONDITION levers during normal operations, and the MANUAL OVERRIDE lever during emergency operation. Fuel flows through the fuel flow transducer on its way to the fuel flow divider and dump valve. The fuel flow transducer converts fuel flow rate into an electrical signal which is then displayed in the cockplt on the Engine instrument System (EIS).

The fuel flow divider and dump valve serves two functions. First, it divides the fuel between the primary and secondary system. Secondly, It directs air from the purge air accumulator into the fuel manifolds to purge them of unused fuel at engine shutdown. A total of 14 fuei nozdes are used with a primary and secondary spray pattern.

POWER c m a LEVER

FLTER BYPASS VALVE - - - MECWMCAL L m a E - HlGH PRESSURE FUEL

KEY LOW PRESSURENK

- O I L - VENTLNE


I Revision 2: February 14, 1995

Figure 7-8. Engine Fuel System

Report No: 01 973-00 1 7-37 I

SECTION 7 rCrPILATUS3. AIRPLANE AND SYSTEMS DESCRIPTION PC12

OIL

Refer to Figure 7-9, Engine Oil System, for system configuration.

The engine oil system consists of pressure, scavenge, and breather systems with the oil tank being an integral part of the engine compressor inlet case. Oil is supplied to the engine bearings, reduction gears, accessory drives, torquemeter, and propeller governor. Oil is also used to cool the bearings. A filler neck with quantity dipstick and cap are located on top of the accessory gearbox. The quantity dipstick is marked in one US quart increments. A visual sight gauge is provided to determine oil quantity without removing the dipstick. If the oil level is in the green range of the sight gauge there is sufficient oil quantity for flight. If the oil level is below the green range, the oil system needs refilling according to the dipstick markings. If the CAWS OIL QTY caption illuminates, the oil level is not adequate for safe engine operation. It is not recommended to start a flight with the oil level below the green range on the sight glass. Total oil capacity is 3.6 US gal (13.6 liters) while usable oil quantity is 1.5 US gal (5.7 liters). The oil tank incorporates a drain plug.

An engine driven gear type pressure pump provides oil to the engine bearings, torquemeter, propeller bearings and reduction gears, and propeller governor. Oil flows from the integral oil tank, through the pick-up screen, to the oil pump. Oil then goes through a pressure regulating valve which regulates oil pressure to between 90 and 135 psi (6.2 to 9.3 bar).,A pressure relief valve opens when pressure exceeds 160 psi (11.0 bar), possibly during cold weather operations. Oil then goes through a cartridge type oil filter assembly, which incorporates a bypass valve and a spring loaded check valve. The bypass valve allows oil to bypass the filter in case the filter becomes clogged, however oil pressure drops to below 90 psi (6.2 bar) when the filter by-pass valve is open. The check valve prevents gravity oil flow into the engine after shutdown and permits the oil filter to be changed without draining the oil tank. Oil is then directed throughout the engine and applicable accessories.

The oil scavenge system incorporates two double element pumps. The oil from the reduction gearbox is pumped directly through the airframe mounted oil cooler. All remaining oil passes through the oil to fuel heat exchanger and, depending on oil temperature, is directed back to the oil tank or through the oil cooler.

When the fuel temperature is low, warm oil flows through the oil to fuel heater. At fuel temperatures above 21" C (70' F) the bypass valve begins to open and at 37" C (98' F) the bypass valve is fully open and the oil bypasses the oil to fuel heater. The scavenge system in the propeller reduction gearbox incorporates a magnetic chip detector that detects foreign matter in the system and causes the CHlP caution light on the CAWS annunciator panel to come on. Post SB 31-005 the CAWS CHlP caution light will illuminate on the ground and in flight. The chip detector also acts as the propeller reduction gearbox oil drain. Optional SB 79- 005 installs a second magnetic chip detector in the accessory gearbox. It is also connected to the CAWS CHlP caution and operates in parallel to the reduction gearbox chip detector.

The breather system allows air from the engine bearing compartments and the propeller reduction and accessory gearboxes to be vented overboard into the right exhaust stub, through the centrifugal breather in the accessory gearbox.

Report NO: 01973-001 7-38


PROPELLER W A F 1

C U FILLER AN0 MPSTICK

TMRUST BEARlffi THERWSTATIC BVPASS W C H E C K VALVE

OIL-TO-FUEL HEATER

R L D U C m CASE SCAVENGE PUMP

REOWTION GEARBOX CWPOEECTOR

- PRESSUREW

0 1 L L E R H D R V C U

0 1 A V E H G E G i l

BREAT+IERAIR

0 TORWEMETER PRESSURE


STARTING

Starting is provided by a combination starterlgenerator unit. Starter function is controlled by the STARTER switch on the pilot's lower left panel. This switch has two momentary positions, ON and RESET. When ON is pressed the Starter Master and the Starting Control Relays close to energize the starter. The starter will automatically disengage thirty seconds (Pre SB 24-004) 1 sixty seconds (Post SB 24-004) after the Condition Lever is moved to the GROUNDIFLIGHT IDLE position. Post SB 80-001 the Starting Control Relay is controlled by the Starter Conlrol Unit and the starter is disengaged when the engine Ng reaches 50%.

The start sequence can be interrupted at anytime by pressing the RESET position of the STARTER switch. When pressed, fhe ground is removed from Ule Starter Master Relay causing it to open and remove electrical power from the starter circuit. The Delay Starling Sequence Relay will automatically reset to zero seconds at the completion of a normal start sequence or after the RESET switch has been pressed.

During cold weather starting the thirty second starting relay (Pre SB 24-004) sequence may be completed before the engine reaches 50% Ng. The starter can be re-engaged by pressing and holding the starter switch to ON. If the starter switch is released it will reset straight away.

For improved engine starting with a cold engine (oil temperature below +5' C), use FLIGHT IDLE position for engine starting. At FLIGHT IDLE more fuel is provided during the start cycle to enhance engine acceleration to idle speed.

For information on the generator function, refer to Electrical System, this section.

IGNITION

Ignition is provided by an ignition exciter and two spark igniter plugs. The ignition exciter is a sealed electronic unit mounted at the engine cowling and is operated by the aircraft 28 VDC system. Two spark igniter plugs, located at the 4 and 9 o'clock positions in the gas generator section, provide the spark to ignite the fueVair mixture.

Ignition is controlled from the cockpit by the IGNITION switch, located on the pilot's bwer left panel. This switch has two positions, ON and AUTO. When set to ON, ignition will occur continuously.

( Pre SB 77-002. When set to AUTO, ignition will automatically activate when the ITT is less than 500' C with normal fuel system pressure and the condition lever forward of the CUT- I OFFIFEATHER position.

Post SB 77-002 and MSN 261 and UP. When set to AUTO, ignition will automatically activate, regardless ot the PCL position, when the ITT is less than 500' C and the Ng is 10% or more. It stops 10 seconds after the ITT is more than 500' C and when the Ng is less than 10%.

Ignition should be manually switched ON when operating in heavy precipitation.

Report No: 01973-001 7-40


S P I L N U S E SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

FIRE DETECTION

The system Is composed of a sensor element and a responder. The sensor Is a stainless steel capllary tube filled with helium and containing a central hydrogencharged core which readily releases hydmgen gas when heated above a temperature threshold. The responder houses both the fire pressure switch and the integrity switch consisting of performed metal diaphragms which snap over center to contact statlonary pins under the effect of gas pressure.

Due to generalized temperature increase over the entire length of the sensor, the helium pressure increases and actuates the fire pressure switch triggering the alarm. Alternatively, when the sensor Is heated up Intensely over a short length, the core material releases hydrogen gas causing a pressure rise and actuates the fire pressure switch. The red ENG FlRE warning light will illuminate whenever the fire pressure switch Is activated. Both the averaging and discrete functions are reversible. When the sensor tube is cooled, the average gas pressure is lowered and the discrete hydrogen gas returns to the core material. The reduction of internal pressure allows the alarm switch to return to its nomal position, opening the electrical alarm switch.

In addition to the pressure activated alarm switch, the Integrity switch Is held closed by the averaging gas pressure at all temperatures down to -55" C. If a detector should develop a leak, the loss of gas pressure would allow the integrity switch to open and activating the system fault cautlon. The amber FlRE DETECT caution llght will illuminate when the Fire Detection system I Is inoperative.

System integrity is checked by pressing the FlRE test switch, located on the Test Panel. When pressed, the availability of electrical power and circuit continuity is checked. Proper system function Is indicated when both the ENG FlRE and FlRE DETECT annunciators illuminate. If the FIRE DETECT annunciator fails to illuminate during the test, the warnlng circuit is already dosed and will not provide proper warnlng.

I

[ CAUTION I DUE TO THE COMPOSiTE CONSTRUCTION OF THE ENGINE COWLING AND THE POSSIBILITY OF TOXIC GASSES, THE AIRPLANE ECS MUST BE SHUTOFF WHEN A FlRE CONDITION IS SUSPECTED.

Issued: February 14,1994

t Revision 2: February 14, 1995 Report No: 01 973-001

7-4 1


ENGINE INSTRUMENT SYSTEM ( EIS )

The Engine Instrument System is a computer controlled system which receives input from the sources shown in Figure 7-10. It displays engine and other system information and provides warnings when certain parameter limits are exceeded. The system comprises two Acquisition and Processor Units (APU) and a Display Unit (DU). Within each APU is an internal power supply and self test function.

Inputs for the displayed parameters are compared in the APU with the limitations for those parameters. If the limitations are exceeded the system illuminates an amber caution or red warning light in the DU. To help identify the exceeded parameter the appropriate LCD display will flash 40 timeslminute when the parameter is in the caution range and 80 time~/min~te when the parameter is in the warning range.

Post SB 77-002 and MSN 261 and UP. Illumination of an amber caution or a red warning on the EIS can be cancelled by pressing the appropriate master CAUTION or WARNING l ih t switch. The appropriate display will continue to flash. This will allow for any further exceeded limitations to be annunciated and shown.

The system self lest is activated by momentarily pressing: - the EIS switch on the System Test Panel located on the left side panel (MSN 101-1 11) - the test switch on the EIS panel (MSN 112 and UP). The EIS self test progressively activates all of the LCD display segments, illuminates the caution and warning lights, and checks the APUs and power supplies for serviceability. If there are no fauls detected after all of the display segments are activated, the displays will return and operate normally. If an internal faun is detected, the fault code will appear on the display before returning to the normal display. If the self test is initiated before engine start and a fault detected the EIS caution light will flash. The caution light will not flash if a fault is detected after the engine is running.

Continuous internal monitoring of the EIS system detects internal failures and automatically selects an alternate signal path for the displayed parameter wlhout alerting the pilot when the engine is running. An internal failure will be signalled to the pilot when the engine is not running by a llashing EIS DU amber caution light. To identify which failure a Self test shouM be initiated. A self test can also be initiated with the engine running in flight to check for serviceability of the EIS system. See Section 2 Limitations for EIS failures not accepted for takeofi.

For aircraft MSN 101-1 11 the EIS failure codes and system status are as follows:

Amber Caution Light flashing

300 Display Unit failure, a double APU failure causes all of the EIS displays to freeze

302 Display Unit power failure, a double power supply failure causes the EIS to go blank and the DU caution light to Mink


SECTION 7 AlRPLANE AND SYSTEMS DESCRIPTION

306 No transmission

566 Failure of one acquisition and processing channel

304 Transmission error

305 Test phase

307 Serial interface or LCD failure

Aircraft MSN 112 and UP have an EIS with the two independent Acquisition and Processing Units (APU) and a Display Unit (DU) combined into a single unit. Each APU driwes a section ol the DU, both APUs acquire and can display the critical parameters of TORQUE and Ill. If a failure occurs in the APU that normally drives the TORQUE and ITT displays, the pilot can select TORQUE and Ill from the other APU by pushing the menu switch on the DU. When the secondary menu is selected the TORQUE and I l l digital values are displayed in place of the OAT and ENDUR displays. The primary analog and digital displays for TORQUE and ITT go blank.

The APUs also process the fuel quantity and fuel flow signals received from the fuel tank probes and provide a wing fuel quantity balancing function and provide continuous monitoring of all the EIS parameters. EIS fault codes are allocated to specifically identify a fault, a full 1 listing of the fault codes can be found in Chap 77 of the Aircraft Maintenance Manual.

If an amber caution light illuminates after a system test before takeoff, system redundancy is lost, but flight is permitted. Maintenance is required as soon as practicable to clear the system fault.

NOTE

A fault of all serial data lines from the APUs to the DU causes the DU to turn off all display modules and suspend parameter updating. No failure code will be presented.


Report No: 01 973-001 7-43

SECTION 7 mPlLATUSW AIRPLANE AND SYSTEMS DESCRIPTION K12

Within the EIS display is a fuel totalizer function. The digital display indicates total fuel remaining, fuel flow, fuel used and endurance (MSN 112 and UP). The fud QTY indicates the total computed fuel quantity in the wings. The fuel USED indicates fuel consumed based on fuel flow vs time (FUH) of engine operation. The ENDUR display (MSN 112 and UP) has a range of 9 hrs and 59 minutes and is computed from fuel fbw during engine operation. The QTY and USED values are stored in non volatile memory at shutdown. After engine start, push the FUEL RESET switch on the DU to re-datum fuel quantity to the actual fuel contents detected by the tank capacitance probes and to reset the fuel USED to zero. When the FUEL RESET switch is pushed, the APUs totals the indicated left and right fuel quantities, stores the new value in memory, and displays quantity on the DU. As fuel is consumed, the fuel QTY and ENDUR values (MSN 112 and UP) will be recomputed based on the current fuel flow.

NOTE

The computed fuel QTY value is based on a previous known quantity and will not indicate the increase in fuel quantity after fuel has been added to the wings until the FUEL RESET switch has been pushed. On aircraft 112 and UP the FUEL RESET switch must be pushed after engine start to update the digital display.

If a failure of the fuel flow sensing system occurs the fuel OTY and USED values are frozen.

I Re-setting of the fuel quantity indication is possible in flight (MSN 101-111, MSN 261 8 UP and EIS units modified to Part No. 975.29.02.015 Model PEFlOA-2C) by pressing the FUEL RESET button when the aircraft is in straight and level flight.

Electrical power is supplied to the APUs and DU from the generator 1 bus and battery bus through the EIS 1 and 2 and the EIS ACQ 1 and 2 circuit breakers (MSN 101-1 11) and the EIS 1 and 2 (MSN 112 and UP). Power to Ihe LCD permanent back-lighting is supplied from the battery bus through the EIS LIGHT circuit breaker.

Report No: 01 973-001 7-44


WPLATUSF SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

OENERATOR 1 INDICATION SWITCH

INDICATION SWKCH

WARNINQ HDWTOR

CAUTION NDWTOR

\-

ENGINE INSTRUMENT DISPLAY "NK \

lo SYSTEMTESTPANEL

Figure 7-10. EIS (MSN 101-1 11) (Sheet 1 of 7)


Report No: 01 973-001 7-45

SECTION 7 WPLATUSW AIRPLANE AND SYSTEMS DESCRIPTION KL-2

WARNHG INDICATOR

CAUTION INOICATOR

ENGINE INSTRUMENT DISPLAY UNIT \

FUEL RESET SWITCH

Figure 7-10. EIS (MSN 1 12-999) (Sheet 2 of 7)


E P I L A N S F SECTION 7 Kl2 AIRPLANE AND SYSTEMS DESCRIPTION

FUEL OW RESET

SYSTEM TEST W E R RESET

BAllERY BUS

+I EIS ACO

I

EIS DU 1

PROCESSOR I PROCESSOR CHANNELA I CHANNEL B

I I t-

I DlSPUY INTERFACE I

Figure 7-10. EIS (MSN 101-1 11) (Sheet 3 of 7)

issued: June 10.1994 Revision 9: September 1. 1999

ENGINE RUNNPla IGNITION &cwc

CMJTROL OUTPUTS

Report No: 01973-001 7-47

SECTION 7 WPLATUSB AIRPLANE AND SYSTEMS DESCRIPTION P(p

CONTROL INPUTS 1 1

BATTERY BUS

BATTERY BUS

Figure 7-10. EIS (Pre SB 77-002 MSN 112-260) (Sheet 4 of 7)

Reporl No: 01 973-001 7-48



CONTROL INPUTS

BATTERY BUS

LIGHT

BAHERY BUS

-

GENERATOR 1

BUS

EIS 2

CONTROL OUTPUTS

IGNITrn <MOnc

FUEL PUMP LH

FUEL PUMP RH

Figure 7-10. €IS (Post SB 77-002 and MSN 261-999) (Sheet 5 o f 7)


Report No: 01 973-001 7-48A



Report No: 01973-001 ' I 7-488

/ Issued: June 10,1994

Revision 9: September 1, 1999 I

-

qPILATUSW SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

EIS CAUTlONS AND WARNINGS

Note 1: Caution or warning inhibited during pre-start and post-flight. Note 2: Caution or warning inhibited during engine start. Note 3: 1 the warning illuminates for a parameter, the caution automatically extinguishes.

Figure 7- 10. EIS (Sheet 6 of 7)

I

PARAMETER

Torque

Ill

During Start Only

Prop RPM above 1000

Ng

Oil Temperature

Oil Pressure

Ng above 72 %

OAT


I

Report No: 01 973-001 7-49

CAUTION RANGE (Digits blink 4OImin)

44.4 to 61 .O psi

800 to 870" C

870 to 1000" C

below 350" C

below 60% (engine run) (MSN 1 12-999)

-40 to 10" C (Note 1 & 2) 105 to 110" C

40 to 60 psi (Note+l

1 &2)

60 to 90 psi (after 5 sec. delay) 135 to 200 psi

below +4" C and Probes de-ice switch OH (Note 2)

WARNINO RANGE (Digits blink 80Imin)

44.4 to 61.0 psi (after 20 sec. delay).

above 61 psi

800 to 870" C (after 20 sec. delay). above 870" C

870 to 1000" C (after 5 sec. delay). above 1000" C

above 104 %

below -40" C (Note 2 )

105 to 110" C (after 10 min. delay). above l i oO c

below 40 psi (Note 1& 2). 40 to 60 psi (after 20 sec. delay)

60 to 90 psi (after 20 sec. delay). 135 to 200 psi (after 20 sec. delay). above '200 Psi

None

SECTION 7 mPLATUSW AIRPLANE AND SYSTEMS DESCRIPTION K12

- - -- --

EIS CAUTIONS AND WARNINGS (CONT'D)

Note Note Note Note

PARAMETER

N~

Ng above 90 %

DC Volts (MSN 101 -1 1 1)

DC Amps (MSN 101-111)

DC Amps Bat (MSN 1 12-999)

DC Amps Gen 2 (MSN 1 12-260)

DC Volt Bat (MSN 261 -999)

M: Volt Gen 2 (MSN 261-999)

1 : Caution or warning inhibited during pre-start and post-flight. 2: Caution or warning inhibited during engine start. 3: If the warning illuminates for a parameter, the caution automatically extinguishes. 4: 'below -60' means more than 60 amps discharge.

Figure 7-10. EIS (Sheet 7 of 7)

Repon No: 01 973-001 7-50

CAUTION RANGE (Digits blink 401min)

below 950 (on ground) (Note 1 & 2) (MSN 101-111)

below 1640 (after 5 sec. delay) 1760 to 1870

below 22.0 V (after 3 sec. delay) (Note 2)

beween 29.5 to 32.2 V

below -60 amps (Note 2 and 4)

below -60 amps (Note 2 and 4)

above 140 amps


above 29.6 V

below 22.0 V (after 3 sec. delay if Ng more than 50%)

above 31.5 V

Issued: June 10.1994 Revision 10: September 1.2000

WARNING RANGE (Digits blink 80imin)

below 950 (on ground. after 10 sec. delay) (Note 1 8 2) (MSN 101-1 11)

below 950 (on ground, after 5 sec. 1 unfeather delay) (Note 1 & 2) (MSN 1 12-999)

1760 to 1870 rPm (after 20 set. delay)

above 1870


above 32.2 V

None

Continuous increase above 10 amps in 7 seconds

None

None

None

SPILATUSF SECTION 7 PC XI1 AIRPLANE AND SYSTEMS DESCRIPTION

TORQUE LIMITER

A torque limiter is installed on the engine at the torque transmitter boss on the forward engine case. Within the unit is a sealed bellows connected directly to the torquemeter oil pressure outlet, a chamber connected to the reduction gearbox to provide gearbox static pressure and lo a drain port on the thrust bearing cover, a balance beam, and a pneumatic pressure orilice

Oil pressure proportional to engine torque is applied through cored passages in the reduction gearbox to the sealed bellows in the limiter body. The bellows is mechan~cally connected to the balance beam and lo the conlrolling spring. With an increase in torque pressure, above the control spring selling, the balance beam adjusts to compensate lor this increase and causes the pneumatic pressure orilice to open and bleed oft Py air. As Py air pressure is bled off, the fuel llow from the FCU is reduced by closing the metering valve, causing engine speed and hence engine torque to decrease until engine torquemeter pressure is balanced by the torque control spring pressure; at this time the Py pressure orifice close.

The torque limiter limits the englne torque to below 44.3 psi. Due to ambient pressure at I altitude and interference w~th the FCU maximum governing speed, maximum torque may not be obtained.

Engine torque will drop approximately 1.7 psi per 10.000 feet of altitude. Above 102% to 104% Ng (maximum Ng limit) the torque will decrease by approximately 2.9 psi. Alter thls point the engine power is limited to maintain 104% Ng.

I The maximum torque droop due to altitude has been considered in the static takeoff and balked landing torque charts.

If the maximum torque according to the torque chart is below llat rating (below the torque limiter setting), the torque has to be set manually by the PCL. Torque limiter operation must always be verified to ensure engine l~m~ts are respected. During the takeoft and the balked landing the PCL does not need to be retracted unless any l~mits are exceeded. The torque (11 below flat rating) and ITT increases are acceptable.

I

Issued: June 10. 1994 Revlsion 7: July 1. 1997

Report No. 01 973-00 1 7-51

SECTION 7 GPILATUSF AIRPLANE AND SYSTEMS DESCRIPTION ?c %)I

INDlCATlONlWARNlNG

Engine indication lor normal operation is d~splayed on the EIS.

MSN 101-111

Battery or generator 1 or generator 2 voltage and current are displayed. The digital display will be lor the system as selected by the two DC IND switches located at the top of the EIS. Generator 1 or generator 2 output will be d~splayed when the GEN lor GEN 2 switch annunciates DC IND. Pressing both switches until both GEN 1 and GEN 2 DC IND annunciators are extinguished selects the battery, or external power if selected, lor display on the EIS. Continuous monitor~ng of the voltages and currents lor close lo limit Caut~ons and out ol l~mit Warn~ngs 1s prov~ded by the EIS.

MSN 112-999

Battery and generator 2 voltages and currents are displayed on two of the EIS displays (BAT and GEN 2). Cont~nuous monitoring ol the voltages and currents for close lo limit cautions and out of limit warnings is provided by the EIS. In flight, an increase in battery current ol 10 amps or more in an interval of less than 7 seconds, will give an EIS warning indication.

The CAWS warninglcautionladvisory lights that indicate engine status are:

INERT SEP lnd~cates door is open. There is a 10 second delay in the indication atter switch aclivation.

DE ICE Indicates a malfunction in the propeller deice or surlace deice

OIL QTY Indicates low engine oil quantity when the engine IS not running.

CHIP Indicates metal particles in the engine oil system when airplane is

I on the ground (and In the air with optional lit).

IGNITION Indicates that engine ign~tion is on

PROP LOW P Indicates propeller has gone below the m~nimum in-flight p~tch (6O) with the airplane not on the ground.

FlRE DETECT lnd~cates a maltunction in the engine fire detection system

ENG FIRE lnd~cates an overtemperature cond~tion and/or possible engine fire.

Report No: 01973-001 7-52

Issued: June 10, 1994 Rev~s~on 7: July 1, 1997

=PLATUS E SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

PROPELLER

GENERAL

Refer to Figure 7-7, Engine and Propeller Controls1 Indications and Figure 7-1 1, Propeller Pitch Mechanism.

The airplane is equipped with a Hartzell 105 in (2.67 m), four blade, variable pitch, full feathering propeller which is driven by the engine power turbine through a reduction gearing. The propeller hub and the four propeller blades are made of aluminum. Each blade incorporates an electric deice boot.

DESCRIPTION

The propeller is powered by the engine through the reduction gearbox. Propeller pitch is adjusted by engine oil pressure regulated through the Propeller GovernorIConstant Speed Unit (CSU). Nominal propeller rpm during all phases of operation is 1.700 rpm, except at low power settings at low speeds where there is insufficient energy available to rotate the prop at 1700 rprn.

The pitch change mechanism is mounted on the propeller front hub and consists of a fixed cylinder, a sliding piston, and a feathering spring. The piston is connected to each propeller blade by a fork assembly which engages a cam follower on the blade root. A counterweight is attached to each blade near its root in such a position that when the propeller is rotating the counterweight is transferred to the blade as a force tending to turn the blade to coarse pitch. The feathering spring within the cylinder also tends to move the blades towards coarse pitch and the feather position.

Oil pressure from the engine oil system is boosted to a higher pressure by a pump in the CSU. Oil pressure is then applied to the rear of the sliding piston, overcoming the force of the feathering spring and counterweights, to move the blades towards fine pitch. Thus, the blade angle is set by controlling the pressure of the oil supplied to the propeller.

In case that neither the CSU nor the overspeed governor limit the propeller speed, the Nf governor will limit the engine power (Np=109%)1853 rprn.

Should the CSU governing system fail, the overspeed governor will operate to limit the propeller speed (Np) to 106% (1802 rpm. The overspeed governor incorporates a feathering solenoid valve which is energized when the Condition Lever is moved to the CUT OFF position, causing the blades to feather. The BATTERY BUS must be powered to enable propeller feathering.

Issued: June 10,1994 [ Revision 5: May 10, 1996

Report No: 01 973-001 7-53

SECTION 7 =PLATUS- AIRPLANE AND SYSTEMS DESCRIPTION r c XII-

BUDE RM)T

FORK ASSEMB

PROPELLER BLADE

OIL PRESSURE

COUNTERWEIGHT

FEATHER

," FINE PITCH

REVERSING RING

"--.

OIL PRESSURE

Report No: 01 973-001 7-54

PITCH

Figure 7-1 1. Propeller Pitch Mechanism


PPLATUSS SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

OPERATION

Refer to Figure 7-1 1, Propeller Pitch Mechanism.

In normal operation the propeller unfeathers after the condition lever is moved to Ground Idle and the engine is accelerating to idle Ng during engine start. On the ground at idle power the propeller rotates at approximately 1060 rpm. When power is increased the CSU will control propeller speed at 1700 rpm. In the air, at low speeds and idle power (F.I.) the propeller rprn may drop below 1700 rpm. The propeller feathers automatically when the condition lever is moved to CUT OFFIFX.

The propeller is reversible for operation in the Ground Operating range during ground operations only. To achieve propeller pitch below the low pitch stop, lift up the triggers on either side of the Power Control Lever (PCL) to clear the idle detent and pull aft. As the PCL moves aft, the propeller blade angle decreases to the maximum reverse blade angle of -17.5".

GROUND OPERATION WITH PROPELLER BELOW 950 RPM IS NOT PERMITTED.

PROPELLER DE ICE

Each propeller Made has an electrically heated boot on the inboard upper and lower leading edge. 28 VDC power supply for the boots is taken directly from the GEN 1. It is supplied to the propeller de-ice boots via a slip ring mounted on the rear of the spinner bulkhead and brush block mounted on a bracket on the engine. Protection against the affects of lightning strike is provided by a set of metal oxide varistors (MOV's) mounted on the brush Mock assenbly. The system is selected by the PROP DE ICE switch on the DE-ICING control panel (pilots lower right switch panel).

A deice controller unit selects power alternately to opposite pairs of blades to minimize the chance of asymmetric ice shedding. When the PROP DE ICE switch is turned ON, the timer selects automatically the appropriated cycle depending on the indicated OAT. The three possible modes are:

Mode 1 (IOAT > 0°C): the timer is in stand by and none of the blades are heated.

Mode 2 (0°C 2 IOAT > -1 6°C): blades 1 and 3 are heated for 45 seconds followed by blades 2 and 4 for 45 seconds, then all blades are off for 90 seconds.

ISSut+d: June 10,1994 Revision 4: November 24, 1995 - - -

Report No: 01 973-001 7-55

SECTION 7 EPILATUSa AIRPLANE AND SYSTEMS DESCRIPTION ?c XII-

Mode 3 (IOAT 5 -16°C): blades 1 and 3 are heated for 90 seconds followed by blades 2 and 4 for 90 seconds.

i The above cycles are repeated until the DE ICING PROP switch is turned OFF.

OAT sensing is by a sensor mounted under the left hand wing. This sensor is termed the controller and presents the principal control signal. A second sensor is mounted in an identical position under the right hand wing. This sensor is termed the comparitor and allows the control sensor to be checked.

I The Propeller De Ice Controller (PDIC) also monitors various system control functions and initiates warnings in the event of detected failures. The following functions are monitored:

- Inhibit input open - Failure of OAT sensor (Open or short sensor or unacceptable difference between

OAT controll sensor and OAT comparitor sensor) - Heater supply voltage out of tolerance - Heater current out of tolerance - Built in test for PDIC internal failure (power supply, oscillator, watchdog etc).

( I f the system is switched ON and the PDIC detects a fully se~iceaMe System, the green CAWS PROP DE ICE caption illuminates continuously. I f the controller detects a failure, the PROP DE ICE caption blinks and the amber CAWS DE ICE caption illuminates. An aural gong will sound.

I Do not operate the propeller de ice system i f the aircraft is on the ground, in order to prevent overheat damage of the heating elements.

The propeller speed is displayed digitally on the EIS.

The CAWS annunciator PROP LOW P will illuminate when the propeller pitch is less than 6' (minimum pitch in flight) and the airplane is not on the ground.

The green CAWS PROP DE ICE advisory caption indicates proper system function. Upon initaition of the system the de-ice timer performs a built in test function lasting 5 seconds. A pre-flight lest is performed in this manner. The amber CAWS DE ICE advisory caption will illuminate and aural gong will sound and the green CAWS PROP DE ICE advisory caption will blink when the system electrical load is outside its limits.

Report No: 01 973-001 7-56

Issued: June 10,1994 Revision 5: May 10. 1996

eP ILATUSI SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

FUEL

GENERAL

Fuel is contained in two integral wing tanks and is supplied to the engine in excess of that required for all ground and flight operations. Each wing tank contains drain valves. The transfer and delivery of fuel is achieved using a motive flow jet pump system and two engine driven pumps (low pressure pump and the FCU high pressure pump). Electric fuel pumps provide pressure only during the engine start sequence and as a standby function when the normal system cannot maintain adequate pressure. Fuel symmetry is maintained automatically by a Fuel Balancing Device.

Refueling is accomplished using over-wing filler caps. Fuel quantity and fuel flow rate are displayed on the Engine Instrument System (EIS). Electric pump operation, low fuel pressure. and low fuel quantity conditions will be indicated on the Central Advisory and Warning System (CAWS) annunciator panel. In an emergency, fuel flow to the engine can be stopped by pulling the FUEL EMERG SHUT OFF handle, located at the aft end of the center console, left of the airplane centerline.

DESCRlPT ION

The fuel storage system includes integral wing tanks, fuel drains, refueling ports, and vents. The main fuel tank is between ribs 6 and 16, forward of the rear and main spars. A collector tank is forward of the main spar between ribs 3 and 6. Fuel drains are located in the lower wing- skins and in the fuel service bay on the left side of the fuselage, left of the nose wheelwell. These fuel drains allow the removal of water and other contaminants during preflight.

Refueling is accomplished through an oveming filler cap located at the outer, upper section of each wing. Each wing has a usable fuel capacity of 200 US gal (758 liters) MSN 101-140. 201 US gal (761 liters) MSN 141 AND UP.

The fuel vent bay allows venting of the fuel system through inward and outward vents located on the lower surface of the outer fuel bay.

Service Bulletin 28-001 installs a check valve in the motive flow line at each collector tank. The check valves stop fuel flow between the left and right wing tanks. The check valve in the motive flow at the engine firewall on MSN 141 AND UP is not installed.

The distribution system transfers fuel from left and right wing tanks and delivers fuel from the collector tanks to the engine fuel control unit. W i i n the wing tank are electric boost pumps, transfer ejector pumps, and delivery ejector pumps. From the wing tank the fuel flows through a fuel filter, maintenance and firewall shutoff valves, an air separator, a low pressure engine driven pump, an oilfiuel heat exchanger, and a high pressure engine driven pump to the fuel control unit.


Report No: 01973-001 7-57


OPERATION

During normal operation with the engine running, fuel is transferred from the wings to the engine by a motive flow system. Fuel under pressure from the low pressure engine driven pump is returned to the wings to provide motive flow through the transfer ejector pump and the delivery ejector pump. The transfer ejector pump transfers fuel from the wing tank to the collector tank. The left and right wing delivery ejector pumps transfer fuel to a common manifold. Fuel then flows through the maintenance shutoff valve and the fuel filter. The fuel filter incorporates a bypass valve in case the filter becomes blocked, and a spring loaded drain valve. Fuel is then directed into the air separator. The air separator passes air in the fuel system to the vent return line and incorporates the fuel low pressure switch. The fuel then passes through Ihe firewall shutoff valve to the low pressure engine driven fuel pump. The firewall shutoff valve is mechanically connected to the FUEL EMERG SHUT-OFF handle in the cockpit. The low pressure engine driven fuel pump includes a pressure relief valve that maintains a fuel pump outlet pressure of 43.5 psi (3 bar). A bypass valve allows for fuel flow around the engine driven fuel pump in the event of a fuel pump failure.

An electric boost pump, located within each collector tank, provides fuel pressure during engine start and is used to maintain system pressure when required. Each boost pump is controlled by a two position (ON or AUTO) switch located on the pilot's lower left switch panel. When set to ON, the boost pump will operate continuously. With the switch set to AUTO (the normal operating setting), the boost pump will operate automatically whenever fuel system pressure falls below 2 psi (0.14 bar). The boost pump will shutoff automatically 10 seconds after Ule fuel system pressure reaches 3.5 psi (0.24 bar). A boost pump is capable of supplying the engine in case the low pressure pump fails.

Fuel supply greater than engine demand is returned from the fuel control unit to the vent bays.

Refer to Engine Fuel System, this section, for engine fuel supply.

Fuel symmetry is automatically maintained by a Fuel Balancing Device when the Fuel Pump switches are set to AUTO. Lefl and right fuel quantities are monitored to detect fuel asymmetry

I exceeding 5% of each wing total fuel capacity (approximately 10.5 US gallons, 2 LCD segments) and will activate the fuel boost pump in the tank with the higher quantity. Fuel booster pump activation is delayed one minute to avoid pump cycling during flight in turbulence. The fuel boost pump will continue to operate until the left and right fuel levels are sensed to be equal. Automatic activation of the fuel boost pumps will only occur when the condition lever is

I out of the CUT-OFF position. To cater for refueling errors, up to 40 gallons (150 liters), up to 6 LCD segments will be automatically handled by the automatic fuel balance system. In the event of a system failure, the fuel load symmetry can be maintained by manually selecting the Fuel Pump switch to ON for the fuel tank with the higher quantity until a balanced fuel condition is restored and then turning OFF the fuel boost pump. During normal operation the pilot should monitor the fuel quantity gauges to verify that the Fuel Balancing Device is operating properly. Normal system operation is indicated by the left and right fuel quantity gauges remaining within 2 LCD segments of each other. (When a difference of 3 LCD segments is observed, the fuel boost pump for the tank with the higher quantity should be tumed ON until the quantities are even. Monitor the fuel quantity gauges for fuel symmetry for the remainder of the flight.)

Report No: 01973-001 7-58



KEY - FUEL TRANSFER

NOTE: RIGHT WlNG SHOWN. LEFl WlNG SlMlUR

Figure 7- 12. Fuel System (Sheet 1 of 5)

Issued: June 10, 1994 ! Revision 5: May 10, 1996

Report No: 01973-001 7-qQ


CAWS P O - - -. - - - - - -

BAlTERY BUS

, - - - - - - - - - - - - - - - - . I I

I I I

I 1 I

I I

I L E ~ LOW FUEL L E ~ FUEL RIGHT FUEL RIGHT LOW FUEL : I I I LEVEL SWITCHES TANK TANK LEVEL SWITC~ES

QUANTITY I I OUANTITY I

PROBES I I PROBES I

I I I

1 I I I I I

I I I

I

I I - I

GEN 1 BUS

I = - - - - - - - - - - - -

INTERMEDIATE FUEL FLOW DEVICE(S) TRANSMITTER

BATTERY

GEN 1

-

ACO 1 ACO 2

- - FUEL SYSTEM INDICATION

I , -

Figure 7-12. Fuel System (MSN 101 to 11 I) (Sheet 2 of 5)

- - - - - - - - - - - - - - - - A -

Report No: 01973-001 7-GO

L--

- - . I


€PQANSW SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

i * I I - I I !---------------------------------__I

FUEL SYSTEM INDICATION

Figure 7-12. Fuel System (MSN 112 to 999) (Sheet 3 of 5)


Reporl No: 01973-001 7-61


RKiHT FUEL PUMP REUY

NOTE: FUEL SYSTEM OPERATION

Figure 7-12. Fuel System (Sheet 4 of 5)

Report No: 01973-001 7-62


9 PILATUS SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

FUEL QUANTITY

FUEL QUANTITY INDICATORS

ENGINE INSTRUMENT ENGINE INSTRUMENT DISPLAY UNIT DISPLAY UNIT (MSN !01.111) IMSN 112-999)

\,

/ FUEL RESET =I rz SWITCH

/ FUEL F I R ~ W A L L FUEL PUMP SWITCHES SHUTOFF VALVE

\

PILOT'S LOWER LEFT PANEL

F i g u r e 7 - 12. Fuel S y s t e m

(Sheet 5 of 5)

Issued: June 1 0 . 1994 Revision 3: September 29. 1995

CAWS

R e p o r t No: 0 1 9 7 3 - 0 0 1

7 - 6 3

SECTION 7 =PILATUS= AIRPLANE AND SYSTEMS DESCRIPTION -PC XC

MSN 101 . 105

Each wlng tank contalns flve capacltance type luel quant~ty probes that are connected to thelr respecl~ve lnlerrnedrate dev~ces

MSN 106- 111

Each wlng tank contalns tour capacltance type fuel quanl~ty probes that are connected to an lnterrned~ate dev~ce

The lnlerrnedlate devlce(s) supply s~gnats lo lhe EIS for an analog ndcat~on of left and r~ght fuel contents and also the control Input to the fuel balancing devlce.

MSN 112-999

Each wing tank conlalns lour capacitance type luel quanllly probes that are connected to the EIS

Power lor the fuel analog lndlcatlons IS taken lrom the BATTERY BUS Fuel level sw~tches n the collector and matn lanks powered lrorn the GEN 1 BUS and BATTERY BUS, cause the CAWS captlons to lllumlnale at low luel levels A fuel llow sensor located forward 01 the FCU sends a s~gnal to the EIS to Indicate luel flow In Ibs per hour The EIS calculates and dlsplays luel quantlty. luel rernalnlng and endurance (MSN 112 and subsequent) The fuel flow lndlcal~on requlres power lrorn Englne Acqu~s~t~on Un~t A Fuel Reset sw~tch IS used to re- dalurn the lotat luel quant~ty and fuel used value o l the totalizer lunctron alter each tlme luel IS

added to the wlng tanks These values are stored In non-volat~le memory when power IS

removed To reset the total~zer, rnomenlarlly press the Fuel Reset sw~tch alter englne starl Ver~ly that lhe luel quant~ty lndlcat~on Increases to the new luel quantlty and the fuel used lndlcat~on 1s reset to zero Refer to Englne lnslrurnent System, th~s sectlon , lor addltbonal ~nforrnal~on

The CAWS captlons thal lndlcate fuel system status are.

L FUEL PUMP, R FUEL PUMP

lndlcates luel boost pump operallon

L FUEL LOW. tnd~cales luel level ~n tank less lhan 160 Ibs (20 US R FUEL LOW gal, 75 I~ters).

FUEL PRESS lnd~cates tuel system pressure less than 2 psi (0.14 bar)

Report No: 01973-001 7-64

Issued: June 10, 1994 Rev~s~on 7: July 1. 1997

r PILATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

ELECTRICAL

GENERAL

The electrical 28 VDC system, consists of the following power sources:

r GEN 1 system: a main generator 28 V, 300 A or a 28 V 400 A option GEN 2 system: a secondary generator 28 V, 115 A Storage battery system: one or two nickel cadmium batteries 24 V 40 Ah or one or two

I lead-acid batteries 24 V 42Ah Optional Emergency Power System (EPS): Refer to Supplement No. 22

An external power socket permits DC power to be provided from a ground power unit. In the event of system failures, automatic switching and load shedding takes place to reduce pilot work load.

Two static inverters provide 26 VAC 400 Hz synchro references for certain avionics equipment.

DESCRIPTION

POWER SUPPLIES

When the engine is running, Generator 1 is the main source of power. It is capable of supplying the full aircraft electrical load. Generator 1 is also the engine starter motor. If the engine STARTER switch is selected ON; GEN 1 is automatically switched OFF.

Generator 2 is a secondary power supply which, in normal operation, supplies only the GEN 2 BUS (and charges the second battery - if installed). If GEN 1 fails, the GEN 2 ensures essential services are supplied, but it is unable to supply the full electrical load. With bat off, GEN 1 OH and caws cautions illuminated, do not operate the GEN 2 reset as this action may result in the loss of all electrical power (Pre SB 24-010 and MSN 101-230).

The battery (two batteries - if installed) provides power for starting the engine and, in case of an engine or double generator failure, it will supply essential electrical systems for 20 minutes if the load is reduced below 60 amps or 30 minutes if the load is below 50 amps. (When two batteries are installed they will supply essential electrical systems for 40 minutes if the load is reduced below 60 amps or 60 minutes if the load is below 50 amps).

On ground the DC system can be powered by an external power unit which is connected under the rear fuselage left side. When external power is connected to the battery, GEN 1 and GEN 2 relays remain open, preventing a direct connection between the external and airplane power sources. To apply external power to the aircraft electrical system, the EXT PWR switch on the overhead panel must be selected to ON. Post SB 24-008 and MSN 231 & UP. An external power control unit is installed which will disconnect the external power unit if the output voltage goes above 29.5 VDC or below 23 VDC.


Report No: 01 973-001 7-65

SECTION 7 aPlLATUSEr AIRPLANE AND SYSTEMS DESCRIPTION PC12

1 CAUTION 1 THE GPU MUST BE CAPABLE OF PROVIDING A MINIMUM OF 500 AMPS FOR ALL SYSTEMS TO BE OPERATED CONTINUOUSLY AND 1,000 AMPS IF ENGINE STARTER OPERATION IS REQUIRED.

BUS BARS

The DC power sources distribute the power via an interconnected system of BUS BARS.

The BATTERY BUS, the GEN 1 BUS and the GEN 2 BUS are connected to their respective power sources through the BAT relay, the GEN 1 relay and the GEN 2 relay, each relay being controlled by a switch on the cockpit overhead panel.

The AVIONIC 1 BUS is connected to the main battery line and the AVIONIC 2 BUS is connected to the main GEN 1 line through relays controlled by switches on the overhead panel.

The NON ESSENTIAL BUS is connected to the main battery line through an automatic load shedding device.

BUS TIE INTERLOCK

To permit both generators to charge the battery (two batteries - if installed) and supply all electrical selvices, GEN 1 is connected to the BAT BUS via the BUS TIE and GEN 2 to the BAT BUS via the GEN 2 TIE. The two BUS TIES have an automatic interlock.

When the engine is running and both generators are on line, the BUS TIE is closed and the GEN 2 TIE is open. GEN 1 then supplies the GEN 1 and BAT busses and charges the battery. GEN 2 supplies only GEN 2 BUS and charges the second battery (if installed).

When both generators are off line (engine off or double generator failure), both the BUS TIE and GEN 2 TIE will close, allowing the battery (two batteries - if installed) or external power unit to supply all BUS bars.

If GEN 1 stops supplying power to the main battery line (GEN 1 fail or BUS TIE open), the GEN 2 TIE will close connecting GEN 2 (and the second battery - if installed) to the main battery line.

NON ESSENTIAL BUS

To protect GEN 2 from overload, the NON ESSENTIAL BUS is provided with an automatic load shedding device which operates when its selector switch is in the AUTO position.

Report No: 01 973-001 1 7-66



When at AUTO, if GEN 1 or an external power unit is supplying the electrical system, the NON ESSENTIAL BUS is connected; in ail other cases the NON ESSENTIAL BUS is disconnected.

If it is required to operate services on the NON ESSENTIAL BUS from either the battery or GEN 2 with no GEN 1 on line, set the bus selector switch to OVRD ON (override on) to reconnect the bus to the main battery line.

I CAUTION I MAXIMUM PERMl?TED CONTINUOUS LOAD FROM GENERATOR 2 IS 115 AMPS. ABOVE 115 AMPS, GENERATOR 2 VOLTAGE WlLL PROGRESSIVELY DECREASE AND THE BATTERY WlLL DISCHARGE.

CIRCUIT BREAKERS

The COOLING, ELECTRICAL HEATING. HYDRAULIC and FLAP systems are all high current consuming systems which are connected directly to the main battery line. The electrical power circuit breakers for these systems are not accessible to the pilot (under cabin floor). The control circuits for the COOLING, ELECTRIC HEATING and HYDRAULIC (landing gear) systems are on the NON ESSENTIAL BUS.

Circuits supplied from the BUS BARS have circuit breakers on color coded panels on the left and right cockpit'walls (see Figure 7-13). The BAT and AVIONIC 1 BUS are green, the GEN 1 and AVIONIC 2 BUS are blue, the GEN 2 BUS is yellow, and the NON ESSENTIAL and AC busses are white. These colors correspond with those on the overhead panel.

Both the BUS TIE and the GEN 2 TIE are overcurrent protection devices which open automatically if the current through them is excessive to protect power supplies and circuits. The BUS TIE open automatically when the continuous current exceeds 220 amps and the GEN 2 TIE opens with a continuous 145 amps. Both TIES can be opened manually and reset, if required, by pulling or pushing the control circuit CBs on the overhead panel.

AC POWER

AC power (26 V and 115 V at 400 Hz) is provided by one of two static inverters. One inverter is powered from the BAT BUS, the other from the GEN 1 BUS. The selector switch is on the overhead panel. AC power is used only for synchro references for avionics systems.

Issued: June 10.1994 Revision 1 1 : March 1, 2003

Report No: 01973-001 7-67

SECTION 7 EPILATUS3 AIRPLANE AND SYSTEMS DESCRIPTION PC12

CONTROLS AND INDICATORS

Overhead Panel

The electrical system is controlled from the cockpit Overhead Control Panel. Switches are provided for the BAT (BAT 1 and BAT 2 - if second battery installed), EXT PWR, GEN 1 and GEN 2; the AVIONIC 1, AVIONIC 2 and N ESNTL busses and an INV selector switch and the BUS TIE and GEN 2 TIE control relay circuit breakers. (With a second battery installed the panel has AMP, VDC and BAT 1, BAT 2 OVERTEMP indicators).

A STBY PWR switch (if installed) when selected provides pawer to certain electrical systems, without the battery or external power circuits being on. Usage is controlled by a time delay relay and is limited to 10 minutes.

Red indicators illuminate to show when a BUS BAR is not powered. An amber indicator illuminates when GEN 2 TIE closes, indicating an abnormal condition. An green indicator illuminates to show that external power is available to the aircraft and an amber indicator illuminates when EXT PWR supplies the aircraft. A GANG BAR enables, with one action, all electrical power to be switched OFF.

I MSN 261 to 400. A GEN 1 - AMPNDC indicator is installed in the cockpit headliner to the left of the overhead panel. A current sensor installed near the generator 1 contactor will detect any current drawn from the generator 1 and send the amps reading to the indicator. The voltage signal is taken directly from the generator 1 output. Power for the indicator is taken from the battery bus through the GEN 1 VIA IND circuit breaker on the BATTERY BUS circuit breaker panel. The backlighting for the indicator is connected to the Power Management Panel backlighting system.

Engine Instrument System (EIS) (MSN 101-11 1)

To monitor electrical power output there is a voltJammeter in the EIS. Two push button switches labelled GEN 1 and GEN 2 at the top of the EIS enable the VIA meter to be selected to display battery (or external power) GEN 1 or GEN 2 output. Depressing the GEN 1 button will cause DC IND also to be shown in the button and will select the VIA meter to GEN 1; similarly for GEN 2. Depressing the button a second time will deselect the generator. When neither button indicates DC IND, the VIA meter shows battery voltage and charge or discharge current; or external power unit supply voltage when an external supply is connected and the EXT PWR switch is ON.

A positive BAT current indicates battery charging rate. A positive GEN current indicates generator output. The GEN 1 and 2 voltmeters indicate output volts before the generator relay and the ammeters show supply current after the generator relay.

Report No: 01973-001 7-68


SECTION 7 WPILATUS W

AIRPLANE AND SYSTEMS DESCRIPTION ?c XII-

PANEL LH FRONT

7

PANEL LH REAR Figure 7-1 3. Electrical Power System (Typical Left Circuit Breaker Panels MSN 101 -1 20)

(Sheet 2 of 10)

Report No: 01 973-001 7-70

Issued: June 10,1994 Revision 8: September I , 1998

c 3PILATUSe SECTION 7 Y -PC XII- AIRPLANE AND SYSTEMS DESCRIPTION I

PANEL RH FRONT

PANEL RH REAR 5 Figure 7-13. Electrical Power System (Typical Right Circuit Breaker Panels MSN 101-120)

t (Sheet 3 of 10) I Issued: June 10,1994 Revision 8: September I , 1998

Reporl No: 01 973-001 7-71

SECTION 7 WPlLATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

~;SN 261 AND UP

PANEL LH FRONT

OPTIONAL EWIPMENT

PANEL LH REAR 5

Figure 7-13. Electrical Power System (Typical Left Circuit Breaker Panels MSN 121 AND UP) (Sheet 4 of 10)

Report No: 01 973-001 7-72


XPILATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

- OPTIONAL

EOUlPMENT

MSN 281 AND ~6 \

MSN 121-280 PANEL RH REAR

Figure 7-13. Electrical Power System (Typical Right Circuit Breaker Panels MSN 121 AND UP) (Sheet 5 of 10)


Report No: 01 973-001 7-73 1


28V DC GENERATOR 1 BUS

7

NOTE: FROM MSN 161 THE TWO INVERTERS ARE COMBINED INTO A DUAL INVERTER.

Figure 7-13. Electrical Power System (Sheet 6 of 10)

Report No: 01973-001 7-74



Figure 7-13. Electrical Power System (Typical Overhead Panel) (Sheet 7 of 10)


Report No: 01973-001 7-75

SECTION 7 WPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION ?@ %I1

DC POMR WlCATlON

SWITCH POSITION ANNUNCIATON VOLTSIMAPS HDICATlON

GEN 1 OFF BAlTERY

GEN 2 OFF

G E N E T

FUEL 8 klf$F3&8

ENGINE INSTRUMENT DISPLAY UNIT (MSN 101-111)

Figure 7-13. Electrical Power System (Sheet 8 of 10)

Repon No: 01973-001 7-76


SECTION 7 FPILATUSZ AIRPLANE AND SYSTEMS DESCRIPTION PC12 1

4

ELECTRICAL POWER MANAGEMENT fJ=jrrg A n m c I N E Y m AYlCIUCl HY

OPTION INTERIOR

Figure 7-13. Electrical Power System - (Overhead Panel With Two Batteries Installed) (Sheet 10 of 10)



Engine Instrument System (EIS) (MSN 112 and UP)

To monitor electrical power output there are two voltlammeter displays on the EIS; BAT and GEN 2. A positive BAT cunent indicates battery charging rate. A positive GEN 2 current indicates generator 2 output. Continuous monitoring of the voltages and currents for close to limit cautions and out of limit warnings is provided by the EIS. In flight, an increase in battery current of 10 amps or more in an interval of less than 7 seconds will give an EIS warning indication.

Wth a second battery installed a BAT 1 and a BAT 2 OVERTEMP red indicator will illuminate to show H the internal temperature of a battery becomes excessive. These indicators are linked to the CAWS BAT HOT amber caption. If this caption comes on, the overhead panel battery over-temperature indicators should be checked to determine which battery is at faul before taking the appropriate action.

OPERATION

Before applying electrical power to the aircraft, ensure that all electrical services are OFF, the NON ESSENTIAL bus selector is at AUTO and the EXTERNAL PWR switch is OFF. In this condition, applying external power to the socket in the rear fuselage, will cause the green indicator to illuminate showing that power is applied and all red bus indicators to illuminate showing that no bus ban are powered.

Turning the battery (batteries - H second battery installed) ON will then cause the BAT, GEN 1 and GEN 2 bus failure indicators to extinguish. Battery vols can be checked on the EIS VIA meter. (With a second battery installed, the battery volts can be checked on the overhead panel). If external power is required, selecting the EXT PWR switch ON, will apply external power to the system and the NON ESSENTIAL bus failure indicator will extinguish. If no external power Is used and services fmm the NON ESSENTIAL BUS are required, select OVRD ON.

~XiiiGiq WHEN OVRD ON IS SELECTED MONITOR BATTERY AMPS AND VOLTS AS IT IS POSSIBLE TO RAPIDLY DISCHARGE THE BATTERY COMPLETELY WITH HIGH CURRENT CONSUMING SERVICES.

Alter engine start, selecting GEN 1 ON will cause the GEN 1 OFF CAWS caption to extinguish, the GEN 1 relay will close allowing GEN 1 to supply the GEN 1, BATTERY, and GEN 2 bus bars. GEN 1 charges the battery (charges the two batteries - if second battery installed). Selecting GEN 2 will cause the GEN 2 TIE to open. GEN 2 then supplies the GEN 2 BUS (and charges the second battery - H installed). GEN 1 supplies all the other busses. Disconnecting the external power from the airplane will cause the overhead panel green indicator to extinguish.


Report No: 01973-001 7-77 I

SECTION 7 *PLATUSW AIRPLANE AND SYSTEMS DESCRIPTION Kl2

AFTER ENGINE START, SELECT GEN 1 ON BEFORE GEN 2. IF GEN 2 IS SELECTED FIRST, THE HIGH BATTERY CHARGE CURRENT MAY CAUSE THE GEN 2 TIE TO OPEN DUE TO CURRENT OVERLOAD.

The voltages and bad or charging current of the battery, GEN 1 and GEN 2 may be obse~ed by selectively operating the VIA meter push buttons on the EIS (MSN 101-1 11).

The output voltages and load or charging current of the main battery and GEN 2 may be observed from the VIA meters on the EIS. (With a second battery installed, the voltage and load or charging current of both batteries may also be obse~ed on the BAT 1, BAT 2 VIA meters on the overhead panel. The voltage and load or charging current of the second battery may be observed from the BAT 2 VIA meter on the overhead panel).

Selecting AVIONIC 1 and 2 switches to ON applies power to the AVIONICS busses and extinguishes their red failure indicators. It is not significant which INVERTER is selected, as long as the INVERTER CAWS caption is not illuminated.

When the electrical system is functioning correctly, all indicator lights on the overhead panel are extinguished.

I MSN 261 to 400. Before engine start with power supply on, the GEN 1- AMPNDC indicator will read approximately zero. After engine start with the generator 1 on-line, the GEN 1 indicator will show the voltage and current drawn from the generator 1. In the event of a generator 1 failure the indicator display will reduce to zero. If the current sensor or the internal power supply of the indicator fails the display will show an E in all the segments. If the power supply to the indicator is not available or too low the display will remain blank.

Report No: 01973-001 7-78

' Issued: June 10,1994 Revision 10: September 1, 2000

EPILATUSW SECT ION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

MALFUNCTIONS, CAUTIONS AND WARNINGS

The BUS TIE interlock system and the NON ESSENTIAL BUS automatic load shedding ' system minimize the actions necessary in the event of system failures.

If GEN 2 fails, the GEN 2 TIE closes automatically and GEN 1 supplies all bus bars. No load shedding is necessary. The amber CAWS GEN 2 caption will illuminate and the overhead panel ambdr indicator will show GEN 2 TIE has closed.

If GEN 1 fails, the GEN 2 TIE closes automatically and the NON ESSENTIAL BUS is automatically switched OFF. The pilot must ensure that the remaining electrical load on GEN 2 does not exceed 115 amps and that the battery is not discharging. The amber CAWS GEN 1 and NON ESSENTIAL BUS captions and the overhead panel red NON ESSENTIAL BUS amber GEN 2 TIE indicators will illuminate.

I Refer to the Emergency Procedures (Section 3) for further information on emergency procedures.

Electrical system status is displayed on the CAWS and overhead control panel. No captions or lights Indicate that the system is functioning correctly (except in the event of a total electrical failure).

The CAWS displays the following WARNINGS and CAUTIONS:

RED WARNING

ESNTL BUS - Indicates voltage on BAT, GEN 1 or GEN 2 Busses is below 18 V.

AV BUS - Avionics bus voltage below 18 V.

AMBER CAUTION

N ESNTL BUS - NON ESSENTIAL BUS voltage below 18 V.

BAT OFF - Battery is OFF LINE.

GEN 1 OFF - GEN 1 is OFF LINE.

GEN 2 OFF - GEN 2 is OFF LINE.

BUS TIE - BUS TIE is open.

INVERTER - Inverter output below 20 v.

BAT HOT - Battery over temperature or thermal runaway (inoperative on aircraft with an optional lead acid batteries installed)


Report No: 01 973-001 7-79

SECTION 7 ~ P I L A T U S W AIRPLANE AND SYSTEMS DESCRIPTION PC12

LIGHTING

INTERIOR

Cockpit lighting consists of internally lit cockpit instruments, glareshield mounted floodlights, control wheel reading lights, and a dome light. Light intensity is controlled by switches and reostats located at the aft end of the center console. Separate intensity control of the pilot, copilot, and center console panels is provided. The overhead and side panel intensity is controlled with the center console switches. The overhead dome light can be set to two preset intensities ot 50% or 100% brightness. The advisory lights are on a fixed dim circuit. The map light on each control wheel is controlled by separate rheostat.

NOTE

Cockpittcabin curtain must be installed and closed during night flight to prevent glare and reflection in the cockpit area.

A switch located on the forward edge of the cabin door (accessible when open) will activate a timer for the cockpit overhead dome light. When pressed the dome light will on for 45 seconds to facilitate night preflight boarding.

A vestibule light illuminates the cabin airstairs and the baggage area has an overhead light. The main cabin is equipped with an overhead flood light system that can be set to 50% or 100% brightness as selected by the cockpit switch. Individual reading lights are provided for each passenger seat and are controlled by a switch in the cockpit and by a switch near each seat.

I Optional SB 33-007 installs an additional light in the baggage compartment. It is operated by a push switch installed on the bulkhead trim adjacent to the cargo door. The light stays on for five minutes when the switch is pushed. For continued lighting the switch must be pushed again.

EXTERIOR

POSITION AND STROBE

Exterior lighting consists of a position and strobe light on each wing tip, a white position light on the tail, a landing light on each main landing gear, a taxi light on the nose landing gear, and a wing inspection light mounted in the left fuselage forward of the cabin door. These lights are controlled by switches located on the pilot's lower panel to the left of the Center console.

I BEACON LIGHTS (Post SB 33-002)

Optional red flashing beacon lights can be installed on the top of the horizontal stabilizer fairing and on the lower center fuselage. They give recognition during ground operation and additional anti-collision protection in flight. The lights are controlled by a three position switch OFF, BEACON, STROBUBEACON located in the External Lighting switch panel to the lower left of

Report No: 01973-001 7-80



the center console. Power for the lights is supplied from the battery bus through the circuit breaker RED BEACON.

RECOGNITION LIGHTS (Post SB 33-001)

Optional recognition lights and power supply units can be installed in the left and right forward outer flap fairings. They provide forward illumination during taxiing and enhance the conspicuity of the aircraft in the traffic pattern or enroute. The lights are controlled by an ONIOFF switch located in the External Lighting switch panel to the lower left of the center console. Power for the lights is supplied from the generator 1 bus through the circuit breaker RECOG LIGHT.

LOGO LIGHTS (Post SB 33-003) I Optional logo lights can be installed under each side of the horizontal stabilizer. They provide illumination of the vertical stabilizer to show the owner's logo. The lights are controlled by a three position switch OFF, LOGO, TAXI/LOGO located in the External Lighting switch panel to the lower left of the center console. Power for the lights is supplied from the battery bus through the LOGO LIGHTS circuit breaker, which is installed in the rear fuselage. Each logo light has two filaments, if the battery is the only power source one filament in each light will illuminate. When the Generator 1 is on-line all four filaments will illuminate. If the Generator 1 fails in flight, two filaments are automatically switched off. The remaining two filaments should be switched off manually. Avoid prolonged use of the logo lights when the aircraft is on the ground without external power.

DUAL FILAMENT NAVIGATION LIGHTS (Post SB 33-009) I Optional dual filament navigation lights can be installed in place of the standard lights. The lights are controlled by an electronic switch unit installed adjacent to each navigation light assembly. When the EXTRENAL LIGHTS NAV switch is set to ON the electronic switch units start a 4.5 minute test sequence. During this time both filaments in each navigation light are on and show as a bright light. This gives sufficient time to check that all the filaments are serviceable. If a light is seen to be flashing, it is an indication that one filament is unserviceable. After the test sequence is completed the secondary filament will go off. It will only come on again if the primary filament fails. Each electronic switch unit has a thermal protection fuse which, if a failure condition occurs and the two lights come on, will open after slx minutes and make the second filament go off.


Report No: 01 973-001 7-81


ENVIRONMENTAL CONTROL SYSTEM

GENERAL

Refer to Figure 7-14, Environmental Control System, for system layout.

The Environmental Control System (ECS) is designed to take engine bleed air, reduce its temperature to that desired, and deliver it to the cabin air distribution system for pressurization and ventilation. The ECS consists of an air cycle system, a distribution system, and a control system. The air cycle system cools a portion of the bleed air and then mixes it with hot bleed air to provide the correct temperature. A firewall shutoff valve can be closed to prevent contaminated air from entering the cabin in the event of an engine compartment fire.

DESCRIPTION

The air cycle system consists of a flow control venturi, a heat exchanger, a cooling turbine, a temperature control valve, a water separator, P3 shutoff valve, a system main shutoff valve, and associated non return valves and control sensors.

The flow control venturi is sized to regulate flow and pressure.

The heat exchanger is an aluminum single pass, crossflow, plate and fin unit. The unit includes one charge air tap to assist the injection of water into the heat exchanger coolant intake. The evaporation of the water on contact with the heat exchanger surface increases the efficiency of the unit.

The cooling turbine is a ball bearing turbo fan and consists of a radial turbine in a stainless steel assembly coupled to an axial flow fan. The turbine casing incorporates a containment ring.

I The Temperature Control Valve is three ported consisting of one inlet and two outlets and driven by a 28 VDC actuator. The valve body and rotating drum are aluminum. The actuator has gearing, limit switches, and magnetic brake to control the reversible series wound motor.

The water separator consists of an aluminum shell containing a coalescor and its support. The coalescor collects moisture from the passing air and forms large droplets which then enter a swirl section, where they are removed by centrifugal force. The separator has a spring loaded poppet valve which allows air to bypass the unit in the event of the coalescor becoming blocked.

The P3 shutoff valve is solenoid operated and allows automatic selection between P3 and P2.5 compressor stages depending on flight condition to maintain the pressure schedule required for cabin pressurization.

Report No: 01 973-001 7-82


W PILATUSE SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

-

ENQINE

KEY - HIGHPRESSURE

LOWPRESSURE

SWITCH 290%

FIREWALL I I I I

I DUCT TEMP 4- -, SENSOR

ECS

I - I

I I I

- TEMPERATURE SELECTOR

AND CONTROLLER

Figure 7-14. Environmental Control System (Pre SB 21-003) (Sheet 1 of 3)


Report No: 01 973-001 7-83

SECTION 7 'BFIPLATUS w AIRPLANE AND SYSTEMS DESCRIPTION PC XII- .

CONDITIONED AIR SUPPLIED FROM ECS

Figure 7-14. Environmental Control System (Pre SB 21-002) (Sheet 2 of 3)

Report No: 01973-001 7-84


b 1PLNUSliT SECTION 7 P K12 AIRPLANE AND SYSTEMS DESCRlPnON

Figure 7-14. Environmental Control System (Post SB 21-003, MSN 181 and UP) , (Sheet 1 of 3)


Report No: 01973-001 7-84A

SECTION 7 W'lLATUSW AIRPLANE AND SYSTEMS DESCRIPTION Kl2 4

4

CONDITK)NED AIR SUPPLIED FROM ECS

Figure 7-14. Environmental Control System (Post SB 21-002, MSN 181 - 320) (Sheet 2 of 3)

, Report No: 01973-001 7-848


=PILATUSm SECTION 7 -PC XII- AIRPLANE AND SYSTEMS DESCRIPTION

CoPILors LOWER LEFT PANEL

CABIN TEMPERATURE INDICATOR ('C)

Figure 7-14. Environmental Control System (Sheet 3 of 3)

Issued: February 14. 1994 Revision 3: September 29. 1995

ECS FIREWALL SHUTOFF VALVE

Report No: 01 973-001 7-05 1

SECTION 7 'EPILATUSS AIRPLANE AND SYSTEMS DESCRIPTION ?c XII-

A Firewall Shutoff Valve enables isolation of the system in emergency conditions such as an engine lire. Operation ol the Firewall Shutoff Valve also opens a ram air scoop on the right luselage underside which introduces ambient ventilation air through the distribution system. This is used in the event of smoke in the cockpit or cabin.

DUE TO THE COMPOSITE CONSTRUCTION OF THE ENGINE COWLING AND THE POSSIBILITY OF TOXIC GASES, THE AIRPLANE ECS MUST BE SHUTOFF WHEN A FIRE CONDITION IS SUSPECTED.

OPERATION

Air is drawn from the P2.5 and P3 compressor bleed ports on the engine casing. This consists ol a single port in the case of the P2.5 connection and two diametrically opposed ports lor the P3 connections. The bleed air will be taken exclusively from the P2.5 port during normal operat~on. However, when the engine is at idle there is insuflicient pressure to maintain cabin pressurization. When the P2.5 bleed air pressure falls below a specific value, a pressure sensor in the bleed air ducting opens the P3 shutoff valve. This creates a back pressure on the non-return valve at the P2.5 port and closes the valve to shut OH the P2.5 bleed. The bleed air lhen passes through the Primary Shuton Valve and the Flow control venturi, which is 1 sized to regulate the bleed air flow rate and pressure.

1 Pre SB 21-003 - The air then passes on to the heat exchanger.

Post SB 21-003, MSN 181 and UP - The air then passes on to the Temperature Control Valve (TCV). At the TCV the bleed air splits where variable amounts are either supplied to the Heat Exchanger or to a mix point downstream ol the Cooling Turbine .

The heat exchanger is cooled by ambient air drawn from a NACA intake in the airplane skin. Cooling airflow is provided by the Heat Exchanger Coolant Fan located downstream of the I heat exchanger.

( Pre Sf3 21-003 - From the heat exchanger, a variable amount of bleed air is passed to the Temperature Control Valve and the remainder ot the bleed air is passed to the Cooling

I Turbine. Post SB 21-003, MSN 181 and UP - From the heat exchanger, the bleed air is passed to the Cooling Turbine.

As the bleed air passes through the Cooling Turbine, its pressure is reduced to delivery pressure and its temperature is, in many cases, close to 0°C. The energy extracted lrom the bleed air is used to power the Heal Exchanger Coolant Fan which is mechanically linked to the turbine by a shaft.

Report No: 01 973-001 7-86


%PILATUS E SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Pre SB 21-003 - The duct downstream of the turbine is the mixing duct where the now-cooled I turbine exhaust air is mixed with air taken from upstream and downstream of the heat exchanger. The mixing proportions are controlled by the Temperature Control Valve (TCV) by regulating the amount of bypass air to be added. The TCV is an electrically operated three port valve which opens first to let through air from the heat exchanger outlet and then, as necessary, opens further to allow higher temperature air from downstream of the Flow control venturi to be added to achieve the desired temperature. The TCV operation is controlled by the Temperature Control System.

Post SB 21-003. MSN 181 and UP - The duct downstream of the turbine is the mixing duct where the now-cooled turbine exhaust air is mixed with un-cooled bleed air directed from the other port of the TCV. The mixing proportions are controlled by the TCV. The TCV is an electrically operated three port valve with one inlet and two outlet ports. Depending on the selected temperature the TCV modulates to either pass air through or bypass the Heat Exchanger and Cooling Turbine. The TCV operation is controlled by the Temperature Control System. d

The Temperature Control System has two modes of operation as selected on the AutoIManual switch. In the Auto mode, TCV operation is controlled by electrical signals from a cabin temperature sensor, located on the cabin headliner, to maintain the selected temperature between 18°C and 30°C. I Pre SB 21-003 - The TCV will open if the cabin temperature is less than desired and close if the temperature is greater than desired. I Post SB 21-003. MSN 181 and UP - The TCV will move to allow more bleed air to bypass the Cooling Turbine if the cabin temperature is less than desired. Conversely it will move to pass more air through the Heat Exchanger and Cooling Turbine if the temperature is greater than desired.

The temperature of the duct downstream is monitored by a temperature sensor and will limit the movement of the TCV as required to keep the duct temperature within the maximum and minimum temperature limits. In the Manual mode, a separate switch marked HOT and COLD directly controls the position of the TCV.

I CAUTION 1 WHEN OPERATING IN THE MANUAL MODE THE SYSTEM MAXIMUM AND MINIMUM TEMPERATURE LIMITS ARE DISABLED AND THERE IS A POSSIBILITY OF DAMAGING THE SYSTEM BY SETTING OUTLET TEMPERATURES OUTSIDE THESE LIMITS. THE MANUAL MODE SHOULD THEREFORE ONLY BE USED WHEN IT IS NOT POSSIBLE TO CONTROL THE ECS AIR OUTLET TEMPERATURE USING THE AUTOMATIC MODE.



From the mixing duct the conditioned air passes through a water separator. Moisture Is removed from the conditioned air and drawn to the heat exchanger and sprayed into the heat exchanger intake. The conditioned air passes through the Firewall Shutoff Valve and the non- return valves to the cabin for distribution. The non-return valves prevent sudden depressurization in the event of a loss of cabin air supply.

The air enters a small plenum where it is distributed to the cockpit and the cabin. Cockpit air is directed to fixed outlets (adjustable outlets - MSN 321 & UP) at the crews feet and to adjustable outlets adjacent to the instrument panel. Air to the cabin is introduced through fixed outlets placed at floor level along bath sides of the cabin.

The ECS will automatically shut down when the engine starter is activated.

\

Cabin air temperature is displayed on the center console fotward of the engine power controls. Overpressure and overtemperature switches are installed to monitor the system. If pressures greater than 33 psi (Pre SB 21-003) 40 psi (Post SB 21-003, MSN 181 and UP) are sensed in the bleed air line downstream of the Flow control venturi, temperatures greater than 290°C in the bleed line upstream of the Primary Shutoff Valve, temperatures greater than 105OC are sensed in the air line downstream of the water separator, or if the Firewall Shutoff Valve is closed, the ECS will automatically shutdown. The CAWS ECS annunciator will illuminate when

I the ECS is automatically shutdown.

HEATING SYSTEM (MSN 101-320)

GENERAL

Refer to Figure 7-14, Environmental Control System, for system layout.

The PC-12 is equipped with an auxiliary electrical heating system which is used to supplement the air cycle system during prolonged low temperature operations such as cruise at high altitude. It can also be used for preheating the cabin on ground.

DESCRIPTION

The system comprises of two 28 VDC heating units each equipped with a 75 mm mixed flow fan. Cockpit and cabin air is drawn through the units and recirculated back into the main ECS distribution system. Each unit is cylindrical in form and contains three heating elements producing a total of 1.5 kwlunit. The system therefore produces a total of 3 kW in addition to that of the air cycle system.

Report No: 01973-001 7-88



The unls are situated adjacent to each other under the cabin floor between frames 18 and 19.

Pre SB 21-002 - One heater is dedicated to the cockpit and the other to the cabin. The fan on the cockpit heater unit scavenges its air from the underfloor area in which it is situated and from two outlets in the left hand side of the cabin roof headliner. The air is passed across the heating elements where its temperature is raised. It then passes through ducting to the cockpii distribution system downstream of the underfloor plenum aft of frame 15. Here it is mixed by means of an ejector type nozzle with conditioned bleed air from the air cycle system. The fan on the cabin heater unit also scavenges its air from the underfloor area and from two outlets in the right hand side of the cabin roof headliner. Air passes across the heating elements where its temperature is raised. It then passes rearwards to the cabin distribution system downstream of the plenum. Here it is mixed with conditioned bleed air by means of an ejector type nozzle. The effect of the units.when operating is to increase the flow rate and temperature of air in the main distribution system and at all of its outlets in the cockpit and cabin area.

Post SB 21-002, MSN 181 and UP - One heater is dedicated to the LH side and the other the RH side of the cabin. The fan on the heater units scavenges its air from the underfloor area in which it is situated. The air is passed across the heating elements where its temperature is raised. It then passes through ducting to either the LH or RH distribution ducts integrated in the cabin sidewall bwer interior panels where it mixes with the condiiioned bleed air. Additional spillage ducts reduce the back pressure on the heater fans by ducting discharge air into the condiiioned bleed air distribution system upstream of the cabin sidewalls.

OPERATION

The system is controlled by a three position rocker switch situated on the pilot's lower left panel to the lett of the pilots control yoke. The HEATING SYSTEM switch positions and power supplies are as follows:

Pre SB 21-002 - OFF which electrically isolates both healers and fans, CABIN which allows selection of the cabin heater and fan only, and CKPT 8 which Selects both heaters and fans together. The cockpit heater and fan can not be operated by itself. The cockpit heater is supplied 28 VDC power from the GEN 1 bus. The cabin heater is supplied from the GEN 2 bus.

Post SB 21-002, MSN I81 and UP - OFF which electrically isolates both heaters and tans. LOW which allows selection of the LH side cabin heater and fan only, and HIGH which selects both heaters and fans together. The RH side cabin healer and tan can not be operated by itself. The cabin RH heater is supplied 28 VDC power from the GEN 1 bus. The cabin LH heater is supplied from the GEN 2 bus.

I

The power for the heater circuits is interrupted when the hydraulic pump is operating. The fans continue to run. This minimizes generator accessory loads during continuous normal operation and prevents generator overload after failure.


Report No: 01973-001 7-89


Each unit is equipped wWh an internal thermal protection syslem which isolates the heater when the element temperature exceeds 200 "C (Pre SB 21-002) 259 "C (Post SB 21.002, MSN 181 and UP). In such a case the fans continue to run until the temperature falls bebw the reset value when the heater is again energized.

The need lor supplemental heating is reduced during engine P3 operation. The heaters and tans are not active on the ground during ground idle mode, during a flight idle descent, or any other time when P3 bleed is being extracted.

When the Vapor Cycle Cooling System (VCCS) is being operated the heaters are isolated but the fans will continue to run. This prevents an excessive current drain on the electrical generation system.

The heater and fan units can also be operated independently of the bleed air system for pre- healing the aircraft before engine start when required. For this the aircraft must be connected to a 28 VDC external power supply.

Thermal protection and interlock with P2.51P3 bleed extraction and VCCS operation is fully aulomatic and requires no pilot input. The units will operale conlinuously and independently of the Cabin Air Temperature Control System when selected. Temperature control of the cabin is a lunction 01 the air cycle system. If the air temperature of the cabin exceeds the selected value the air cycle system will compensate by moving to a cooler position by passing more air through the syslem cold air unit. The mixed air temperature at the main distribution system outlets will then become colder.

There are no special operating procedures for the system. The units can be selected when desired. The cabin air temperature display on the center console gives a good indication of when they are required. i t is recommended that the heaters be selected on when the cabin air temperature is observed to fall to 18 OC. However, selection at cabin air temperatures above 18" C wlll help in preventing an initially cold cabin.

( System tunction may be monitored on the GEN 1 (MSN 101-111) and GEN 2 DC current indrcation on the EIS Display Unit.

Pre SB 21-002 - When the CKPT position is selected, an increase of approximately 60 amps 1 will be observed on the GEN 1 (MSN 101-1 11) output. When the CAB B position is selected,

an increase of approximately 60 amps will also be observed on the GEN 2 output.

Post SB 21-002 - When the LOW position is selected. an increase of approximately 60 amps will be observed on the GEN 2 output. When the HIGH position is selected, an increase of approximately 60 amps will also be observed on the GEN 1 (MSN 101-1 11) output.

Report No: 01973-001 7-90


WPLATUSW SECTION 7 K P AIRPLANE AND SYSTEMS DESCRIPTION

HEATING SYSTEM (MSN 321 Q UP and aircraft with modified heating system)

OENERAL

Refer to Figure 7-14A. Environmental Control System, for the system layout.

The PC12 is equipped with an auxiliary electrical heating system, whlch is used to supplement the air cycle system during prolonged tow temperature operations such as cruise at high altihrde. It can also be used for pre-heating the cabin on the ground.

DESCRIPTION

The system comprises two 28 VDC heating units each equipped with a 75 mm mixed flow fan. Each unit is cylindrical in form and contains two heating elements producing 1.625 kW1unit. The system therefore produces 3.25kW in addition to that of the air cycle system. The units are situated under the cabin floor, one is dedicated to heating the cabin and the other to heating the under floor avionics bay. The cabin heater is supplied 28 VDC power by the GEN 1 bus and the under floor heater is supplied by the GEN 2 bus.

The under floor heater is located between frames 21 and 22. The fan scavenges its air supply from the general under floor zone, through a wire mesh inlet grill, and passes it over the heating element where its temperature is raised. The air is then distributed along the length of the under fbor avionics bay by way of a longitudinal distribution duct.

The cabin heater is located between frames 29 and 30. The fan draws its air supply from the cabin, through a grill in the rear floor step. The heated air is then ducted directly to the ECS distribution duct in the cabin sidewall and augments the ECS airflow. The airflow created by the cabin heater is effective in equalizing the temperature throughout the cabin.

Both heater units are equipped with an internal thermal protection system, which isolates the heater when the element temperature overheats. In the event of an over heat, the fans continue to run and the relevant CABIN or UIF HEATER circuit breaker, located in the rear left circuit breaker panel on the pilots side, will trip. The heater will remain isolated until the temperature falls within the heater allowing the circuit breaker to be reset by the pilot.

The power for the heater element circuits is interrupted when the hydraulic pump or cooling system (VCCS) is operating. The under floor fan continues to run, the cabin fan is inhibited. This minimizes generator accessory loads during continuous normal operation and prevents generator overload.

The heating capacity of the system is reduced while the engine is operating at P3 bleed in flight. The cabin heater and fan are inhibited while airborne and P3 bleed is extracted, the under floor heater and fan remain operating. While on the ground (WOW valid) the cabin

Issued: June 10, 1994 Revision 10: September 1,2000


heater and fan continue to operate when P3 is extracted. During engine start and for 10 seconds following engine start both heaters and fans are inhibited.

The function of the power inhibits are fully automatic and require no pilot input. Thermal protection, once tripped, will require pilot action to reset.

OPERATION

The Auxiliary heating system is controlled by a combination of the ECS switch and the CABlN HEATING switch, situated on the pilot's lower left panel to the left of the pilots control yoke. Selecting the ECS switch to AUTO or MANUAL will automatically enable the under floor heater. Selecting the CABlN HEATING switch to AUTO will enable the cabin heater.

When selected, both heater elements are automatically controlled by temperature sensors located in the under floor and the rear cabin. The sensor outputs are processed along with pilot inputs, power supply condition inputs and heater thermal safety switch condition to enable or inhibit the heater and fan functions.

With the ECS switch selected to AUTO, the under floor fan runs continuously and the heater element is switched on when the under floor sensor reads below + 5°C and is switched oft above + ll°C. The cabin fan runs continuously when the CABlN HEATING switch is selected to AUTO and the heater element is switched at +/- 1°C about a nominal cabin temperature of 24°C. This switching temperature can be adjusted from 21°C to 27°C by using the HEAT CAB potentiometer located in the maintenance panel. It is recommended that the potentiometer be normally set to the mid (24°C) position.

Both heaters and fans can be operated independently of the ECS bleed air system for preheating the aircraft before engine start when required. Switching the CABlN HEATING to AUTO will select both heaters on when the aircraft is connected to a 28 VDC external power supply.

There are no special operating procedures for the system. Temperature control of the cabin is principally a function of the air cycle system (ECS). The CABIN TEMP display on the center console gives an indication of the cabin air temperature. It is recommended that the CABlN HEATING switch be kept at AUTO and the ECS TEMPERATURE control be used to adjust the cabin temperature. With the cabin heater operating, the recommended ECS TEMPERATURE control settings, based on the IOAT, are:

- 20°C IOAT, set the control to the mid (12 o-clock) position - 40°C IOAT, set the control to the MAX (hot) position

System function may be monitored on the GEN 1 AMP indication on the overhead panel and GEN 2 DC A indication on the EIS Display Unit.

Report NO: 01973-001 7-908


ePL-MUSS SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

When CABIN HEATING is selected ON the GEN 1 output will increase by approximately 60 amps. The under floor heater can be checked by monitoring the GEN 2 output. A reading of more than 50 amps indicates that the heater is on.

FOOT WARMER SYSTEM (OPTIONAL OR POST SB 21-004)

The system comprises a 28 VDC 1kW heater installed forward above the cockpit floor. Ducting connects the heater to foot outlets at the pilot and copilot position. A rocker type switch HEATING FOOT is installed on the pilot's lower left panel. It has the positions ON and OFF. Power supplies are 28 VDC from the powerline to the heater relay and from the non essential bus through the FOOT WARMER circuit breaker to the switch.

OPERATION

The foot warmer system operates from the aircraft electrical power or from external power. When the HEATING FOOT switch is set to ON 28 VDC is supplied to the heater relay. The relay is energized and the heater and fan operates. The heated air is sent by the fan to the pilot and copilot foot outlets. if the temperature of the heater becomes too high the thermal protection switch operates and de-energizes the heater relay.


Report No: 01973-001 7-90C

SECTION 7 EPILANSW AIRPLANE AND SYSTEMS DESCRIPTION K P 4

1

COhlDKWED AIR SUPPLIED F R W ECS

I FOOT WARMER FIREWALL

OPTION AN0 1 SHUTOFF VALVE

4

FOOT OUTLET I

MOT OUTLET 1 OPENCLOSED

LEVER 1

PANEL 0un.m

t

EMERGENCYW AIR IMET

4

N W R E N R N VALVE

CABH SIDEWALL CABIN SIDEWALL DISTRIBUIION WCT RlOHT

UlFLOOR DlSTRIBUTKX(

1 DUCT

I

Figure 7-14A. Envi ronmenta l Control System (MSN 321 & UP) (Sheet 1 o f 2

Report No: 01 973-001 7-900

Issued: June 10, 1994 Revision 10: September 1,2000

'ISPILNIJSW SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

FOOT WARHER OPTION AND SB 21-004 SHUTOFFVALVE

Figure 7-14A. Environmental Control System (MSN 321 & UP) (Sheet 2 of 2) I


Report No: 01973-001 7-90E

SECTION 7 'BCPYATUSW AIRPLANE AND SYSTEMS DESCRIPTION K12


Report No: 01973-001 7-90F



COOLING SYSTEM (OPTIONAL)

GENERAL

The Vapor Cycle Cooling System (VCCS) (when installed) is designed to operate on the ground from a 28 VDC power cart or aircraft electrical power when the engine is operating. The electric motor driven system provides a means of precooling cockpit and cabin areas prior to and during passenger boarding, providing comfort prior to engine start.

The system may be operated during ground operations and up to an altitude of 25,000 feet.

This type of system removes a large percentage of the moisture as well as dust and pllen particles from the cabin air.

DESCRIPTION

Refer to Figure 7-15, Vapor Cycle Cooling System.

A refrigerant gas is the media which absorbs heat and rejects heat from the cabin air. By continuous recirculation of cabin air, heat is absorbed in the evaporator modules and transferred to the outside through the system condenser.

The system is provided with safety interlock devices to prevent component damage andlor excessive power drain from the aircraft electrical system. The evaporator modules are equipped to prevent coil icing at all ambient conditions.

Cabin temperature control is by varying the airflow through each evaporator module rather than cycling the refrigerant compressor. Airflow is controlled by the flight crew. The cabin is cooled by air ducted from the two evaporators located just forward of the aft pressure bulkhead and exhausted through individuals outlets down the left and right sides of the cabin. A third evaporator, located between the other two, exhausts air directly into the cabin.

The cockpit is cooled by individual outlets located in the overhead panel. These outlets receive air ducted from the two evaporators in the cabin.


Report No: 01 973-001 7-91 1


OPERATION

When the system is activated, an electric motor drives the compressor at constant speed and capacity which compresses the refrigerant gas to high pressure. The hot, high pressure gas then passes through the condenser coil where it is cooled and condensed into a warm liquid at constant pressure. The heat removed from the fluid is exhausted overboard through a vent in the right rear tail section aft of the pressure bulkhead. The warm liquid from the condenser is then routed into a receiver-dryer container where the liquid and any remaining gas are separated and any moisture in the liquid is absorbed. The warm dry, high quality liquid is then routed to the evaporator module expansion valve where the high pressure liquid is expanded to a low pressure. The large expansion process creates a super cool liquid which passes through the evaporator coil and absorbs heat from the warm cabin air. The cooled air is returned to the cabin. The gas, now warm, is returned to the compressor to repeat the cycle.

Moisture removed lrom the cabin air by each evaporator drains into a small holding tank below the rear baggage floor panel. The water is held in the tank until the cabin differential pressure is low enough for the the tank outlet valve to open allowing the water to drain ov8rboard.

The VCCS operation is controlled by three switches located on the pilot's lower lefl panel; COOLING SYSTEM, FLOOD FAN, and VENTS FANS.

The COOLING SYSTEM switch has three positions. The OFF position removes power from both the compressor and the evaporator fans. The RECIRC position will supply power to the evaporator fans only for recirculation of the cabin air. The ON position will supply power to the fans and the compressor.

Cabin air temperature control is accomplished by selecting evaporator fan speed. The evaporator fan speed is controlled by the FLOOD FAN and the VENTS FANS switches. Each switch has two positions, LOW and HIGH. When the FLOOD FAN switch is set to LOW, power is removed from the center evaporator fan. When the FLOOD FAN switch is set to HIGH, the center evaporator fan blows cool air directly into the cabin. The VENTS FANS switch controls the speed of the evaporator tans which blow cool air into the left and right overhead ducts. Individual outlets in the overhead panel are adjustable for local temperature control at each seat location.

When the VCCS is operating, a blue COOL annunciator will illuminate on the CAWS panel. The GEN 1 DC Indication will increase by approximately 100 amps for compressor and evaporator fans operation.

Report No: 01973-001 ( 7-92


=PLATUS= -PC xll- SECTION 7


FAN AIRFLOW

CABIN WERHEAO DUCT

FAN AIRFLOW

EVAPORATOR I EL0

PRESSOR1 CONOENSER

CONDENSER INLET

HlGU PRESSURE LKIUID

LOW PRESSURE W

VENTFAN

FLOOOFAN

I Figure 7-15. Vapor Cycle Cooling System

Issued: February 14, 1994 : Revision 3: September 29. 1995

Report No: 01 973-001 7-93 1

SECTION 7 =PLATUS= AIRPLANE AND SYSTEMS DESCRIPTION PC XI^


GENERAL

Control of the cabin pressure is obtained by regulating the rate of exhaust of the outflow air that is provided for pressurization and ventilation of the airplane cabin. Operation of the Cabin Pressurization Control System (CPCS) is entirely pneumatic, except when switching between ground and flight operating modes and during manual inlight depressurization. Mode switching and inflight depressurization are accomplished through solenoid operated valves (Pre SB 21- 001)1 a solenoid valve (Post SB 21-001 and MSN 141 AND UP) controlled by the weight on

/ wheels switch, a microswitch on the engine condition lever and the DUMPIAUTO switch on the I ECS control panel. The CPCS will maintain the selected cabin altitude up to a maximum

pressure differential of 5.75 psi, equivalent to a 10.000 foot cabin altitude at a cruising altitude of 30,000 feet.

DESCRIPTION

Refer to Figure 7-16, Cabin Pressurization, for system layout.

The CPCS consists of a cabin outflow valve controller, two cabin outflow/safety valves, an auxiliary volume tank, a check valve, a ejector valve, a manual control valve, two solenoid

I valves (Pre SB 21-001)l one solenoid valve (Post SB 21-001 and MSN 141 AND UP), a condition lever switch, a depressurization switch, a weight on wheels switch, and indication of cabin altitude, cabin altitude rate of change, and pressure differential. Airplane electrical power is supplied to the system for the mode switching, solenoid operation, cabin altitude warning on the CAWS, and outflow valve controller lighting.

The outflow valve controller is pneumatically connected to the system through three ports: VALVE, TANK, and VACUUM. The controller VALVE port is connected to the outflow valve control chamber to provide control reference pressure for operation of the outflow valve. The controller TANK port is connected to the auxiliary volume tank, located near the controller. The VACUUM port is connected to the ejector valve which provides a source of low pressure air. The controller contains a chamber that houses an absolute pressure bellows and a rotating actuator. This chamber is vented to the cabin through the cabin air sense port. Rotating the cabin altitude selector knob on the face of the controller to the desired altitude, causes the actuator to rotate and compress or extend the bellows depending on the direction of knob rotallon. The bellows position controls the reference pressure applied the outflow valve control chamber. The cabin rate control selector knob on the face of the controller adjusts the position of the rate control valve. Rotating the knob counterclockwise decreases rate of change while rotating the knob clockwise increase the rate of change. When the arrow is at 12 0' clock the rate of cabin altitude change is approximately 500 fpm. The rate of change can be adjusted from approximately 100 fpm to 2000 fpm.

Report No: 01973-001 7-94


=PlLATUS E SECTION 7 -PC XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Figure 7-16. Cabin Pressurization (Sheet 1 of 2)


Report No: 01 973-001 7-95

SECTION 7 WPIL ATUS W AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1 *

4

COPILO~S LOWER LEFT PANEL

\ , \. 9 1%

I

ECS FIREWALL snmw VUVE

Figure 7-16. Cabin Pressuriratiarr (Sheet 2 of 2)

Report No: 01 973-001 1 7-96


5 P L A T U S E SECTION 7 ?c XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Low pressure needed for operation of the controller is provided by the ejector valve, which is connected to the controller VACUUM port through a check valve and a normally open solenoid valve (Pre SB 21-001). The ejector valve is also connected to the safety valve through a 1 normally closed solenoid valve to provide the low pressure to operate the valve for inflight depressurization and unpressurized ground operation. Engine bleed air flows through the ejector valve inducing low pressure to be applied to the controller. The check valve installed in the vacuum line behveen the ejector valve and the controller prevents high pressure bleed air from entering the control system in the event that the ejector valve exhaust becomes blocked.

The manual control valve is connected on one side to the control line from the controller to the outflow valve. The other side of the manual control valve is vented to atmospheric pressure. Opening the manual control valve reduces the pressure in the outflow control chamber, opening the outflow valve. The manual control valve controls the cabin pressure from the unpressurized condition to the maximum differential with the outfiow valve controller inoperative.

The auxiliary volume tank is a small sealed chamber that adds volume to the controller rate chamber to increase the accuracy of the rate of change control.

There are two cabin oufflowlsafety valves located under the cabin floor. These valves are identical except for the installation and intended function. The outflow valve is connected to the controller and controls cabin pressure. The safety valve is connected to the vacuum line from the ejector valve through a normally closed solenoid to open for inflight depressurization and unpressurized ground operation. Both valves are vented to atmospheric pressure. Each valve senses control chamber-to-atmosphere pressure differential and modulates to maintain the selected pressure differential.

Pre SB 21 -001

Two solenoid valves are installed in the CPCS. One valve (solenoid valve 1) is installed in the low pressure line from the ejector valve to the controller and is normally open. The other valve (solenoid valve 2) is installed in the safety valve control line and is normally closed. These solenoid valves are controlled by the pressurization switch, AIR/GND switch, or the microswitch on the condition lever. When these switches are activated, vacuum is applied to the safety valve to depressurize the cabin. Electrical power for the mode switching is from the ECS circuit breaker on the battery bus.

Post SB 21-001 and MSN 141 AND UP

A solenoid valve is installed in the CPCS. It is installed in the safety valve control line and is '

normally closed. It is controlled by the pressurization switch, AIWGND switch, or the microswitch on the condition lever. When these switches are activated, the solenoid is energized, the solenoid valve opens and vacuum is applied to the safety valve to depressurize the cabin. Electrical power for the mode switching is from the ECS circuit breaker on the battery bus.

Issued: June 10.1994 Revision 5: May 10, 1996

Report No: 01973-001 7-97


OPERATION

Pre SB 21-001

During normal ground operation with the pressurization switch in AUTO and the condition lever in the Ground ldle position, solenoid valve 1 will close and solenoid valve 2 will open when electrical power is applied to the system. Aller engine start and engine bleed air is supplied to the ejector valve, the outflow valve will remain closed while the safely valve will open. The airplane will be unpressurized. Immediately before takeoff, set the desired cabin rate of climb and set the outflow controller to 500 feet above the airplane cruise altitude.

When the condition lever is moved to the flight idle position prior to takeoff, electrical power is removed lrom both solenoids. Solenoid valve 1 will open and vacuum is applied to the controller. Solenoid valve 2 will close, removing the vacuum from the safety valve and allowing it to move to the closed position. At this time, the controller begins to apply a low reference pressure to the oufflow valve control chamber.

I Post SB 21-001 and MSN 141 AND UP

During normal ground operation with the pressurization switch in AUTO and the condition lever in the Ground ldle position, the solenoid valve will open when electrical power is applied to the system. The airplane will be unpressurized. Before selecting the ECS on, set the desired cabin rate of climb and set the outflow controller to 500 feet above the airplane cruise altitude. After engine start and engine bleed air is supplied to the ejector valve, the outflow valve and the safety valve will open

/ When the condition lever is moved to the flight idle position prior to takeoff, electrical power is j removed from the solenoid. The solenoid valve will close, removing the vacuum from the safety , valve and allowing it to move to the closed position. As before, the controller applies a low I reference pressure to the outflow valve control chamber.

As the airplane altitude increases, the cabin altitude increases at the selected rate of change until the preselected cabin altitude is reached. Cabin altitude is then maintained until the cabin altitude controller is set to a different cabin altitude, the airplane descends below the selected cabin altitude, or the airplane climbs to an altitude where the maximum pressure differential is exceeded.

Should the airplane climb to an altitude where the cabin-to-abnosphere pressure differenlal exceeds the calibrated settings on the outflow/safety valves, the pressure relief fIJn~ti0n overrides the automatic pressure control function, and the cabin altitude rate of change begins to track the airplane altitude rate of change.

Report No: 01973-001 7-98

Issued: June 10,1994 Revision 5: May 10. 1996

=PLATUS= SECTION 7 PC XI^ AIRPLANE AND SYSTEMS DESCRIPTION

NOTE

If the airplane climbs at a high rate with a low cabin rate selected on the controller, the positive pressure relief function may activate before the airplane reaches the selected cruise altitude.

When operating at a cabin altitude that produces a pressure differential near the maximum limit and a climb is initiated, select a higher cabin altitude on the outflow valve controller so that the cabin altitude is controlled by the rate control and not by the positive pressure relief function of the outflow valve.

Pre SB 21 -001

I f an unpressurized condition is desired for an emergency (smoke in the cabin), select the DUMP position on the pressurization switch to apply electrical power to the solenoid valves. Solenoid valve 1 closes, removes the vacuum from the controller and closes the outflow valve. Solenoid 2 opens, allows vacuum to be applied to the safety valve and opens the valve. Cabin air is quickly exhausted to the atmosphere, depressurizing the airplane cabin. An alternate means to depressurize the cabin is by closing the Bleed Air Firewall Shutoff Valve located at aft end of the center console.

Repressurizing the airplane cabin after depressurization is accomplished by setting the pressurization switch to AUTO to remove electrical power from the solenoid valves. Solenoid valve 1 opens and applies a vacuum to the controller. Solenoid valve 2 closes and removes the vacuum from the safety valve, causing the valve to close. Automatic control functions return and the CPCS operates on the rate-of-change control until the selected cabin altitude is regained.

Post SB 21 -001 and MSN 141 AND UP

If an unpressurized condition is desired for an emergency (smoke in the cabin), select the DUMP position on the pressurization switch to apply electrical power to the solenoid valve. The solenoid valve opens, allows vacuum to be applied to the safety valve and opens the valve. Cabin air is quickly exhausted to the atmosphere, depressurizing the airplane cabin. An alternate means to depressurize the cabin is by closing the Bleed Air Firewall Shutoff Valve located at aft end of the center console.

Repressurizing the airplane cabin after depressurization is accomplished by setting the pressurization switch to AUTO to remove electrical wwer from the solenoid valve. The solenoid valve closes and removes the vacuum from the safety valve, causing the valve to close. Automatic control functions return and the CPCS operates on the rate-of-change control adjusting the position of the outflow valve poppet untl the selected cabin altitude is regained.

When atmospheric pressure exceeds cabin pressure and there is a reduction of cabin air inflow, a negative cabin differential exists across the inner diaphragm of both the outflow and

Issued: June 10, 1994 Revision 5: May 10,1996

Report No: 01973.001 7-99


safety valves. When the control chamber-to-atmosphere pressure differential is sufficient, poppet valves open and allow air at atmospheric pressure to flow into the cabin until the pressures are equal. This condition may occur during a rapid descent as the airplane descends through the selected cabin altitude. Selecting a lower cabin altitude before a descent is commenced, will allow the cabin altitude to be controlled by the rate controller and prevent an unpressurized condition.

During descent and in preparation for landing, select a cabin altitude approximately 500 feet above landing field elevation and select a cabin rate of change that will bring the cabin altitude to the new setting before the airplane descends to selected cabin altitude. This reselection should be accomplished far enough in advance to prevent the airplane from descending through the cabin altitude, which may occur because of a low cabin rate of change selection. As the airplane descends through the preselected cabin altitude, the cabin becomes unpressurized and follows the airplane rate of descent to touchdown. On approach check that the cabin is depressurized. At touchdown the AIRIGND switch closes.

Pre SB 21 -001

After landing when the condition lever is set to Ground Idle, electrical power is applied to the solenoid valves. Solenoid valve 1 closes, removing the vacuum from the controller and closes the outflow valve. Solenoid 2 opens, allows vacuum to be applied to the safety valve and opens the valve.

Post SB 21 -001

After landing when the condition lever is set to Ground Idle, electrical power is applied to the solenoid valve. The solenoid valve opens, allows vacuum to be applied to the safety valve and opens the valve.

The cabin altitude selector displays the selected airplane cnrise altitude and the corresponding cabin altitude at the maximum pressure differential. The actual cabin altitude, cabin altitude rate of change, and cabin pressure differential are displayed on the copilots lower panel.

It is possible to operate at a cabin differential pressure above the normal operating range if the cabin altitude controller was improperly set by the pilot or i f a system failure has occurred. A 'dead band' exists between the normal operating maximum differential pressure (5.75 psid) and the pressure at which the system relief valves open (6.35 psid). This "dead band' is indicated on the cabin differential pressure gauge by an amber arc between 5.75 and 6.5 psid. A red radial at 6.5 psid indicates the maximum permissible cabin pressure and is the maximum switching point set for the cabin pressure differential warning switch.

The CAB PRESS annunciator on the CAWS will illuminate when the cabin altitude exceeds ) 10.500 * 200 feet or when the maximum cabin pressure differential is exceeded.

Issued: June 10.1994 Revision 5: May 10. 1996

L mPILATUSW SECTION 7

I PC12 AIRPLANE AND SYSTEMS DESCRIPTION

OXYGEN SYSTEM

GENERAL

The airplane is equipped with an emergency oxygen system for use by the crew and passengers in the event of contaminated air being introduced into the cabin or a loss of pressurization with a rapid descent to tower altiiudes.

The pilot and copilot masks are supplied with quick-donning diluter-demand masks which are permanently connected to outlets in the cockpit sidewalls.

A constant flow mask is provided at each passenger seat location in the cabin. In the Corporate Commuter interior configuration the nine masks must be connected to the bayonet outlets In the cabin sidewall before flight by the flight crew for flights above 10,000 ft. In the executive interior configuration the masks (the number is dependant on the interior variation) are located in boxes in the arm rests and are permanently connected for all flights. No connection actin is required by the flight crew or passengers.

DESCRIPTION

An oxygen cylinder, made of composite material, is located in an external compartment in the right side of the fuselage forward of the main wing (outside the pressure area) from which the oxygen system is serviced and replenished (Refer to Section 8 for servicing instructions).

Pre SB 35-001 - Attached to the cylinder head is a manual isolation valve to perm& cylinder removal and installation, a connection to the ground charging valve, a connection to the contents pressure gauges, and a connection to the over-pressure relief valve.

Post SB 35-001, MSN 181 and UP - Attached to the cylinder head is an isolation valve to permit cylinder removal and installation. The valve is connected by a push pull cable to a handle in the cockpit allowing the system to be isolated while the aircraft is on the ground. The valve is connected to the aircraft supply, ground charging valve, the contents pressure gauges and the over-pressure relief valve.

Two gauges are provided, one in the service bay and one on the lefl cockpit side panel forward of the Test Panel. Overpressure protection is provided by a relief valve in the form of a green rupture disc located in the fuselage skin above the service bay door. This disc is designed to rupture at 2775 +50/-0 psi, discharging the cylinder contents overboard. Disc integrity is checked during the preflight inspection. If found ruptured and the contents pressure gauge indicates zero, proper maintenance must be performed on the system before flights above 10,000 R altitude.

When filled, the storage cylinder should be charged to 1850 psi (127.6 bar) at 20" C. with a minimum pressure of 265 psi (18.3 bar) for proper flow to the masks. A pressure reducing valve, adjacent to the oxygen cylinder reduces the oxygen pressure to a nominal 70 psi, prior


! Revision 10: September 1,2000 Report No: 01973-001

7-101

SECTION 7 CPILATUSP AIRPLANE AND SYSTEMS DESCRIPTION P(12

to entering the cabin. This is for safety reasons and to avoid excessive flow through the masks.

Two crew full-face masks of the diluter demand type are located In boxes on the front of the cockpit bulkhead behind each crew member. They are permanently connected to outlets in the cockpit sidewalls. Each mask which is of the diluter-demand type, is equipped with a microphone and an ONIOFF - AIRMIX/100% selector valve. Oxygen is provided to the crew masks at all times regardless of the Oxygen control valve position. Each mask has a PRESS TO TEST button and a flow indicator that shows when proper pressure is supplied to the mask. Turning the PRESS TO TEST button counterclockwise to the emergency position will supply 100% oxygen at a slight overpressure.

Post SB 35-001, MSN 181 and UP - The oxygen shut-off handle is mounted to the right of the centre console. It is connected by a push pull cable to the isolation valve on the cylinder head. While the aircraft is on ground the handle is normally in the OFF position isolating the cylinder from the system and preventing prolonged leakage fmm the crew masks. Before engine start and as the first action associated with the oxygen system, the handle should be moved to the ON position.

The Oxygen Control Valve, located in the left cockpit sidewall, has three positions to control the operation of the passenger distribution system. The OFF position stops the flow to the passenger outlets. The ON position permits flow to the passenger masks. The AUTO position will permit flow to the passenger masks when a pressure switch senses a cabin altitude above 13,500 feet +I- 500 feet.

In the Corporate Commuter configuration the passenger constant flow oxygen masks are stored under or near each seat position. For flights below 10,000 ft altitude the masks need not be connected to the outlets in the lower cabin sidewalls. In the event of an emergency requiring oxygen use, the passengers are instructed to connect the mask bayonet type connector to the outlets themselves. For flights above 10.000 ft altitude the mask must be connected to the outlets by the flight crew before flight. When disconnected, the outlets are spring loaded closed to prevent oxygen leakage.

In the executive interior configuration the passenger constant flow oxygen masks are stowed under covers placarded OXYGEN MASK INSIDE in the cabin sidewall armrests. The masks are permanently connected to the outlets irrespective of the type of operation and flight altitude. The mask stowage compartments are located near to the seats. The masks have a red tape band which must positioned to show from the cover in the direction accessible to the seat occupant. A placard PULL TAPE FOR OXYGEN MASK is attached to the armrest near each oxygen mask cover. Optional SB 35-003 installs an oxygen mask in the lavatory. The mask is connected to the passenger oxygen system and is stowed in a box attached to the top of the lavatory sidewall. A visible red tape band is pulled to release the oxygen mask.

In aircraft with an optional three seat bench installed, an oxygen mask is stowed below each seat behind a cover in the front pedestal of the bench seat. A red tape band attached to the mask shows from the panel and assists in the removal of the mask, in the event of it being required for use.

Report No: 01973-001 7- 102

issued: June 10,1994 Revision 1 1 : March 1,2003


OPERATION

I WARNING 1 TO PREVENT POSSIBLE FREEZING AND MALFUNCTIONING OF SYSTEM, MAKE SURE THAT SYSTEM IS ONLY SERVICED WlTH APPROVED, AVIATION GRADE OXYGEN.

TO PREVENT POSSIBLE EXPLOSION AND/OR FIRE, MAKE SURE ALL OIL AND GREASE IS KEPT AWAY FROM OXYGEN SYSTEM COMPONENTS.

SMOKING IS STRICTLY PROHIBITED ANY TIME OXYGEN IS IN USE.

OILY, F A l l Y OR GREASY SUBSTANCES, INCLUDING SOAPS, LIPSTICK. AFTER SHAVE LOTION, MAKE-UP ARE CAPABLE OF SPONTANEOUS COMBUSTION ON CONTACT WlTH OXYGEN.

Normal system operation is with the three-position Oxygen Control Valve in the AUTO position, to provide oxygen immediately in the event of a depressurization. The crew will then don their own masks and order the passengers to don their masks. The masks in an executive interior aircraft can easily be removed from their stowage by pulling the red tape band showing from the cover marked OXYGEN MASK INSIDE. Oxygen flow to the cabin is verified by the oxygen pressure switch activating the CAWS annunciator PASS OXY.

The ON position will be selected by the pilot, in the event ol smoke or fumes being present in the cabin, and also to test the passengers masks on the ground before passenger boarding. The OFF position will be selected if the aircraft is being flown without passengers or is taken out of service for an extended time in order to conserve oxygen.

NOTE

When a full oxygen supply is stored, it will supply two crew and nine passengers for a minimum of ten minutes, in which time a descent from 30.000 ft to 10,000 ft is performed. Refer to the Oxygen Duration Chalt in Section 4 to determine the minimum oxygen supply required for the number of occupants when operating at less than full oxygen pressure.

As the oxygen system is an emergency system, normal usage will consist only of periodic mask testing (both crew and passengers masks require testing) and of checking, and topping up, if necessary, the storage cylinder.


Report No: 01 973-001 7- 103

SECTION 7 WPLATUSW AIRPLANE AND SYSTEMS DESCRIPTION Kl2

INDICATION I WARNING

Oxygen system pressure is indicated on a gauge on the left cockpit sidewall fonvard of lhe Test Panel. A pressure switch downstream of the Oxygen Control Valve will activate the CAWS annunciator PASS OXY when proper oxygen pressure is supplied to the passenger masks.

OPTIONAL LARGER CAPACITY OXYGEN SYSTEM

The system has a 1965 liter gaseous oxygen cylinder installed in the top left side of the rear hrselage compartment, behind the rear pressure bulkhead. The cylinder head isolation valve is secured in the open position. System shut off, when the aircraft is on the ground, is by a rotary valve connected to the cable from the oxygen shutolf handle on the center console. The rotary valve is installed between frames 16 and 17 on the right side of the fuselage. A pressure transducer installed near the oxygen cylinder sends a pressure signal to the pressure gauge on the lelt side of the cockpit. The oxygen replenishment point comprising a charging valve and a system pressure gauge is installed at the bottom of the rear fuselage compartment. The system overpressure protection burst disc indicator is installed on the left side of the rear fuselage.

System controls and operation are the same as for the standard system. The system with full oxygen pressure will meet the Canadian Operational CAR 605.31 and CAR 605.32 requirements. Refer to the Oxygen Duration Chart in Section 4 to determine the minimum oxygen supply required for the number of occupants when operating at less than full oxygen pressure.

COCKPIT ARRANGEMENT

The cockpit is of a conventional layout in that all of controls, switches, and instruments are readily accessible to the pilot for single pilot operation. The overhead panel contains the switches to control the electrical bus distribution system and the sidewalls contain the circuit breaker panels. The instrument panel contains: the flight, navigation, and engine instruments; avionics; pressurization and cabin heat controls; and the lower panel on the left side contains the electrical switches. The center console contains the CAWS annunciator panel, EFlS control units, trim indicator, engine power controls, flap selector, cockpit and cabin lighting controls, emergency landing gear handpump, and the fuel and ECS firewall shutoff controls.

DESCRIPTION

The overhead panel contains the switches to control the DC and AC electrical power generation and distribution to the various busses. A bus distribution diagram is shown on the panel with the switches in the appropriate places to help identity system design, operation, and maltunction. The individual busses and their associaled circuit breakers are located in the lert and right cockpit sidewalls. The arrangement of the overhead switches is similar to the location of its associated bus in the sidewalls. For instance, the generator 2 switch is on the

Report NO: 01 973-001 1 7-104 Issued: June 10,1994

Revision 9: September 1,1999

ePILATUSZ SECTION 7 K12 AIRPLANE AND SYSTEMS DESCRIPTION

left side of the overhead panel which corresponds to the generator 2 bus circuit breakers on the left side circuit breaker panel.

Aft of the left side circuit breaker panel is the Test Panel. The different system and annunciator test circuits are controlled here. Individual test circuits are described with their associated systems within this section, except for the LAMP test switch. When this switch is pressed, all of the annunciator lamp bulbs are functionally checked. This includes the CAWS, landing gear indicators, flap overspeed light within the flap gauge, and the master warning and caution lights. The lamp function may be checked at any time without interfering with system function.

Forward of the Test Panel is the oxygen pressure gauge. The clock and parking brake handle are located forward of the left side circuit breaker panel below the instrument panel. The clock is powered directly from the Battery Direct bus.

The lefl side of the instrument panel contains the flight instruments for the pilot. These indude the airspeed, attitude, directional, alitude, and rate of climb instruments. The airspeed indiiator provides data for various system control inputs and has a variable Mmo needle for the constant Mach maximum airspeed limit above 15.200 feet. Exceeding this maximum airspeed limit (Vmo or Mm) will trigger the overspeed aural warning alert. Also. if the maximum speed for the current flap setting is exceeded. a red warning light will illuminate on the face of the flap indiiator. To the left of the airspeed indicator is the standby attitude gyro. This gyro is power from the Battery Bus. Refer to the Electronic Flight Instrumentation System (EFIS) within this section for the description of the attitude and directional instruments. A single instrumented aircraft has a standby altimeter lower left of the instrument panel.

The right side of the instrument panel is available for optional copilot instruments.

The let! center of the instrument panel contains the Engine Instrument System (EIS). Refer to EIS wlhin this section for the description of this system. The right center of the instrument panel contains the audio selector panel and the communication and navigation radios. Refer to Section 9. Supplements for description and operation information lor optional avionics systems.

The lower panel on the pilot's side contain electrical switches for engine and fuel systems, cabin fans, avionics, exterior lights, deice and anti ice systems, and the landing gear selector and position indicators. The bwer panel on the copilot's side contain the ECS and pressurization controls.

The center console contains the CAWS annunciator panel, EFlS control units, trim indicator. cabin air temperature gauge, the trim and flap interrupt and alternate power switches, and the engine power controls and flap lever. Further aft will be the cockpit and cabin lighting controls. The ECS and fuel firewall shutoff valve controls and the emergency landing gear handpump' can be found on the alt vertical surface of the console.


Report No: 01973-001 7- 105

SECTION 7 'BEPLmSW AIRPLANE AND SYSTEMS DESCRIPTION 36 %I1

I. SYSTEM TEST PANEL 2. LEFTSMWAU RUR 3. LEFl SlDEWALl FRONT 4. PILOTS LOWER LEFT 5. PILOTS LOWER RIO(T 6. PLOFSWSTRYENI 7. OVERHEAD a OURESMIEU)

ta RYWTBDmALLREAR 13 COPLOTSLOWER RKUIT 14 COPLOPS LOMR LEFT 15. CENTERLOWER 16 CENTEIICONOQE

Figure 7-17. Typical Cockpit Layout (Sheet 1 of 2)

Report No: 01973-001 1 7-106 Issued: Februaly 14,1904

Revidon 3: September 29,1995


1. SYSTEM TEST PANEL 2. OXYGEN PRESSURE INDICATOR 3. EFl SIDEWALL CWUIT BREMER PANES 4. CLOCK 5. PARKMa BRAKE MANME 6. 2nd ALTIMETER SINGLE INSTRUMENTED A I R C W 7. OWCT vl- b v ) w- 8. MAW(UIBEA 0. SUNVISOII 10. STANDBY ATTITUDE GYRO $1. NAVKi*TK)N WSTRWNTS 12. F U P POSITtON IWICATOR 13. AIRSPEED INDICATOR 14. MASTER CAUTION AND WARNING LIGHTS 1: PI. VERTICAL SPEED PRESELECTOR

ii: En3i 16. CONTROL WHEEL 19. VERTICAL SPEED INDICATOR E. SZWETER

ii t& RANGE NAVIGATION RADIO a. AUTOPILOT MOM CONTROLLER :: W&EZIC COMPASS

%, A ~ T < ~ T S n OVERHEAD ELECTRICAL CONTROL PANEL 28 SUNVISOR 29 AUDIO SELECTOR PANEL 10. COMMVNICATON I NAVlWTlON 31 AIRSPEED MICATOR !f. p S / E R CAUTION AND WAANINQ LIGHTS - - . - . .- . 34. ALTWETER 35. AKiHT SIDEWALL CIRCUIT BREAKER PANELS 36. RUDDER PEDALS

37. RUMXR PEDAL W S T M E N T HANOLE 38. PRESSURIUTICN CONTROLS I INDICATORS 39. CONTROL WnEEL 40. VERTICAL SPEED MMGATOR

::. a* ii: €76 MFD 44 ECS CONTROLS 45. CAWS 46. RESERVED

52. ALTERNATE STAB TAM SWITCH S. F W SELECTOR 54. CONDITION LEVER 55. COCKPIT1 CABIN LIGHTING CONTROLS 56. ECS FIREWALL SHUTOFF CONTROL 57 FUEL FIREWALL SHUTOFF CONTROL 9. EMERGENCY LANDING GEAR HANDPUMP 59. POWER CONTROL LEVER 60. MANUAL OVERRIDE FUEL CONTROL 61. TRlM AND F U P SYSTEM INTERRUPT SWITCH1 62. TRIPLE TRlM INDICATOR 63 LANDING GEAR SELECTOR. POSIT ION INDlCAl

AN0 WNWNG SILENCER 64. PILOTS LOWER RIGHT SWITCH PWEL 65. PILOTS LOWER LEFT SWITCH PANEL 68. RUDOER PEDALS 67 RUDMR PEDAL AOJUSTMENT HANDLE 68 M I C P W E JACKS 69. GANG BAR M. UTILITY LIGHT 71. COPILOTS LOWER R W T PANEL

Figure 7-17. Typical Cockpit Layout (Sheet 2 of 2)


Report No: 01973-001 7-107

SECTION 7 ?PILATUSS AIRPLANE AND SYSTEMS OESCRIPTION ?c XI1

CENTRAL ADVISORY AND WARNING SYSTEM (CAWS)

GENERAL

Refer to Figure 7-18, Central Advisory and Warning System (CAWS) Annunciator Panel, for more informat~on on the warning/caution/advisory lights.

DESCRIPTION

The Central Advisory and Warning System (CAWS) annunciator panel integrates the control and dlsplay functions of aircraft systems status into a single unit. The CAWS annunciator panel contains lights indicating warning, caution, and advisory conditions. The CAWS annunciator panel is located at the lower center section of the instrument panel.

A warning light is red and indicates a condition that requires an immediate corrective action by the pilot.

A caution light is amber and indicates a condition that requires a pilots attention but not an immediate reaction.

Advisory lights are green and blue. A blue advisory light indicates an operating system that has high electrical power requirements or automatic system.

Red MASTER WARN and amber MASTER CAUTION lights are positioned directly in front of the pilot and copilot, above the EADl on the instrument panel. They alert the crew to changes in status on the CAWS annunciator panel. Any condition that causes a red or amber CAWS light to come on also causes the applicable MASTER WARN or MASTER CAUTION light to come on. An audible tone will sound through the overhead speaker andlor headset(s) anytime a MASTER WARN or MASTER CAUTION light comes on. Pushing the applicable MASTER WARN or MASTER CAUTiON light will extinguish that light. The CAWS warning or caution light that triggered the MASTER WARN or MASTER CAUTION light will remain on.

The individual lamps of the annunciators can be checked by pressing the LAMP test switch on the System Test Panel located on the left cockpit sidewall.

The CAWS has a continuous Built In Test that checks the CAWS function. When an error in the CAWS is detected, the MASTER CAUTION light will blink and the CAWS can be reset by pressing the MASTER CAUTION.

Report No: 01973-001 7-108

Issued: June 10, 1994 Revision 4: November 24, 1995

=PLATUS= SECTION 7 -PC XII' AIRPLANE AND SYSTEMS DESCRIPTION

SYSTEM TEST PANEL \

.- Figure 7-18. Central Advisory and Warning System (CAWS) Annunciator Panel

(Sheet 1 of 5)

Issued: June 10,1994 Revision 4: November 24,1995

Report No: 01 973-001 7-109

SECTION 7 SPILATUS" AIRPLANE AND SYSTEMS DESCRIPTION PC XII-

Figure 7-18. Central Advisory and Warning System (CAWS) Annunciator Panel (Sheet 2 of 5)

Index

1

2

3

4

5

6

7

8

9

10

11

12

13

Report No: 01 973-001 7-110


Description

Indicates main entry door and/or handle is not locked.

indicates cargo door and/or handle is not locked.

Indicates cabin altitude above 10.700 R. or if the maximum pressure differential is exceeded.

Indicates a disparity between LH and RH AIWGND switch inputs to the Stick Pusher computers.

Indicates propeller has gone to a low pitch (below minimum in-flight pitch) with aircraft not on the ground.

Indicates autopilot and/or auto trim failure.

Indicates voltage of the Bat, Gen 1, or Gen 2 busses less than 18 VDC.

Indicates avionic busbar 1 or 2 voltage less than 18 VDC.

indicates stabilizer trim is unsafe for takeoff (on ground only).

Indicates low engine oil quantity (engine not running).

Indicates overtemperature condition andlor possible engine fire.

Indicates autopilot pitch and aileron servos disengaged.

Indicates that generator 1 is off-line.

Nomenclature

PASS DOOR

CAR DOOR

CAB PRESS

AIWGND

PROP LOW P

Alp TRIM

ESNTL BUS

AV BUS

STAB TRIM

OIL QTY

ENG FIRE

A/P DISENG

GEN 1 OFF

Color

Red

Red

Red

Red

Red

Red

Red

Red

Red

Red

Red

Amber

Amber



Issued: June 10,1994 Revision 10: September 1,2000 ,

I

Report No: 01 973-001 7-1 1 1

Index

14

15

16

17

18

19

20

21

22

23

24

25

26

Color

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Amber

Nomenclature

GEN 2 OFF

BUS TIE

PUSHER

FIRE DETECT

PUSHER ICE MODE

BAT OFF

INVERTER

BAT HOT

FLAPS

CHIP

N ESNK BUS

L FUEL LOW

FUEL PRESS

-

Descrlption

Indicates that generator 2 is off-line.

Indicates generator 1 bus tie isolation relay is open.

Indicates Stall Warning/Stick Pusher System malfunction.

Indicates a malfunction in the engine fire detection circuit.

Indicates that the pusher computer is set to ICE mode.

- Indicates the battery is off-line.

indicates inverter output less than 20 VAC.

Indicates battery over-temperature or temperature sensor disconnected (inoperative on aircraft with an optional lead acid battery installed)

Indicates an asymmetrii condition, motor over temperature, or computer failure.

Indicates metal particles in the engine oil system.

Indicates non essential busbar voRage less than 18 VDC.

Indicates fuel quantity in left wing tank has reached 20 US gal (75 liters).

lndicates fuel system pressure is less then 2 psi (0.14 bar). Light goes off when fuel system pressure is greater than 3.5 psi (0.24 bar).

SECTION 7 WPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12


Index

27

28

29

30

31

32

33

34

35

36

37

Report No: 01973-001 7-1 12

Issued: June 10, 1994 Revision 1 1 : March 1,2003

Nomenclature

HYDR

ECS

AOA DE ICE

R FUEL LOW

L FUEL PUMP

PASS OXY

DE ICE

INERT SEP

PROBES DE ICE

R FUEL PUMP

IGNITION

Color

Amber

Amber

Amber

Amber

Green

Green

Amber

Green

Green

Green

Blue

Description

Continuously illuminated in flight indicates low hydraulic pressure.

Continuously illuminated on the ground indicates that the hydraulic pump cycled too often during flight and requires maintenance.

Indicates Environmental Control System malfunction. ---- Indicates AOA deice malfunction or DE ICING PROBES switch set to OFF (3 minute delay).

Indicates fuel quantity in right wing tank has reached 20 US gal (75 liters).

Indicates left fuel boost pump is operating.

Indicates adequate pressure of oxygen to the passenger masks

Indicates a deicing malfunction with the propeller or surface deice.

On when door is fully open.

Indicates probe deicing is operating.

Indicates right fuel boost pump is operating.

Indicates that engine ignition is operating.

=PLATUSE SECTION 7 ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Figure 7-1 8. Central Advisory and Warning System (CAWS) Annunciator Panel (Sheet 5 of 5)


Index

38

39

40

4 1

42

-

Report No: 01 973-001 7-1 13

Color

Green

Green

Green

Blue

Blue

Nomenclature

DE ICE BOOTS

WSHLD HEAT

PROP DE ICE

COOL

AIP TRIM

Description

On steady with boots deice operating and pressure sequence correct. Blinks with continuous amber DE ICE after pressure sequence failure.

Indicates windshield heat is operating.

On steady with prop deice operating. Blinks with continuous amber DE ICE after system failure.

Indicates vapor cycle cooling system operating (optional).

Indicates autopilot pitch trim activation (autotrim function) (MSN 101 -120). lndicates autopilot trim activation (autotrim function) (MSN 121-999).

SECTION 7 =PILATUSF AIRPLANE AND SYSTEMS DESCRIPTION -PC %I1 y 4

PlTOT STATIC SYSTEM

GENERAL <

4 The pilot static system obtains dynamic and static pressure to operate instruments for flight inlormation.

t

DESCRIPTION

P~tot pressure, for the airspeed indicator is sensed by a heated pilot head which is installed on the boltom of the right wing and is carried through lines within the wing and fuselage to the gauge on the instrument panel. Static pressure for the airspeed, encoding altimeter, vertical speed, as well as the cabin altimeter differential pressure indicator, is sensed by two static ports, one each side of the rear fuselage aft of the rear pressure bulkhead. They connect to a single line lead~ng to the instruments. The dual pickups balance out the differences in static pressure caused by slight side slips or skids. Also, a tapping is taken from the static line to supply the cabin altimeter differential switch. From the pitot and static lines, tappings are taken to supply the autopilot air data computer.

If one or more of the pitot static instruments malfunction, the systems should be checked for dirt, leaks or moisture. The holes in the sensors for pitot and static pressures must be fully open and free trom blockage. Blocked sensor holes will give erratic or zero readings on the instruments. The static line may be drained by a valve located fuselage bottom right side, aft of the rear pressure bulkhead. The pitot system can be drained by a valve located in the right bottom wing root.

The heaters lor the pilot head and static ports are controlled by the DE ICING PROBES switch located on the pilot's lower right swrtch panel. Electrical power lor healing and CAWS indication IS supplied through the PlTOTlSTATlC DE ICE and PlTOT DE ICE circuit breakers on the I generalor 2 bus. Refer to the Probe Deice Schematic in Figure 7-19.

4


Load current sensing of the pitot and static heating circuits is indicated by the illumination of the CAWS annunciator PROBES DE ICE (green). It a malfunction is sensed, the PROBES DE ICE annunciator w~ll go OH.

Report No: 01 973-001 7-1 14


4

=PlLATUSF SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

BATTERYBUS

-FROBES

GENERATOR 2 BUS I

II PITOT PRESSURE

STAT lC PRESSURE

Figure 7-19. Typical Pitot and Static Systems


Report No: 01 973-001 7-115

SECTION 7 EPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

STALL WARNING 1 STICK PUSHER SYSTEM

GENERAL

The airplane is equipped with a stick shaker-pusher system to improve aircraft handling in the low speed flight regime by preventing the airplane from inadvertently entering a stall condition. The stick shaker-pusher system contains two Angle-of-Attack (AOA) sensors, two computers, a single stick shaker, a single aural warning device and a single stick pusher. The two computers are connected in such a way that either computer can, independently, provide stall warning (stick shaker and aural warning) but both computers are required to actuate the stkk pusher.

DESCRIPTION

Refer to Figure 7-20, Stall WarninglStick Pusher System, for system operation.

The left and right Stick Pusher Computers are each provided power from the Battery and Generator 1 bus. Each computer receives inputs from its respective AOA vane and AIWGND switch. Both computers receive inputs from the engine torque, flap position, and self test. From these various inputs, each computer independently determines the 'Defined Angle of Attack" for stall warning (aural stall warning and stick shaker activation), stick pusher activation, and stick pusher disengagement following an actual push.

The stick pusher, shaker, and the aural stall warning are disabled on the ground through the AIRIGND switches, except for the self test function. The stick pusher is inhibited for 5 seconds after lift-OM. The shaker and the aural stall warning are operative immediately atter lift-off.

The stick pusher actuator has a built-in g-switch which inhibits the stick-pusher when the airplane's normal acceleration becomes less than 0.5 g. The output torque of the stkk-pusher actuator is electronically-limited to have a force of 60 to 65 Ibf on the control .wheel. A slip- clutch on the stick-pusher capstan allows control on the elevator with a force of 85 to 90 Ibf on the control wheel, in the event of stick-pusher jam. The force on the control wheel is defined when the longitudinal control is pulled to Y4 of its travel. This allows the pilot or copilot to override the stick-pusher in the instance of an inadvertent operation.

I Each outboard control wheel horn is equipped with a PUSHER INTR push switch providing a means to quickly disengage the stick pusher actuator in the event of an inadvertent operation.

When operated in PUSHER ICE MODE (to provide protection in icing conditions), all the shaker and pusher actuating points are shifted down by 8" AOA. The pusher ICE mode is set when the propeller de-icing system is switched ON and the inertial separator is set to OPEN. When both pusher computers are set in ICE mode, the amber CAWS PUSHER ICE MODE caution is activated. An aural gong will sound. If only one computer is set in ICE mode, or if no computer is set in ICE mode while conditions for ICE mode are present, the amber PUSHER caution is activated. An aural gong will sound.

Report No: 01973-001 7-116


=PLATUSI SECTION 7 -PC XII- AIRPLANE AND SYSTEMS DESCRIPTION

The system is provided with a self test function that can be activated at any time by pressing and holding the PUSHER switch located on the Test Panel. The amber PUSHER and amber PUSHER ICE MODE annunciator will remain illuminated until the self test is passed.

The system must be tested when the airplane is on the ground before takeoff. With the engine operating at a minimum of 5 psi torque, set the flaps to 15" and press and hold the PUSHER test switch. The following sequence will be observed; PUSHER ICE MODE caution, aural stall 1 waming with stick shaker for 2 seconds followed by a 1 second pause, aural stall waming with stick shaker for 2 seconds followed by a 1 second pause, aural stall waming with stick shakers, stick pusher and PUSHER ICE MODE caution until test switch is released. During the pusher operation, check for proper pusher interruption when each DISC switch is pressed. If the test switch is pushed and the test sequence does not occur and/or the PUSHER annunciator remains illuminated, the system has failed the self test and further flight before maintenance is not approved. If the test switch is pushed without the engine operating above 5 psi torque and the flaps are not set to IS0, the PUSHER annunciator will remain illuminated with the aural warning and the test sequence will not occur.

The system function may be tested in the air anytime the engine is operating with the flaps at any setting. Press and hold the test switch and observe the following sequence; PUSHER ICE MODE caution, aural stall waming with stick shaker for 2 seconds followed by a 1 second pause, and the aural stall warning with stick shaker for 2 seconds. The pusher will not activate when the system is tested inflight. If the test switch is pushed and the test sequence does not occur and/or the PUSHER annunciator remains illuminated, the system has failed the self test.

STALLS MUST BE AVOIDED WHEN THE STICK PUSHER IS INOPERATIVE. EXCESSIVE WING DROP AND ALTITUDE LOSS MAY RESULT DURING STALL WITH FLAPS DOWN AND/OR WHEN POWER IS APPLIED.

The AOA vanes and mounting plates are electrically heated by internal heating elements. AOA vane and mounting plate heat is controlled by the PROBES DE ICE switch located on the pilot's lower right switch panel. Refer to Figure 7-20, Stall WaminglStick Pusher System for system schematic.

Issued: June 10,1994 i Revision 4: November 24.1995

Report No: 01 973-001 7-1 17

SECTION 7 =PLATUS= AIRPLANE AND SYSTEMS DESCRIPTION -?c XII- ,

OPERATION

The vane attached to the AOA probe aligns itself with the relative airflow. As it moves, it positions a wiper unit in the probe. This wiper unit adjusts the electrical output to its respective

'

pusher computer. As the airplane approaches the artificial stall (5 to 10 knots before pusher actuator), the stick shaker and the aural stall warning will activate when one of the AOA pusher compulers senses the defined angle of attack for stall warninglstick shaker activation. If the stall warnings are ignored and the approach to stall is continued, the stick pusher will activate , when both AOA pusher computers sense the defined angle of attack for stick pusher activation. The stick shaker and aural stall warning remain active during pusher operation.

Pusher operation will be stopped when either AOA computer senses an angle of attack lower 1

than the angle of attack required to active the pusher or when the airplane acceleration is less . than 0.5 g.

I Activation of the stick shaker disengages the autopilot if engaged, in order to give full authority I

to a possible stick pusher activation. The autopilot can be manually reconnected alter the angle of attack is reduced and the stick shaker has ceased operation.

IF ACCELERATED STALLS ARE PERFORMED IN THE LANDING CONFIGURATION WITH HIGH POWER AND SIDESLIP, A RAPID PITCH-DOWN MAY RESULT WITH AN ALTITUDE LOSS OF UP TO 500 FEET.

Activation of the stick shaker disengages the autopilot if engaged, in order to give full authority to a possible stick pusher activation.

Rep011 No: 01973-001 7-118


mPILANSE SECTION 7 P(12 AIRPLANE AND SYSTEMS DESCRIPTION

INDlCATlON I WARNlNG

A digital serial output, from the left hand computer, provides the data for the FASTISLOW pointer on the EFlS EADI. (Refer to EFlS section for more information). I In the instance of disparity between the LH and RH weight-on-wheels inputs, the Central Advisory and Warning System annunciator AIWGND and PUSHER are activated to warn the pilot of the malfunction.

The stick pusher computers have an internal-fault monitoring system which will illuminate the I CAWS PUSHER annunciator when one of the following events occur:

a built-in test failure a push signal from only one computer that is longer than 3 seconds no output toque during a push if either of the pilot or copilot DISC switches is pressed ii the aircraft normal acceleration is below 0.5 g for longer than 3 seconds.

A malfunction in either pusher computer activates the Central Advisory and Warning System annunciator PUSHER to warn the pilot about a system malfunction and the pusher becoming inoperative.

The stkk shaker and aural stall warning devices may still be operational i f the stick pusher is inoperative.

The CAWS annunciator PROBES DE ICE will illuminate when the DE ICING PROBES switch is set to ON. The CAWS annunciator AOA DE ICE will illuminate when a malfunction is sensed in the AOA vane or mounting plate heater circuits (current sensing).

The CAWS PUSHER ICE MODE caution will illuminate when the propeller de-ice system is set ON and the Inertial separator is set OPEN. In the ICE mode, the shaker and pusher activation points are coming 8" earlier than in the NORMAL mode and the FASTISLOW pointer is set for a 15" flap landing.

If the Flap Control and Warning Unl detects a flap asymmetry or a twist and the flap angle is greater than 2" it will make the CAWS FLAP caution come on and will send a sisnal to the stick pusher computers. This will set the stick pusher computers to a 'safe' mode irrespective of the actual flap position. The 'safe' mode will make the stick pusher operate at approximately 5 kts higher airspeed for the failed flap position. The setting of the stick pusher computers to the 'safe' mode will also make the CAWS PUSHER caution come on 10 seconds after the FLAPS caution, to annunciate the condition.

Issued: June 1 0, 1994 Revision 10: September 1,2000

Report No: 01 973-001 7-119

SECTION 7 WPILATUSW , AIRPLANE AND SYSTEMS DESCRIPTION PCP 4

3

1

,

I

1

<

1

1

ON OPEN

I I COPILOT

INTR

STffiK PUSHER SWITCH

ACTUATOR BATTERY

BUS

I I NOTE: "OR" SWITCH - ACTIVATED BY

PUSHER EITHER COMPUTER - SYS "AND" SWITCH - ACTIVATED BY - MHCOMPVTE*I 1

Report No: 01973-001 7- 120

Figure 7-20. Stall WarningIStick Pusher System (Sheet 1 of 3)

Issued: June 10,1994 Revision 1 1 : March 1,2003 ,


BATTERY 8VS NON ESSEMUL BUS

POWERLINE

POWEWE

CB 418 CONTROL CONTROL

t I

LHAOAPLATE FAIL FAIL RH AOA PLATE CONTROUER CONTROLLER

1 1 i LH M A PLATE RH AOA PLATE

WATER HEATER

--------- -------- - LHAOAVANE RH ADA VANE

HEATER HEATER - CAWS

I 1 W C

31 LHAOALOAD . FAIL FAIL RH AOA LOAD

SENSOR SENSOR

2w DC

NOTE MSN 101.280 THE GENERATOR 2 RELAY MUST BE ENERGIZED FOR THE AOA PLATE RELAY TO OPERATE

Figure 7-20. Stall WarningIStick Pusher System (Sheet 2 of 3)


Report No: 01973-001 7-121


SYSTEM TEST PANEL

Repon No: 01973-00 1 7-122

Figure 7-20. Stall ~arkn~lSlick Pusher System (Sheet 3 of 3)

issued: June 10,1994 Revision 9: September 1,1999

'CPAATUS F SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

PNEUMATIC WING DEICE SYSTEM

GENERAL

Inflatable neoprene boots are installed on the leading edges of the wings and horizontal tail surfaces. Their purpose is to inflate and dispense any ice which may accrete on their surface during flight in atmospheric icing conditions. When not in use, the boots have a vacuum applied to prevent partial inflation while in flight.

I

DESCRIPTION

The airplane is equipped with inflatable pneumatic deicing boots fixed to the leading edges of the wings ( two boots per wing- inboard and outboard) and the horizontal stabilizer. Air bled from the 3rd stage of the engine compressor section, is routed to the regulator-reliever valve of nominal 14 psi regulating pressure, then through a water separator to the ejector flow control valves. These valves, which are solenoid-operated, port air pressure to the deicing boots in a prescribed sequence: - first to the horizontal stabilizer deicer, then to the lower portion of the inboard wing deicers, the upper portion, the lower portion of the outboard wing deicers, and finally the upper portion. Progression through this sequence is controlled by an electronic Timer/controller and monitored by low pressure sensing switches in each line, which are linked to annunciator lights in the CAWS.

When pressure is not being applied to the deicer boots a small airflow is allowed to pass through the ejector valves to impose a vacuum in the lines to the deicing boots. This provides a negative air pressure at the boots ensuring the airfoil contour is maintained.

The pneumatic deice boot consists of a smooth neoprene and fabric blanket containing small spanwise deicer tubes. Each wing deicer has two air connections - one for the tubes on the lower surface and one for the tubes on the upper surface. The smaller boots on the horizontal stabilizer have one connection only.

The deicer boots are cement-bonded to the airfoil leading edges and have tapered edges to conform smoothly with the wing contour. The deicer boots are approximately 0.075 inches thick and have a 'neoprene' surface.

The water separator is located upstream of the ejector control valves. Its function is to remove any condensation from the system and consists simply of a set of vanes which introduce a rotational swirl to the air that removes entrained water through centrifugal forces. A drain connection is fitted to the bottom of the housing to vent the moisture overboard.

The pressure-reliever valve consists of a spring and poppet valve which, at the required pressure, will open to allow air to pass from the inlet to the outlet port. The nominal regulating pressure is 14 psi. It also has an integral relief valve relieving at 18 psi.

Issued: June 10,1994 Revision 4: November 24, 1995

Report No: 01973-001 7-123

SECTION 7 =PLATUSE AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

OPERATION

Refer to Figure 7-21, Pneumatic Wing Deice System.

In the OFF mode the system applies a continuous vacuum to the de-ice boots while the engine is running. The system is initiated by moving the switch on the DE-ICING control panel (pilots lower right switch panel) labelled BOOTS to either 3 min or lmin position. This activates the timer. When activated the timer actuates each ejector flow control valve (EFCV) in the prescribed sequence, for eight seconds. The time to inflate and deflate all of the de-icer units is thus 40 seconds. There is then a dwell period of 20 seconds (if the 'one minute cycle' has been selected) or of 140 seconds (if the 'three minute cycle' has been selected) before the inflation sequence is repeated.

Pressurization of each de-icer will cause the pressure switch to close, maintaining constant illumination of the green CAWS caption 'DE ICE BOOTS". Operation of the wing boots can also be obsewed directly during ground checkout or from the airplane cabin. If the control System is de-activated during a de-icing cycle, the cycle will be completed prior to system shut-down.

( CAUTION I OPERATION OF THE PNEUMATIC WING DEICE SYSTEM IN AMBIENT TEMPERATURES BELOW -40' C OR ABOVE 40' C MAY CAUSE PERMANENT DAMAGE TO THE DEICER BOOTS.


Proper system function is indicated by constant illumination of the CAWS green DE ICE BOOTS annunciator which takes an input from pressure switches in each pneumatic supply line at the inlet to each boot. Shouid the supply pressure during an inflation sequence fall below a prescribed value min 4 0 psi, indicating failure, then the CAWS DE ICE BOOTS caption will blink and the amber caption DE ICE is illuminated. An aural gong will sound.

After failure of the de-icing boots, the aircrew should prepare for departure of icing conditions

I as soon as possible.

Report No: 01973-001 7-124

Issued: June 10,1994 Revision 5: May 10, 1996 ,


L J J

PILOTS LOWER RIOHT PANEL

Figure 7-21. Deicing System (Sheet 1 of 2)


Report No: 01973-001 7-125

SECTION 7 FPILATUS W AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

GEN 2 BAT OEN 1 BUS BUS BUS

$

BOOTS CAWS DE lCE 2

DEICING 0 BOOTS SWITCH

DE ICE TIMER CAWS

3UINI1 MU4

1

1

4 1

EJECTOR VALE PRESSWE SWITCH

HORIZONTAL STABILIZER 4 v

i

I m

EJECTOR VALVE MWRE SWITCH

I I Figure 7-21. Deicing System

(Sheet 2 of 2)

Report No: 01973-001 Issued: June 10,1994

7-126 Revision 5: May 10,1996

OUTBOARD UPPER

-

INBOARD LOWER

v

4 1 v

I I I m

I

EJECTOR VALVE PAESSWE M C H

INBOARD UPPER

i

4 v

I m I m

EJECTOR VALVE PRESSURE SWITCH

OUTBOARD LOWER

r v

I I I m

EJECTOR VALVE PRESSLIRE SWITCH


COMFORT FEATURES

GENERAL

Passenger comfort is provided for by an ECS and a pressurization system. Additional comfort can be provided on the ground with the vapor cycle cooling system (when installed). The fans installed at the rear of the cabin can be used to increase the general air circulation around the cabin. The switches for the fans are on the pilot's lower left switch panel.

CABIN FEATURES

GENERAL

The PC-12 has a large cabin that offers a flexible interior configuration for passenger and cargo loading. There are two basic cabin configurations, a Corporate Commuter and an Executive interior. Variations to the two basic configurations are continually being developed. refer to Section 2 for the variations that have been approved. See Section 6 for passenger seat locations, combi conversions and cargo loading information.

Divider walls are installed behind the pilot and copilot seats and a curtain or door fits between the walls to form a division between the cockpit and cabin.

A fire extinguisher is located on the forward side of the cabin divider behind the copilot seat. Full operating instructions are given on the side of the extinguisher. I SB 25-008 modifies the passengerlcrew door to permit a thermal blanket to be fitted on the inside of the door. The blanket can be easily fitted and removed (instructions are on the blanket), its color matches the color of the interior. It is installed on the inside of the door when closed for flight. It covers the steps on the door to help maintain the cabin temperature during long high flights. The door operating placards that are on the inside ol the door are duplicated in the same format and attached to the outside of the blanket.

CORPORATE COMMUTER INTERIOR

The standard Corporate Commuter Interior consists of two crew seats plus seating for up to nine passengers. The baggage compartment is situated at the rear of the cabin and a baggage net must be installed at frame 34 when baggage is stowed. An optional three seat bench can be installed at the rear of the cabin in place of seats 7.8 and 9. This provides a larger area for baggage stowage which is then secured with the larger baggage net (Post SB 25-010). An optional bulkhead and curtain assembly (Post SB 25-017) can be installed at tram8 32 in front of the larger baggage net.

EXECUTIVE INTERIOR

The standard executive interior aircraft consists of two crew seats plus executive seating for


Report No: 01973-001 7-127

SECTION 7 iEPlLATUSP AIRPLANE AND SYSTEMS DESCRIPTION PC12

EXECUTIVE INTERIOR 1

The standard executive interior aircraft consists of two crew seats plus executive seating for six passengers. The two forward passenger seats 1 and 2 face rearwards and the remainder face forwards. The baggage compartment is situated at the rear of the cabin and a bagsage , net must be installed at frame 34 when baggage is stowed.

* Folding tables installed in the cabin sidewalls extend between the seats. Ashtrays, cupholders, table and overhead lighting switches are provided in the sidewall armrests adjacent to each seat. Individual reading lights and air outlets are installed in the headliner panel above each , seat position.

A toilet compartment is installed in the front right hand side of the aircraft. The forward wall of , the toilet compartment forms the cabin divider. Left and right storage cabinets are installed, the left cabinet fits against a small divider behind the passenger door and the right cabinet fits f

against the toilet compartment rear wall.

An optional wardrobe can be installed in place of the toilet compartment. The wardrobe is used . to hang up to six coats and stow small items of luggage. A safety net must be installed inside , the wardrobe above the stowed luggage. A small drawer is installed at the bottom of the wardrobe.

An optional stereo/CD unit can be installed in the right hand cabinet. Headphone jack points are provided adjacent to each seat. The antenna for the stereoJCD unit is installed under the upper dorsal fairing. 4

Passenger information no smokingJfasten seat belt illuminated signs are installed on the rear of left cabin divider and above the baggage compartment. The signs are turned on and off by the pilot using the switch installed in the electrical overhead panel.

An optional three seat bench can be installed in place of the rear cabin seats. This can provMe 4

a larger area for baggage stowage which is then secured with the larger baggage net (Post SB 25-010). An optional bulkhead and curtain assembly (Post SB 25-017) can be installed at frame 32 in front of the larger baggage net. i

COMBllCARGO INTERIOR

A Combi or a full cargo interior can be made by the removal of passenger seats from both the Corporate Commuter and Executive Interior aircraft. Cargo net attachment points are installed in the cabin walls at frame positions 24 and 27. Baggage net attachment points are installed at frame 34. Cargo restraining nets can be installed at the attachment points and allow lightweight cargo to be loaded without being secured with tie-down straps. A cargo securing kit contains the necessary items for the securing of heavyweight cargo.

Report No: 01973-001 7-128


aPLATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

GENERAL AVIONICS INSTALLATION

AUDIO PANEL

GENERAL

The audio panel provides the capability to select transmitter and audio either though the cockpit speakers or headphones. Crew intercom and cabin PA are also available. A secondary audio microphone amplifier is provided within the audio panel and is selected by setting the ALT AUDIO switch, located on the pilot's lower left panel, to the ALT position. Aural warnings can be heard in the headphone audio and loudspeakers.

An optional second audio control panel can be installed in the copilot's instrument panel. The operating mode is the same as that of the pilot's. Each audio panel has access to the transmitters and the intercom lines. A summing amplifier is installed to match certain warning audio signals for the two audii panels.

AVIONICS

Refer to Figure 7-22, Avionics.

The standard avionics suite includes two communication/ navigation radios, a transponder, a DME, an ADF, and a marker beacon. The radios are panel mounted, except the DME, and are centrally located on the instrument panel. Provisions for additional optional equipment are included in the standard avionics harness.

The Attitude and Heading Reference System (AHRS). Electronic Flight Instrument System (EFIS) and auto flight system are described on the following pages.

Optional SB 23-003 introduces commlnav transceivers with 8.33kHz channel spacing, which is mandatory for aircraft that operate at high altitudes in European airspace.

Refer to the appropriate manufacturer's pilot's guide for information on individual avionics equipment.

Refer to Section 9, Supplements, for information concerning limitations and operating instructions for optional avionics installations.

Issued: June 10,1994 Revislon 10: September 1,2000

Report No: 01973-001 7- 129

SECTION 7 €PILATUSP AIRPLANE AND SYSTEMS DESCRIPTION PC12

A-iZ p TEMP p~~ TRIM TRIM

STATIC KEY PITOT - DATABUS

Figure 7-22. Typical Avionics

Repolt No: 01973-001 7- 130


3PILATUSW SECTION 7 PC12 AIRPLANE AND SYSTEMS DESCRIPTION

AlTlTUDE AND HEADING REFERENCE SYSTEM

GENERAL

The Attitude and Heading Reference System (AHRS) supplies pitch, roll, heading and turn rate information to h e EFlS EADl and EHSl and the Radio Magnetic Indicator (RMI). The system maintains accurate indications in all aircraft attitudes. The AHRS is installed in the avionics bay and the magnetic sensing unit is within the right wing.

DESCRIPTION

The computer processes information from a Sensor Module containing a Fiber Optic Gyro (FOG) a two-axis Electrolytic Tilt Sensor, and from a Flux valve. The system requires 28 V DC power from either the AVIONICS 1 or AVIONICS 2 bus AHRS circuit breaker and 26 V AC from the inverter through the AHRS HDG (for RMi) and AHRS ATT (for Wx Radar) circuit breakers. The DGISLAVE switch, located on the pilot's lower left switch panel, enables full gyro operation with CW and CCW slave capability.

OPERATION

When one of the AVIONICS switches is turned ON, the system automatically goes through an alignment process. During the first 15 seconds, phase 1, self-test functions and coarse alignment are performed. The duration of alignment phase 2 depends upon airplane movement. Under static conditions, no ground movement or during straight and level flight. phase 2 align'ment is completed after a further 15 seconds (30 seconds after power application) and attitude and heading references are valid. I f excessive motion is detected during alignment, the AHRS switches to the moving alignment sequence and completion of the phase 2 alignment will take additional time. On completion of the alignment process, the ATTITUDE FAIL and HDG flags will disappear from view on the EFlS displays and the RMI.

During normal operation, the aircraft has complete freedom to maneuver in all axis without risk of the AHRS toppling. When in straight and level flight, the system continually re-datum's the indications, at a slow rate, to earth vertical and magnetic North.

During continuous orbiting maneuvering with LCR-92 AHRS the attitude and heading will not re-datum and after approximately 20 to 30 minutes the attitude and heading information is unreliable and warning flags are set. If the warning flags are set, a reset can be performed by flying straight and level for 10 to 60 seconds. If the AHRS does not recover after 1 minute of straight and level flight, a hardware failure has to be suspected. To avoid this situation and to maintain AHRS accuracy, fly the aircraft straight and level for 1 minute after each 15 minutes of continuous maneuvering.

Two switches on the pilot's lower left panel control the source of heading information displayed. The DGISLAVE switch selects a free gyro mode when set to DG or AHRS heading information when the switch Is set to SLAVE. The CWICCW switch slews the heading card

Issued: June 10, 1994 Revision 1 1 : March 1,2003

Report No: 01 973-001 7-131

SECTION 7 XPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION PC12

clockwise (CW) or counterclockwise (CCW) when the DGISLAVE switch is set to DG. For normal operation set the DGISLAVE switch to SLAVE. DG mode is recommended when magnetic abnormality is detected and compass re-datum is required.

BUILT-IN TEST EQUIPMENT (BITE)

The AHRS provides attitude, rate of turn and heading information regardless of aircraft maneuver. BITE detects any system errors that occur and causes the appropriate flag to appear on the EADl or EHSl indicating a system failure.

INDICATION / WARNING

During normal operation the following faults may be detected and displayed:

AlTITUDE FAIL (red) Appears on the EADl when attitude information is unreliable and horizon display blanks out.

HDG FAIL (red) Appears in the EHSl and the RMI when heading information Is unreliable.

I CAUTION 1

AN ENGINE RELIGHT CAN GIVE A TRANSIENT DC VOLTAGE DROP, CAUSING THE AHRS TO LOSE ITS REFERENCE, FORCING THE SYSTEM TO REALIGN WHEN THE VOLTAGE IS RESTORED.

SECOND AHRS INSTALLATION (OPTIONAL)

An optional second AHRS provides an alternative source of aircraft attitude, rate of turn and heading information.

The system comprises a second computer installed in the avionics bay, a second magnetic sensing unit installed in the right wing and a selector push switch installed on the pilots instrument panel.

The second AHRS information can be selected by pressing the push switch to AHRS 2. Pressing the switch to AHRS 2 also transfers the autopilot operation, the RMI heading source and the weather radar stabilization data to the second AHRS.

If the AHRS selection is changed the autopilot will automatically disconnect. It can be re- engaged manually.

YAW RATE SENSOR INSTALLATION (OPTIONAL)

An optional remote mounted sensor when Installed, provides angular rate information, displayed independent of AHRS function on the EFlS EHSl display unit(s).



YPILATUSI SECTION 7 P(12 AIRPLANE AND SYSTEMS DESCRIPTION

STANDBY ATTITUDE INDICATOR

GENERAL

A second attitude indicator gyro is located on the instrument panel to the left of the airspeed indicator.

DESCRIPTION

The attitude gyro provides a second source of pitch and roll information. This attitude gyro is a self contained vertical axis gyroscope, mounted in a pitch gimbal, which is mounted in a roll gimbal. The indicator is capable of operation through 360" of airplane pitch and roll displacement. The gyroscope incorporates stops on the inner (pitch) axis to limit pitch axis freedom to +I- 05" relative to the roll gimbal. Pitch range of 360" is accomplished using controlled precession when the pitch stops are contacted.

OPERATION

The attitude gyro operation is continuous whenever the Battery Bus is powered.


An OFF flag will appear when power is removed from the gyro.

EMERGENCY LOCATOR TRANSMITTER

DESCRIPTION

NARC0 ELT 91 0

The Emergency Locator Transmitter (ELT) is installed in the rear fuselage. It is connected to an antenna which is installed on the top of the fuselage below the dorsal fairing and has a battery pack that must be replaced after a specified time. The ELT transmits on the international distress frequencies of 121.5 and 243.0 MHz.

An optional remote control switch and monitoring indicator light panel is installed on the pilot's lower left instrument panel. Power to the remote control panel is provided from the BATTERY BUS through the ELT IN0 circuit breaker.

Post SB 25-020 and MSN 322 through 400. A remote control switch and monitoring indicator 1 light panel is installed on the pilot's lower left instrument panel. Power to the remote control panel is provided from the BAT DIRECT BUS through the ELTICLOCK circuit breaker.


Report No: 01973-001 7-133


I KANNAD 406 AF ELT (Post SB 25-022)

An alternative ELT is installed which will also transmit at 406 MHz. An ELT remote control panel is installed on the pilot's lower left instrument panel. On the panel there is a switch with the positions TEST I RESET. ARMED and ON, and an indicator light.

SB 25-029 installs a smart connector in the ELT wiring hamess. With an ELT 406 AF installed the smart connector is programmed with the aircraft identity data. If there is a change to the aircraft identity the smart connector must be re-programmed at an approved service center.

SB 25-030 installs a navigation interface module adjacent to the ELT In the rear fuselage. It has a 28 VDC power supply from the Battery Direct Bus, receives aircraft position information from the GPS and is connected to the ELT. The interface module also has an ONIOFF switch and an indicator. This navigation interface module can only be installed with an ELT 406 AF.

OPERATION

I NARC0 ELT 910

In the normal ARM mode the ELT is automatically operated at a specified g force by an internal g switch. The ELT will continuously transmit for up to 50 hours and can be stopped by setting the switch on the unit to OFF. The ELT can be operated in a manual mode by selecting the switch to ON. It will transmit on the same frequencies as the normal mode. Manual mode

I must only be used in an emergency or during a test. The optional remote control panel indicator light will flash when the ELT switch is set to ON.

I KANNAD 406 AF ELT (Post SB 25-022)

The ELT is installed in the aircraft with the switch at the ARM position this also makes the remote control panel active. For flight the remote control switch must be in the ARMEO position. In the ARMEO mode the ELT is automatically operated at a specified g force by an internal g switch. The ELT will continuously transmit at 121.5 and 243.0 MHz for up to 48 hours and it will also transmit a digital message at 406 MHz every 50 seconds for the first 24 hours. With the Post SB 25-030 navigation interface module installed the aircraft position is also transmitted as part of the digital message at 406 MHz.

In an emergency the remote switch can be selected to ON. The ELT will then immediately start the distress signal transmission. The red indicator will come on.

In the case of accidental transmission, the ELT can be reset by either selecting the remote switch to RESET or the switch on the ELT unit to OFF.

I The remote switch TEST position is used to check the battery voltage and transmission power of the ELT for maintenance purposes.

Report No: 01 973-001 7-134


'PILATUSS SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

ELECTRONIC FLIGHT INSTRUMENTATION SYSTEM EFS 40150

GENERAL

The Bendix King Electronic Flight Instrumentation System ( EFlS ) is an electronic navigation display system. The system receives data from the AHRS, angle of attack, and navigation systems. The data is displayed to the pilot on two multimode color displays.

The following information is applicable to both pilot and copilot (if installed) EFlS installations and the EFlS MFD (if installed).

DESCRIPTION

The EFIS symbol generator receives information from flight systems either directly or through data converter adapters. Data such as aircraft attitude, heading, rate of turn and respective valid flags are received from the Attitude and Heading Reference System ( AHRS ), and angle of attack information are received from the SPWU. Navigational data from the various systems, either directly or through data converters is also received by the symbol generator. The information is processed and displayed graphically on two multicolored Display Units (DU). A reversionary mode select switch is dedicated to EFlS 1, EFlS 2 (if installed) and EFlS MFD (if installed) in order to allow composite mode selection if one tube or a partial symbol generator failure occurs. This function allows the restoration of critical flight data.

The symbol generator monitors the system for faults using continuous Built In Test (BIT) and, if errors are detected, are displayed on the DU's. A self test is available by pushing the TEST REF switch which initiates a self test of the system the result of which is displayed on the DU's.

Two DU's are installed in the instrument panel. The DU's are electronically identical but differ physically, with a slip-skid inclinometer installed on the upper DU. During normal operation, the upper DU is used as an Electronic Attitude Director Indicator ( EADl ) and the lower DU is used as an Electronic Horizontal Situation Indicator ( EHSl ).

The EFlS control panel is installed in the center console. This control panel is for the selection of display mode, format and navigation data, brightness control of the display units, course pointer, and heading bug selection.

The EFlS is supplied with DC power from both the AVIONICS 1 and AVIONICS 2 bus. This prevents a total loss of the system in the event of a single DC bus failure. Similarly, system AC power can be provided from either of two switchable static inverters ( BAT or GEN ).


I Revision 3: September 29, 1995 Report No: 01 973-001

7- 135

SECTION 7 -PLATUS= AIRPLANE AND SYSTEMS DESCRIPTION PC XII-

ELECTRONIC AlTITUDE DIRECTOR INDICATOR

The Electronic Altitude Director Indicator ( EADl ) displays aircraft flight attitude from roll, pitch, yaw and heading data provided by the AHRS. Refer to Figure 7-23, EADI.

Airplane Symbol: An airplane symbol in the center of the display is used as a reference point.

Sky Pointer: The roll altitude scale is marked at 10, 20, 30, 45, and 60 degrees. Roll indication opposite direction of turn. Recovery from a turn is in the direction of the sky pointer.

Speed Pointer: The speed pointer on the LH side of the display is contrdled by the AOA system and is visible at all times. AN is displayed to indicate that pointer is referenced to me angular system.

A Glideslope ( GS ) scale and deviation pointer on the RH side and a runway symbol towards the boltom of the picture is displayed when an ILS frequency is selected and the EHSl course pointer is within 105' of the aircraft heading. Beyond 105" BC will be annunciated for backcourse sensing.

Marker Beacon: The marker beacons are annunciated at lhe boltom LH side of the display when a marker beacon signal is detected.

Rate of Turn Pointer: A scale and pointer at the boltom of the display provides rate of turn information based on yaw rate corrected for angle of bank.

Inclinometer: The inclinometer installed on lhe EADl presents aircraft slip and skid.

Heading Tape: Aircraft heading is shown on a heading tape at the sky/ground border.

Lateral Deviation: Lateral deviation scale and bar are for the primary navigation sensor.

Rising Runway: Present during ILS approach. The centerline represents the fly to command. The rising runway symbol will start increasing in size at 200 feet and continue to increase in size to 0 feet, if the radar altimeter is installed.

Radar Altimeter The radar altimeter display is composed of three elements; (if installed) radar altimeter (RA) height, decision height (DH) set and

decision height (OH) alert. In the upper right corner of the display, a field of white alphanumeric data provides radar

Report No: 01973-001 1 7-136

Issued: June 10, 1994 Revision 3: September 29. 1995

I i =PLATUSW SECTION 7 r -PC XII AIRPLANE AND SYSTEMS DESCRIPTION

HEU)Na TAPE

\

\ COURSE DEVUTlON

DISPLAY RATE OF

TURN SCAG


1

Figure 7-23. EADl

Report No: 01 973-001 7-1 37


altimeter height and annunciation. In the lower right comer of the display, a green alphanumeric field displays the selected OH. This field is active only when the DH set knob is active or when the radar altimeter is providing valid height data. To the right and above center of the display, a yellow DH alert is displayed in a black box outlined in yellow. When the radar altimeter height is equal to or less than the selected decision height, the OH annunciator will be displayed combined with an aural alert. When first activated, the DH annunciator will flash for 10 seconds. The OH symbol is a large yellow DH on a black background enclosed in a yellow box, located to the right of the pitch scale and above the horizon line when in normal attitude.

ELECTRONIC HORIZONTAL SITUATION INDICATOR

The Electronic Horizontal Situation Indicator (EHSI) can provide 360' compass HSI. 120" arc sectored HSI and navigation map displays. The type of display is selected on the EHSI control panel located on the center console. Refer to Figure 7-24, EHSI.

360" Compass Display

Normal Compass A 360q rotating white compass scale indicates the airplane Card Display: heading referenced to the white triangle heading index (lubber

line). The comoass scale is divided in 5' increments with the 10'.divisions approximately twice as tong as the 5' marks. Fixed 45' index marks are adjacent to ?he compass scale. Compass heading is referenced to magnetic north.

Symbolic Airplane: The symbolic airplane provides a visual reference of the airplane position in relationship to the deviation bar.

Navigation Source A vertical three letter alphanumeric readout, located on the left Annunciation: side of the display, indicates the navigation selecting system

selected as the primary navigation sensor.

Green annunciation indicates a NAV 1 system and yellow indicates a NAV 2 system. These cdor codes apply to the NAV source annunciator, CRS pointer and CDI, CRS line in MAP mode, CRS readout, distance, ground speed readout, and time to station.

NOTE

A failure to NAV 2 when LOC 2 is the primary nav sensor is indicated by the removal of the nav display and flagged with a red X. The primary nav sensor annunciator will revert to VOR 2.

, Report No: 01973-001 i 7-138

Issued: June 10,1994 Reviston 3: September 29, 1995


NORMAL COUPASS CARD MSPUY

HEADING SELECT BUG

SELECTED COURSE DISPUV

\ I / DISTANCE

GROUND SPEED AM) TIME

:OURSE POINTER 1

. . . ... .. .. . . NAVIGATION -

SOURCE ANNUNCIATION

i

BEARINQ POINTER MSPUY

b ' I I

COURSE DEVIATION DISPLAY

Figure 7-24. EHSl (Sheet 1 of 2)


BEARING POlNTER DISPLAY

Report No: 01 973-001 7-139 '

SECTION 7 =PLATUSS AIRPLANE AND SYSTEMS DESCRIPTION -PC %I1 4

SELECTED COURSE DISPLAY

lCRSl SPEED OR TIME AWlUNCUTOR

/

PRIMARY NAVIGATION

SOURCE ANNUNCIATION

SYSTEM ANNUNCIATOR

Blank

GLIDESLOPE MSRAV

losl

BEARING POINTEMME

B E A R M POINTER SELECT ANNUNCIATOR SELECT ANNUNCUTOR

DUE Blank DISTANCE MS%E

Figure 7-24. EHSl (Sheet 2 of 2)


PLATU US^ SECTION 7 PC XII- AIRPLANE AND SYSTEMS DESCRIPTION

Heading Select Bug: An orange heading bug is manually rotated around the compass scale by the heading select knob on the control panel. A digital readout of the selected heading is displayed in the lower right hand comer. Once set, the heading bug rotates with the compass card. The heading bug is used to indicate desired heading and provides selected heading reference for autopilot steering.

The auto sync feature allows the heading bug to be centered under the lubber line by pulling out on the heading select knob.

Course Pointer: The course pointer is rotated about the compass scale by the course select knob. Once set, the course pointer rotates with the compass card. It is used to indicate the desired navigation course to be flown.

The auto sync feature allows the course and digital course readout to slew to the direct course. to the selected navaid, or active waypoint.

In the upper left corner of the display, an alphanumeric readout of course pointer annunciates the letters CRS and indicates the selected navigation course in degrees.

Course Deviation The course deviation scale, bur white dots evenly spaced on Scale: the symbolic airplane, provides a reference for the course

deviation bar to indicate the centerline of the selected navigation or localizer course in relation to the symbolic airplane.

The following represents the lateral deviation scale for different navigation sources:

MODE ANGULAR (DEGREES) LINEAR (MILES) DEV DEV

VOR 1 dot 5.0 deg 2 dots 10.0 deg

ADF 1 dot 7.5 deg 2 dots 15.0 deg

GPS


1 dot 2.5 NM 2 dots 5.0 NM

Report No: 01973-001 7-141

SECTION 7 €PLATUS- AIRPLANE AND SYSTEMS DESCRIPTION ?C XI1

TotFrom Indicator: A white arrow head is displayed near the center of the EHSl (not illus. in with head pointing toward the course pointer (TO) or toward the Fig. 7-23) tail or the course pointer (FROM). It indicates the selected

course is to or from the station or waypoint. The TOIFROM indicator is not displayed unless a valid signal is received or during ILS operation.

Distance and The EHSl provides three distance displays; the upper right Ground Speed corner, lower left below the #I bearing pointer sensor Display: annunciator, and lower right below the 12 bearing pointer

sensor annunciator. In the upper right comer, an alphanumeric readout annunciates distance in nautical miles from the aircraft to selected DM€ station in VOR mode or to waypoint in GPS mode. Below the distance readout is an alphanumeric readout of the aircraft ground speed in knots or time to station in minutes and is selectable by pressing the TSTIREF button.

When the selected bearing pointer source has DME data associated with it, the distance information will be displayed below the bearing pointer source annunciator.

In the event that the DME station is out of range or not operational, or if tor any reason the DME receiver is operational but not providing computed data, the distance will be dashed in the original color. If the DME receiver is indicating an internal fault, is being tuned by another receiver, or is turned off, the distance will be dashed in red. When DME is flagged, the ground speed and time to station display is removed.

Bearing Pointer The rotating blue # I RMI Pointer points in the direction of the Display: selected sensor ground station or waypoint.

The rotating magenta #2 RMI Pointer points in the direction of the selected sensor ground station or waypoint.

No pointer will be displayed if a localizer frequency is selected for that sensor.

NOTE

If a selected bearing sensor ground station is out of range Or signals are not being received properly, the RMI pointer assigned to that bearing source is not displayed. If the selected bearing sensor has DME distance associated with it and is valid, the distance data will remain valid. Failure of the RMI source is flagged by a red X and the source letter appearing in red.

Report No: 01973-001 ( 7-142


SPL#TUSW SECTION 7 ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

Course Annunciation: Magnetic course is automatically displayed in VORROC, GPS. and ADF.

Glide Slope Display: If the selected primary sensor is an ILS, the vertical scale will appear on the right side when the selected inboard course is within 105 degrees of the aircraft heading. Beyond 105" BC will be annunciated for backcourse sensing.

The vertical two letter identiiier in the pointer annunciates the deviation source. GS will be annunciated if the source is an 11s.

Failure of the glide slope receiver is flagged by a red X and the letters GS appearing in red.

Nav Map Display

The EHSl provides two basic types of map; a 360 degree map display about the aircraft and an approximately 85 degree sectored map display in front of the aircraft. Options to be displayed on the map include waypoints and navaids. The type and amount of data presented on the map will depend on the interfacing equipment. When coupled with GPS, waypoints, flight plan identification, airports, and navaids may be may be selected for display by pushing the TSTIREF button. Refer to Figure 7-25, Map Mode Display.

The following will address ody those areas of the EHSI MAP Mode which are different from the standard compass presentations previously described.

MAP 360 Compass The operation of the compass card remains the same in the Card Display: map modes as in the standard EHSl display. To provide more

usable display area for map waypoints and navaids, the 5 and 10 degree tic marks have been reduced in size. The compass card radius is unchanged.

Symbolic Aircrafl The size of the symbolic aircraft is reduced to provide a less Disptay: cluttered display as the map graphic data is added.


Report No: 01973-001 7-143

SECTION 7 =PLATUSE AIRPLANE AND SYSTEMS DESCRlPTlON -PC %I1

Selected Course The alphanumeric course select readout in the upper left comer Display: of the display functions the same in the map mode as in the

standard EHSl mode. The standard EHSl selected course is removed from the center of the display. The selected course pointer is replaced with the course line. If the selected waypoint or vortac is within map range, it will be displayed with a movable course line drawn through its center. As the selected course is changed, the course line will rotate about the referenced point. If the selected primary nav sensor is an approach approved No. 1 sensor, the inbound TO course line is green and the outbound FROM course line is white. If the selected primary nav sensor is GPS the inbound TO course line is cyan (light blue). Any time the No. 2 sensor is selected as the primary nav sensor, the inbound TO course line is yellow.

Course Deviation The stationary white deviation scale along the bottom of display Display: provides reference for the course deviation bar to indicate

position of airplane in relation to selected navigation course. To provide back course CDI needle reversal annunciation. ILS map mode only, the deviation scale center triangle will point to the bottom of the display and annunciate a BC when the selected course is 105 degrees or more from the aircraft heading.

TWFROM Display: To the right of the alphanumeric course select, a TO or FR will replace the standard EHSl TOlFROM pointer when in non ILS map modes.

Bearing Pointer Standard EHSl bearing pointers will display when the selected Display: bearing source does not have DME distance associated with it

or when the distance is not within map range. Once the source is within map range, the associated map symbol is displayed as the bearing pointer is removed.

Range Ring Display:

A light blue range ring located between the symbolic aircraft and the outside of the map compass scale aids in determining distance of navaids in relation to the aircraft. Off the right wing of the symbolic aircraft adjacent to the range ring is the range ring distance. The range ring represents half the distance to the outer range ring compass scale. The available ranges are 5,10,20,40,80,160,240,320 and 1000 NM.

Report No: 01973-001 , 17-144

Issued: June lo, 1994 Revision 3: September 29,1995

=PLATUS= SECTION 7 PC XU- AIRPLANE AND SYSTEMS DESCRIPTION

SELECTED COURSE MSPUY

SYMBOLIC AIRCRAFT

MAP S O COMPASS

CARD DJSPUY

- RAffiE RHG

BSPLAY

COURSE DEVlATlON MSPIAY

Figure 7-25. Map Made Display


Report No: 01973-001 7-145


ARC Display

The expanded ( ARC ) format provides an enlarged display for increased resolution to NAV data due to the enlarged compass scale presentation. Refer to Figure 7-26. Arc Display.

The following will address only those areas of the EHSl ARC Mode which are different from the standard compass presentations previously described.

Heading Bug: The heading bug operation is the same in all display modes. The only difference is in appearance, the digital readout will change to the cdor orange. This display occurs when the heading bug moves out of view and the heading display will appear on the left or right side of the compass scale, whichever is closest.

Course Deviation The rotating white deviation scale operates the same in Scale: all modes. The deviation scale is slightly reduced in size and is

positioned at the bottom of the display.

Course Pointer: The course pointer may not be in view but the CRS readout will be displayed in the upper left corner.



u =PLCTTUS- SECTION 7 B -PC XII- AIRPLANE AND SYSTEMS DESCRIPTION

HEADING SELECT BUG

\ COURSE DEVIATION DISPLAY


I Revision 3: September 29, 1995

Figure 7-26. Arc Display

Report No: 01973-001 7-147

SECTION 7 ''PLATUSS AIRPLANE AND SYSTEMS DESCRIPTION 'PC XII

EFlS CONTROL PANEL

Refer to Figure 7-27, EFlS Control Panel. The EFIS Control Panel is located on the center console and provide the following functions.

NAV SENSOR During installation, the system was programmed with SELECT: the type and quanbty of each piece of interfacing equipment.

Of all the equipment interfaced to the system, only a few sensors are usable for navigation. The system creates and maintains in permanent memory a list of the usable navigation sensors.

The NAV push button is used to select the primary nav sensor which is annunciated on the left side of the display. A press of the NAV sensor select button sequentially selects the next available sensor from the list of those installed. DM€ information in the upper right comer, selected course, course pointer, and deviation bar are referenced to !he selected primary nav sensor.

The following is a list, in order, of the possible primary navigation sensors that may be of interfaced with the EFlS 40150:

VOR ( L W G I S )

GPS

ADF

1-2 SYSTEM The 1-2 button is used to cycle between primary navigation SELECT: sensor system #I and 12 for display. The primary NAV system

selected is annunciated as sensor, sensor 1, or sensor 2 on the EHSI. For example, if the VOR 1 is being displayed and the 1- 2 button is pressed, VOR 2 will become the displayed sensor. If only one sensor is installed, the display will not cycle and the sensor annunciation will not show a system number.

Report No: 01973-001 I 7-148


= P k A T U S ~ SECTION 7 -PC XII- AIRPLANE AND SYSTEMS DESCRIPTION

NUMBER 1 SYSTEM The bearing pointer select button works similar to the NAV BEARING POINTER sensor select button. A press of the bearing pointer button SELECT: sequentially selects the next available sensor for display. The

bearing pointer select list contains only those sensors which are associated with the bearing. If the selected sensor has distance information paired with it, that distance will also be displayed in the lower left-hand corner along with the sensor annunciation.

The following is a list, in order, of the bearing pointer sensors that may be interfaced with the EFS 40150:

DECLUlTER (no number one bearing pointer information is displayed)

VOR 1

GPS

ADF

Only those sensors installed in this aircraft and interfaced will be selectabie for use and display.

NUMBER 2 SYSTEM The bearing pointer select button works similar to the NAV BEARING POINTER sensor select button and the Number 1 Bearing Pointer Select SELECT: button. A press of the bearing pointer bunon sequentially

selects the next available sensor. The bearing pointer sensor list contains only those sensors whim have bearing associated with them. If the selected sensor has distance information paired with it, that distance will be displayed in the lower right hand comer along with the sensor annunciation.

The following is a list, in order, of the bearing pointer sensors that may be interfaced with the EFS 40150:

DECLUlTER (no number two bearing pointer information is displayed)

VOR 2

GPS

ADF

Only those sensors installed in this aircraft and interfaced will be selectable for use and display.


t Revision 3: September 29, 1995 Report No: 01973-001

7-149 (

SECTION 9 mPILATUSW AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

HSI 360 DEGREE: Two different 360 degree display formats: standard HSI compass rose and HSI NAV MAP mode. Each press of the HSI button sequentially button will sequentially select the possible display formats.

The display selection list may include the following:

MODE SELECTION

HSI COMPASS ROSE WITH NAV DISPLAYS

HSI NAV MAP

ARC MODE SELECTION:

The ARC mode provides the pilot with a large scale view of the CDI by presenting an approximate 85 degree sector display of the compass.

The EHSl has two possible ARC sectored display formats: standard HSI compass rose and ARC NAV MAP mode. A press of the ARC button will sequentially select the possible display formats.

A press of the ARC button while in the 360 degree mode will result in an ARC presentation of the same format. For example, i f the 360 NAV MAP mode was being displayed and the ARC button was pressed, the resulting display format would be ARC NAV MAP. A press of the HSI button while in the ARC mode will change the display to the standard HSI compass rose.

The ARC display selection list may include the following:

ARC COMPASS ROSE WITH NAV DISPLAY

ARC NAV MAP

RANGE SELECTION: RANGE DOWN - A press of the RANGE DOWN button selects the next lower range to be displayed while in the NAV MAP mode of operation. Once the lowest selectable range is reached, the RANGE UP button must be used for a range change.

RANGE UP - The operation of the RANGE UP button is similar to the RANGE DOWN except it selects the next higher range to be displayed while in the NAV MAP mode of operation.

Report No: 01 973-001 1 7-150 Issued: June 10.1994


%PLATUSs SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

DH SET: To set Decision Height, pull out and turn the DH knob. Turning (if installed) clockwise will increase the Decision Height selected, turning

counter clockwise will decrease the Decision Height. The Decision Height range is from OFF to 2,500 feet and will be displayed in one foot increments to 500 feet and then in 10 foot increments to 2,500 feet. Once the Decision Height is selected, push in the DH knob to lock the selected DH altitude. I f DH is set to off, the DH annunciator will not be displayed.

TST:

REF:

The TSTIREF button performs the function of SELF TEST display. To activate the SELF TEST, press and hold the TSTIREF button for 3 seconds. The self test processing time may last up to 5 seconds depending upon the particular functions being performed by the symbol generator. Upon completion of self test, a test pattern annunciating pass or fail will be displayed until the TST button is pressed once again.

When GPS MAP display has been selected, pressing the button for 1 second will annunciate the present map format. If the displayed format is desired, no additional action is required. If a different format is desired, sequence through the list by pressing the button for 1 second for each format until the desired annunciation (FPL ID, AIRPORT, or NA'JAIDS) is displayed. In either case, approximately 10 seconds after the last button press, the map format annunciation will be removed. The alphanumeric readout of the DME can be changed between ground speed in knots to time to station by pressing the TSTmEF button.

RALT TST: Pressing the RALT TST push button provides a discrete output (if installed) to the Radar Altimeter initiating its self test function.

COURSE SELECT Rotation of the COURSE SELECT knob allows the course KNOB: pointer on the EHSl to be rotated to the desired course.

The Control Panel provides a DIRECT TO feature. Pulling the COURSE SELECT knob will cause the course pointer and digital course readout on the EHSl to slew to the direct course to the selected navaid or active waypoint.


Report No: 01 973-001 7-151

SECTION 7 =PILATUSm AIRPLANE AND SYSTEMS DESCRIPTION 'PC XII'

EHSl DISPLAY SYS REF CHANGE

HSI 360" MODE ARC MODE OR SELECTION SELECTION EFlS SELF TEST

i

DECISION HEIGHT RADIO ALTITUDE DISPLAY UNIT CONTROL BRIGHTNESS CONTROL 4

\ 1

COURSE SELECT KNOB

Report No: 01973-001 7-152

NO 2 SYSTEM BEARING 1 . 2 SYSTEM POINTER SELECT SELECT

HEADING SELECT KNOB

Figure 7-27. EFlS Control Panel


TEPLATUS= SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

HEADING SELECT Rotation of the HEADING SELECT knob allows the KNOB: heading bug on the EHSl to be rotated to the desired heading.

The Control Panel provides a HEADING SYNC feature. Pulling the Control Panel HEADING SELECT knob will cause theheading bug on the EHSl to slew to the present aircrab heading (lubber line).

DISPLAY UNIT The BRT knob is a split knob which allows individual control of BRIGHTNESS the EADl and EHSl display brightness. CONTROL:

MULTl FUNCTION DISPLAY (IF INSTALLED)

Display

The Multi Function Display (MFD) is a secondary EFlS display providing the pilot with the ability to cross check the primary EFlS displays, preview or set up for a potential navigation scenario, and still function as a primary weather radar, navigation map display or checklist source.

The MFD display tube is identical to the EHSl display tube and is located in the center instrument panel. The actual presentation on the MFD is similar to that on the EHSl for navigation purposes. Refer to Weather Radar, this section, for information concerning radar display. The MFD k contrded by two separate cmtrd panels located in the center lower panel of the cockpit.

When the MFD course select knob is not active, a bar the color of the CRS annunciator wiH appear above the Selected Course Display in the upper left corner of the MFD display. The MFD course annunciator is then referenced to the pilot's EHSl selected course.

Issued: June 10,1994 Revision 3 September 29,1995

Report No: 01 973-001 7-153


Control Panel

Refer to Figure 7-28, Multi Function Display Control Panel. Various buttons on the MFD control panel have identical functions as the buttons on the EFlS Control Panel. Refer to EFlS Control Panel, this section, for more information.

The MFD Control Panel provides the following additional functions:

Course Select The Course Select Knob will function identical to, but KnoWButton: independent of the one on the EFlS Control Panel when the

CRS SEL button is activated. This provides independent course selection on the MFD referenced to the same or different navigation source.

When the CRS SEL button is not activated, the MFD course will be the same as on the EHSl unless the selected navigation source on the MFD is providing DTK course guidance. A bar the color of the CRS annunciator wiU appear above the Selected Course Display.

ARC Button:

Joystick:

The ARC button will allow the selection of a radar only display. This display will not contain navigation or map information.

The joystick is used to generate and move single waypoints on the display unit. These waypoints can be entered into the KLN 90A.

With the GPS selected as the Primary Navigation Source, initial movement of the joystick will create a standard white waypoint symbol ahead of the airplane symbol on the half range ring on the current heading. Direction and speed of the waypoint symbol is commanded by joystick displacement from the center position. LaVLon coordinates of the waypoint will be displayed in the lower center of the display. The coordinates of the waypoint may be entered into the GPS by pressing the ENT button whever the GPS is in the active OBS mode.

ENT Button: The LaVLon coordinates of the joystick created waypoint will be transferred to GPS when pressed. The ENT button must be pressed while the coordinates are displayed on the MFD.

TCAS ONLY Button: This button activates the TCAS traffic-only display on the traffic overlay along with NAV map and/or weather information, i f the system is installed. Refer to Section 9 Supplements, for information concerning limitations and operating instructions for optional avionics installation.

Report NO: 01973-001 ( 7-154



CHECKLIST DATA PORT \

Figure 7-28. Multi Function Display Control Panel

Issued: June 10, 1994 Revision 3: September 29. 1995

Report NO: 01973-001 7-155 1

SECTION 7 =PILATUS= AIRPLANE AND SYSTEMS DESCRIPTION -PC XII-

CHK LlST Button: The CHK LlST button accesses the checklist function. The checklist function of the MFD may be accessed prior to engine start by activating the STBY power switch on the Overhead Panel. When activated, the MFD display and control panel functions will be active for 10 minutes. The timer cycle may be interrupted turning the Battery Master Switch ON and OFF one time. Individual checklists may be created on a personal computer and loaded into the MFD Control Panel through the Checklist Data Port.

THE USERllNSTALLER IS RESPONSIBLE FOR THE CHECKLIST CONTENTS. IT IS ALSO THE RESPONSIBILITY OF THE OWNEWOPERATOR TO UPDATE THE CHECKLIST BASED UPON THE LATEST REVISION OF THE PILOT'S OPERATING HANDBOOK. THE COMPUTER SOFIWARE REQUIRED TO CREATE THE CHECKLIST IS AVAILABLE FROM BENDIXIKING.

Report No: 01 973-001 7-156


mPILATUSE SECTION 7 PC XI1 AIRPLANE AND SYSTEMS DESCRIPTION

WEATHER RADAR (IF INSTALLED)

Display

Weather radar information may be displayed with or without navigation data presentation on the EFlS or the MFD. When selected, a vertical view of the weather return along the Track Line may be displayed on the MFD. Refer to Figure 7-29, MFD Weather Only Display.

Radar Mode: Indicates present operational mode - OFF, TST, SBY, WX. WXA, and MAP.

Track Line: A dashed line drawn from the center of the symbloic aircraft to the compass card or the outer range ring is activated and slewed by pressing one of the TRK buttons.- The track line will automatically dedutter if it remains stationaw for more than 15 seconds. Pressing the VP button will adtivate the vertical profile feature of the radar at the angle of the track line.

Track Line A digital readout in the upper left hand comer displays the Location present position of the track line relative to the nose of the Annunciation: airplane.

Select Range: Distance scale for radar display in nautical miles.

Range Rings: Evenly spaced dashed lines for distance scale.

Symbolic Aircrafl: Provides a visual reterence of the aircraft position in relationship to the weather display.

Tilt Angle: Indicates present antenna vertical angle.

Profile Angle: Indicates present position of the vertical profile radar display relative to the nose of the airplane.

Altitude Lines: Three horizontal altitude lines are present on the VP display. The center line represents the actual airplane altitude. Equally spaced above and below the center line are two lines a certain number of feet above or below the airplane. The number associated with these lines varies with the selected range to compensate for the radar beam width at the various ranges. The altitude is expressed in thousands of feet from the center line.


I Revision 3: September 29. 1995

Report No: 01973-001 , 7-157

SECTION I ePLATUSW AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1 4

TRACK LINE LOCAT)[)E(

AnwNaATION

SELECT FIANOE

- M E RINGS /-

TRAW L W

/-

RAOUI MOLE

I TILT ANGLE

\ SYMBOLIC URCRAFI

1 WEATHER ONLY DISPLAY SELECTED

Report No: 01973-001 1 7-158

Figure 7-29. MFD Weather Radar Display (Sheet 1 of 2)

Issued: June 10,lggq Revision 3: September 29, 1995

=PLATUS= SECTION 7 -PC x F AIRPLANE AND SYSTEMS DESCRIPTION

PROFILE ANGCE

ALTITUDE LINES

SYMBOUC AIR-

RADAR' WM

R A N a RlNOS

VP MODE SELECTED

Figure 7-29. MFD Weather Radar Display (Sheet 2 of 2)


Report No: 01973-001 7-159 1

SECTION 7 EPILATUS W AIRPLANE AND SYSTEMS DESCRIPTION ?c XI1

Contrd Panel

Weather radar d~splay is controlled by a panel located on the center lower panel ot the cockpit. Refer to Figure 7-30, Radar Contrd Panel.

The Radar Control Panel provides the following functions.

Radar Function Rotary switch selects the desired operating condition. Selector Switch:

ON - Selects the condition of normal operation, allowing for weather detection or other modes of operation. Radar emission exists in the ON position. However, if a radar mode is not selected for display on at least one indicator the ART is placed in STANDBY.

TST - Selects the test mode and the test pattern will appear on the indicator. The antenna will scan without radar emission.

STBY - Selects the standby mode for system warm-up and when the system is not in use. After 30 seconds in this mode, the system is in a state of readiness. No radar emission occurs and the antenna is parked in the down position.

OFF - Removes power from the ART.

WX Button:

WXA:

GND MAP:

VP:

Report No: 01 973-001 1 7-160

Selects the Weather mode when pressed. Areas of high rainfall will appear in magenta color. WX will appear on the indicator when a weather mode is selected.

Selects the Weather Alert mode when pressed. Magenta areas of precipitation will flash between magenta and black. WXA will appear on the indicator when a weather mode is selected.

Selects Ground Mapping mode, disables the weather alert feature, and activates the gain control. MAP will appear on the indicator when a weather mode is selected.

Selects and deselects the vertical profile modes. either the VP only modes or the horizontal and vertical (VP) weather presentation. Selecting one of the VP mode will not change the selected mode of operation. Once in VP mode, weather modes may be changed as desired. VP will engage from GND MAP mode but the NAV function will be disabled.


I =pkATUS= SECTION 7 B 'PC X+ AIRPLANE AND SYSTEMS DESCRIPTION

Figure 7-30. Radar Control Panel


Report No: 01973-001 7-161

SECTION 7 mPILATUSS AIRPLANE AND SYSTEMS DESCRIPTION -PC %I1

TRK Buttons: When pressed provides a yellow azimuth line and a digital display of the azimuth line placement left or right of the airplane nose. For vertical profile (VP) operations. the track bunons perform the two functions:

1. Prior lo engaging VP, the appropriate button (left or right) is used to place the track line at the desired azimuth angle to be vertically scanned (sliced). When VP is engaged, the slice will be taken at the last position of the track line, whether it is visible or not. If the track line has not been selected after power has been applied to the system and VP is selected, the slice will be taken at 0" (directly in front of the airplane).

2. Continuously holding the TRK button will result in the system slicing in 2" increments.

GAIN Control Knob: Manual gain control becomes active only when GND MAP is selected. Gain is automatically set for all other modes.

Tilt Control Knob: Permits manual adjustment of the antenna tilt angle to a maximum of + I - 15" in order to obtain the best indicator presentation. The tilt angle is displayed in the lower left comer of the display.

Lightning Button: Selects the Lightning display when pressed if a weather mode (if installed) is selected for display.

Lightning Detection and Display (if installed)

The Stormscope System detects lightning activity 360 degrees around the aircraft up to a distance of 200 nautical miles. The system processes the lightning data and transmits it to the EFIS in the form of cells (areas) of electrical discharge (not individual discharges). The locations of up to 63 cells are displayed, together with the intensity level of each cell. There are three levels of intensity, level 1 is the lowest rate of electrical discharge and level 3 is the

highest.

The lightning display mode is selectedldeselected on the weather display by pressing the Lightning button on the Radar Control Panel. The lightning mode status is shown on the EflS display in a data field above the weather radar data fields. The lightning data field shows a lightning bolt followed by ON (lightning mode selected). OFF (lightning mode deselected) or FLT (fault). The default condition is lightning mode OFF at system power up.

With lightning mode ON, up to 63 grey colored lightning cell symbols can be displayed on the

EFIS. There are three lightning cell symbols to show the different levels of lightning intensity.

Report No: 01973-001 7- 162


=PILATUSr SECTION 7 -PC %I1 AIRPLANE AND SYSTEMS DESCRIPTION

The level 1 (lowest rate of discharge) symbol is a lightning bolt with no arrowheads. Level 2 has one arrowhead. Level 3 (highest rate of discharge) has two arrowheads. The lightning symbols can appear anywhere in the 360 degrees of display area (unlike weather radar information).

A pushswitch marked STORMSCOPE CLEAR is used to erase the lightning data from the Stormscope System memory. This removes the accumulated lightning cell symbols from the EFlS display in order to start a new display.

OPERATION

Refer to the BendixlKing EFS 40150 Pilot's Guide for detailed operating information.

When the main DC distribution system is energized, the EFlS is switched on by selecting the AVIONICS 1 and AVIONICS 2 switches to ON. Various flags may be annunciated on the displays after initial power on, indicating systems which are not immediately ready for use.

At initial power on the system conducts a self test. The displays should be clear of waming flags shortly after power on. System testing can be initiated using the test switch.

In the event of a DU or EADl symbol generator failure, a combined ( composite ) display of EADl and EHSl information may be displayed on the remaining operational DU. This is accomplished by setting the appropriate EFlS switch to the CMPST position. The navigation information added is the standard deviation scale. selected course and heading, distance information, selected navigation sensor, and the TOIFROM display. This format presents the pilot with a familiar display that requires minimal transition time when it is selected for use. The function of the control panel, navigation guidance presentation and color coding remain the same as in normal operation.

The pilot's EFlS switch is located on the pilot's lower left panel and the copilot's EFlS switch (if installed) is located on the copilot's lower right panel.

A composite display may be selected on the MFD (if installed) by setting the EFlS MFD switch, located on the copilot's lower right panel, to the CMPST position. This display is the same as the pilot's EFIS composite display when mode is selected.


Report No: 01973-001 7-163

SECTION 7 WPLATUSW AIRPLANE AND SYSTEMS DESCRIPTION P(l2

4

COURSE POINTER

\

W / DISTANCE MSPUY

/ \ DEVIATK*I OlSPUY

SELECTED COURSE SELECTED H W W O DISPLAY DlSPUY

Figure 7-31. Composile Display - Approach

Report No: 01973-00 1 1 7-164 ~ssued: June 10,1994


b FPLATUS* SECTION 7 t Kl2 AIRPLANE AND SYSTEMS DESCRIPTION

TO/FAOM DISPLAY

Figure 7-32. Composite Display - Enroute


Report No: 01 973-001 7-165


INDICATION i WARMNO

During normal operation lhe fdlowing faults may be detected and displayed automatically.

SO ( Red ): EFiS information in general is unreliable. Revert to backup instruments.

CP ( Red ): EFlS contrd panel switch stuck for more than 10 seconds. If the switch fails, the display will remein all currently selected conditions. Continue with current Nav system selection or revert to backup instruments.

Blank CRT: Display Unit failure. Switch to DU CMPST mode.

A red cross an any display indicates that the information is unreliable.

SG ( YeUow): Symbol Generator fan has failed. Once annunciated, the faulty DU will continue to operate for at least 30 minuteg H the rated ambient temperature is not exceeded. Reduce display information to a minimum to extend operating time.

DU ( Yellow ): Display Unit fan has failed. Once annunciated, the faulty DU will continue to operate for at least 30 minutes if the rated ambient temperature is no1 exceeded. Reduce display information and brighhem level to a minimum to extend operating time.

Refer to the BendixMng Pilors Guide for additional warning cautkm lags.

Report NO: 01973-001 I 7-166 Issued: June lo, 1094

Revision 3: September 29,1995

XPL ATUSF SECTION 7 C PC XI1 AIRPLANE AND SYSTEMS DESCRIPTION

k

I AUTOPILOT

GENERAL

The autopilot installed in the PC-12 aircraft is the BendidKing KFC 325 Digital Automatic Flight Control System (AFCS). The limitations presented in Section 2 of this Handbook are pertinent to the operation of the KFC 325 AFCS as installed. The Automatic Flight Control System must be operated within the limitations specified. Refer to the appropriate sections of this Handbook for the Emergency and Normal Procedures associated with this installation.

The KFC 325 Digital AFCS has 3 axis control: pitch, roll and yaw

The KFC 325 Digital AFCS has an automatic electric pitch trim system which provides pitch autotrim during autopilot operation. When the pitch autotrim function is in operation, a signal is sent to the triple trim indicator to illuminate the pitch trim light (if this version of indicator is installed). The autotrim system is designed to withstand any single inflight malfunction. Trim faults are visually and aurally annunciated.

The KFC 325 Digital AFCS has an automatic rudder trim relief function which provides directional trim during yaw damper and autopilot operation. When the rudder autotrim function is in operation, a signal is sent to the triple trim indicator to illuminate the rudder trim light (if this version of indicator is installed).

No aileron autotrim function is available.

Vertical autopilot functions include Altitude Select and Vertical Speed modes.

A lockout device prevents autopilot engagement until the system has been successfully preflight tested.

DESCRIPTION

t MODE CONTROLLER

The KFC 325 AFCS operation is controlled by a KMC 321 Mode Controller located at the top of I the center panel. Refer to Figure 7-33. Autopilot Mode Controller. Autopilot mode selection

provides the following functions.

r HDG: Alternately engages and disengages the Heading Select mode. Heading information is received from the Heading Bug on the EHSI. Depressing HDG will activate the Flight Director in Heading mode.

NAV: Alternately engages and disengages the Navigation mode. Depressing NAV will activate the Flight Director. The Flight


Report No: 01973-001 7-167

SECTION 7 EPILATUSI AIRPLANE AND SYSTEMS DESCRIPTION ?c %I1

1

APR:

YD:

AP:

Director will command tracking of the coupled navigation *

receiver based on the EHSl selected primary navigation I

source. Glideslope coupling is inhibited in the NAV mode.

Alternately engages and disengages the Approach mode. Depressing APR will activate the Flight Director. APR mode will capture and track selected EHSl primary navigation sensor with approach accuracy. Glideslope coupling is allowed in the APR Capture or Track mode. BC is automatically engaged and disengaged. Back Course functions identically to the Approach mode except that the autopilot response to the localizer signal is reversed. Glideslope coupling is inhibited in the Back Course Approach mode.

Wilh the EFS 40150, Back Course is determined from aircraft heading and the Course Pointer. The APR pushbutton , activatesldeactivates the Back Course mode.

Alternately engages and disengages the Yaw Damper and rudder trim relief independent of the autopilots pitch and roll axes

Alternately engages and disengages the autopilot. Yaw Damp is automatically activated when the autopilot is engaged, 4

however, Yaw Damp remains engaged if AP is pressed again.

Controls the vertical axis of the autopilot. The rocker switch function is dependent upon the autopilots active mode. Depressing and holding the switch up or down results in the following:

In Pitch Attitude Hold mode the vertical trim switch adjusts the pitch attitude at a rate dependent upon the current airspeed.

In Altitude Hold mode the vertical trim switch adjusts the altitude which the autopilot is holding. Trim control up and down operates at a rate of up to 500 feet per ,

minute. This does not affect the altitude selected and displayed on the Altitude Select.

In Indicated Airspeed Hold mode lhe vertical trim switch adjusts the indicated airspeed reference at a constant rate of three quarters of a knot per second (0.75 kVsec).

In Vertical Speed Hold mode the vertical trim switch adjusts the vertical speed at a rate of one hundred feet per minute for each second the switch is held.

Report No: 01973-001 7-168


~ P R A T U S ~ SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

ALT:

IAS:

FD:

SOFT RIDE:

HALF BANK:

TEST:

Issued: June 10,1994 I Revision 5: May 10.1996

Figure 7-33. Autopilot Mode Contrdler

Altemately engages and disengages the Altitude Hold mode. Altitude Hdd commands the aircraft to maintain the pressure altitude existing at the moment of selection. The pilot must correct for altimeter changes during flight to insure barometrically corrected altitude.

Altemately engages and disengages the Indicated Airspeed Hold mode. This mode commands the aircraft to maintain the Indicated Airspeed existing at the moment of selection. The aircraft pitch command is varied by the Flight Director to maintain the selected airspeed during changing air conditions, power andlor configuration changes.

Engages the Flight Director in Pitch Attitude Hold mode and Wings Level mode. The pitch attitude of the Flight Director is synchronized to the current aircraft pitch attitude. Pressing FD when the Flight Director is engaged will disengage all Flight Director modes if the autopilot is not engaged.

Engages the Soft Ride mode. Soft Ride mode decreases the autopilot gains thus decreasing the aggressiveness of the autopilot resulting in a more comfortable ride in turbulent conditions. Routine use of this mode during all flight conditions will result in less than optimum autopilot performance. Soft Ride mode is automatically disengaged when the Approach mode is activated.

Engages the Half Bank mode. The autopilots maximum commanded bank angle is reduced to one half the normal value. This mode is provided to increase passenger comfort. This mode is automatically disengaged when the Approach mode is activated.

The Preflight Test Sequence is initiated when this button is momentarily pressed. The test includes illumination of all annunciator lights, testing of rate, acceleration and trim monitors, and testing of the computers internal logic circuits. The AP annunciator will flash for approximately six seconds

Report No: 01973-001 7-169 )

SECT ION 7 AIRPLANE AND SYSTEMS DESCRIPTION

accompanies the Hashing AP annunciation. The Autopilot mode will not be enabled until the Preflight Test has been successfully passed.

ALTITUDE I VERTICAL SPEED PRESELECT

The KAS 297C operation is controlled by a Controller located in the pilot's instrument panel. Refer to Figure 7-35. Altitude I Vertical Speed Preselect Controller. Mode selection provides the following functions.

VERTICAL SPEED I Concentric knobs allow selection of altitude or vertical speed. ALTITUDE SELECT The small inner knob has two positions: KNOB

IN - Altitude is displayed. When rotated, the small inner knob selects altitude in 100 foot increments with roll over into the 1000 digits. The large outer knob selects 1000 foot increments with roll over into the 10.000 digits.

ARM

I ENG

OUT - Vertical Speed is displayed. When rotated, the small inner knob selects vertical speed in 100 feet per minute increments. The large outer knob selects vertical speed in 1000 feet per minute increments up to a maximum of 5000 feet per minute.

Engages the autopilot Altitude Arm mode when depressed while the selected altitude is displayed. Altitude Select (ARM) mode will cancel the autopilot Altitude Hold (ALT) mode if autopilot Altitude Hold is already engaged. Glideslope coupling will cancel Altitude Select (ARM) mode. The engagement of Altitude H d d by the pilot with the Autopilot Mode Controller will cancel the Altitude Select (ARM) mode. Altitude Select (ARM) mode allows selection of a new altitude without deactivating the ARM.

NOTE

The display of the selected altitude is required to activate the Altitude Select mode. This assures pilot verification of altitude before activation.

Engages the Vertical Speed Hold mode when depressed while the selected vertical speed is displayed. When depressed with no vertical speed value selected, the Vertical Speed HoM mode is engaged and is synchronized to the current vertical speed of the aircraft. The synchronized vertical speed is momentarily displayed.

Report NO: 01973-001 1 7-170


-P(L#TUSW SECTION 7 L ?C XI1 AIRPLANE AND SYSTEMS DESCRIPTION

OPERATION

AUTOPILOT

Emergency and Normal Procedures are detailed in Sections 3 and 4 of this Handbook.

The following conditions will cause the autopilot to automatically disengage:

Pilot related input

Rdl rates in excess of 10" per second will cause the autopilot to disengage except when the CWS switch is held depressed.

Manual trim engage pilot or copilot, FD and operational modes remain engaged.

Alternate trim Mi .

Activation of Trim Interrupt.

System related input

Power failure.

Internal Flight Control System failure.

Pitch rates in excess of 5" per second will cause the autopilot to disengage except when the CWS switch is held depressed.

Accelerations outside of a t1.6 g to +0.3 g envelope (1 .O g's being normal for straight and level flight). Disengagement will take place regardless of whether or not the CWS switch is activated.

The presence of an EFlS ATTMDG flag. The flight director will also disengage.

Stall warning (approaching stall condition).

The airplane AVIONICS 2 (AV 2) SWITCH function is unchanged and can be used in an emergency to shut off electrical power to all flight control systems while the problem is isolated.

PITCH LIMITS

IAS mode +15" -10" PAHNS +20° -10" APR (GS) +20° -20° ALTfALT CAPT rate limited

The following circuit breakers are used to protect the following elements of the King KFC 325 Autopilot:

Avionic Bus 2 AIP Disc, AIP Trim Adapter, AIP 26 V AC Bus NP (ref voltage only)


Report No: 01973-001 7-171

SECTION 7 BiPLATUSW AIRPLANE AND SYSTEMS OESCRlPTlON ?c %I)


Emergency and Normal Procedures are detailed in Sections 3 and 4 of this Handbook.

CONTROL WHEEL STEERING

Mounted on each inboard control wheel horn is the Control Wheel Steering (CWS) switch. Pressing the CWS pushbunon disengages the autopilot sew0 clutches and allows manual control of the aircraft. Upon release of the pushbunon, the autopilot clutches reengage and the autopilot Idlows the new vertical reference if no lateral or vertical mode was selected or re- track the previous engaged AP mode.

I f the Flight Director is not engaged when the CWS pushbutton is depressed, the Flight Director will be activated in the Pitch Attitude and Wigs Level Hold mode. The Flight Director command bar will synchronize the aircraft to the attitude present upon pushbutton re)ease. Re- synchronisation is possible by pressing the CWS button again.

AUTOPILOT DISCONNECT

Mounted on each outboard control wheel horn is the Autopilot Disconnect switch. When momentarily pressed, it disengages the autopilot and yaw damper and cancels aN operating Flight Director modes. A tone will sound upon successful autopilot disconnect. Autometic pitch trim function is inhibited and the second pitch trim motor reverts to alternate pitch trim f~nction.

MANUAL TRIM ENGAGE

Mounted on each outboard control wheel horn is Ulo manual trim engage switch. When momentarily pressed, it disengages the autopilot but leaves the yaw damper and ell selected modes including the flight director engaged. A tone will sound upon successful autopilot disconnect. Rudder trim relief function will be operational as long as the yaw damper is activated.

Report No: 01973-001 1 7-172

issued: June 10,1994 Revision 3: September 29, 1995

~PICATUSB SECTION 7 ?c XII- AIRPLANE AND SYSTEMS DESCRIPTION

INDtCATION I WARNING

AUTOPILOT

The engaged and armed autopilot modes are annunciated on the Mode Controller and in the upper area of the EADI. Refer to Figure 7-34, Autopilot Mode Annunciations. The following annunciations are illuminated on the Mode Controller ( EADI annunciations are shown in Figure 7-34).

NAV:

llluminates when the Heading Select mode is engaged by depressing the Heading pushbutton.

llluminates when the Navigation mode is engaged by depressing the NAV pushbutton and normally sequenced through NAV ARM. Nav mode can be used with VOR or GPS Navigation sensors.

NAV ARM: illuminates when Nav mode is called for by the NAV pushbutton and the course needle deflection exceeds the capture requirements of the Nav mode. While capture requirements (needle displacement and rate of needle displacement) are exceeded. the system will remain in the Arm mode. When the requirement is achieved, the autopilot will capture and track the course needle.

APR:

APR ARM:

BC:

YD:

AP:

ALT:

llluminates when the Approach mode is engaged by depressing the APR pushbutton and normally sequenced through Approach Arm.

llluminates when the Approach mode is called for by Ihe APR pushbutton and the Course Needle exceeds the capture requirements of the Approach mode. Heading mode may be used to intercept the desired course while the autopilot is Approach Armed.

llluminates when the Back Course mode is engaged. Back Course automatically activates the Approach ArmlCapture mode and illuminates the respective annunciator.

llluminates when the Yaw Damp is engaged by depressing Ihe YD or AP pushbutton.

Illuminates when the autopilot is engaged by depressir~g the AP pushbutton.

Illuminates when the Altitude Hold mode is engaged by depressing the ALT pushbulton or by automatic sequencing through Altitude Capture when using the KAS 297C Altitude Preselect System.

I Issued: June 10, 1994 Revision 3: September 29. 1995

Report No: 01973-001 7-173

SECTION 7 S PILATUS= AIRPLANE AND SYSTEMS DESCRIPTION ?c XII-

IAS:

FD:

SR:

HB:

TRIM:

Illuminates when the Airspeed Hold mode is engaged by I

depressing the IAS pushbutton. fl

llluminales when the FD pushbutton is depressed or by default when any Flight Director mode is engaged. The Flight Director mode is Wings Level and Pitch Attitude Hold. *

llluminates when the Soft Ride mode is engaged by depressing the SR pushbunon. Soft Ride mode can be activated only when the autopilot is engaged. !

llluminates when the Half Bank is engaged by depressing the HB pushbunon. Half Bank can only be activated when the autopilot is engaged.

llluminates continuously in the absence of trim power or il Ihe system has not been preflight tested. An audible warning accompanies the annunciator during a trlm lault.

NOTE

A flashing mode annunciator indicates a loss of the selected (mode) source or a unreliable source.

EFlS

All armed or captured modes are duplicated in the EFlS EADI display. The following additional autopilot annunciators are presented in the EFlS EADI.

PTRM: (red) Illuminates when the autopilot monitors a pitch trim failure.

ROLL: (yellow) Illuminates when the autopilot monitors an aileron mistrim.

AP: (red) Illuminates when the autopilot monitors a system fault and momentar~ly flashes on autopilot disengagement.

CAWS

Additional autopilot annunciations are provided on the CAWS panel.

A'P TRIM: (red) Indicates an autoplot and/or auto trim failure. Warning is also present prior to successful passing of the autopilot preflight test.

AP DISENG (amber) Indicates that the autop~lot p~tch and a~leron servo IS disengaged During CWS ac1ivatcon w~th an engagcd autopilot, disengage annunciation IS inhibited AIP DISENG caution IS delayed 3 seconds In order to prevent aural warning conflict with the autop~lot system A/P DISENG caution goes olf after 30 seconds MSN 181 8 UP.

I AIP TRIM. (blue) Indicates that the autopilot trim is act~ve (autotrim function)

Report No: 0 1973-001 7-174

Issued: June 10, 1994 Revision 7: July 1, 1997


PTAU (pch ~r) - m ~ w a w m ARMED LATERAL - lh a r r . n t umwi EWED LATERU MOW. - R* surm rmd m Iha SR Lsld rhmn olYh Im.tr.l mod. is dnpl.y.d on r**M Whbn Ihm n p . g . d W u m m o d . i s ~ m g . . n aulolmanual t r~m has la~lmd mod. p ~ s lrom Armed to Couplbd. th8 PoruM.- l . rsnm~poubmum

MMhOnmoWvpblh-bW mKi. NAV. APR. LOC. BC. ROL a Un* noLLW*mmhry - .~ l r u h p o u m n n d c h u ~ n m p m maawm m ~ m h . G R W d ~ r * r m annunsodnulr are NAV. APR. LOC BC or

ENGAG€D VERTICUMDD€ - lh M*1 bmlc&mod.mdnp*y.dmpnn M

S + l l a o I I w h ) - m ~ n g m b n m u r u h m r m w p c w m n MT.UlC

wbnavar Boll R~da mod. IS (All~ludm Cnplurm). 115. VS GS m4 W

a u l ~ ~ doWng.9.mbnt. Ihm .AP. annunc8atlon IMndrmnrrh(lnd mumnrHmy.d - YO (yaw dampar) - r -'"pnnJ*n(h. yaw d l W U 18 bU9.p.d Upon yaw dampar

zx%2"TrZ vconh

CWS (control whoa1 .(smnnp) - ts aapmvmd m weon whha Ihm Conllot Whmal Stbmrlg Mod# IS ac11va lhm p~ lo l game m m d m n M o l k w . l m h l S mOQ.

/ / ARMED VERTICAL MODE - 1- Cunml wmd wrtul mod. as CwplY.d n m a When tne mod* goar lrom Amad lo carp*d.thmmuM.Unnsmuptoch. bnaaoad vmrltcal mod* wsfloon and

I Figure 7-34. Autopilot Mode Annunciations


Report No: 01973-001 7-175

SECTION 7 €PLATUS* AIRPLANE AND SYSTEMS DESCRIPTION ?c %(I


The engaged and armed modes are annunciated on the Controller and in the upper area of the EADI. Refer to Figure 7-36, Altitude I Vertical Speed Preselect Annunciations. The following annunciations are illuminated on the Controller ( EADI annunciations are shown in Figure 7-35).

VS: Illuminates when the Vertical Speed Hold mode is engaged by depressing the ENG pushbutton.

ALERT: Illuminates 1000 f 50 feet prior to the selected altitude and extinguishes 200 f 50 feet prior to the selected altitude. The ALERT annunciator will momentarily illuminate when the selected altitude is reached. Anytime the aircraft is more than 200 * 50 feet and less than 1000 f 50 feet from the selected altitude, the annunciator is illuminated. An aural tone accompanies the ALERT annunciator illumination.

ALTITUDE l Indicates the selected altitude from 100 to 50.000 feet or the VERTICAL SPEED selected vertical speed ranging from zero to 5,000 feet per

minute up or down. Altitude is displayed while Ule small inner selector knob is in the 'IN' position. Selected vertical speed is displayed when the small inner selector knob is in the 'OUT' position. Rotating the concentric knobs change the selected altitude or vertical speed.

NOTE

Attention is required to determine it the number selected is selected vertical speed or selected altitude, depending on the position of the small concentric knob. However, the display will start to flash after 5 seconds if it is displaying a selected veacal speed.

VERTICAL SPEED Indicates whether the selected vertical speed is up or down. UP I DOWN CARET

ARM: Indicates that the Altitude Select mode is armed to capture the 68JeCted altitude. The ARM pushbunon activates the Altitude Select mode or whenever the adjustment knobs on the KAS 297C are rotated for a new altitude and the autopilot light director (FD) associated with or without other modes are engaged.


WPILATUSler SECTION 7 ?C %I1 AIRPLANE AND SYSTEMS DESCRIPTION

Figure 7-35. Altitude I Vertical Speed Preselect Annunciations

CAPT: Illuminates when the KAS 297C has switched the Flight Director from the active Pitch mode to Altitude Capture (CAPT) mode. The Altitude Capture mode occurs prior to the point the Flight Director engages Altitude Hold. The point at which the Flight Director initiates Capture varies with vertical speed. The higher the rate of altitude change, the sooner Altitude Capture becomes active. At a low rate of altitude change, the activation of the Altitude Capture mode and the transfer to Altitude Hold occur almost simultaneously.

FT / MIN: Indicates FT I MIN when in Vertical Speed mode. Displays fl when in Altitude Hold mode.


Report No: 01973-001 7-177

SECTION 7 '%EPILATUSw SO XI1

I

AIRPLANE AND SYSTEMS DESCRIPTION .(

4


Report No: 01973-001 7-178

ISSUed: June 10,1994 Revision 3: September 29. 1995


Handling, Servic. & Maint.

EPILAIUSS SECTION 8 K12 HANDLING, SERVICING, AND MAINTENANCE

SECTION 8

HANDLING, SERVlClNG AND MAINTENANCE

TABLE OF CONTENTS

Subject Page

GENERAL

IDENTIFICATION PLATE

AIRPLANE INSPECTIONS

AIRPLANE INSPECTION PERIODS AIRPLANE SCHEDULED INSPECTIONS COMPONENT LIFE POLICY

PREVENTATIVE MAINTENANCE

MODIFICATIONS OR REPAIRS

SERVICE BULLETINS AND SERVICE LETTERS

GROUND HANDLING

TOWING PARKING

MOORING

JACKING

SINGLE WHEEL JACKING MAIN JACKING POINTS LEVELLING

SERVICING

BATTERY


Report No: 01973-001 8-1

SECTION 8 BePLATUSW HANDLING, SERVICING, AND MAINTENANCE K12

Subject

SERVICING (CONT'D)

ENGINE OIL Oil Replenishment Procedure Complete Oil System Replenishment FUEL SYSTEM Refueling Precautions Fueling Procedure Fuel Contamination Fuel Anti-Ice Additive

LANDING GEAR - TIRES

I LANDING GEAR - BRAKES

HY DRAULlC SYSTEM

LUBRICATION POINTS

OXYGEN SYSTEM

REPLENISHMENT PROCEDURE

CLEANING AND CARE

WlNDSHlELDMllNDOWS Windshield (Glass) Side Windows (Acrylic)

EXTERIOR PAINT SURFACES

DEICING BOOT CARE

PROPELLER CARE

LANDING GEAR CARE

ENGINE CARE

INTERIOR CARE

EXTENDED STORAGE



ePLATUSF SECTION 8 ?c XI1 HANDLING, SERVICING, AND MAINTENANCE

GENERAL

This section contains factory-recommended procedures for proper ground handling and routine care and servicing of the PC-12 airplane. It also identifies certain inspection and maintenance requirements that must be followed i f the airplane is to retain its performance and dependability. It is recommended that a planned schedule be followed for lubrication and preventive maintenance based on climatic and flying conditions which may be encountered.

All correspondence regarding the airplane must contain a reference to the manufacturer's serial number (MSN) and be addressed to:

PILATUS AIRCRAFT LTD. CUSTOMER SUPPORT DEPARTMENT. CH-6371 STANS, SWITZERLAND

Tel: 41-41-619 64 54 Fax: 4141-610 33 51 Telex: 866 202 PlLAT CH

Pilatus Aircraft Ltd. cannot accept responsibility for continued airworthiness of any airplane not maintained in accordance with the information contained within this section or the Airplane Maintenance Manual (AMM).

IDENTIFICATION PLATE

An identification plate is located on the lower left side of the fuselage aft of the cargo door. This plate displays the manufacturer's name, model designation, serial number (MSN), date of manufacture and the FOCA and FAA type certificate numbers.

Certain regulations may require an identification plate that displays the airplane registration number. This identificalon plate is located in the empennage.

AIRPLANE INSPECTIONS

AIRPLANE INSPECTION PERIODS

As required by regulations, all civil airplane must undergo a complete inspection annually (each twelve calendar months). In addition to the required annual inspection, the manufacturer also requires 100 hour inspections and Time Limited Inspections.


Report No: 01973-001 8-1

SECTION 8 EPILATUSS HANDLING, SERVICING, AND MAINTENANCE ?C XI1

Other inspections may be required by the issuance of airworthiness directives or service bulletins applicable lo the airplane, engine, propeller and components. It is the responsibility of the operator to ensure compliance with all applicable airworthiness directives and, when the inspections are repetitive, to take appropriate steps to prevent inadvertent non-compliance.

AIRPLANE SCHEDULED INSPECTIONS

As required by regulations, the airplane must be the subject of a complete Annual lnspection each 12 calendar months of operation. The manufacturer also requires 100 Flying Hour Periodic Inspections and Time Limited Inspections. A complete Annuall100 Hour lnspection includes all the items of the Time Limited lnspection which are due at the specified time.

The inspection intervals are based on nonal usage of the airplane under average environmental conditions. Airplane operated in extremely humid tropics, or in exceptionally cold, damp climates, salt-laden conditions may need more frequent inspections for wear, corrosion and lubrication. Under these adverse conditions, the 100 Hour lnspection should be done in compliance with lhe inspection sheets at a more frequent interval. The owner or operator can then set his own inspection interval based on field experience.

The 100 Hour lnspection interval should never be exceeded by more than 10 hours, which can be used only I additional time is required to reach a maintenance center. Any extension of the 100 hour interval must be subtracted from the following inspection interval. For example, if a 100 Hour Inspection is at 1 10 hours. the next is due 90 hours later at 200 hours.

The owner or operator is responsible for complying with any local regulations. The owner or operator is primarily responsible for maintaining the airplane in an airworthy condition, including compliance with Airworthiness Directives. It is further the responsibility of the owner or operator to make sure that the airplane is inspected in confonity with the inspection sheets.

lnspection sheets have been prepared to assist the owner or operator in meeting the foregoing responsibilities. They include, together with the inspection requirements, lists of tools, equipment, parts and materials which are necessary to do the inspections. The inspections sheets are not intended lo be all-inclusive, for no such sheets can replace the good judgment of a qualilied mechanic in the pelfonnance of his duties. As the one primarily responsible for the airworthiness of the airplane, the owner or operator should select only qualified personnel to maintain the airplane.

Detailed information of systems and sub-systems on the airplane can be found in the relevant chapters of the AMM. Reference is made to the topics in this manual and Pilatus issued Service Bulletins for inspection, repair, removal and installation procedures called for in the inspection sheets. It is Vle responsibility of the owner or operator to make sure that mechanics inspecting the airplane have access to these documents as well as the inspection sheets.

The 100 Hour lnspection and Annual lnspection sheets list the maintenance and structural significant items for inspection and state the level of inspection required.

Report No: 01 973-001 8-2

Issued: February 14, 1994 Revision 1: June 10. 1994

WPLATUS= SECTION 8 PC XII- HANDLING, SERVICING, AND MAINTENANCE

The Time Limited Inspection sheets list items that require inspecting at intervals that are different from the 100 hour and annual inspections.

COMPONENT LIFE POLICY

The AMM Section 4 contains the Ainnrorthiness Limitations which specify Life Limit and Inspection Intervals for major components of the airplane.

The AMM Section 5 contains the time limits for overhaul and replacement of components based on average usage and environmental conditions. The stated time limits do not constitute a guarantee that the component will remain in service until this time as the environmental conditions that the component is operated in cannot be controlled by the manufacturer.

PREVENTIVE MAINTENANCE

Pilots operating airplane should refer to the regulations of the country of registry for information on preventive maintenance that may be performed by pilots.

The holder of a Pilot Certificate may perform certain preventive maintenance described in FAR Part 43. This maintenance may be performed only on an airplane which the pilot owns or operates and which is not used to carry persons or property for hire, except as provided in the applicable FAR'S. Although such maintenance is allowed by law, each individual should make an analysis as to whether he/she has the ability to perform the work.

Pilatus Aircraft Ltd should be contacted for further information, or for the required maintenance whkh must be accomplished by appropriately licensed personnel. All other maintenance required on the airplane should be accomplished by the appropriately licensed personnel.

The aircraft has Computer Aided Testing (CAT) connectors which are installed in the maintenance test panel on the right side of the f l i t compartment. They are the central access point for ground maintenance to do aircraft system tests using either a portable computer or a maintenance box. Serious flight safety implications could result if equipment is connected to the CAT connectors during flight. The protective CAT connector caps must be installed during flight and all test equipment must be removed from the aircraft.

If maintenance is accomplished, an entry must be made in the appropriate logbook. The entry should contain:

The date the work was accomplished. Description of the work. Number of hours on the airplane. The certificate number of pilot performing the work. Signature of the individual doing the work.


Report No: 01973-001 8-3

SECTION 8 mPLATUSW HANDLING, SERVICING, AND MAINTENANCE K12

MODIFICATIONS OR REPAIRS

It is essential that the Airwonhiness Authorities of the country of registry be contacted prior to any modifications to the airplane to ensure that the aiworthiness of the airplane is not violated. Modifications or repairs to the airplane must be accomplished by licensed personnel.

SERVICE BULLETINS AND SERVICE LElTERS

Pilatus Aircrait will issue Service Bulletins and Service Letters from time to time whiih will be sent to owners, service centers and distributers. Service Bulletins should be complied w9h promtly and depending on their nature material and labor allowances may apply, this aspect will be addressed in the Planning Information section of the bulletin. Senrice Letters give information on product improvements, changed part numbers or discussion on field problems. Service Bulletin and Service Letter Indexes are issued periodically to provide a complete listing of all issued bulletins and letters.

GROUND HANDLING

TOWING

The use of a towing arm whiih attaches to lugs on the nose leg is the recommended method of towing the airplane over prepared, hard, even ground. The towing arm should incorporate shock absorbers to prevent damage to the airplane. The steering arm provided for this airplane is a steering bar extension to the tail stand. When not in use the components of the towing arm are stowed inside the rear fuselage cone accessible through the battery door.

When towing the airplane, a qualified person should sit in the cockpit ready for immediate braking action, in the event that the towing arm becomes uncoupled. The movement of the towing vehicle should always be started and stopped slowly to avoid unnecessary shock loads. When towing in a congested area, two helpers should watch the wing tip and tail clearances.

In any towing operation, especially when towing with a vehicle, do not exceed the nose gear maximum tow limit angle either side of center, or damage to the nose gear will result. The maximum tow limit angle is ind i ted by a placard on the nose strut. If the airplane is towed or pushed over a rough surface during hangaring, watch that the normal cushioning action of the nose gear does not cause excessive vertical movement of the tail and the resulting contact with low hangar doors or structure. A flat nose tire will also increase the tail height.

In the event that towing lines are necessary, ropes should be attached to the main gear struts as high as possible without contacting brake lines or wire harness. The lines should be long enough to clear the nose andlor tail by not less than 20 feet. A qualified person should occupy the pilot's seat to maintain control of the airplane by the use of the nose wheel steering and brakes.


-PILAWS W SECTION 8 ?C %!I HANDLING, SERVICING, AND MAINTENANCE

Issued: Feb~ary 14,1994 Revision 1: June 10.1QQ4 -- - -

Figure 8-1. Aircraft Towing

Report No: 01973-001 8-5

SECTION 8 mPILATUSW HANDLING, SERVICING, AND MAINTENANCE ?c %)I

PARKING

In nsmal wealher conditions, the airplane can be parked on any firm surface, headed'into wind (11 possible) and the parking brake applied, or wheel chocks in place, or both. Parking for long periods should be done with wheel chocks in place and the parking brake released. Jnstall cockplt control locks. Blanks and covers should be fitted at any time the airplane is parked for an extended time or overnight. The airplane should be moored i f it is b be parked in Ule open for long periods and weather conditions are unfavorable. In extreme conditions, the airplane should be parked in a hangar, as structural damage can occur in high winds, even when moored correctly.

MOORING

The airplane should be moored for immovability, security and protection. The following procedures should be used for the proper mooring of the airplane (Ref. Fig. 8-2 and 8-3):

Head the airplane into wind, where possible.

Retract the flaps.

Close the inertial separator.

Install cockpit control locks.

Chock the wheels.

lnslall the blanks and covers.

Install the propeller anchor.

Secure tiedown ropes to the wings at approximately 45' and tail tiedown points at a maximum of 25" angle to lhe ground.

Fit the propeller boots, and attach to the nose landing gear, to prevent engine wind milling.

Report No: 01973-001 8-6

Issued February 14, 1994 Revision 2: February 14, 1995

3PILATUSW SECTION 8 PC12 HANDLING, SERVICING, AND MAINTENANCE

Figure 8-2. Blanks and Covers (Sheet 1 of 2)


Report No: 01973-001 8-7

SECTION 8 HANDLING, SERVICING, AND MAINTENANCE

Figure 8-2. Blanks and Covers (Sheet 2 of 2)

Report No: 01973-001 8-7A




Issued: February 14,1994 Revision 1 1 : March 1, 2003

Report No: 01 973-001 8-70 1

SECTION 8 ePILATUS3s HANDLING, SERVICING, AND MAINTENANCE ?c %I1 4

I CAUTION 1 USE BOWLINE KNOTS, SQUARE KNOTS OR LOCKED SLIP KNOTS. DO NOT USE PLAIN SLIP KNOTS.

MAKE SURE PROPELLER ANCHOR IS PROPERLY INSTALLED TO PREVENT POSSIBLE ENGINE DAMAGE DUE TO WINDMILLING WITH ZERO OIL PRESSURE.

NOTE

When using rope of a non-synthetic material, leave sufficient slack to avoid damage to the airplane should the ropes contract. Hemp ropes contract significantly in high moisture conditions.

NOTE

Additional preparations for high winds include using tiedown ropes from the nose landing gear.

Report No: 01 973-001 1 8-8 Issued: F e b ~ a r y 14, 1994 Revision 1: June 10. 1994

EPILATUSW L SECTION 8 C ?c XI1 HANDLING, SERVICING, AND MAINTENANCE

APPPKIX w FROM VERTlCU

GROUND MOORING POINTS

RECESSED RI*j m rn AMI CONCFlElE MOOrPMj B L r n

I Rgun 8-3. Airplane Mooring

Issued: February 14, 1994 ! Revision 1: June 10,1994

SECTION 8 EPRATUS= HANDLING, SERVICING, AND MAINTENANCE ?c XII-

JACKING

SINGLE WHEEL JACKING

To assist in wheel and brake maintenance, both the two main wheels and the nose wheel can be jacked, independently, using a bottle jack and an adapter (Ref. Fig. 8-4). The adapters are shaped to accept the piston of a bottle jack. It is advisable that when jacking the nose wheel up, the tail support should be fitted in the rear main jacking pad as a precautionary measure.

Chock the other two tires before single wheel jacking to prevent airplane movement.

I AIRPLANE JACKING

The airplane is equipped with two main jacking points and a combined tail jacking padlmooring point (Ref. Fig. 8-5). The two main jacking points are located on the wing bottom surface just outboard of the fuselage and the tail jacking pad is located on the fuselage bottom surface just forward of the empennage.

Hydraulic jacks are used at the main jacking points to raise and lower the airplane. The tail jacking point is used to maintain the airplane in a level attitude during lilting. When the airplane is raised or lowered, the airplane tail is also progressively raised or lowered accordingly.

ATTACH BALLAST TO THE TAIL JACKING POINT TO PREVENT ANY POSSIBLE REAR FUSELAGE UPWARDS MOVEMENT, WHILE THE AIRPLANE IS ON JACKS.

Refer to the Aircraft Maintenance Manual Chap 7 for procedures on lifting and lowering the complete airplane and information concerning the amount of ballast to be attached to the tail jacking point.

NOTE

When jacking the airplane outdoors, use the tiedown for provisions for the wing and tail as described in Figure 8-3.

LEVELLING

Longitudinal and lateral levelling of the airplane is achieved by positioning a spirit level along or 1

across one of the seat rails in the aft fuselage area. This task is normally done in conjunction w~th raising the airplane on the three main jacks for weighing. setting of landing lights and fuel system calibration.

Report No: 01 973-001 8-1 0


~ P I L A T U S ~ SECTION 8 -PC XC HANDLING, SERVICING, AND MAINTENANCE

Figure 8-4. Single Wheel Jacking


Report No: 01973-001 8-1 1

SECTION 8 BPLATUSB HANDLING, SERVICING, AND MAINTENANCE ?c XI1

I

TAIL JACKING MOORING POINT

MAIN JACK

Figure 8-5. Main Jacking Points

Report No: 01973-001 8-12


=PLATUS= SECTION 8 ?@ XII- HANDLING, SERVICING, AND MAINTENANCE

SERVICING

In addition to the inspection per i is (detailed in Airplane Inspection) and the pre-flight inspections provided in Section 4 of this Handbook, complete servicing instructions are detailed in the AMM Chapter 12-00-00. The following sub-paragraphs give an overview.

BATTERY

Access to the battery (batteries - if second battery installed) is gained by opening the hinged panel (31AB) located on the rear fuselage bottom surface. The battery (batteries - if second battery installed) must be regularly maintained in accordake with the AMM. The operator must also make sure that the battery vents pipes which extrude from the fuselage, just aft of the hinged panel, are free of dirt and any sign of corrosion. In the event that corrosion or a blockage is found, a maintenance shop visit e required, as this situation - if left unchecked - could lead to explosive pressure being reached within the battery which could jeopardize airplane safety.

SB 24-008 installs an external power control unit which will allow the battery (batteries - if second battery installed) to be charged on the ground. With an external power unit connected and operating set the EXT PWR and BAT (BAT 1 or 2 - if second battery installed) switches to ON to ground charge a battery. The battery must be vented during ground charging operations, refer to the AMM Chap 24 for instructions.

ENGINE OIL

Oils specified for use in the PT&A-67B engine oil system are listed in the Pran and Whitney (P&WC) SB No. 14001. latest revision.

If operating conditions are such that the engine will be subjected to frequent cold soaking at an ambient temperature of -18°C or lower, the use of PWA521, Type II oil (5cs) (viscosity) oil (Type II) is recommended. The engine oil dipstick is marked MAX HOT, MAX COLD. ADD US QUARTS. 1,2.3,4.5.6. The term HOT refers to the engine condition when the engine has been shutdown. COLD refers to the engine condition when the engine has been shutdown lor 12 hours or more. Ideally, the engine oil tank level should be checked and replenished, as required. within 10 minutes of shutdown.


Report No: 01973-001 8- 13

SECTION 8 WPlLATUS %l HANDLING, SERVICING, AND MAINTENANCE ?c XI1

A visual sight gauge is provided to allow the oil level to be checked without removing the dipstick. If the oil level is below the green band on the sight gauge Ule dl level has to be checked with the dipstick.

THE GREEN MARKS ON THE FILLER SLEEVE AND THE DIPSTICK MUST BE ALIGNED WHEN THE DIPSTICK IS INSTALLED.

NOTE

The usual oil level is when the dipstick shows one to two quarts below maximum. Oil above this level can be vented overboard.

NEVER REPLENISH THE OIL IN A COLD ENGINE, AS THIS CAN RESULT IN OVERFILLING OF THE SYSTEM. MOTOR THE ENGINE AND RECHECK THE OIL LEVEL BEFORE ADDING OIL TO THE SYSTEM.

MAKE SURE THAT THE OIL IS OF THE CORRECT TYPE. DO NOT MIX BRANDS, SPECIFICATIONS OR TYPES OF OILS. IF OILS ARE ACCIDENTALLY MIXED, DRAIN AND FLUSH THE COMPLETE SYSTEM AND REFILL WITH THE APPROVED OIL.

TO PREVENT OIL DRIPPING FROM THE DIPSTICK AND CONTAMINATING EQUIPMENT, HOLD A PIECE OF ABSORBENT LINT-FREE MATERIAL UNDER THE DIPSTICK DURING REMOVAL.

Report No: 01973-001 1 8-11 Issued: February 14, 1994 Revision 1: June 10, 1994

SPILATUSW SECTION 8 ?c %)I HANDLING, SERVICING, AND MAINTENANCE

OIL REPLENISHMENT PROCEDURE

Open the left engine access panel and secure open with the sttuts.

Release the locking mechanism and remove the filler cap/dipstlck assembly from the Aller neck on the Rler neck on the accessoly gealbox.

Replenish the OH according to HOTICOLD condtion of the engine.

ReinstaN the RNer capldipstick assembly and engage the locking mechanism.

Close the access panel.

COMPLETE OIL SYSTEM REPLENISHMENT

Refer to the AMM for the Complete 011 System Replenishment procedure.

MAKE SURE THE FILLER CAPlDlPSTlCK IS PROPERLY ENGAGED AND LOCKED AFTER REPLENISHMENT.

Issued: Febiualy 14,1994 Revlsion 1: June 10,1994

SECTION 8 BPILATUSW HANDLING, SERVICING, AND MAINTENANCE K12

FUEL SYSTEM

The left and right wing fuel tanks are gravity filled through openings on the upper surface. The tanks should always be kept full between flights to reduce explosive vapor space and condensation. Allowance should be made for expansion to minimize venting of fuel if ambient temperature is expected to rise markedly. Approved fuels conforming to PIWC spec i t i t ins are to be used.

CHECK FUEL SUPPLY VEHICLE FOR CORRECT FUEL GRADE AND TYPE. USE AN APPROVED WATER DETECTION KIT TO CHECK FOR WATER CONTAMINATION.

I CAUTION I AS THE ANTI-ICING ADDITIVE IS NOT ALWAYS INDICATED ON THE FUELLING INSTALLATION PLACARD, CHECK WlTH THE FUEL SUPPLIER TO MAKE SURE THE FUEL CONTAINS AN APPROVED ANTI-ICING ADDITIVE.

IF IT IS KNOWN THAT THE AIRCRAFT WILL FLY IN AMBIENT TEMPERATURES OF LESS THAN 0" C AND IF THE FUEL DOES NOT CONTAIN AN ANTI-ICING ADDITIVE, ONE MUST BE BLENDED WlTH THE FUEL DURING FUELING.

NOTE

There are two fuel tank drain valves on the lower surface of each wing and one on the front left of the fuselage, aft of the nose wheel well.


qPLATUSW SECTION 8 K12 HANDLING, SERVICING, AND MAINTENANCE

REFUELING PRECAUTIONS

During refuelingldefueling operations, the following anangements must be complied with:

Refuel and defuel only in a well ventilated area.

Do not allow open flame or smoking in the vicinity of the airplane while refueling.

Do not replenish the oxygen system during refueling or defueling.

Do not operate airplane electrical or radio equipment while refueling.

High frequency pulse transmissions in the vicinity of the airplane represents a fire hazard.

During all refuelingldefueling operations, fire fighting equipment must be available.

FUELING PROCEDURE

Make sure the fuel supplied is checked for type, grade and freedom from contamination.

Make sure that the refueling vehicle is grounded.

Ground the vehicle to the airplane (attach the vehicle grounding lead to the nose landing gear).

Remove external power, 1 connected

Make sure all electrical power is OFF.

Connect the grounding cable from the nozzle to grounding point next to the fuel cap.

Open the wing fuel cap and insert the nozzle, after first making sure that the filler mule is clean.

Add fuel. Allow the fuel to settle when topping-off the fuel tank. Remove the fuel nozzle and disconnect the grounding cable. Secure the filler cap.

Repeat the procedure for the other wing tank.

Remove the vehicle grounding cab$ from the airplane.

Clean up any fuel spillage (Use a water hose il excessive).

Check all switches OFF.

Set the Battery switch to ON and check the fuel quantity gauges for correct indication.

Reset the fuel totalizer.

Set the Battery switch to OFF.


Report No: 01 973-001 8-17 (

SECTION 8 EPLATUSW HANDLING, SERVICING, AND MAINTENANCE K12

FUEL CONTAMINATION

Fuel contamination is usually the result of foreign material present in the fuel system. This foreign material can take many forms, i.e. water, sand, dirt, microbes or bacterial growth. In addition, additives that are not compatible with the fuel used can cause the fuel to became contaminated.

Jet fuel contains some dissolved, suspended water and is a fuel contamination concern. The quantity of water that can remain in solution will depend upon the temperature of the fuel. Dissolved water cannot be removed by a filter during a fuel selvice but will be released from suspension as the fuel temperature decreases, as during flight. These supercooled water droplets only need to contact sold contaminates or receive an impact shock to change into ice crystals. If a sufliiient quantity of water drops out of suspension and changes to ice, a blocked filter may result.

Before the first flight of the day and after each refueling, use a clean container and drain at least one sample of fuel from each tank drain valve to determine if contarninants are present (and that the airplane has been fueled with the proper fuel). If contamination is detected, drain all fuel drains points until all contamination has been removed. If after repeated sampling. evidence of contamination still exists, the fuel tanks shoukl be completely drained and the fuel system flushed. Do not fly the airplane with contaminated or unapproved fuel.

In addition, operators who are not acquainted with a particular airfield should be assured that the fuel supply has been checked for contamination and is properly filtered before allowing the airplane to be serviced. Also, fuel tanks should be kept full between flights, provided weight and balance considerations will permit, to reduce the possibility of water condensing on the walls of partially filled tanks.

Report No: 01973-001 I 8-10


=PILAWSF SECTION 8 PC12 HANDLING, SERVICING, AND MAINTENANCE

FUEL ANTI-ICE ADDITIVE

A fuel anti-ice additive, conforming to MIL-DTL-27686 or MIL-DTL-85470 specification, may be added to the fuel as the airplane is being refueled. Blend the additive in accordance with the following procedure:

I Calculate the quantity of anti-Icing additive required based on the quantity of fuel to be added. Refer to the CAUTION below. I Remove the cap containing the tube and clip assembly from the "HI-FLO PRIST" blender, model PHF-204.

Attach the pistol on the collar, press the tube into the button, and clip the tube end to the fuel nozzle.

Pull trigger firmly to ensure full flow, then lock into place.

Start flow of additive when fueling begins. Refueling rates should be between 30 and 60 gallons per minute.

Do a water drain check before the first flight of the day.

THE FUEL SYSTEM ANTI-ICING ADDITIVES CONTAIN ETHYLENE GLYCOL MONOETHYL ETHER WHICH IS HIGHLY TOXIC. THESE PRODUCTS MUST BE HANDLED WITH EXTREME CARE. AVOID ALL DIRECT CONTACT WlTH SKlN AND CLOTHING. ANY CLOTHING ACCIDENTLY CONTAMINATED BY SPLASHING SHOULD BE PROMPTLY REMOVED AND THE SKlN WASHED WlTH SOAP AND WATER. PREVENT CONTACT WlTH EYES AND AVOID INHALATION OF VAPORS. IF CONTACT IS MADE WlTH THE EYES THEY SHOULD BE FLUSHED WITH WATER FOR 15 MINUTES. CONSULT A PHYSICIAN AS RAPIDLY AS POSSIBLE AFTER ALL CONTACT CASES.


Report No: 01 973-001 8-19

SECTION 8 EPILATUSW HANDLING, SERVICING, AND MAINTENANCE PC12

I CAUTION I THE ADDITIVE CONCENTRATION BY VOLUME SHALL BE A MINIMUM OF 0.06% AND A MAXIMUM OF 0.15%.

THE CORRECT MIX OF ANTI-ICING ADDITIVE IS IMPORTANT. CONCENTRATIONS OF MORE THAN 0.15% BY VOLUME WILL CAUSE DAMAGE TO THE PROTECTIVE PRIMER AND SEALANTS OF THE FUEL TANKS AND TO THE SEALS IN THE FUEL SYSTEM AND ENGINE COMPONENTS. CONCENTRATIONS OF LOWER THAN 0.06 VOL % MAY NOT BE ENOUGH TO INHIBIT ICE FORMATION.

MAKE SURE THAT THE ADDITIVE IS DIRECTED INTO THE FUEL STREAM. START ADDITIVE FLOW AFTER THE FUEL FLOW STARTS AND STOP THE ADDITIVE FLOW BEFORE THE FUEL FLOW STOPS. DO NOT ALLOW CONCENTRATED ADDITIVE TO CONTACT THE INTERIOR OF THE FUEL TANKS OR EXTERIOR PAINTED SURFACES.


For maximum sewice, keep tires inflated to the proper pressures. All wheels and tires are balanced before original installation, and the relationship to tire and wheel should be maintained upon reinstallation. Unbalanced wheels can cause extreme vibration in the landing gear; therefore, in the installation of new components, it may be necessary to re-balance the wheels with tires mounted. When checking the tire pressures, examine the tires for wear, cuts, bruises and slippage.

Nose Wheel Tire

Wheel type - BFG PN3-1501 Tire size - 17.5 x 6.25-6, 8PR, TL (160 mph) Tire Pressure - 60 +3 -0 psi (4.1 +0.2 -0 bar) Max. castor rotation - +I- 60" free (+I- 12' Nose Wheel Steering)

Main Wheel Tires

Wheel type - BFG PN3-1543 or 1543-1 (Post SB 32-013) I Tire size - 8.50-10, 8PR. TL (160 mph) Tire pressure - 55 +3 -0 psi (3.8 +0.2 -0 bar)

Refer to the AMM for the alternative types of tires that can be installed.

Report No: 01 973-001 8-20


=PlLATUSI SECTION 8 PC12 HANDLING, SERVICING, AND MAINTENANCE

LANDING GEAR - BRAKES

The fluid level should be checked periodically or at a scheduled maintenance event and replenished as necessary. Each brake assembly incorporates a brake lining wear indicator. As the brake pads wear, the pin will be pulled into the piston housing. When the system is pressurized and the pin is flush with the piston housing, the brake linings must be overhauled.

Refer to the AMM for complete information on the type of hydraulic fluid, servicing the fluid level and brake inspection and replacement.

HYDRAULIC SYSTEM

MSN 231 and UP. With the cargo door open, the hydraulic fluid level can be checked on the visual indicator installed on the hydraulic pressure tank in the wing root.

Servicing of the hydraulic system should only be performed by approved personnel with the correct maintenance equipment in accordance with procedures in the AMM. It is normally not required between annual inspections.

LUBRICATION POINTS

Proper lubrication is essential for trouble-free operation of mechanical components. Lubricants and dispensing equipment must be kept clean. Use only one lubricant in a grease gun or oil can. After lubrication, clean off all excessive grease or oil to prevent dust and dirt build-up.

The frequency of application may be increased for a particular type of operation or if excessive wear is experienced. For lubricating instructions, locations and lubricants refer to the AMM, Chapter 12.


Report No: 01 973-001 8-21

SECTION 8 =PILATUSW HANDLING, SERVICING, AND MAINTENANCE PC XI1

OXYGEN SYSTEM

( The standard oxygen system replenishment is carried out at a hinged service panel (1 1BR) on the right side of the fuselage, forward of the wing leading edge. The service panel is fitted with an oxygen replenishment valve and a system pressure gage. The gage is marked from 0 to 2000 psi, with a red zone from 1850 to 2000 psi. A charge pressure/temperature chart is installed on the inside of the service panel.

The larger capacity oxygen system replenishment is carried out at a hinged service door (31AB) on the bottom of the fuselage, rear of the wing trailing edge. An oxygen service panel is installed inside of the rear fuselage on the forward frame. The service panel is fitted with an oxygen replenishment valve and a system pressure gage. The gage is marked from 0 to 2000 psi, with a red zone from 1850 to 2000 psi. A charge pressureltemperature chart is also installed on the service panel.

I Replenishment Procedure

MAKE SURE THAT THE AIRPLANE IS FITTED WlTH A GROUNDING CABLE AND IS PROPERLY GROUNDED. THE OXYGEN CART MUST BE ELECTRICALLY BONDED TO THE AIRPLANE.

DO NOT OPERATE THE AIRPLANE ELECTRICAL SWITCHES OR CONNECTIDISCONNECT GROUND POWER DURING OXYGEN SYSTEM REPLENISHMENT.

DO NOT OPERATE THE OXYGEN SYSTEM DURING REFUELINGlDEFUELING OR ANY OTHER SERVICING PROCEDURE THAT COULD CAUSE IGNITION.

INTRODUCTION OF PETROLEUM BASED SUBSTANCES SUCH AS GREASE OR OIL TO OXYGEN CREATES A SERIOUS FIRE HAZARD. USE NO OIL OR GREASE WITH THE OXYGEN REPLENISHMENT EQUIPMENT.

ALWAYS OPEN SHUT-OFF VALVE SLOWLY TO AVOID GENERATING HEAT AND REPLENISH THE SYSTEM SLOWLY (MINIMUM TIME 6 MINUTES).

Report NO: 01973-001 8-22


=PILATUS= SECTION 8 -PC x F HANDLING, SERVICING, AND MAINTENANCE

REPLENISHMENT OF THE OXYGEN SYSTEM SHOULD ONLY BE CARRIED OUT BY QUALIFIED PERSONNEL.

Obtain the outside air temperature. (OAT). A fully charged cylinder has a pressure of 1850 psi at an temperature of 20°C. Filling pressures will vary depending upon the ambient temperature in the sewice bay and the temperature rise due to the compression of the oxygen. If the airplane is or has been parked outside in the sun. the temperature inside the fuselage will be appreciably higher than ambient. Figure 8-5 1 lists the required charging pressures for a range of temperatures.

Open the oxygen sewice panel 11BR on aircraft with the standard oxygen system. Open the service door 31AB on aircraft with the larger capacity oxygen system. I Hold the thermometer close to the oxygen cylinder. I Make sure the thermometer indication is constant. Make a note of the indication. I Refer to the temperaturelpressure graph for the correct oxygen cylinder pressure. I If the pressure on the service panel gage is low, fill the oxygen cylinder. I Make sure the area around the service panel charging valve is clean. Remove the cap from the charging valve.

Make sure the oxygen supply hose is clean and connect it to the charging valve. I Sbwly pressurize the oxygen cylinder to the correct pressure. I Close the oxygen supply and let the cylinder temperature become stable. I Monitor the oxygen pressure on the gage and fill to the correct pressure if necessary. I Release the pressure in the oxygen supply hose and disconnect from the charging valve.

Install the cap on the charging valve. Make sure the work area is clear of tools and other items.

Close the service panel 1 1 BR or the service door 31 AB. I Issued: February 14, 1994 Revision 8: September 1. 1998

Report No: 01973-001 8-23

SECTION 8 'F%PPATUSW HANDLING, SERVICING, AND MAINTENANCE ?c %I1 4

4

Figure 8-5. Oxygen Charging Pressures

Report NO: 01 973-001 I 8-24 Issued: February 14, 1994 Revision 1: June 10, 1994

'T$PM.ATUSW SECTION 8

?c XI1 HANDLING, SERVICING, AND MAINTENANCE

CLEANING AND CARE

WINDSHIELDBIDE WINDOWS

REMOVE WRIST-WATCHES, RINGS AND OTHER JEWELRY FROM HANDS AND WRISTS BEFORE CLEANING THE SlDE WINDOWS.

WINDSHIELDS AND WINDOWS ARE EASILY DAMAGED BY IMPROPER HANDLING AND CLEANING TECHNIQUES.

DO NOT USE SOLVENTS, FUELS, DETERGENTS, ALCOHOL, ACETONE OR THINNERS TO CLEAN THE SlDE WINDOWS.

TRANSPARENT PLASTICS LACK THE SURFACE HARDNESS OF GLASS. EXERCISE CAUTION WHEN CLEANING ALL THE SlDE WINDOWS TO AVOID SCRATCHING OR SCORINQ TRANSPARENCIES.

The following procedures provide informatbn reganling cleaning and sewicing of windshields and widows. Improper cleaning, or use of unapproved cleaning agents, can cause damage lo these surfaces. As a preventive measure, do not park the airplane where it might be subjected to dred contact with or vapor from: methanol, denatured alcohol, gasolne, benzene, xylene, MEK, acetone, ca&n tetrachloride, lacquer thinners, commeFciel or household window cleaning sprays, paint strippers or other types d solvents. Do not park airplane near a paint- spray shop.

Issued: Febluary 14,1994 Revision I: June 10,1894

SECTION 8 aPLATUSW HANDLING, SERVICING, AND MAINTENANCE ?c XI1

Windshield (Glass)

Place the ailplane inside a hanger or in a shaded area and allow lo cool from the heat of the sun's rays.

Using clean (preferably mnning) water, flood the surtace. Use bare clean hands, with no jewelry, to feel anddislodge any dirt or abrasive materials.

Using a mild soap or detergent (such as dsh washing liquid) in water, wash the surface. Again, use only the bare hand to provide tubbing force. (A clean lint-free doth may be used to bander the soap solulion to the surtace, but extreme care must be excised to prevent scralchiig the surface.)

Rinse the surface thoroughly with clean fresh water and dry wilh a dean doth or damp chamois leather.

Side Windows (Acrylic)

Flush wilh clean water to remove loose dust elc.

Wash the side windows using a soft sponge, warm water and soft soap mlution.

Rinse wiUl dean water and diy with a damp chamois leather.

Use an appropriate transparency cleaner to remove any grease, smears, etc., sUJl adhering to the side windows.

NOTE

Rubbing transparencies with a dry cloth wil cause scratches and the build-up of an electrostatic charge which attracts dust Where an electrostatic chaw is present gently pat the area with a damp chamois lealher to remove the &awe and any accumulated dust.

Repoll No: 01973-001 8-26

Issued: February 14,1894 Revision 1: June 10, 1994

'9saoPLPFTUSS SECTION 8

?C XI1 HANDLING, SERVICING, AND MAINTENANCE

EXTERIOR PAINT SURFACES

The airplane should be washed with a mild soap and water solution. Harsh abrasives or alkaline soaps or detergents could make scratches on painted or plastic surfaces or cause corrosion of metal. Cover areas where cleaning solutions could cause damage.

Exterior Recommended Cleaning Agents:

Mild soap or approved detergent.

Jet MULSO 2 (TURCO product) or equivalent.

To wash the airplane, use the following procedure:

NOTE

To prevent water from entering the pitoVstatic systems, the pitot tube openings and .the static ports should be blanked off. Exposed night control bearings should be protected prior to washing.

Rush away loose dirt with water.

Apply cleaning solution with a soft cloth, a sponge or a soft bristle btush. Do not allow the solution to dF/ before washing on. To remove exhaust stains, allow the solution to remain on the sutface longer.

To remove shlbbom dl and grease, use a cloth dampened with naphtha.

Rinse all sutfaces thoroughly.

Polish and seal the surfaces with a wax polish.

NOTE

Any good automothre wax may be used to preserve the painted surfaces. Soft lint-tree deaning cloths should be used to prevent scratches when deaning or polishing. A heavier coating of wax on the leading surfaces will reduce the abrasion problems in these areas, but see also paragraph 'DEICING BOOT CAREm.

Issued: June 10,1994 Revislon 1: June 10, 1994 - - -

Report NO: 01973-00 t 8-27

SECTION 8 -PLATUS W HANDLING, SERVICING, AND MAINTENANCE ?c %I)

DEICING BOOT CARE

The wings, T-tail, and propeller deicing boots have a special electrical~onductive coaling lo bleed off static charges which cause radio interference and may perforate the boots. Fueling

'

and other servicing practices should be done carefully to avoid damaging the conductive coating or tearing of Ule boots.

To prolong Ihe life of the deicing boots, lhey should be washed, with a mild soap and water , solution, rinsed with clean water, and serviced on a regular basis in accordance with the inst~ctions in the AMM. Keep the boots clean and free from oil, grease and olher solvents *

which cause neoprene to swell and deteriorate.

Report NO: 01 973-001 8-28

Issued: June 10,1994 Revision 1: June 10, 1994

WPLAflJSW SECTION 8

?c %I1 HANDLING, SERVICING, AND MAINTENANCE

PROPELLER CARE

Pmpefler care consists of checking the propeller area for leaks and damage, this also includes any damage to Ute propeller hub and deicing boots. Inspect the visible hub parts daily for surface damage. Look lor evidence of gtease and or oil leaks. Inspect the propeller blades, daily, lor scratches and gouges in the leading or trailing edge, or on the blade lace and camber surlaces.

NOTE

Any scratch or gouge, in a metal blade, must be repaired before next fflght. Otherwise, fatigue cracks may start, and the blade may fail in flight. Scratches and gouges in the outer 18 ln(457 mm) of the Made propeller diameter are especially critical because this is the area of highest vibratory stress.

Should any damage, scratches or gouges be found, obtain a qualified opinion prior to flight.

LANDING GEAR CARE

Before deaning the landing gear, place a plastic cover or similar material over the wheel and brake assembly.

Place a catch-pan under the gear to catch the waste.

Spray or bmsh the gear area with solvent or a mixture of solvent and degreaser, as desired. Where heavy grease and dirt deposits have collected, it may be necessary to brush the areas sprayed, in order to clean them.

Allow the solvent to m a i n on the gear from five to ten minutes. Then rinse the gear with additional solvent and allow to dry. I1 necessary help the drying process with a gentie blast d compressed air.

Remove the plasUc cover and the catchpan fmm the wheel.

Lubricate the gear in accordance with the Lubrication Chart in the AMM.

Issu~~: June 10,1994 Revision 1: June 10, 1994

---

SECTION 8 mPLATUSIB HANDLING, SERVICING, AND MAINTENANCE ?c XI1

ENGINE CARE

The engine exterior and compartment may be cieaned, using a suitable solvent Most efficient cleaning is done using a spry-type cleaner. Before spry cleaning, make sure the proteclon is afforded for components which might be adversely affected by the solvent. Refer to the AMM lor proper lubrication ol controls and components after engine cleaning.

INTERIOR CARE

The cockpit area should be lrequently vacuum-cleaned. lnslrument and side panels may be cleaned with a lint-lree cloth dampened wilh petroleum solvent (white spirit).

Seat hamesses that have been soiled may be deaned by gently scrubbing with a son brush, water and an approved soap. Alternatively, an offiially approved detergent emulsion may be . used when dluted in the proper proportions. Seats may be cleaned with a mild sdution of warm water and commercially-available upholstery deaner (as per manufachrrer's instructions).

Dust and loose dirt should be picked up regularly with a vawum-deaner. Stained carpets should be cleaned with a non-flammable dry deaning carpet shampoo which should be kept as dry as possible and again vacuumed.

1

Blot up any spilled liquid on the seats promply with cleansing tissue or rags. Do not pat the spot; press the blotting material finnly and hold it for several seconds. Continue blotting until no more liquid is taken up. Scrape off any sticky materials with a dul Mfe, then spotclean the area with a commercial cleaner, following the manufacturer's insttwtions.

Headliners, side panels and paint wok should be cleaned with a lint-free cloth dampened with ,

a mild soap and water mixture. Oil and grease can be removed with a sponge and common household detergent and then wiped dry with a clean rag.

Oxygen masks assemblies should be deaned with a suitable ol-free disinfedant. and then wipe dirt or foreign particles from the unit with a clean dry lint-free cloth.

Report No: 01973-001 8-30

Issued: June 10,1994 Revision 1: June 10,1994

WP1LA7USBB SECTION 8

PC XI1 HANDLING, SERVICING, AND MAINTENANCE

EXTENDED STORAGE

Probnged outof-service care applies to all airplane w h i i wil not be flown for an indefinite pefiod (less than 60 days) but which are to be kept ready-to-fly, with the least possible preparation. I f the airplane is to be stored temporarily, or indefinitely, reference must be made to the AMM for the proper storage procedures, which are all time related and classified as ~ o w c

stwe 1 Up to 7 days. Stage 2 7 to 30 days. Stage 3 30 to 90 days. Stage 4 More than 90 days.

Stages 1 and 2 are considered as flyable storage status.

No special sewice care is required for Stage 1 other than the airplane is moored and properly grounded, all covers and blanks are fitted, and that the fuel tanks are full. The engine may be left in an inactive state, with no presewabn protection, provided the engine is sheltered, humidity is not excessively high, and the engine Is not subjected to extreme temperature changes that would produce condensation. Where possible, cover the windshield with a light cotton dust cover.

Stage 2 storage, beglns alter Stage 1 (7 days) has elapsed, and includes placing desiccant bags and humidly indicators in the engine exhaust stubs and behind the exhaust stub coven. A suitable means must be provided to view the humidity indicators with the stub coven iI'IStaIled. An engine inactive lor up to 28 days requires no preservation provided all engine openlngs am sealed off and relathre humidity in the engine is maintained at less than 40%.

At 7 day intervals:

Check the Urn pmssures.

Drain any water lm the fuel system.

Check the humidity Indcetor, in the engine exhaust stubs,and replace the desiccant bags, if the humidity is in excess of 4096.

Issued: June 10,1994 Reviskn 1: June 10,1994


At 14 day intervals:

Movs the airplane to prevent flat areas on the tires. Ma* the tires with tape to ensure the tires are placed approximately 90 degrees from their previous position.

Stage 3 storage should be a planned situation, when lhe time difference can be foreseen but following on from the Stage 2, the engine fuel system would need to be presetved in accordance with the PBWC CMM.

At 30 day intervals:

Drain all fuel drain points and check for water accumulation. Prolonged dorage of the airplane will result in a water build-up in the fuel which 'leeches out' the EGME fuel additive. An indication of this is when an excessive amount of water accumulates at lhe fuel Win points. The concentration can be checked using a differential refractometer, but, it is imperahive that the technical manual for differential refractometer be fobwed explicitly when cheddng the additive concentration.

Slage 4 is a definite planned exercise, when deterioration of the airplane must be considered. An engine inactive for over 90 days in the airframe, or removed for long term storage, must In addition to the Stage 3 procedure, have the engine oil drained and filled with preserving oU in accordance with the PBWC CMM. Remove the battery and regularty check its state of charge.

To return lhe airplane to service, refer to the AMM for specific instructions.

Report No: 01 973-001 8-32

Issued: June 10,1994 Revision 1: June 10,1994 4

1

r PILOT'S OPERATING I HANDBOOK

Supplements

SECTION 9 SUPPLEMENTS

SECTION 9

SUPPLEMENTS

TABLE OF CONTENTS

Supp No. Subject Subject Report No.

GENERAL I

1. BendixMing KLN 90A GPS Navigation System 01 973-00119-1

2. BendiVKing RDS 82 VP Weather Radar 01 973-00119-2

3. BendixMing KHF 950 Communications System 0 1 973-001 19-3

4. BendixlKing KLN 908 GPS Navigation System 01 973-00119-4

5. Argus 5000 Electronic RMIlMoving Map Display 01 973-00119-5

BendixIKing Traffic Collision and Avoidance System 01 973-00119-6 CAS 66A TCAS 1

PC-1 2 Registered in Austria 01 973-00119-7

PC-1 2/45 1973-00119-8

BendiVKing RDR 2000 Weather Radar 1973-00119-9

Operations in Cold Conditions 1973-00119-10

PC-1 2 and PC-1 2/45 Registered in Canada 1973-00119-11

Engine Conditioning Trend Monitoring (ECTM) 1973-00119-12

Dual Channel Audii Controller System 1973-00119-1 3

Dual Air Traffic Control Transponder System 1973-00119-14

Ground Proximity Warning System (GPWS) 1973-00119-15

PC-12 and PC-12/45 Registered in Republic of South Africa 1973-00119-16

Mechanical Copilot Instrumentation 1973-00119-17

PC-1 2 Registered in France 1973-00119-18

Operation with GPS KLN 900 Navigation System 1973-00119-1 9

Issued: February 14, 1994 Revision 11 : March 1,2003

Report No: 01973-001 9-i

SECTION 9

SUPPLEMENTS

I Supp No. Subject

Second Pitot Static system

MSN 294 Specific Equipment

Emergency Power System

Multifunction Display (KMD 850)

Enhanced Ground Proximity Warning System (EGPWS)

Not applicable

Not applicable

Not applicable

Not applicable

Pilot's Relief Tube

Report No: 01973-001 1 9-ii

Report No.



GENERAL

This section provides information in the form of supplements for the operation of the airplane when equipped with optional equipment or systems which are not installed on the standard airplane. All of the supplements are FOCA Approved and are a permanent part of this Handbook.

The information contained in each supplement applies only when the specific equipment or system is installed in the airplane.

SUPPLEMENTS LIST OF EFFECTIVE PAGES

Supp. No. Page No. Rev. No. Supp. No. Page No. Rev. No.

9-01-1 thru 9-01-4 9-02-1 thru 9-02-4 9-03-1 thru 9-03-4 9-04-1 thru 9-04-1 2 9-05-1 thru 9-05-4 9-06-1 thru 9-06-6 9-07-1 thru 9-07-8 9-08-1 9-08-2 and 9-08-3 9-08-4 and 9-08-5 9-08-6 9-08-7 9-08-8 9-08-9 9-08-1 0 9-08-1 1 9-08-1 2 9-08-13 thru 9-08-1 7 9-08-1 8 thru 9-08-24 9-08-25 thru 9-08-30 9-08-31 9-08-32 thru 9-08-37 9-08-38 9-08-39 t h ~ 9-08-48 9-08-49 thW 9-08-58 9-08-59 9-08-60 and 9-08-61 9-08-62

Issued: February 14,1994 Revision 1 1 : March 1. 2003

9-08-63 and 9-08-64 9-08-65 9-08-66 and 9-08-67 9-08-68 9-08-69 thru 9-08-78 9-08-79 thru 9-08-81 9-08-82 thru 9-08-85 9-08-86 thru 9-08-89 9-08-90 9-08-91 and 9-08-92 9-08-93 thru 9-08-98 9-09-1 thru 9-09-4 9-10-1 thru 9-10-6 9-1 1-1 9-1 1-2 9-1 1-3 thru 9-1 1-5 9-11-6 9-11-7 thru 9-11-12 9-12-1 and 9-12-2 9-12-3 and 9-08-4 9-13-1 thru 9-13-3 9-14-1 thru 9-14-3 9-15-1 thW 9-15-4 9-16-1 thm 9-16-4 9-17-1 thru 9-17-5 9-18-1 9-1 8-2 9-1 8-3 and 9-18-4 9-18-5 t h ~ 9-18-7

Report No: 01 973-001 9-00-1


SUPPLEMENTS LIST OF EFFECTIVE PAGES

Supp. No. Page No. Rev. No. Supp. No. Page No. Rev. No.

9-1 8-8 and 9-1 8-9 9-18-10

19 9-19-1 thru 9-19-21 20 9-20-1 t h ~ 9-20-10 21 9-21-1 thru 9-21-4 22 9-22-1 t h r ~ 9-22-4 23 9-23-1 thru 9-23-6 24 9-41-1 thru 9-24-8 29 9-29-1 thru 9-29-3

Reparl No: 01 973-001 9-00-2


SECTION 9 SUPPLEMENT 1



SUPPLEMENT NO. 1 FOR

BENDM I KING KLN 90A GPS NAVIGATION SYSTEM

Thls supplement must be attached to the Plot's Operating Handbook and FOCA Approved Alrplane fllght Manual when the BenMing KLN 9OA Global Posilionlng System (GPS) is Installed in accordance with Pilatus Drawings. The information contahed herein supplements or supemedes the information in Ihe basic Pilot's OperaUng Handbook and FOCA Approved Airplane Flight Manual only in those m a s listed. For limitations, procedures end perfonence Information not contained in this supplement consult the baslc Pilot's Operating Handbook a d FOCA Appnwed Ahplane FII*t Manual.

Federal Offbe for Civil Aviation [FOCA) of Switzerland

Issued: June 10,1994 Revlskn 1: June 10,1994 -

Report No: 01973401 9-01 -1 I


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the BENDIXIKING KLN 90A GPS Navigation System installed in accordance with FOCA Approved Pilatus drawings.

SECTION 2 - LIMITATIONS

GPS equipment is limited to supplemental navigation in accordance with the cerhilcation limitations. The pilot must actively monitor approved operational equipment (e.0. IFR fit), required for the kind of operation.

KLN 90A Pilot's Guide must be immediately available to the pilot whenever supplemental navigation is predicated on the use of lhe GPS system. The operational revision Status (ORS) of the Pilot's Guide must match the ORS level of lhe KLN 9OA.

IFR navigation is restricted to enroute and terminal operation predicated on the use of the KLN 90A GPS system.

IFR approaches predicated on h e use of the KLN 90A GPS system are not approved.

IFR navigation is prohibited unless the pilol verifies that the date base is cumnt

The required standard IFR navigation systems appropriate to the route of flight must be available and operational for IFR navigation.

Outside of the area 74'N to 6 0 3 , magnetic variation must be manually entered.

The following placard is located on the instrument near the KLN 90A:

SECTION 3 - EMERGENCY PROCEDURES

When GPS hbrmatiin is flagged, utilize remaining operational appmved navigation equipment as required.

Report No: 01973-001 9 0 1-2

Issued: June 10.1994 Revision 1 : June 10,1994


SECTION 4 - NORMAL PROCEDURES

Normai operating procdtres are presented in the latest revision of the BendbdKing KLN 90A Pibls Guide.

GPS navigation data may be displayed on the navigation Instruments when GPS is selected.

GPS system message alert is duplicated on the EFIS, if GPS is selected as the Pdmary Navigation Source. The MSG annunciation wiil appear on the left side of the display above the Primary Navigation Source Annunciation.

SECTION 5 - PERFORMANCE

SECTION 6 - WEIGHT AND BALANCE

Factory installed optional equipment k included in the licensed weight and balance data in Sectbn 6 of the basic Pilot's Operating Handbook and FOCA Approved Airplane flight Manual.

SECTION 7 - DESCRIPTION

EQUIPMENT OPERATION AVAILABILITY

The Global Positioning System (GPS) consists of 21 satelites and three spares circling the earth twlce daily. These sateRites am positioned in six orbital paths approximately 10,900 NM ebove the earth. Each satellite contlnuoudy transmils timing (position) waveforms, orbital conection data, dock conectkn, and once each hour an updated almanac. The GPS receiverlcomputer unit determines present position by decoding end deltacalculation of bi- faced unprobted CIA oode horn different satellites. Far continuous three-dimensional positkn determining, at least 4 satellites must be Msible' with a mkrhnum signal strength to noise ratio. With only 3 sate9ites avaliable to the unit, two dimensional navigation is possible by using an Air Data Computer or Encoding Altkneter.

Issued: June 10, 1994 Revidon 1: June 10,1994 - -

Report No: 01973-001 9-01 -3 1


EQUIPMENT DESCRIPTION

The basic installation consists of lhe panel mounted unit, a data cartridge, a omnidredknal flat micro strip antenna with internal preamplifier and at least one source of altilude input and intellaces for CDI, RMI, and EFIS. The actual interlace will be determined according to the equipment installed.

The BendixlKing KLN 9OA GPS is a panel mounted, fast sequencing MulUSaf L1 frequency (1575.42 MHz) satellite tracking. CIA coded GPS area navigation system. An opUonal Jeppesen NorVl America or Jeppesen International database &ridge, updated every 28 days, is available. The database cartridge is an electronic memory containing hformation on airpow, navaids, intersections, special use airspace and other items of value to the pilot The database may also be updated with the use of an IBM XTlAT or compalible personal computer using the RS232 serial data bus port, present on the maintenance panel.

The BendixlKing KLN WA GPS has the ability to create and store up to 26 flight plans, each containing as many as 20 way points. An addilonal250 user generated way points may be stored as USER, NAVAID or AIRPORT data and which may be used in creating flight plans. The BendixlKing KLN WA GPS navigation computer supplies bearhg and range to and from any way point, vertical navigation features, area and allitude warnings and Hight (trip) planning. For stand alone applications, all navigation data inckrdng CDI, map presentation, and messages are presented on the display of the KLN QOA. Standard navigation data is provided to CDI, RMI and EFIS.

NOTE

Detailed operating instructions or technical information for the Bendix/King K I A 90A GPS may be obtained by conurlllng Ihe latest available revision of Plot's Guide.

Report No: 01 973-001 9-01-4

Issued: June 10,1994 Revision 1: June 10, 1994





BENDM I KING RDS 82 VP VERTICAL PROFILE WEATHER RADAR

This wrpplement must be attached to the Pilots Operating Handbook and FOCA Appmved Ahplane Right Manual when the BendWKimg RDS 82 VP Vertical Pmlile Weather Radar Is instalid in accordance with Pdatus Drawings. The Information conlakred herein supplements or supersedes the information in the basic Plbts Operating Handbook and FOCA Approved Alrplane Flight Manual only h those areas Wed. For Umllatbns, pmerhres and perlonnance information not contained in this supplement, consul the bask Pilot's Operating Handbook and FOCA Approved Ahplane Rlghl Manual.

by:

I- June 10 , lW iM4slon 2: February 14,1995

Report No. 01973-001 942-1


SECTION 1 - GENERAL

This supplement supplies lhe infomalion necessary for We operalon of We airplane when the BENDIXIKING RDS 82 VP Vertrcal Profile Wealher Radar is installed in accordance wilh FOCA Approved Pilatus drawings.


Do not operate the radar during refueling operations or in the vicinity of t ~ c k s or containers accommodating flammables or explosives.

Do not allow personnel within 15 feet (4.5 meters) of area being scanned by antenna when system is transmitting.

Vefical profile (VP) weather display is not allowed on the EHSl dsplay (mode inhibited).


No changes to the Emergency Procedures described in the basic Pilot's Opemting handbook.


Preflight and normal operating procedures are presented in the latest reviSi0II of the BendidKing RDS 82 VP Vertical Profile Radar Plots Guide.

TESTS INVOLVING THE RADIATION OF RF ENERGY BY THE RADAR ANTENNA MUST NOT BE MADE WHILE THE RADAR ANTENNA IS DIRECTED TOWARD CLOSE-BY LARGE METAL OBJECTS SUCH AS HANGARS, DOORS, OR THE INSIDE OF A HANGAR. USE TEST MODE OR TURN THE INDICATOR FUNCTION SWITCH TO OFF WHERE APPLICABLE.

Report No: 01973-001 902-2

Issued: June 10,1994 Revision 2: February 14,1995



No change.


Factory installed optional equipment is included in the licensed weight and balance data in Section 6 of the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.


The RDS 82 VP Weather Radar installation consists of a ReceiverKransrnilter unit in a pod located at the right wing tip. Weather indication is provided on the pilots EFlS and supplemented on the co-pilots EFlS (if installed), the EFlS MFD (if installed) andlor the independent WX indicator.

I

NOTE

Detailed operating instructions or technical information for the B e n d i i n g RDS 82 VP Vertical Profile Weather Radar may be obtained by consulting the latest available revision of the Pilot's Guide.

Issued: June 10,1994 Revlslon 2: February 14,1995

Report No: 01 973-001 9-02-3

SECTION 9 '%EPILATUSW SUPPLEMENT 2 PC %(I 4

4


Report No: 01973401 942-4

Issued: June 10,1994 Revision 2: Febtuary 14,1995 I





BENDIX / KING KHF 950 COMMUNICATIONS SYSTEM

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the BendixMing KHF 950 Communications System is installed in accordance with Pilatus Drawings. The informallon conteined herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane flight Manual.

Federal Office for Civil Aviation (FOCA) of Switzerland Sect i i for Type Certification

Date of Approval:

1 0. April 199

Issued January 31,1995

-

Report No: 01 973-001 9-03.1

SECTION'O 'IEPLATUS= SUPPLEMENT 3 ?C xb

SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the BENDlWKlNG KHF 950 Communications System is installed in accordance with FOCA Approved Pilatus drawings.

SECTlON 2 - LIMITATIONS

The KHF 950 Pilot's Guide (KPN 006-08343-0002 or later revisions) must be immediately available to the pilot whenever this HF Communication System is used.

Do not operate the HF Communications System when ground power is connected.


No changes to the Emergency Procedures described in the basic Pilot's Operating handbook.


Preflight and normal operating procedures are presented in the latest revision of the . BendidKing KHF 950 Pilot's Guide.

Report No: 01 973-001 9-03-2

Issued: January 31,1995



No change.

SECTION 6 -WEIGHT AND BALANCE

Factory installed optional equipment is included in !he licensed weight and balance data in Section 6 of the basic Pilors Operating Handbook and FOCA approved Airplane Flight Manual.


The High Frequency (HF) communication system gives long-distance voice cornmwdcation between aircraft and a ground station. The system consists of a control unit, a receiver, a coupler and an antenna.

De!ailed operating instructions or technical information for the ~ e n d i x ~ i k KHF 950 Communications System may be obtained by consulting the latest available version of the Pilors Guide. KPN 006-08343-0002 or latest available revision.

Issued: January 31.1995

I

Report No: 01 973-001 9-03-3

SECTION 9 'EPLATUSW I

SUPPLEMENT 3 ?c %!I 4 4


Report No: 01973-001 9-03.4






BENDIX I KING KLN 90B GPS NAVIGATION SYSTEM

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the BendixlKing KLN 908 Global Positioning System (GPS) is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement. consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

/& Approved by:

Federal Office for Civil Aviation (FOCA) of Switzerland Section for Type Certilication

Date of Approval:

2 8. April 1998

Issued: March 13, 1995 Revision 2: 10 February, 1998

Report No: 01973-001 9-04- 1

SECTION 9 WPILATUS€ SUPPLEMENT 4 ?c XI1

SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the BENDIWKING KLN 908 GPS Navigation System is installed in accordance with Modification Approval Sheet PIL 12/341009.


A. The KLN 90B GPS Pilot's Guide. PIN 006-08773-0000, dated December, 1994 (or later applicable revision) must be immediately available to the flight crew whenever navigation is predicated on the use of the system. The Operational Revision Status (ORS) of the Pilot's Guide must match the ORS level annunciated on the Self Test page.

B. IFR Navigation is restricted as follows:

(1) The system must utilize ORS level 20 or later approved revision.

(2) The data on the self test page must be verified prior to use. Verify valid altitude data is available to the KLN 90B prior to flight.

(3) IFR en route and terminal navigation is prohibited unless the pilot verifies the currency of the data base or verifies each selected waypoint for accuracy by reference to current approved data.

(4) Instrument approaches must be accomplished in accordance with approved instrument approach procedures that are retrieved from the KLN 90B data base. The KLN SOB data base must incorporate the Current update Cycle.

(a) lnslrument approaches must be conducted in the approach mode and RAlM must be available at the Final Approach Fix.

(b) APR ACT (approach active) mode must be annunciated at the Final Approach Fix.

(c) Accomplishment of ILS, LOC, LOC-BC and LDA approaches are not authorized.

(d) When an alternate airport is required by the applicable operating rules. it must be served by an approach based on other than GPS navigation.

Report No: 01973-001 9-04-2

Issued: March 13,1995 Revision 2: '10 February. 1998


(e) The KLN 90B can only be used for approach guidance if the reference co-ordinate datum system for the instrument approach is WGS-84 or NAD-83. (All approaches in the KLN 906 data base use the WGS-84 or the NAD-83 geodetic datum.)

(5) The use of the KLN 906 to perform Basic RNAV (8-RNAV) operations in the designated European airspace is limited as follows:

(a) The system must have the GPS XPRESS software update with the following display on the STA 3 page:

HOST 01 515-0046 RCVR 01621-0001

(b) The KLN 906 data base must incorporate the current update cyde.

(c) Given a GPS constellation of 23 satellites or less (22 or less when the KLN 906 incorporates automatic pressure altitude aiding), the availability of RAIM must be confirmed for the intended flight (route and time). Dispatch for B-RNAV must not be made in the event of predicted continuous loss of RAIM of more than 5 minutes for any part of the intended flight.

(d) Radar coverage must be available for the route to be flown.

(6) Traditional approved navigation equipment (e.g. VOR, DM€. ADF) appropriate to the rwte of flight must be installed and operational in accordance with the applicable operational rules.


ABNORMALPROCEDURES

A. If the KLN 90B GPS information is not available or invalid, utilize traditional operational navigation equipment as required.

8. If the alarm limit is exceeded during 6-RNAV operation, revert to an alternative means of navigation.

C. If a 'RAIM NOT AVAILABLE" message is displayed while conducting an instrument approach, terminate the approach. Execute a missed approach if required.

D. If a 'RAIM NOT AVAILABLE" message is displayed during 8-RNAV operation, navigation may continue using KLN 90B provided cross-check with VOR. DME and

Issued: March 13.1995 Revision 2: 10 February. 1998 3 of 12

Report No: 01973-001 9-04-3


NDB information shows adequate airplane positioning performance (RNPS). Otherwise, revert to an alternative means of navigation

E. Refer to the KLN 90B Pilot's Guide, Appendices B and C, for appropriate pilot actions to be accomplished in response to annunciated messages.


I WARNING I

FAMILIARITY WITH THE EN ROUTE OPERATION OF THE KLN 908 DOES NOT CONSTITUTE PROFICIENCY IN APPROACH OPERATIONS. DO NOT ATTEMPT APPROACH OPERATIONS IN IMC PRIOR TO ATTAINING PROFICIENCY IN THE USE OF THE KLN SOB.

OPERATION

(A) Normal operating procedures are outlined in the KLN 90B GPS Pilot's Guide, PIN 006-08773-0000, dated December, 1994. (or later applicable revision).

(8) During B-RNAV operation, traditional navigation equipment (e.g. VOR. OME and ADF) should be selected lo available ground based aids so as to allow immediate cross- checking or reversion in the event 01 loss of GPS navigation capability.

I (C) Before joining and during operation on an RNAV route, correct operation of the RNAV system shall be established. This includes that:

I - the routing is in accordance with the clearance, and

I - the navigation performance accuracy meets RNP5.

SYSTEM ANNUNClATORSlSWlTCHESlCONTROLS

1 (A) GPS Status (MSG) annunciation - The message (MSG) annunciation on the EFlS display, will flash to alert the pilot of a situation that requires attention. Press the MSG button on the KLN 900 GPS to view the message. (Appendix B of the KLN 90B Pilors Guide contains a list 01 all of the message page messages and their meanings).

( (B) The waypoint (WPT) annunciation on the EFlS display, will be activated prior to reaching a waypoint in the active flight plan. The KLN 908 GPS will provide navigation along a curved path segment to ensure a smooth transition between two adjacent legs

Report No: 01973-001 9-04-4

Issued: March 13,1995 4 01 12 Revision 2: 10 February, 1998


in the flight plan. This feature is called turn anticipation. Approximately 20 seconds prior to the beginning of turn anticipation the WPT annunciator will flash, going solid upon initialization of the turn. and extinguishing upon turn completion.

TURN ANTICIPATION IS AUTOMATICALLY DISABLED FOR FAF WAYPOINTS 'AND THOSE USED EXCLUSIVELY IN SIDISTARS WHERE OVER-FLIGHT IS REQUIRED. FOR WAYPOINTS SHARED BETWEEN SIDISTARS AND PUBLISHED EN ROUTE SEGMENTS (REQUIRING OVER- FLIGHT IN THE SIDISTARS), PROPER SELECTION ON THE PRESENTED WAYPOINT PAGE IS NECESSARY TO PROVIDE ADEQUATE ROUTE PROTECTION ON THE SIDISTARS.

(C) GPS omni bearing or leg (CRS OBS I LEG) course switcWannunciator - Used to select the basic modes of KLN 908 operation, either:

1) single waypoint with omni - bearing course (08s) selection through that waypoint (like a VOR)

2) automatic leg sequencing (LEG) between waypoints.

NOTE

Either LEG or OBS will illuminate during system self test depending upon switch position.

01 (D) EFlS course control . knob - Provides analog course input to the KLN 908 in OBS when the GPS source is selected on the EFIS. When other than GPS navigation sources are selected. GPS course selection in OBS mode is digital through the use ol the controls and display at the KLN 908.

(E) GPS approach (APR ARMIACT) switcWannunciator - Used to a) manually select or deselect approach ARM (or deselect approach ACT) and b) annunciate the stage of approach operation either armed (ARM) or activated (ACT). Sequential button pushes

Issued: March 13. 1995 Revision 2: 10 February, 1998 5 of 12

Report No: 01973-001 9-04-5


if in ACT would first result in approach ARM and then approach a m cancelled. Subsequent button pushes will cycle between the armed state (if an approach is in the I

flight plan) and approach arm cancelled. Approach ACT cannot be selected manually.

Leftlright steering information is presented on the EFlS as a function of the navigation source selection on the EFlS control panel.

AUTOPILOT COUPLED OPERATION

The KLN 900 may be coupled to the autopilot. The autopilot approach mode (APR) should be used when conducting a coupled GPS non-precision approach, according to the procedure indicated.

NOTE

Select autopilot HDG mode for DME arc intercepts. Using NAV or APR for coupled DME arc intercepts can result in excessive overshoots.

APPROACH MODE SEQUENCING AND RAlM PREDICTION

NOTE

The special use airspace alert will automatically be disabled prior to flying an instrument approach to reduce the potential lor message congestion.

(A) Prior to arrival, select a STAR if appropriate from the APT 7 page. Select an approach

I and an Initial Approach Fix (IAF) from the APT 8 page.

Report No: 01973-001 9-04-6

Issued: March 13, 1995 Revision 2: 10 February. 1998


NOTE

Using the right hand outer knob, select the ACT (Active Flight Plan Waypoints) pages. Pull the right hand inner knob out and scroll to the destination airport, then push the inner knob in and select the ACT 7 or ACT 8 page.

To delete or replace a SID, STAR or approach, select FPL 0 page. Place the cursor over the name of the procedure, press EN1 to change it, or CLR then EN1 to delete it.

(6) En route. check for RAIM availability at the destination airport ETA on the STA 5 page.

NOTE

RAIM must be available at the FAF in order to fly an instrument approach. Be prepared to terminate the approach upon loss of RAIM.

(C) At 30 NM lrom the FAF:

(1) Verify automatic annunciation of APR ARM.

(2) Note automatic dbar scaling change lrom i 5.0 NM to i 1.0 NM over the next 30 seconds.

(3) Update the KLN 908 altimeter bar0 setting as required.

(4) Internally the KLN 906 will transition from en route to terminal integrity monitoring.

(D) Select GPS NAV 5 page or applicable EFlS presentation, to fly the approach procedure.

(1) If receiving radar vectors, or need to fly a procedure lurn or holding pattern, fly in OBS until inbound to the FAF.

Issued: March 13. 1995 Revision 2: 10 February, 1998

Report NO: 01973-001 9-04-7


NOTE

OBS navigation is TO-FROM (like a VOR) without waypoint sequencing.

(2) NoPT routes including DME arc's are flown in LEG.

I IFAF) to M

NOTE

HOG mqde for DMF arc ioterceots. NAV or APR coupled DME arc intercepts can result in excessive overshoots (aggravated by high ground speeds a d o r intercepts from ioslde the arc).

FLYING FINAL OUTBOUND FROM AN OFF-AIRPORT VORTAC ON AN OVERLAY APPROACH; BEWARE OF THE DME DISTANCE INCREASING ON FINAL APPROACH, AND THE GPS DISTANCE-TO-WAYPOINT DECREASING, AND NOT MATCHING THE NUMBERS ON THE APPROACH PLATE.

(E) . At or belore 2 NM from the FAF inbound:

(1) Select the FAF as the active waypoint, if not accomplished already.

(2) Select LEG operation.

(F) Approaching the FAF inbound (within 2 NM.):

(1) Verily APR ACT.

(2) Note automatic dbar scaling change from * 1.0 NM to * 0.3 NM over the 2 NM inbound to the FAF.

Report No: 01973-001 9-04-8

Issued: March 13, 1995 Revision 2: 10 February. 1998


(3) Internally the KLN 908 will transition from terminal to approach integrii monitoring.

(G) Crossing the FAF and APR ACT is M annunciated:

(1) Do not descend.

(2) Execute the missed approach.

(H) Missed Approach:

(2) Navigate to the MAP (in APR ARM if APR ACT is not available).

NOTE

There is no automatic LEG sequencing at the MAP.

in accordance with the published missed approach procedure, I verify or change the desired holding fix and press ENT.

GENERAL NOTES

The data base must be up to date for non-precision instrument approach operation.

Only approach can be in the active flight plan at a time.

If the destination airpdrt is the active waypoint at the time ol the instrument approach selection, the active waypoint will shift automatically to the chosen IAF.

Checking RAIM prediction for your approach while en roule using the STA 5 page is recommended. A self check occurs automatically within 2 NM of the FAF. APR ACT is inhibited without RAIM.

Data cannot be allered, added to or deleted from the approach procedures contained in the data base. (DME arc intercepts may be relocated abng the arc through the SUPER NAV 5 or the FPL 0 pages).

Issued: March 13,1995 Revision 2: 10 February. 1998

SECTION 9 rPlLATUSP SUPPLEMENT 4 ?C XI1

3

Some approach waypoints do not appear on the approach plates (including in some instances the MI I

Waypoint suffixes in the flight plan:

I IAF

1 F AF

m MAP

h missed approach holding fix. I

The DME arc IAF (arc intercept waypoint) will be:

(a) On your present position radial off the arc VOR when you bad the IAF into the flight plan.

(b) The beginning of the arc if currently on a radial beyond the arc limit.

To adjust the arc intercept to be compatible with a current radar I

vector, bring up the arc IAF waypoint in the SUPER NAV 5 page scanning field, or applicable EFlS GPS presentation, or under the cursor on the FPL 0 page, press CLR, then ENT. Fly the arc in LEG. Adjusl the HSI or CDI course pointer with reference to the desired lrack value on the SUPER NAV 5 page, or applicable EFlS GPS presentation (it will flash to remind you). Leftfright dbar inlormalion is relative to the arc. Displayed distance is along the arc but direct to the active waypoint. If desired. select NAV 2 page for digital OME arc distance to and radial from the relerence VOR.

The DME arc IAF identifier may be unfamiliar. Example: D098G where 098 stands lor the 09E0 radial off the referenced VOR, and G is the seventh lener in Ihe alphabet indicating a 7 DME arc.

APR ARM to APR ACT is automatic provided:

(a) You are in APR ARM (normally automatic).

Report No: 01973-001 9-04- 10

Issued: March 13. 1995 10 of 12 Revision 2: 10 February. 1998


(d) Within 2 NM of the FAF.

(e) Outside of the FAF.

(1) Inbound to the FAF.

(g) RAlM is available.

Direct-To operation between the FAF and MAP cancels APR ACT. Fly the missed approach in APR ARM.

Flagged navigation inside the FAF may usually be restored (not guaranteed) by pressing the GPS APR button changing from ACT to ARM. Fly the missed approach.

The instrument approach using the KLN 908 may be essentially automatic starting 30 NM out (with a manual barometer setting update) or it may require judicious selection ot the OBS and LEG modes.

APR ARM may be cancelled at any time by pressing the GPS APR button. (A subsequent press will reselect it.)

SECTION V - PERFORMANCE

No change.



Issued: March 13.1995 Revision 2: 10 February. 1998

Report No: 01973-001 9-04- 1 1



The KLN 906 GPS panel mounled unit contains the GPS sensor, the navigation computer, a CRT display, and all controls required to operate the unit. It also houses the data base 1

cartridge which plugs directly into the back of the unit. ,

The data base cartridge is an electronic memory containing inlormation on airports, navaids. intersections, SID's, STAR'S, inslrumenl approaches, special use airspace, and other items of value to the pilot.

Every 28 days, BendidKing receives new data base information from Jeppesen Sanderson. This inlormation is processed and downloaded onto the dala base canridges. BendixIKing makes these data base cartridge updates available to KLN 90B GPS users.

Provided the KLN 906 GPS navigation system is receiving adequate usable signals, it has I

been demonstrated capable of and has been shown lo meet the accuracy specifications of:

VFRJIFR supplemental en route oceanic and remote, en route domestic, terminal, and non- precision instrument approach (GPS, Loran-C, VOR, VOR-DME, TACAN. NDB. NDB-OM€, RNAV) operation within latitudes bounded by 74" North and 60° South using the WGS-84 (or NAD 83) coordinate relerence datum in accordance with the criteria of AC 20-138. AC 91-49, and AC 120-33. Navigation data is based upon use of only the global positioning system (GPS) operated by the United States.

NOTE

Airplane using GPS for oceanic IFR operations may use the KLN 900 to replace one of the other approved means of long- range navigation. A single KLN 90B GPS installation may also be used on short oceanic routes which require only one means ol long-range navigation.

Report No: 01 973-001 I 9-04-12

Issued: March 13.1995 Revision 2: 10 February. 1998

=PLATUSE SECTION 9 -PC XII SUPPLEMENT 5




ARGUS 5000 ELECTRONIC RMWOVING MAP DISPLAY

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane, Flight Manual when the ARGUS 5000 Electronic RMllMoving Map Display is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by: &$& Federal Office for Civil Aviation (FOCA) of Switzerland

Section for Type Certification

Date of Approval:

6. Dez. 1995

Issued: September 30. 1995

- -

Report No: 01 973-001 9-05-1

SECTION 9 =PILATUSS SUPPLEMENT 5 -?c xr

SECTION 1 - GENERAL

Thts supplement supplies the information necessary for the operation of the airplane when the ARGUS 5000 ELECTRONIC RMIIMOVING MAP DISPLAY is installed in accordance with FOCA Approved Pilatus drawings.


The eventide ARGUS 5000 PILOT'S GUIDE, PIN 5009, Rev. 4 (or later revision) must be immediately available to the flight crew whenever navigation is predicated on the use of the ARGUS 5000.

The ARGUS 5000 utilization is limited for GPS andlor ADF navigation display.

If opttonal RMI adapter is installed, the ARGUS 5000 use may be extended for VHF navigation dtsplay.

The ARGUS 5000 moving map presentation must not be utilized for conducting GPS guided instrument approaches or departures.

IFR navigation must not be predicated on the use of ARGUS 5000 unless it's data base is current. The currency of the ARGUS 5000 data base must be verified before departure.


If sensor information is intermittent or lost, utilize alternate navigation equipment as required.

SECTION 4 - NORMAL PROCEDURES I

Normal operating procedures are presented in the eventide ARGUS 5000 PILOT'S GUIDE, PIN 5009, Rev 4 (or later revision).


No change.

1

I Report No: 01973-001 9-05-2

Issued: September 30. 1995

I

L =PW.ATUS= SECTION 9 I

r -PO XC SUPPLEMENT 5

1 SECTION 6 - WEIGHT AND BALANCE

1 . Factory installed optional equipment is included in the licensed weight and balance data in

t Section 6 of the basic Pilot's Operating Handbook and FOCA approved Airplane Flight Manual.

I t SECTION 7 - DESCRIPTION

L INTRODUCTION

The Eventide Avionics ARGUS 5000 Moving Map Display is a pictorial navigation instrument which provides visual reference of aircraft position relative to landing facilities, navigational aids and special use airspace. The ARGUS 5000 requires present geodetic position and other navigation information from a Global Positioning Satellite system (GPS). Map graphics and bearing displays of ARGUS 5000 relate to the airplane's stabilized direction system. The standard ARGUS 5000 displays an ADF bearing pointer and digital magnetic bearing readout from compatible ADF receivers or indicators. With an optional RMI Adapter, the ARGUS 5000 may display bearing pointer(s) and digital magnetic bearing readouts(s) from compatible VHF navigation receivers. The ADF MODE or the RMI MODE may be used for IFR approaches, independent of other navigation systems. The bearing pointer(s) may be superimposed on the map graphics screen. Optional Flight Planning software is available for the ARGUS 5000 to allow users to add, delete, and activate user-entered waypoints and flightplans directly on the ARGUS 5000.

SYSTEM DESCRIPTION

k

P Reference must be made to the Eventide ARGUS 5000 PILOT'S GUIDE, PIN 5009, Rev. 4 (or later revision) for descriptive and operational details.

' The bask ARGUS 5000 System comprises a Pilot andlor Co-Pilot Indicator.

The indicators receive digital and analog sensor inputs from the airplane systems and process ' the data for presentation on the display unit(s). The system inputs which may be processed include AHRS, NAV (VOR, ADF, LNAVNNAV).

1 The controls are used to select the desired display format. NAV source data and modes.

Issued: September 30,1995 t

Report No: 01973-001 9-05-3


Report No: 01973-001 9-05-4


Issued: September 30, 1995

--- - -





BENDIX / KING TRAFFIC COLLISION AND AVOIDANCE SYSTEM CAS 66A TCAS I

This supplement musl be allached lo the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when BendidKing Traffic Collision and Avoidance Syslem (CAS 66A TCAS I) is Installed in accordance wilh Pilalus Drawings. The Informalion conlained herein supplements or supersedes Ihe information in the basic Pilot's Operaling Handbook and FOCA Approved Airplane Fiighl Manual only in those areas listed. For lirnilallons, procedures and performance inlormalion no1 conlained in lhis supplement, consul1 the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office lor Civil Aviation (FOCA) of Switzerland

Date 01 ~pprov'al:

2 4. JULl 2001

Issued: May 1,1996 Revision 1 : April 2, 2001

Report No: 01973-001 ' 9-06-1

SECTION 8 'EPILATUSS SUPPLEMENT 6 PC12

SECTION 1 - GENERAL

This supplement supplies the information necessaly for the operation of the airplane when the BENDIXIKING Traffic Collision and Avoidance System - CAS 66A TCAS I is installed in

I accordance with Modification Approval Sheet PIL 12/34/01 7 or PIL 12/34/109.


The BendidKing CAS 66A TCAS I Pilot's Guide (BendidKing part no. 006-08746-0000 Rev. 0, Dated 5/93 or later revision) must be readily accessible to the flight crew when operating the CAS 66A TCAS I system.

The pilot should not manoeuvre the aircraft based on the traffic display only. The traffic display is intended to assist in visually locating the traffic. The traffic display lacks the resolution necessary for use in evasive manoeuvring.

I Maximum Intruder display limit on the KMD 850 or EFlS MFD is set to 20.

TCAS 1 is unable to detect any intruding aircraft without an operating transponder. TCAS 1 can select and track aircraft with either ATC operating Mode A. C or S.


No change.


NOTE

IF ATTITUDE SOURCE, ALTITUDE SOURCE AND/OR RADIO ALTIMETER IS INOPERATIVE, TCAS WILL BE INOPERATIVE.

1. Normal TCAS operation is described in the BendidKing CAS 66A TCAS I Pilot's Guide (BendidKing part no. 006-08746-0000 Rev. 0, Dated 5/93 or later).

2. TCAS I Non-Threat Traffic:

Report No: 01 973-001 9-06-2

AURAL

None

Issued: May 1, 1996 Revision 1 : April 2, 2001 ,

VISUAL

An open white diamond on the traffic display.

DEFINITION

Intruder relative altilude is greater than rt 1200ft or distance is beyond 5 NM range.

CREW RESPONSE

None, not considered as threat.


3. TCAS I Proximity lntruder Trafflc:

4. TCAS I Traffic Advisory (TA):

AURAL

None

THE TRAFFIC DISPLAY IS ADVISORY ONLY. DO NOT MANOEUVRE THE AIRCRAFT BASED ONLY ON TRAFFIC DISPLAY INFORMATION. ATTEMPT TO VISUALLY ACQUIRE THE INTRUDER OR CONTACT ATC BEFORE YOU MANOEUVRE THE AIRCRAFT.

VISUAL

A filled white diamond on the traffic display.

NOTE

AURAL

TRAFFIC TRAFFIC

Aural alerting is enabled at 600 11. AGL climbing and inhibited at 400 ft. AGL descending.

DEFINITION

Intruder relative altitude is within f 1200ft and within 5 NM range.

VISUAL

A filled yellow circle on the traffic display.

DEFINITION

Intruder time to closest point of approach is 15 to 30 sec.

NOTE

CREW RESPONSE

None, not considered as threat.

CREW RESPONSE

Conduct visual search for the Intruder. If successfui, maintain visual acquisition to ensure safe operation.

In most situations no manoeuvre will be necessary to maintain safe separation. Manoeuvre only if it becomes apparent safe separation will not be maintained.

a. Attempt to visually acquire the Intruder aircraft and maintaintattain safe separation in accordance with regulatory requirements and good operating practice.

Issued: May 1, 1996 Revision 1 : April 2, 2001

Report No: 01 973-001 9-06-3


b. If the lntruder aircraft is not visually acquired, air traffic control should be contacted to obtain any information that may assist concerning the lntruder aircraft.

c. Minor adjustments to the vertical flight path consistent with air traffic requirements are not considered evasive manoeuvres.

NOTE

Traffic Advisories (TA's) can be expected to occur during normal flight operation. Generally, TA's will occur more frequently in terminal areas during amval, and less frequently during departure and en route operations. In the vast majority of these cases, the aircraft displayed will be safely separated and there will be no need for pilots to initiate any avoidance manoeuvres.

NOTE

Evasive manoeuvres (rapid change in pitch, roll, normal acceleration, thrust or speed) should only be conducted after visual acquisition of the lntruder and then only when necessary to achieve or assure sate separation.


No change.



Report No: 01973-001 9-06-4

Issued: May 1,1996 Revision 1 : April 2, 2001 ,



The TCAS I is an on-board traffic alert and collision avoidance system which includes a display. The system detects and tracks other (Intruder) aircraft by interrogating their transponders. From the transponder replies, TCAS I determines range, bearing and (if the lntruder is equipped with a Mode C or S transponder) relative altitude. Intruders equipped with a Mode A transponder do not provide altitude information. With this data, the TCAS I uses standard algorithms to determine the threat of collision. When a possible collision hazard exists, the TCAS I issues a visual and aural Traffic Advisory (TA) to the flight crew. The TCAS I will not detect aircraft which have no operating transponder.

The TCAS I Is a single system installation consisting of one TCAS I processor, one top- mounted bearing anten&, one bottom-mounted bearing antenna and the EFlS MFD or KMD 850. Aural alerts are available through the headphones and cockpit speaker. I For Aircraft with an EFlS MFD, system control is through the EFlS MFD CP469A control panel. For aircraft with a KMD 850 installed system control is through the KMD 850. The traffic display I is informative only, displaying area traffic without attempting to provide any form of conflict resolution.

If an Intruder gets to within 15 to 30 seconds of a projected Closest Point of Approach (CPA) andlor meets other range and closure criteria, it is then considered a potential threat and an aural and visual TA is issued with the voice message, "TRAFFIC TRAFFIC". This assists the pilot in achieving visual acquisition of the threat traffic.

TCAS I Is intended as an aid to the see and avoid concept. Once an lntruder is visually acquired, it is the pilot's responsibility to manoeuvre as necessary to maintain safe separation.

TCAS I does not Incorporate the sophisticated sensors, bearing accuracy or track rate computations incorporated in TCAS II or TCAS Ill that are necessary for evasive manoeuvring (rapid change in pitch, roll, normal acceleration, thrust or speed). In general, TCAS I does not provide adequate information for pilots 10 determine rellably which horizontal or, in some cases, vertical direction to move to increase separation, and there is some likelihood that such manoeuvres will actually result in reduced separation.

Issued: May 1,1996 Revision 1 : April 2, 2001

Report No: 01973-001 9-06-5

SECTION 9 WPILATUSB SUPPLEMENT 6 PC12 f

Report No: 01973-001 9-06-6


Issued: May 1, 1996 Revision 1 : April 2, 2001

- - - -





PC-12 REGISTERED IN AUSTRIA

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating the PC-12 in Austria. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Federal O f f i far Civil Aviation (FOCA) of Switzerland Section for Type Certification

Date of Approval:

2s' Jc~e r i

Issued: April 10.1996 Revision 1: May 15.1996

Report No: 01973-001 9-07-1



KINDS OF OPERATION EQUIPMENT LIST

SYSTEM I EQUIPMENT VFR VFR IFR IFR ICING DAY NIGHT DAY NIGHT

Fllght Instruments:

Sensitive Altimeter Atiitude lndicator (EADI) Attitude lndicator (Self contained) Rate of Turn lndicator (EADI) Directional lndicator (EHSI) AHRS Vertical Speed lndicator Clock

Emergency Equipment:

Fire Extinquisher Cockpit Fire Extinquisher Cabin First Aid Kit

CARGO LIMITATIONS

When a cabin cargo net is installed the cabin fire extinquisher must be positioned to be accessable forward of the cargo net.

PLACARDS

On Interior Emergency Exit:

On Interior Emergency Exit Handle Cover and Handle:

Report No: 01973-001 9-07-2

Issued: April 10,1996 Revision 1: May 15, 1996 .


F b a Interior Cabin Door:

AUSGANG DARF WAHREND DEM FLUG NlCHT

HEBEL ZIEHEN UND DREHEN

AUF DER ElNSTlEGSTREPPE

In lnterior Cabin Door and Cargo Door:

TRIEBWERK NlCHT GE~FFNET WERDEN

;sued: April 10,1996 levision 1: May 15, 1996

Report No: 01973-001 9-07-3

SECTION 9 =PILATUSP SUPPLEMENT 7 ?c %[I

On Interior of Cargo Door:

FLUG NlCHT ENTFERNT WERDEN

On the leA and right cockpit side panel:

r

COCKPIT FIRE EXTINGUISHER LOCATED BEHIND CO-PILOTS SEAT. CABIN FlRE EXTINGUISHER LOCATED BEHIND REAR RIGHT HAND SEAT, OR IN COMB1 CONFIGURATION IN FRONT OF FREIGHT NET.

At each seat position on cabin sidewall:

DER FEUERLOSCHER BEFINDET SICH HINTER DEM RECHTEN HINTEREN SlTZ ODER VOR OEM FRACHTNETZ

At rear of cabin on freight bay frame:

In the cabin on the Frame 16 bulkhead divider:

DER ERSTE-HILFE-KASTEN BEFINDET SlCH HINTER OEM PlLOTENSlTZ

Report No: 01973-001 9-07-4

Issued: April 10, 1996 Revision 1 : May 15, 1996

WPkATUSW L SECTION 9 1 ?C %)I SUPPLEMENT 7

1

1 SECTION 8 - HANDLING, SERVICING AND MAINTENANCE

Aircraft Serial No: v

L

Mark X in box when Service Bulletins are incorporated on the aircraft.

- Crack Inspection of the

issued: April 10,1996 t Revision 1: May 15. 1996 - - - - -

006

Report No: 01973-MH 9-07-5

24-002 Re-routing of the Sterter-Generator and Generator 2 power return cables

.


SECTION 9 - SUPPLEMENTS

Aircraft Serial No:

Mark X in box which Supplements are applicable to the aircraft.

Report No: 01973-001 9-07-6

Supp. No

1

2

3

4

5

6

7

Issued: April 10,1996 Revision 1 : May 15,1996

Subject

BendidKing KLN 90A GPS

BendidKing RDS 82 VP Weather Radar

Bendiming KHF 950 Communications System

BendixlKing KLN 908 GPS

Argus 5000 Electronic RMllMoving Map Display

BendixIKing Traffic Collision and Avoidance System CAS 66A TCAS 1

PC- 12 registered in Austria

Applicable

X

Page

9-01-1

9-02-1

9-03-1

9-04-1

9-05-1

9-06-1

9-07- 1 *


SECTION 10 - SAFETY AND OPERATIONAL TIPS

For the location of the emergency equipment installed in the aircraft refer to the following figure:

Figure 10-1 Exits and Emergency Equipment Location

Issued: April 10,1996 Revision 1: May 15, 1996

Report No: 01 973-001 9-07-7


Report No: 01 973-001 9-07-8


Issued: April 10, 1996 Revision 1: May 15, 1996

- --





PC-12/45

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating the PC-12 at a weight of 4500 kg. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office for Civil Aviation (FOCA) of Switzerland


Report No: 01 973-001 9-08-1


SECTION 1 - GENERAL

MAXIMUM WEIGHTS

Maximum Ramp Weight



Maximum Zero Fuel Weight

Maximum Cargo Weight

Baggage Area

Cabin Area

TYPICAL AIRPLANE WEIGHTS

9965 Ib (4520 kg)

9921 Ib (4500 kg)

9921 Ib (4500 kg)

9039 1b (4100 kg)

400 Ib (1 80 kg)

3300 Ib (1 500 kg)

Empty Weight 5468 Ib (2480 kg) '

Useful Load 4454 Ib (2020 kg)

'Empty weight of standard airplane without 9 passenger seats and cabin floor covering. .

SPECIFIC LOADINGS

Wing Loading

Power Loading

Report No: 01 973-001 9-08-2

Issued: June3,1996 Revision 1 : Dec 6, 1996





I

AIRSPEED

Maximum Operating Maneuvering Speed - V,

9921 Ib (4500 kg)

9480 Ib (4300 kg)

WEIGHT LIMITS

MARKING

Green Arc

White Arc

Maximum Ramp Weight



Maximum Zero Fuel Weight

KCAS

161

158

Issued: June 3,1996 Revision 1 : Dec 6, 1996

KlAS VALUE OR RANGE

91 to 236

64 to 130

9965 Ib (4520 kg)

9921 Ib (4500 kg)

9921 Ib (4500 kg)

9039 Ib (41 00 kg)

KlAS

158

1 55

REMARKS

Normal operating range. Lower limit is maximum weight stall speed in the clean configuration (V,). Upper limit is the maximum operating speed (V,dM,).

Full flap operating range. Lower limit is maximum weight stall speed in landing configuration (Vso). Upper limit is maximum speed with full flaps extended (VFE).

Report No: 01 973-001 9-08-3

SIGNIFICANCE

Do not make full or abrupt control movements above this speed.


CENTER OF GRAVITY LIMITS

NOTES

Weight Pounds (kilograms)

9921 (4500)

81 58 (3700)

7938 (3600)

661 5 (3000)

5733 (2600)

Straight line variation between points given.

The datum is 118 in (3.0 m) forward of firewall.

Forward Limit A.O.D.: In. I M

232.20 1 5.898

224.13 I 5.693

220.75 1 5.607

It is the responsibility of the pilot to ensure that airplane is loaded properly.

I Aft Limit

A.O.D.: In. I M

240.94 1 6.120

I

242.99 1 6.1 72

242.99 1 6.172

225.47 15.727

See Section 6, Weight and Balance for proper loading instructions.

FLIGHT LOAD FACTOR LIMITS

Flight load limits with flaps up

Flight load limits with flaps down

Report NO: 01973-001 9-08-4


I


KINDS OF OPERATIONAL EQUIPMENT LIST

In addition to the Flight Instruments listed in the POH, the following are required for the PC- 12/45.

SYSTEM I EQUIPMENT VFR VFR IFR IFF7 ICING DAY NIGHT DAY NIGHT

Radar Altimeter . 1 1 1 1 1 2nd AHRS or Yaw Rate Sensor Installation 0 0 1 1 1

OTHER LIMITATIONS

PC-12/45 AIRCRAFT

The aircraft must be operated in accordance with this Supplement when the Identification Plate installed on the lower left side of the fuselage aft of the cargo door defines the Model Designation as PC-1 2/45.

CARGO LlMrT'ATlONS

Maximum Freight Load 3300 Ibs (1500 kg)

Issued: June 3, 1996 , Revision 6: March 1,2003 Report No: 01 973-001

9-08-5


PLACARDS

On the instrument panel:

On the main landing gear doors:

I On the left and right cargo door frame:

Report No: 01 973-001 9-08-6




3.2 AIRSPEEDS FOR EMERGENCY OPERATIONS

All airspeeds shown are with airplane in clean configuration under ISA conditions.

A. Operating Maneuvering Speed (Vo):

9921 Ib (4500 kg) 158 KlAS

9480 Ib (4300 kg) 155 KlAS

8. Best Glide (Propeller feathered):

9921 Ib (4500 kg) 1 14 KlAS

C. Landing Approach Speeds with ice accretion on the airframe:

After failure of: Minimum Approach Speed:

Pneumatic Deice Boots (flap position limit 0") 134 KlAS

AOA Probe Deke and/or

108 KlAS

Pitot and Static Probe Deice 108 KlAS andlor

Pusher Ice Mode (flap position limit 15") 108 KlAS

D. Balked Landing (Go Around)

After failure of:

Pneumatic Deice Boots (flap position limit 0")

(TOtPwr, flaps O", Pusher Ice Mode) 134 KlAS

3.8 EMERGENCY DESCENT

3.8.2 MAXIMUM RANGE DESCENT - AFTER ENGINE FAIL

4. Speed

Issued: June 3.1996

114 KlAS for 9921 Ib (4500 kg) (for best glide speed see para 3.2.6)

Report No: 01973-001 9-08-7


3.9 EMERGENCY LANDING

3.9.1 GUDE DISTANCE AND SPEED

4. Best g l i speed 11 4 KlAS for 9921 Ib (4500 kg) (for best glide speed see para 3.2.8)

3.9.2 FORCED LANDING (ENGINE CUT-OFFIFEATHER)

6. Best glide speed 114 KlAS for 9921 lb (4500 kg) (for best glide speed see para 3.2.8)


I 3.9.8 LANDING WITHOUT FLAPS

I 3. Final approach speed 118 KlAS

I 3.9.9 DITCHING

I 4. Final approach speed 84 KlAS

3.19 DEICE SYSTEM

3.19.2 BOOT DEICE FAILURE IN ICING CONDITIONS

1 1. Landing approach Keep minimum landing approach speed above 134 KIAS.

The CAUTlON referring to EFlS AOA last slow pointer indicafin is M be disregarded.

3.19.5 AOA PROBE DEICE FAILURE IN ICING CONDITIONS

9. Landing approach Keep minimum landing approach speed above 108 KlAS or shaker activation speed, whichever is highest.

3.19.7 PUSHER ICE MODE FAILURE IN ICING CONDITIONS

9. Landing approach Keep minimum landing approach speed above 108 KIAS.

Report No: 01 973-001 9-08-8

Issued: June 3, 1996 Revision 3: Sep 1, 1998



4.2 AIRSPEEDS FOR NORMAL OPERATIONS

Airspeeds for normal operaUons are listed bebw. Unless othennise noted, all ahspeeds are based on a maximum takeoff weight of 9921 Ib (4500 kg) at sea level under ISA standard day conditions.

Takeoff (VR): Flaps 15" 79 KlAS Flaps 30" 73 KlAS

Recommended Climb Speed with Flaps retracted and Pusher Ice Mode 130 KlAS

I

Maximum Operating Maneuvering Speed (VO) (9,921 Ibl4,500 kg) 158 KlAS

Landing Approach Speed : Flaps 0" 118 KlAS Flaps 15" 98 KlAS Flaps 300 89 KlAS Flaps 40" 84 KlAS

with residual ice on the airframe Flaps 15', Pusher Ice Mode 108 KlAS

Balked Landing (Go-Around): TOIPwr. Flaps 15" 95 KlAS TOIPwr. Flaps 30" 85 KlAS TOIPwr, Flaps 40" 80 KlAS TOJPwr, Flaps 15", Pusher Ice Mode 108 KlAS

Maximum Demonstrated Crosswind for Takeoff and Landing (not a limitation): Flaps 0' 30 kts Flaps 15' 25 kts Flaps 30" 20 kts Flaps 40" (landing only) 15 kts

4.14 BEFORE LANDING

4.14.1 APPROACH CHECK

8. AOA CENTERED

AOA failure. Refer to Section 4-2 Airspeeds

[ CAUTION ]

ON LANDING APPROACH AFTER PUSHER ICE MODE FAILURE OR AFTER AOA DEICE FAILURE THE EFIS AOA

Issued: June 3.1996 Revision 4: Sep 1, 1999

Report NO: 01973-001 9 01 98 9-08-9


ON LANDING APPROACH AFTER PUSHER ICE MODE FAILURE OR AFTER AOA DEICE FAILURE THE EFlS AOA FAST SLOW POINTER WILL NOT BE CORRECT AND SHOULD NOT BE USED AS REFERENCE.

4.142 FINAL CHECK

3. AOA CENTERED

AOA failure. Refer to Section 4-2 Airspeeds

4.15 BALKED LANDING (GO-AROUND)

5. Climb airspeed - Pusher Normal Mode 95 KIAS - (Pusher Ice Mode 108 KIAS) - Boot lailure 134 KlAS

4.21 NOISE LEVEL

K=AO Annex 16. Chapter 10

Swiss VEL

FAR Pan 36, Appendii G

4.23 FLIGHT IN ICING CONDITIONS

The minimum recommended speeds for icing encounters and with residual Ice on the airframe are :-

- Climb, Flaps Oa, Pusher Ice Mode = 130 KIAS Holding Pattern, Flaps 0" - 140 KlAS to 170 KlAS - Landing Approach, Flaps 15", Pusher Ice Mode = 108 KlAS - Landing Approach, Flaps Oa, Boot Failure Pusher Ice Mode = 134 KlAS - Balked Landing (Go-Around) Flaps 15" Pusher Ice Mode - 108 KlAS - Balked Landing (Go-Around) Boot labre Flap 0' Pusher Ice Mode * 134 KlAS

Issued: June 3,1996 Revisbn 3: Sep 1.1998


SECTION 5

PERFORMANCE

TABLE OF CONTENTS

Subject

STALL SPEED

FIG. 5-12. STALL SPEEDS KlAS - FLIGHT IDLE POWER (STANDARD UNITS) FIG. 5-13. STALL SPEEDS KlAS - FLIGHT IDLE POWER (METRIC UNITS) FIG. 5-14. STALL SPEEDS KCAS - FLIGHT IDLE POWER (STANDARD UNITS) FIG. 5-15. STALL SPEEDS KCAS - FLIGHT IDLE POWER (METRIC UNITS)

TAKEOFF PERFORMANCE

FIG. 5-16. STATIC TAKEOFF TORQUE FIG. 5-17. ACCELERATE-STOP DISTANCE - FLAPS 30' (STANDARD UNITS) FIG. 5-18. ACCELERATE-STOP DISTANCE - FLAPS 30' (METRIC UNITS) FIG. 5-1 9. TAKEOFF GROUND ROLL - FLAPS 30" (STANDARD UNITS) FIG. 5-20. TAKEOFF GROUND ROLL - FLAPS 30" (METRIC UNITS) FIG. 5-21. TAKEOFF TOTAL DISTANCE - FLAPS 30" (STANDARD UNITS) FIG. 5-22. TAKEOFF TOTAL DISTANCE - FLAr '; 30" (METRIC UNITS) FIG. 5-23. ACCELERATE-STOP DISTANCE - FLAPS 15" (STANDARD UNITS) FIG. 5-24. ACCELERATE-STOP DISTANCE - FLAPS 15" (METRIC UNITS) FIG. 5-25. TAKEOFF GROUND ROLL - FLAPS 15" (STANDARD UNITS) FIG. 5-26. TAKEOFF GROUND ROLL - FLAPS 15" (METRIC UNITS) FIG. 5-27. TAKEOFF TOTAL DISTANCE - FLAPS 15" (STANDARD UNITS) FIG. 5-28. TAKEOFF TOTAL DISTANCE - FLAPS 15" (METRIC UNITS)

CLIMB PERFORMANCE

FIG. 5-29. MAXIMUM CLlMB TORQUE FIG. 5-30. MAXIMUM RATE OF CLlMB - FLAPS 30" (STANDARD UNITS) FIG. 5-31. MAXIMUM RATE OF CLlMB - FLAPS 30" (METRIC UNITS) FIG. 5-32. MAXIMUM RATE OF CLlMB - FLAPS 15" (STANDARD UNITS) FIG. 5-33. MAXIMUM RATE OF CLlMB - FLAPS 15" (METRIC UNITS) FIG. 5-34. MAXIMUM RATE OF CLlMB FLAPS 0" (STANDARD UNITS) FIG. 5-35. MAXIMUM RATE OF CLlMB - 0" FLAPS (METRIC UNITS) FIG. 5-36. CRUISE CLlMB AIRSPEED SCHEDULE FIG. 5-37. RATE OF CLIMB - CRUISE CLIMB (STANDARD UNITS) FIG. 5-38. RATE OF CLlMB - CRUISE CLlMB (METRIC UNITS) FIG. 5-39. TIME TO CLlMB - CRUISE CLlMB (STANDARD UNITS)

Page

Issued: June 3.1996 Report No: 01973-001 9-08-1 1


Subject

FIG. 5-40. TlME TO CLlMB - CRUISE CLIMB (METRIC UNITS) FIG. 5-41. FUEL USED TO CLlMB - CRUISE CLlMB (STANDARD UNITS) FIG. 5-42. FUEL USED TO CLIMB - CRUISE CLIMB (METRIC UNITS) FIG. 5-43. DISTANCE TO CLIMB - CRUISE CLlMB (STANDARD UNITS) FIG. 5-44. DISTANCE TO CLlMB - CRUISE CLIMB (METRIC UNITS)

CRUISE PERFORMANCE

FIG. 5-45. MAXIMUM CRUISE POWER FIG. 5-46. LONG RANGE CRUISE FIG. 5-47. MAXIMUM ENDURANCE CRUISE FIG. 5-48. SPECIFIC AIR RANGE (7000 LB) FIG. 5-49. SPECIFIC AIR RANGE (8000 LB) FIG. 5-50 SPEClFlC AIR RANGE (9000 LB) FIG. 5-51. SPECIFIC AIR RANGE (9800 LB) FIG. 5-52. HOLDING TlME AND FUEL

DESCENT PERFORMANCE

FIG. 5-53. TlME TO DESCEND FIG. 5-54. FUEL USED TO DESCEND (STANDARD UNITS) FIG. 5-55. FUEL USED TO DESCEND (METRIC UNITS) FIG. 5-56. DISTANCE TO DESCEND FIG 5-57. POWER-OFF GLIDE TlME (STANDARD UNITS) FIG. 5-58. POWER-OFF GLIDE TlME (METRIC UNITS) FIG. 5-59. POWER-OFF GLIDE DISTANCE

BALKED LANDING

FIG. 5-60. BALKED LANDING TORQUE FIG. 5-61. RATE OF CLlMB - BALKED LANDING (STANDARD UNITS) FIG. 5-62. RATE OF CLlMB - BALKED LANDING (METRIC UNITS)

LANDING PERFORMANCE

FIG. 5-63. LANDING TOTAL DISTANCE - FLAPS 40" (STANDARD UNITS) FIG. 5-64. LANDING TOTAL DISTANCE - FLAPS 40' (METRIC UNITS) FIG. 5-65. LANDING GROUND ROLL - FLAPS 40" (STANDARD UNITS) FIG 5-66. LANDING GROUND ROLL - FLAPS 40" (METRIC UNITS) FIG. 5-67. LND TOTAL DlST WlTH REVERSE THRUST (STANDARD UNITS) FIG. 5-68. LAND TOTAL DlST WlTH REVERSE THRUST(METR1C UNITS) FIG. 5-69. LAND GRND ROLL WlTH REVERSE THRUST(STANDARD UNITS) FIG. 5-70. LAND GROUND ROLL WlTH REVERSE THRUST(METR1C UNITS)


Report No: 01973-001 9.08- 12

Page

Issued: June 3. 1996 ~evision 1 . Dec 6. 1996


' Page 5-1 new para 1

When landing with flaps set to less than 40°, the total landing distances will be increased by the following factors:

C

Issued: June 3,1996

- - -

FLAP SElllNG

0"

15"

30"

Report No: 01973-001 9-08-13

FACTOR

1.83

1.31

1.22


EMMPLE: W E M 3550 KG

NOTES: RAPS - 40' STALL IS WINED BY PUSHER ACTlVATlON ANGLE OF BAN< 45 DEG LANolNG GEM POSmON NO EFFECT S T W SPEED 67 K I M

WElGKT - KG l ape ANGLE OF B M - DEG


STATIC TAKEOFF TORQUE

TOROUE WILL INCREASE WlTH INCREASING AIRSPEED

PROPELLER SPEED 1700 RPM INERTIAL SEPARATOR CLOSED MAXIMUM TORQUE REDUCTlON WlTH EXAMPLE:

INERTIAL SEPARATOR OPEN : ALTITUDE 4000R - 1.2 PSI IN NON ICING CONDITIONS OAT 28 'C - 2 1 PSI IN ICING CONDITIONS ENGINE TORWE 41 PSI


I Figure 5.16. Static Takeoll Torque

Reporl No: 01973-001 9-08- 18

Issued: June 3, 1996 Rev~s~on 2. July 1 , 1997

AJ yr m V) < c

G rD

5'1 ~ G g C m ACCELERATE-STOP DISTANCE - FLAPS 30' E w r . E (STANDARD UNITS) .-;;; 2 ;;% '?7 WAMPLE: w - KTrmDE - J ?

6WOn OAT

I 18 'C

R ASSOCIATED CCMKIONS. WEIGHT 0

ni6 LB P M R G H O P AT 1.1 VS, HEMW CCM'0NEN-r 2 CC)*)KKM LEMR AT GROUND IDLE 8 KT

4 UP= CaRONEM 1 % Fl€LD SURFACE: TARMAC S ACCELERATE-STOP DISTAN~ 2540 n

r 11 lD-0

"5 X C =$

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

I Figure 5- 18. A

ccelerate - Stop D

~stance - Flaps 30" (metric units)

Repon N

o: 0 1973-00 1 Issued: June 3. 1996

9.08-20 20 01 98

Revislon 2: July 1. 1997

TAKEOFF GROUND ROLL - FLAPS 30' (STANDARD UNITS)

MAMPCE = LE Vn - KUS U T W M W K ) F I

ASSOClATED COH)illONS 6400 58 OAT 18 'C LR OFF AT 1 t vs, 7300 bt ~~YEIGH~ n i 6 LE REFER TO THE SPEED SCHEDUE TABLE BZOO 66 HEAOWMWMWNENT 8 K T RUNWAY SURFACE. T W W 9100 70 UPnU m N I 1 X

_ lmw . 73 TAKEOFF OROUM AOU 1 3 9 FT

- W D

$2 M a c

z w + g ,g! 5? 5

tD

: 0 4 % 0 2 - s C

M 2 O- n g

z I

W W 2 - 3 2 -. 5 " 5 $ a r

2 5 a . . L: s'? - W - - - 4

W ( D (Dw + m

TAKEOFF GROUND ROLL - FLAPS 30" (METRIC UNITS)

WEIOHT-'- KG V, - KIAS K ~ V M 6000 FT

ASSOCIATED COF(D'TIONS 2900- 58 OAT 18 'C

LlFl OFF AT 1 1 Vs, 3300 WEIGHT JSOO KG 62

REFER TO THE SPEED SCMDULE TABLE 3700 66 HEADWIND WWKUdENT 8 KT

RUNWAY SURFAa TARMAC 4100 70 U P H U COMPONENT 1 5.

73 TAKEOFF r x n N O ROU 410 M

TAKEOFF TOTAL DISTANCE - FLAPS 30' OVER 50 FT OBSTACLE; (STANDARD UNITS)

UVhRLE WEIGHT - LB --v;: KIM-' V, - KlAS AITKUDE - - r n R

ASSOCIATED CONDmONS. 64Gfl- - - 58 73 OAT 18 'C LIFT UT AT 1 1 Vs, 7300 62 78 W l M 7716 LB CBSTACLE AT 1.3 Vs, 8200 66 82 HWWIND COMWNENT 8 KT REFER TO M SPEED SCHEDULE TABLE SlOO 70 87 UPHILL COMPONENT 1 X FIELD SURFACE TARM4C ._'!'%? . 5' -_ St _ TAKEOFF TOT& DISTANCE 2054 R

OUTSIDE AR TUlrPERAW9E - 'C h€lGlfl - LB wlm COhRONNT - KT 90PE - %

TAKEOFF TOTAL DISTANCE - FLAPS 30" OVER 15 M OBSTACLE; (METRIC UNITS)

ExNRLE WEIGHT -KG 1 < - KIAS vm - KIAS NTTTUH 60W FT

ASXCIATED CCNDKK3NS 2WO 58 73 OAT 18 'C

LIFT OFF AT 1 1 Vsl 3300 62 78 WEIGHT XdXl KG

OBSTACLE AT 13 Vs, 3700 66 82 HEAMNINO COWWENT 8 KT REFER TO THE SPEED scumu 4100 70 87 W U L L P O N N T 1 %

FIELD SURFACE TARMAC 4500 73 91 TAKEOFF TOTAL DISTANCE 625 M

OUTSIDE AIR TEMPERA- - %

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for Info on effect of Icing

Fiu

re 5-23. A

ccelerate - Stop Distance - Flaps 15" (standard units)

I Issued: June 3,1996 R

evision 5: Sep 1,2000 R

eport No: 01 973-001

9-08-25

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

I See FLIG

HT IN

ICIN

G C

ON

DITIO

NS


Figure 5-24. Accelerate - Stop D

istance - Flaps 15" (metric units)

Report N

o: 01 973-001 9-08-26

Issued: June 3, 1996 R

evision 5: Sep 1,2000

TAKEOFF GROUND ROLL - FLAPS 15' (STANDARD UNITS)

SECTIO

N 9

SUPPLEM

ENT 8

I See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para tor into on effect of lclng

Figure 5-26. Takeofl Ground R

oll - Flaps 15" (metric units)

Report N

o: 01973-001 9-08-28

Issued: June 3,1996 R


WPLA

TUSW

SEC

TION

9

P(12 SU

PPLEMEN

T 8

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for Info on effect of lclng

Figure 5-27. Takeoff Total Distance - Flaps 15" (standard units)


evision 5: Sep 1, 2000 R

eport NO

: 01973-001 9-08-29

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

See FLIGH

T IN IC

ING

CO

ND

ITION

S para for Info on effect of lclng

Figure 5-28. Takeon Total Distance - Flaps 15' (m

etric units)

Repon N

o: 01 973-001 9-08-30



=PnATUSI , SECTION 9 ?c %I1 SUPPLEMENT 8

MAXIMUM CLIMB TORQUE PROPELLER SPEED 1700 RPM INERTIAL SEPARATOR CLOSED MAXIMUM TORQUE REDUCTION WITH EXAMPLE:

INERTIAL SEPARATOR OPEN : ALTITUDE 5000 n - 1.2 PSI IN NON ICING CONDITIONS OAT 29 'C - 2.1 PSI IN ICING CONDITIONS ENGINE TORQUE 33 PSI



I

Figure 5-29. Maximum Climb Torque

31 of 98 Report No: 01973-001

9-08-31

SE

CTIO

N 9

=PILATUS= S

UP

PLE

ME

NT 8

?c %

I1

Figure 5-30, Maximum

Rate ol C

limb - Flaps 30" (standard units)

Report N

o: 01973-001 Issued: June 3.1996

9-08-32 32 O

f 98 R

evis~on 2: July 1, 1997

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

Figure 5-31. Maxim

um R

ate of Clim

b - Flaps 30" (metric units)

Issued: June 3. 1996 R


eport No: 01973-00 1

9-08-33

SECTIO

N 9

5PIL

AT

US

F

4

SU

PP

LEM

EN

T 8 ?c M

I SEC

TION

9 S

UP

PLE

ME

NT 8

Figure 5-32. Maxim

um R

ate 01 Clim

b - Flaps 15" (slandard units)


Issued: June 3.1996 9-08-34

34 of 98 R

evision 2: July 1. 1997 I

3PIL

AT

US

E

SE

CTIO

N 9

?c %I1

SU

PP

LEM

EN

T 8

Figure 5-33. M

aximum

Rate of C

limb - F

laps 15" (metric units)

I Issued: June 3. 1996 R

evision 2: July 1. 1997 I

Report N

o: 0 1973-001 9-08-35

SECTIO

N 9

FPILATU

S W

SU

PP

LEM

EN

T 8

?C X

I1

Report N

o: 9-08-36

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for info on enect of icing

Figure 5-34. Maxim

um R

ate of Clim

b - Flaps 0" (standard units) Issued: June 3,

Rev~sion 2: July 1

,

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for Info on effect of Icing

Figure 5-35. M

aximum

Rate of C

limb - F

laps 0" (metric units)

I Issued

June 3. 1996 R

ev~s~

on

2:

July 1, 1997 R

eport No: 01973-001

9.08-37

SECTION 9 SPILATUS W SUPPLEMENT 8 ?c %I1

CRUISE CLIMB AIRSPEED SCHEDULE

ASSOCIATED CONDITIONS: PROPELLER SPEED 1700 RPM INERTIAL SEPARATOR CLOSED

I Figure 5-36. Cruise Cllmb Airspeed Schedule

Report No: 01973-001 9-08-38

Issued: June 3, 1996 38 of 98

RATE OF CLIMB - CRUISE CLIMB (STANDARD UNITS)


EXAMPLE: ALTITUDE 7000 FT OAT 22 'C AIRCRAFT MIGHT 8600 LB RATE OF CLIMB 1400 FPM

-60 -40 -20 0 20 40 60 11OOO loo00 9000 8000 7000 6000

OUTSIDE AIR TEMPERATURE - 'C AIRCRAFF MIGHT - LB

SE

CTIO

N 9

%PILATUSW

S

UP

PLE

ME

NT 8

?c XI1

l

Figure 5

-38 R

ate of Clim

b - Cruise C

limb (m

etric units)

Report N

o: 01973-001 9.08-40

Issued: June 3, 1996 R

evision 2: July 1. 1997

TlME TO CLIMB - CRUISE CLIMB (STANDARD UNITS)

ASSOCIATED CONDITIONS: MAXIMUM CLIMB POWER LANDING GEAR RETRACTED FLAPS UP AIRSPEED SCHEDULE: SEE RELEVANT TARE

EXAMPLE: ALTITUDE 25000 FT OAT -30 'C AlRCRAn WEIGHT 7950 LB TIME TO CLIMB 18 MIN

OUTSIDE AIR TEMPERATURE - 'C AIRCRAFT WEIGHT - LB

TlME TO CLIMB - CRUISE CLIMB (METRIC UNITS)


EXAMPLE: ALTITUDE 25000 R OAT -30 'C AIRCRAFT WEIGHT 3600 KG TIME TO CLIMB 18 MIN

OUTSIDE AIR TEMPERATURE - 'C AlRCRAFl WEIGHT - KG

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

1 Figure 5-41. Fuel U

sed to Clim

b - Cruise C

limb (standard units)

I Issued: June 3, 1996 R

evision 2: July 1. 1997 b

Report N

o: 01973-001 9-08-43

SE

CTIO

N 9

r PILA

TUS

5 S

UP

PLE

ME

NT 8

?c X

I1 ,

Figure 5-42. Fuel Used to C

l~mb

- Cruise C

limb (m

etric un

~ts)

Repon N

o: 01 973-001 9-08-44


ev~s~

on

2: July 1. 1997

- OUTSIDE AIR TEMPERATURE - 'C AIRCRAFT WEIGHT - LB

i7iz < c DISTANCE TO CLIMB - CRUISE CLIMB Ern 1 9" (STANDARD UNITS) ? 2 3 1 L r n

C-w ASSOCIATED CONDITIONS: EXAMPLE:

1 :- n MAXIMUM CLIMB POWER - a 6'

ALTrmM 25000 FT LANDING GEAR RETRACTED OAT -30 'C 1 ~ 2 ? 5

(0 " 2 FLAPS UP AIRCRAFT WEIGHT 7950 LB

I 0 AIRSPEED SCHEDULE: SEE RELEVANT TABLE DISTANCE TO CLIMB 54 NM

111 W w "5 X C =;G

DISTANCE TO CLIMB - CRUISE CLIMB (METRIC UNITS)

ASSOCIATED CONDITIONS: EXAMPLE:

MAXIMUM CLIMB POWER ALTITUDE 25000 FT

LANDING GEAR RETRACTED OAT -30 'C

FLAPS UP AIRCRAFT WEIGHT 3600 KG

AIRSPEED SCHEDULE: SEE RELEVANT TABLE DISTANCE TO CLIMB 53 NM

OUTSIDE AIR TEMPERATURE - 'C AIRCRAFT WIGHT - KG

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

Figure 5-45. Maxim

um C

ruise Pow

er (Sheet 1 01 4

)


evision 2: July I, 1997 R

eport No

: 01973-001 9-08.47

MAXIMUM CRUISE POWER NOTE: IOAT. TORCUE AND FUEL FLOW W E D ON 8000 b 13619 kg)

PLATU USE SECTION 9 ?c XI1 SUPPLEMENT 8

MAXIMUM CRUISE POWER W O T E : W W T . T ~ A U O F U E L ~ W W O N R l o b ~ ~ ~


Issued: June 3,1996

- -

Report No: 01973-001 9-08-49

SECTION 9 =PILATUSE SUPPLEMENT 8 ?c %)I

MAXIMUM CRUISE POWER N O ~ ~ T . T O R Q U E ~ F U f L F L O Y Y B I 8 L D O N U O O I b ( W I I W


Report No: 01973-001 9-OR-rin

Issued: June 3,1996

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

Figure 54

6. Long R

ange Cruise (S

heet 1 of 4)

Issued: June 3.1996 R

eport No: 01973-001

9-08-51

1 S

EC

TION

9 S

PIL

AT

US

S

SU

PP

LEM

EN

T 8

?c XI)

i



1

Report N

O: 01973-001

0-lW

.53

Issued: June 3,1996

4

LONG RANGE CRUISE NomlmruraD#am.(wam

MAXIMUM ENDURANCE CRUISE mm:wanraItD~mlrImcowrrurr

MAXIMUM ENDURANCE CRUISE msmmmu#rrror1wKnco#rm

MAXIMUM ENDURANCE CRUISE W Q R : ~ T l f D * * U ~ I l W m c o r r r T u n


W P L E : ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 1300 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 234 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.58 NMb

SECTION 9 WPILATUSE SUPPLEMENT 8 SC XII 1 1


EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TOAWE 27 PSI FLAPS UP TRUE AIRSPEED 239 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.595 NM&

80 1w 120 140 160 180 200 220 240 260 280 300

TRUE AIRSPEED - KT

Figure 5-48. Specific Air Range - 7000 Ib (Sheet 2 of 3) I

Report No: 01973-001 9-08-60

Issued: June 3. 1996


SPECIFIC AIR RANGE WEIGHT 7000 Ib (3175 kg) - ISA+20g

EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 243 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.6 NhMb

80 100 120 140 160 180 200 220 240 260 280 XI0

TRUE AIRSPEED - KT

Figure 5-48. Spec~fic Air Range - 7000 Ib (Sheet 3 of 3) I

Issued: June 3. 1996 Report No: 01973-001 9.08-61


FLAPS UP TRUE AIRSPEED 232 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.575 NMllb


TRUE AIRSPEED - KI



WAMPLE: ASSmlATED COM)mONS: PRESSURE KTrmDE 15000 FEET LANDING GEAR f E W M E D TORQUE 27 PSI FUPS UP TRUE AIRSPEED 237 KT INERTIAL SEPARATOA aosm SPECIFIC AIR RAW 0 . 5 ~ ~ ram

TRUE AIRSPEED - KT

Figure 549. Specific Air Range - 8000 Ib (Sheet 2 of 3)

Issued: June 3,1996

-

Report No: 01973-001 9-08-63

SECTION 9 .3PILATUS eP SUPPLEMENT 8 ?c %I1

3

SPECIFIC AIR RANGE WEIGHT 8000 Ib (3629 kg) - ISA+2O0

WAMRE: ASSOCIATED COM)fTKJNS: PRESSURE MrrmOE 15000 FEET UMllJG GEAR RETRACTED TORQUE 27 PSI FUPS UP TRUE AIRSPEED 241 KT WRTW SEPARATOR CLOSED SPECIFIC MR RMKiE 0.59 NIWb

80 1M) 120 140 180 180 200 220 240 280 280 300

TRUE AIRSPEED - KT


Report No: 01 973-001 1 9-08-64

Issued: June 3,1996

-


SECTION 9 SPILATUS= SUPPLEMENT 8 PC XI1 1 4



FLAPS UP TRUE AIRSPEED 234 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.58 NWlb

TRUE AIRSPEED - K1

Figure 5-50. Specif~c Air Range - 9000 Ib (Sheet 2 of 3)

Report No. 01973-001 9-08-66

Issued: June 3. 1996 66 ol98

EPILATUSS SECTION 9 b C PC XI1 SUPPLEMENT 8

SPECIFIC AIR RANGE WEIGHT 9000 Ib (4082 kg) - ISA+2O0

EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 1- FEET LANDING GEAR RETRACTED TORQUE 27 PSI FLAPS UP TRUE AIRSPEED 239 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.585 NMln

Figure 5-50. Specific Air Range - 9000 Ib (Sheet 3 of 3) I

Issued. June 3, 1996 Report No: 01973-001 9-08.67



EXAMPLE: ASSOCIATED CONDITIONS: PRESSURE ALTITUDE 15000 FEET LANDING GEAR RETRACED TORQUE 27 PSI

FLAPS UP TRUE AIRSPEED 228 KT INERTIAL SEPARATOR CLOSED SPECIFIC AIR RANGE 0.565 NMllb

TRUE AIRSPEED - KT

1 F~gure 5-51. Speciflc Alr Range - 9800 Ib (Sheet I of 3)

Report No: 0 1973-00 1 1 9-08 68

Issued. June 3. 1996 Rev~slon 2: July I , 1997



EXAMPLE: ASSOCIATED WNDllK)NS: PRESSURE ALTITUDE 15000 FEET LANDING GEM RETRACTED TORQUE n PSI FLAPS UP TRUE AIRSPEED 232 KT MRTIM SEPARATOA a- SPECKC AIR RANOE 0.575 ram

eo im 120 140 180 180 200 220 240 280 280 rW)

TRUE AlRSPEEb - KT

I

Figure 5-51. Specific Air Range 9800 Ib (Sheet 2 of 3)

Issued: June 3,1996 Report No: 01 973-001 9-08-69


SPECIFIC AIR RANGE WEIGHT 9800 Ib (4445 kg) - ISA+20'

WAMPLE: ASSOCIATED CONDITIONS: PRESSUAE ALTITUDE 15000 FEET UMI~GEMRE~RACTED TORQUE n PSI RAPS UP TRUE AIRSPEED 236a W T W SEPARATOR CLOSED SPECIFIC AIR R A W € 0.58 NMlb

TRUE AIRSPEED - KT


Report No: 01973-001 9-08-70

Issued: June 3.1996

-

WPLATUSW SECTION 9 b ?C XI1 SUPPLEMENT 8

HOLDING TIME AND FUEL W I N G GEAR RETFUCTED - W S UP ISA STANDARD DAY EXAMPLE: ARSPEED 150 KlAS HOCDING TIME 2.6 HR POWER FOR LEVEL FLIQHT ALT~TUDE ~OOOO n ~ T I A L SEPARATOA UOSED FUEL REQUIRED 925 LB

See FLIGHT IN ICING CONDITIONS para for Info on effect of iclng

Figure 5-52. Holding Time and Fuel

Issued: June 3,1996 Report No: 01973-001 9-08-71

FUEL USED TO DESCEND (STANDARD UNITS)

ASXIATE0 COPDIT- U H l l N G 0 E A R R E T R A C T U ) - F U P S U P POHlER AS REQUIRED TO DESCENO AT 2000 FPM MACH 0.48 OR 238 KLAS. WHlCHEMR IS LOWER

KTlTUOE 25000 FT OAT -30 'C AlRCRAnWEM mLB FUELUSE0 79 LB

-60 40 -20 0 20 40 60 l l M K ) l O O O O g O O O 8 0 0 0 7 0 0 0 6 0 0 0 S O O O

UJl!SlDE AIR TEMPERATURE - 'C A#KIUFT WIGKT - LB

FUEL USED TO DESCEND (METRIC UNITS)

ASSOCIATED CON)mONS: W I N G OEAR RETRACTU) - U S UP POWER AS REWIRED TO MSCEND AT 2000 FPM MACH 0.48 OR 236 KW, W H l ~ V E A IS LOWER

.60 4 -20 0 20 40 so 5 0 0 0 4 5 4 0 m 3 5 0 0 3 0 0 0 2 5 0 0 2 m

OUTSB]E AIR TEMPERATURE - C A l R C R K T W W - K G

WPLATUSP SECTION 9 C ?@ XI1 SUPPLEMENT 8

DISTANCE TO DESCEND ASSOClArrD CONDmONS: IANDlNG0EARRETRACrrT)--UP POWER AS REOOlRED TO U(AMPLE: DESCEND AT 2000 FPM ALmUOE 25000 FT AIRSPEED: MACH 0.48 OR 236 K W , OAT -32 'C WHICHEVER IS L M R DISTANCE 58 W

Figure 5-56. Distance to Descend

Issued: June 3,1996 Report No: 01973-001 9-08-75

POWER-OFF GLIDE TIME (STANDARD UNITS)

' KCAS'

~ 7 E D COM)TTW: W20 114 M E : PCrmEROFF gOU) 110 K T r m M 25000 FT PRO?ElLER FEATHERED 6\60 105 M T -30 'c WING OUR RETRACED 7280 #) , AIR- W€W 69K) LB

RAPSUP 6400 92 alDE TIME 32 MIN

mu, c m 0 0 t: z: 5 0)

1 85 x2 '"r

SE

CTIO

N 8

SU

PP

LEM

EN

T 8

Figure 5-58. Power-off G

lide Time (m

etric units)

Issued: June 3,1

996

Report N

o: 01 973-001 9-08-77

SECTION 9 =PLATUS= SUPPLEMENT 0 ?c %!I

POWER-OFF GLIDE DISTANCE (VALID FOR ALL AIRCRAFT WEIGHTS)

ASSOCIATED WNDIT- BkO -

POWER OFF 9MO PROPEUER FEATHERED 8160 3700 UNDWG GEAR RElFUCTED 7280 FUPSUP 6(00

.

EXAMPLE: ALTITUDE 25000 R OAT -30 'C OLlDEDSTANC€ 65NM

4 -20 0 20 60


Report No: 01973-001 ' 9-08-78

Figure 5-59. Power-off Glide Distance

Issued: June3.1996


BALKED LANDING TORQUE PROPEUER SPEED 1700 RPM AIRSPEED 80 KlAS INERTIAL SEPARATOR CLOSED MAXIMUM TORQUE REDUCTION WITH EXAMPLE:

INERTIAL SEPARATOR OPEN : ALTITUDE 8000 n - 1.2 PSI IN NON ICING CONDITIONS OAT 13 'C - 2.1 PSI IN ICING CONDITKINS ENGINE TORQUE 39.9 PSI

46

44

42

rz a. I

W

38 ' E W z a

36 5

34

32

30 -60 -40 -20 0 20 40 60


Figure 5-60. Balked Landing Torque

Issued June 3. 1996 Revision 2- July I . 1997

Report No. 0 1973 00 1 9-08 79

SE

CTIO

N 9

S

UP

PLE

ME

NT 8

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS

para for info

on

effect of icing

Figure 5-61. Rale ol C

limb - B

alked Landing (slandard units) '

Report N

o: 01 973-001 9-08-80


evtslon 2: July 1. 1997

RATE OF CLIMB - BALKED LANDING (METRIC UNITS)

ASSOCIATED CONDITIONS: TAKEOFF POWER LANDING GEAR WENDED FLAPS 40' AIRSPEED 80 KlAS

EXAMPLE: ALTITUDE 5000 OAT 22 AIRCRAFT WEIGHT 3400 RATE OF CLIMB I Inn

R 'C KG FPM

LANDING TOTAL DISTANCE - FLAPS 40" FROM 50 FT; (STANDARD UNITS)

ASOUATED COH)lTlOI4S MLRLE APPROACH AT 13 Va. W E I M - LB - \UP - KIAS UlltWE boo0 FT REFER TO T I E SPEED SCHEME TABLE 6400 67 OAT 18 'C AVERAGE BRAKING TECHNIOUE 7300 72 WEW n l 6 LB O R O W lDlE A J T R TOUCH O m 8 x 0 76 tEADWM) U)MWFENf 8 KT R W A Y SWAGE. TARMAC 91W W UPHILL COMPONNT 1 %

SEE SECTION 2 . LIMllA7Xr(S 1MX)O &( LAMHNG TOTAL 1970 Ff

SE

CT

ION

9

SU

PP

LEM

EN

T 8

Sw

FLIGH

T IN IC

ING

CO

ND

mO

NS

para for Into on effect of Icing

Figure 5-64. Landing Total Distance - Flaps 40' (m

etric units)

Issued: June 3,1

99

6

Report N

o: 01 973-001 9-08-83

SE

CTIO

N 9

S

UP

PLE

ME

NT 8


oll - Flaps 40° (slandard units) R

eport No: 01973-001

9-08-84 Issued: June 3.1996

84 of 98

LANDING GROUND ROLL - FLAPS 40' (METRIC UNITS)

UULRE. LLnnJLE won OAT 18 .C WEW UQ) KG w C O U * N N T 8 K T LRHU camxw'r 1 %

LAH)NOOROUH)FKU W M

--

b0 4 -2P 0 20 40 W 5 S Q ) 5 m 0 * Y r ) U l O O 3 5 4 ) 3 0 0 0 2 5 ~ a m O 0 10 20 30 0 2 4

0UlSIL)E AIR EWERATLRE - C WEW - KG W a W C M N T - I ( T S O P E - X

LANDING TOTAL DISTANCE WITH REVERSE THRUST - FLAPS 40" FROM 50 FT; (STANDARD UNKS)

SSOUATED COEMIX*IS: -..- EXAMRE: APPKmCH AT 1.3 Vs, WIGHT - vU - K ~ S U l n a - - man RFVI TO THE SPEED SCHEWLE TABLE 6UIO 67 OAT I0 'C

AMRW Wffi TEOHKWE m n WEIM 7 1 6 LB

F U REVERS€ TH\UST AFTER T O W DOHM WOO 76 HEADWIND l x f w O N E ~ 0 KI

R W A Y W e : 1- 91W m LR(LL C a W N N T 1 %

SEE 8- 2 - UMTATK)FIS 10000 E = . UH)IE(O TOT& D l S T u a 16w OFT

J J ~ n P E ag 9 . 3

3. g 3

(I, V 0 = 3 o_ ~8 , (I, " 2

$ 2

p ' - -1

0 5 =

8 ,. 8 3 % ; 1 $ E Z 3 : 0

d 2 " Z 5. @ g 5 9 -

LANDING TOTAL DISTANCE WITH REVERSE THRUST - FLAPS 40" FROM 15 M; (METRIC UNITS)

TED C O H ) m MvRcwH AT 1.3 v,, R F E R T O T H E S P E E D S W 3 U T I B L E A m - wffi lw iNuJE FULMMRSETtmUSTmERTOUCHoom RUMNAY SWICACE: 1- SEE SECrKW 2 - UIITATIONS

ExMmE MIGHT - KG : V~ - Kt&, *LllI'UDE boo0 Fl

2WO s7 OAT 18 'C 5300 72 YKlOHl WQ KG 3MO 76 HEADWm Ca)JWNENr 8 KT 4100 , 80 , URULCCkPMM 1 %

M- LAWINO TOT& DISTANCE 513 M -'500---- -

80 -40 .a 0 a0 40 s o u m * s a , r m , 3 5 m m o 2 S C a 0 10 20 30 0 2 4

CUTSIC€ UR T E ~ R I T U R E - % WKWI - KG WIH) C(WDOHNT - KT SLOPE - %

SE

CTIO

N 9

SUPPLEM

ENT 8

See FLIG

HT IN

ICIN

G C

ON

DITIO

NS



oll with the use 01 R

everse Thrust - Flaps 40° (standard units)

Reporl N

o: 01 973-001 9-08-88

Issued: June 3, 1996 88 of 98

Revision 1 : D

ec 6. 1996

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

See FLIGH

T IN IC

ING

CO

NW

TION

S para tor Into on effect of lclng

Fgure 5-70. Landing Ground R

oll with the use of R

everse Thrust - Flaps 40" (metric units)


evision 1 : Dec 6. 1996

89 of 98 R

eport No: 01973-001

9-08-89



STALL SPEEDS

When operating in PUSHER ICE MODE, the stick pusher computer automatically reduces the shaker and pusher settings by 8" AOA. With operational deice boots this results in an increase of the stall speed at the maximum takeoff weight of 11 kts with 0" flaps and 9 kts with 15" flaps.

NOTE

Flap position is limited to maximum 15' in icing conditions with operational deice boots.

With failed deice boots and ice accretion on the lilting surlaces an increase of the stall speed at the maximum takeoff weight of 14 kts with 0' flaps is to be expected.

NOTE

Flap position is limited to maximum 0' in icing conditions with failed deice boots.

The wings level stall speeds at the maximum takeoff weight of 9921lb (4500 kg) and with flight idle power are summarized in the following table:

ENGINE TORQUE

FLAPS

0"

15'

When the engine inlet inertial separator is open and during flight at altitudes above 5000 11, the maximum torque available can be reduced by up to 1.2 psi in non icing conditions, and up to 2.1 psi in icing conditions. 1

Report No: 01 973-001 9-08-90

Non icing

Icing conditions

Pneumatic boots failure

Non icing

Icing conditions

Issued: June 3,1996 1

STALL SPEED

KfAS

91

102

105

74

83

KCAS

93

104

107

76

85


TAKEOFF PERFORMANCE

When taking-oft in or into known icing conditions, the flaps must be set to 15' and the rotation speed increased by 9 KIAS. The speed at 50 R (15 m) height will be correspondingly increased by 12 KIAS. As a result, the takeoff ground roll distance will be increased by 29%. The takeoff distance will be increased by 42% and the accelerate-stop distance by 30%.

The takeoff ground roll and takeoff total distances given in the performance charts wiil be longer with operations on slush or snow covered runways.

CLIMB PERFORMANCE

During flight in icing conditions, the maximum rate of climb can be reduced by up to 1100 FPM. After failure of the pneumatic deice boots, the maximum rate of climb can be reduced by up to 1200 FPM below 15000 11 and by up to 1400 FPM at higher altitudes.

HOLDING ENDURANCE

During holding flight in icing conditions, a higher engine torque is required to maintain level flight. Increases in engine fuel flow between 25% and 50% are to be expected with respect to non icing conditions.

LANDING PERFORMANCE

After icing encounters and with visible ice accretion on the airframe, the landing is performed with 15' flaps and an approach speed of 108 KIAS. As a result, the landing ground roll distance can be increased by up to 909'0 without reverse thrust applied, and by up to 80% with reverse thrust applied. The landing total distances will correspondingly be longer by up to 55% without reverse thrust applied, and by up to 45% with reverse thrust applied.

After failure of the airframe pneumatic deice boots in icing conditions, the landing is performed with 0" flaps and an approach speed of 134 KIAS. As a result, the ground roll distance can be increased by up to 160% without reverse thrust applied, and up to 140% with reverse thrust applied. The landing total distances wiil correspondingly be longer by up to 90% without reverse thrust applied, and by up to 75% with reverse thrust applii.

BALKED LANDING PERFORMANCE

During flight in icing conditions, the balked landing rate of climb with 15' flaps can be up to 100 FPM lower than in non icing conditions with 40' flaps. After failure of the pneumatic deice boots the rate of climb can be up to 500 FPM lower.


Issued: June 3, 1996 Revision 5: Sep 1,2000

Report No: 01 973-001 9-08-91


Figure 6-8. Example Loading Form (Sheet 1 of 2)

Report No: 01973-001 I 9-08-92 Issued: June 3,1996

Revision 5: Sep 1,2000

=PILATUS= SECTION 9 PC12 SUPPLEMENT 8

23. Add Moment Due to Gear

--

Figure 6-8. Loading Form (Sheet 2 of 2)

Issued: June 3,1996 Revision 6: March 1,2003 93 ,of 98

Report No: 01 973-001 9-08-93

SE

CTIO

N 9

=P

ILA

TU

SW

S

UP

PLE

ME

NT 8

PC12

Figure 6-9. C. G

. Envelope (S

heet 1 of 2)

Report N

o: 01 973-001 Issued: June 3,1996

1 9-08-94 94 of 98

Revision 6: M

arch 1,2003

SE

CTIO

N 9

SU

PP

LEM

EN

T 8

Figure 6-9. C. G

. Envelope

(Sheet 2 of 2)


eport No: 01973-001

Revision 6: M

arch 1. 2003 95 of 98

9-08-95 1

SECTION 9 =PILATUSW SUPPLEMENT 8 PC12

SECTION 8 - HANDLING, SERVICING AND MAINTENANCE


Main Wheel Tires

I Wheel type - BFG PN3-1543 or 1543-1 (Post SB 32-013) Tire size - 8.50-10, 1OPR. TL (160 mph) Tire pressure - 60 + 3 - 0 psi (4.1 + 0.2 - 0 bar)

Report No: 01973-001 9-08-96




Issued: June 3,1996 . Revision 6: March 1,2003 97 of 98

Report No: 01973-001 9-08-97 1

SECTION 9 =PILATUSW SUPPLEMENT 8 PC12

Report No: 01973-001 I 9-08-98



aPILATUSE SECTION 9 -PC %I1 SUPPLEMENT 9




BENDIX I KING RDR 2000 WEATHER RADAR

This supplement must be anached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the BendidKing RDR 2000 Weather Radar is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Issued: May 1. 1996 Revision I : Nov 20,1996

Section for Type certification

Date of Approval:

2 2. Jan. 1997

Report No: 01973-001 9-09-1

SECTION 9 xPILATUS3 SUPPLEMENT 9 ?c %I1

SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the / BENDlXlKlNG RDR 2000 Weather Radar System is installed in accordance with FOCA

Approved Pilatus drawings.


Do not operate the radar during refueling operations or in the vicinity of trucks or containers accommodating flammables or explosives.

Do not allow personnel within 15 feet (4.5 meters) in front of antenna when system is transmitting.

Vertical profile (VP) weather display IS not allowed on the EHSl display (mode inhibited) but can be used on the EFlS MFD or the independent weather radar indicator (whichever is installed).


No changes to the Emergency Procedures described in We basic Pilot's Operating handbook.


Preflight and normal operaling procedures are presented in the Bendixfiing RDR 2000 Weather Radar Pilot's Guide KPN 006-08755-0001 latest revision.

TESTS INVOLVING THE RADIATION OF RF ENERGY BY THE RADAR ANTENNA MUST NOT BE MADE WHILE THE RADAR ANTENNA (S DIRECTED TOWARD CLOSE-BY LARGE METAL OBJECTS SUCH AS HANGARS, DOORS, OR THE INSIDE OF A HANGAR. USE TEST MODE OR TURN THE INDICATOR FUNCTION SWITCH TO OFF WHERE APPLICABLE.

NOTE

Operation of the BendixlKing RDR 2000 is inhibited automatically by the air/ground logic when the aircraft is on ground.

Report No: 01973-001 9-09-2

Issued: May 1.1996 Revision 1: Nov 20. 1996

mPKATUSW SECTION 9 -?c %I1 SUPPLEMENT 9


No change.




The RDR 2000 Weather Radar installation consists of a Receiver~Transmitter unit in a pod located at the right wing tip and a weather radar control panel. Weather indication is provided on the pilots EFlS and supplemented on the co-pilots EFlS (if installed) and the EFlS MFD or the independent weather radar indicator (whichever is installed). Vertical profile modes can be activated on the EFlS MFD or the independent weather radar indicator (whichever is installed).

NOTE

Detailed operating instructions or technical information for the BendixJKing RDR 2000 Weather Radar may be obtained by consulting the latest available revision of the Pilot's Guide.

Issued: May 1. 1996 Revision 1 : Nov 20,1996

Report No: 01 973-001 9-09-3


Report No: 01973-001 9-09-4


Issued: May 1,1996 Revision 1 : Nov 20, 1996




SUPPLEMENT NO. 10

OPERATIONS IN COLD CONDITIONS

This supplement must be anached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating the PC-12 and PC- 12/45 in cold conditions with SB 25-002 modifications installed. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations. procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Olfiie for Civil Aviation (FOCA) of Switzerland Section for_Type Certification

Date of Approval:

Issued: Nov 20.1996 Report No: 01 973-001 9-10-1


SECTION 1 - GENERAL

This supplement provides the informalion necessary lo operate the PC-12 and PC-12/45 a~rcralt in cold temoeratures with a balterv and an enaine heater svstem installed (Post SB 25-


OTHER LIMITATIONS

OPERATIONS IN COLD CONDITIONS

Ambient ground temperature 0 lo -1 5' C Battery heater required Ambient ground temperature -15" C and below Battery, engine and supplementary

cabin heater required. External engine blanket recommended.

A cabin underfloor temperature of - 15" or warmer is required prior lo takeoff.

The aircraft must be clear of deposits 01 snow, ice and lrost lrom the lining and control surfaces immediately prior to takeoff.


GENERAL

If the aircrafl is to be parked outside for a period of time and the ambient ground temperature is expected lo be:

010-15°C connect a 110 V AC ground power supply to the balery heater connector. -15" C and below connect a 1 10 V AC ground power supply to the balery and engine heater

connectors. Put a blanket cover over the engine. Put a supplementary heater in the center of the cabin.

PRE FLIGHT INSPECTION

Switch OH and disconnect 110 V AC ground power supply to banery, engine and supplementary heaters. Remove blanket cover lrom engine and supplementary heater from the cabin.

ENGINE STARTING

It is recommended to use the external power procedure lor engine starting, using a ground power unit capable of supplying 1000 Amp 28 V DC power.

Report No: 01973-001 9-10-2

Issued: Nov 20.1996

SPILATUSE SECTION 9 PC XI1 SUPPLEMENT 10

After engine start at cold temperatures of below -15' C, maximum cabin heating should be selected and the temperature of the underfloor avionic bay monitored by pressing the button near the cabin temperature indicator, to observe a minimum temperature of above -15" G prior to commencement of flight.

In the event ot the underfloor temperature sensor being inoperative, a period of 3 minutes per 10" below the ambient temperature of -15' should be allowed alter the selection of maximum cabin heating, to ensure that the underfloor equipment is brought up to an ambient operating temperature ol above -15" C prior to commencement of flight.

SECTION 7 - SYSTEM DESCRIPTION

Description

A belt type 110 V AC heating element is wrapped around the outside of the battery. A temperature sensor is attached to the aircraft skin and a wiring harness connects the temperature sensor to the battery heating element and the external power receptacle. The wiring harness has an additional connector installed lor the connection of a supplementary cabin heater. The receptacle is installed in the lower rear fuselage. A power on indicator light is installed near the receptacle.

When 110 V AC power is supplied to the external receptacle the indicator light will come on and power is supplied to the temperature sensor and supplementary heater connector. The temperature ssnsor will supply power to the battery heating element when the skin temperature ol the aircraft goes below 0° C.

ENGlNE HEATER SYSTEM

Description

Four 110 V AC wrap around type heating elements are installed on the engine at the following locations:

on the LH side of the reduct'in gearbox on the RH side of the reduction gearbox on the LH side of the accessory gearbox on the underside of the fuel control unit.

A wiring harness routed down the left side of the engine connects the heating elements to an external power receptacle. The receptacle is installed in Ihe lelt lower front fuselage. A power on indicator light is installed near the receptacle.

Issued: Nov 20. 1996 Report No: 01973-001 9- 10-3


Operation

When 110 V AC power is supplied to the external receptacle the indicator light will come on and power is supplied to the four engine heating elements. An insulated engine cover is placed over the engine to asslst in heat retention in the engine bay.

SUPPLEMENTARY HEATER

Supplementary ceramic element safety heater of 1500 W power is placed in the center of the aircran cabin to provide heating. The cabin heater is connected to the connector on the banery heating element wiring harness.

A temperature sensor is installed under the cabin floor between frames 17 and 18. A press bulton is installed near the cabin lemperature indicator on the cockpit center panel. When Ihe button is pressed the cabin temperature indication is changed from showing the cabin temperature to the cabin under floor avionic bay temperature.

SECTION 8 - HANDLING, SERVICING AND MAINTENANCE

SERVICING

At each aircraft inspection examine the battery and engine heating elements for damage and the wiring harnesses for security of attachment.

BATTERY SERVICING

When using the battery heating system for long periods of time it is recommended that the lead acid banery water level be checked at least once a week.

When removing and installing the battery from the battery box take care not to damage the heating element on the edges of the box. Small tears in the element can be repaired with R N silmne. If any of the element wire is exposed the element should be replaced.

Report No: 01973-001 9- 10-4

Issued: Nov 20.1996


AIRPLANE EQUIPMENT LIST

This list itemizes the equipment installed and required lor operations in cold weather conditions.

b Issued: Nov 20. 1996 Report No: 01973-001

9-10-5

1

ARM m

MASS

kg

VENDOR1 PILATUS PART NO.

Tanis TAS200-03 Tanis TAS300-15

Tanis PLT Tanis CVR 619

Jet Aviation 160-4N Or 160-2N

975.15.16.102

ITEM

SERVICING

Banery heater installation Engine heater installation Power on indication lght (2) Insulated engine cover 1 lOV power extension outlet

Supplementary heater Recommended 'Windmere' or similar Temperature sensor

Mark 'X' if

installed

ATA Code

12

SECTION 9 FPLATUSB SUPPLEMENT 10 ?C XI1


Report No: 01973-001 9-10-6

Issued: Nov 20, 1996





PC-12 AND PC-12/45 REGISTERED IN CANADA

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating the PC-12 and PC- 12/45 in Canada. The information contained herein supplements or supersedes the Information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations. procedures and performance information not contained in this supplement. consult the bask Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office foydivil Aviation (FOCA) of Switzerland S io for T, pe Certification u. LLL

Date of Approval: - \

issued: March 5,1997 Revision 3: March 13.2001

Report No: 01973-001 9-1 1-1



I Service Bulletin 04-002 lists the aircraft tasks which must be done prior to the registration of the aircraft and the use of this Supplement in Canada.

PC-1 2 CENTER OF GRAVITY LIMITS

PC-12145 CENTER OF GRAVITY LIMITS

Issued: March 5, 1997 Revision 1 : July 1,1997


P t PLACARDS

Luminescent placards on lnlerior of Cabin Door:

EXIT DO NOT OPERATE IN FLIGHT

TO OPEN DOOR LIFT LATCH ( ROTATE HANDLE

Issued: March 5. 1997 Report No: 01973-001 9-11-3


- - -

Lurnlnescent placard on lnler~or of upper LH srdewall near Cabin Door

( EXlT - ) Lum~nescent placard on Interlor of upper RH s~dewall near Emergency Exlt.

( EXIT * ) Luminescent placards on Interlor of Emergency Exit:

( EXIT )

Lum~nescent placards underneath cover on Emergency Exit:

Report No: 01973-001 9-11-4

Issued: March 5. 1997

~ P H A T U S ~ PC XII


Luminescent placards on rear of LH forward bulkhead:

FIRE EXTINGUISHER LOCATED ON RIGHT HAND SIDE BEHIND 1 CO-PILOT SEAT

LOCATED IN THE UN-PRESSURISED REAR

OPERATING INSTRUCTIONS: ARM: TO ARM ELT FUNCTION ON: TO ACTIVATE ELT FUNCTION OFF: TO DE-ACTIVATE ELT FUNCTION

Placard on cover plates under crew seats:

I NO STORAGE AREA j

Issued: March 5, 1997 Report No: 01973-001 9-1 1-5




3.15.1. ESNTL BUS

A. Battery Bus Failure

The following system remains operative:

Low airspeed warning


4.4 BEFORE ENGINE START

4.4.1. PROCEDURE

Additional item:

9. Before first flight of the day:

Trim interrupt switch INTR

Pitch trim switch OPERATE. Check trim interuppted

Trim interrupt switch NORM and guarded


Refer to POH Section 6, lnterior configurations and then to the applicable Interior Code No. for seat bcations, permitted seat Part Nos. that can be installed, seat weight and moment charts and seat occupant moment charts.

The following lnterior Configurations are approved for use in Canada:

Corporate Commuter Interior Code STD-9S (9 single seats)

Corporate Commuter lnterior Code STD-6s-38 (6 single seats and a 3 seat bench)

I Executive lnterior Code EX-6s-2 (6 single seats)

Executive lnterior Code EX-4s-3B (4 single seats and a 3 seat bench)

I Executive Interior Code EX-4s-STD-4S (4 single executive seats and 4 single standard seats)

Issued: March 5, 1997 Revision 3: March 13,2001

SE

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Issued: March 5, 1997

Report N

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9-11-10 l0

0f 12


SECTION 7 - AIRPLANE AND SYSTEMS DESCRIPTION

FLIGHT CONTROLS

INDICATION / WARNING SYSTEM

Page 7-7

The Central Advisory and Warning System (CAWS) annunciator panel inCll~deS a STAB TRlM warning light. The illumination of this warning captcon will illuminate when the aircraft is on the ground for more than 60 seconds and the trim position is unsafe for takeoff.

ELECTRICAL

CIRCUIT BREAKERS

Page 7-70

The AIS LOW WARN ctrcuct breaker on the LH front ccrcuit breaker panel IS deleted


INDICATIONANARNING

Page 7- 100

End of second paragraph additional sentence.

A red radial at 10.500 feet on the cabin pressure altimeter indicates the maximum permcss~ble cabcn pressure altitude.

CENTRAL ADVlSORY AND WARNING SYSTEM (CAWS)

Page 7-1 10

Issued. March 5 . i997

Index

9

12

Report No: 01973-001 9-11-11

Nomenclature

STAB TRIM

AIP DISENG

Color

Red

Amber

Description

Indicates stabilizer trim is unsafe for takeoff (comes on when aircraft is on ground lor more than 60 sec).

Indicates autopilot pitch and aileron servos disengaged (goes off after 30 secs).


AUTOPILOT

INDICATION 1 WARNING

Page 7- 174

A. P DISENG caut~on IS delayed 3 seconds In order to prevent aural warnlng conll~cl wrth the autop~lot system The cautlon will go off alter 30 seconds.

Reporl No: 01 973-00 1 9-11-12

Issued. March 5. 1997


PILors OPERATING HANDBOOK AND



ENGINE CONDITIONING TREND MONITORING (ECTM)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane FligM Manual when the Engine Conditioning Trend Monitoring system is installed in accordance with Pilatus SB 77-001 Revision 2. The information contained herein supplements or supersedes the information in

I the bask Pibt's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Offke for Civil Aviation (FOCA) of Switzerland Section for Type Cenlication

~ a i e of Approval:

Issued: Sep 8,1997 Revision 2: August 29,2000

Report No: 01973-001 9-12-1


- -

SECTION 1 - GENERAL

This supplement supplies the information necessary lor the operation of the airplane when the Engine Conditioning Trend Monitoring (ECTM) system is installed in accordance with the Pilatus I Se~ice Bulletin 77-001 Revision 2.


If the ECTM system is required by the Local Authorities for recording, data downloading and ECTM analyzing the local procedures must be followed.


ENGINE CONDITION TREND MONITORING (ECTM)

Description

The system is a TrendCheck passive system that acts as a recording device and exceedence monitor. It counts engine cycles, records engine running and flight times, identifies and measures hot starts and engine limit exceedences, records the maximum value observed by each sensor during flight and records automatically ECTM data when a stabilized condition exists lor 2 minutes. Stabilized conditions are defined by a preset altitude, torque and ITT band together with a maximum allowable change of these three parameters, aircraft in the air, inellial separator closed and the ECS not on P3. The maximum number of ECTM recordings per flight and the minimum time between two recordings can be defined according to the aircraft

1 operation. Torque, ITT. Ng. Np, engine oil pressure, fuel flow, ambient pressure and intake pressureltemperature data is supplied to a processor installed in the engine compartment. The processor also receives inputs from the LH weight on wheels switch, the inertial Separator bypass door actuator and the ECS high pressure SOV. Power to the processor is supplied from the GEN DIR BUS. An ECTM switch and a data down load socket are installed on the left crossbar panel in the flight compartmenl. All the internal data stored in the non-volatile memory can be downloaded to a ~ r 0 ~ n d based computer system for further analysis, action and system calibration and set-up. Accurate records of exceedences and engine trend data for the aircraft can be maintained. A TrendCheck stand alone ground station can be used with the instruction manual to calibrate the system.

Operation

When electrical power is switched on, the ECTM switch light will come on for 3 seconds during the self test. H the switch light remains on, the self test has failed or the processor memory is full. The system is not operational. If the light switch flashes continuously the processor memory is 85% full andlor an exceedence has been logged. The light will go off if the switch is pressed or the engine is started. The processor should be downloaded and reset. If not downloaded the light will flash again alter each Stan or shutdown.

Report NO: 01 973-001 9-12-2

Issued: Sep 8,1997 Revision 2: Auaust 29.2000


If the switch is pressed with the engine not running, the light will flash while the system carries out a 5 second loop test.

To take a manual recording the ECTM switch can be pressed when stabilized cruise conditions are reached. The ECTM switch will flash for the duration of the recording (set to 20 seconds).

Every time the system takes an automatic ECTM recording, the ECTM switch will flash for the duration of the recording (set to 20 seconds).

Stabilized cruise conditions are with level flight near max. cruise power setting (POH). landing gear and flaps up, inertial separator closed and the ECS not on P3. 1

I

Issued: Sep 8,1997 Revision 1: May 20. 1998

Report No: 0 1973-00 1 9-12-3


le3PILATUS3 PC XII 1

Report No: 0 1973-00 1 9-12-4


Issued: Sep 8, 1997 Revision 1: May 20. 1998


PILOT'S OPERANNO HANDBOOK AND



DUAL CHANNEL AUDIO CONTROLLER SYSTEM (NORTHERN AIRBORNE TECHNOLOGY AMS 44)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when a dual channel NAT AMS 44 Audio Controller System is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

J&, Approved by:

Federal Office for Civil Aviation (FOCA) of Switzerland Section for Type Certification

Date of Approval:

2 4. April 1998

Issued: January 19, 1998

- -

Report NO: 01973-001 9-13-1


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the dual channel NAT AMS 44 Audio Controller System is installed in accordance with Modification Approval Sheet PIL 12/23/039.


The NAT AMS 42/44 Dual Channel Audio Controller Manual. Section 3.8 to 3.14.1 (Rev 1.00. Dated 1 1 -Jan-95 or later applicable revision) must be readily accessible to the flight crew when operating the dual channel NAT AMS 44 Audio Controller System.


Audio System Failure Select EMER on red mode switch.

The emergency mode limits radio signal reception to pilot's and copilot's headsets. The transmitting function is limited to the pilot's hand mike.


No change.


No change.



SECTION 7 DESCRIPTION

GENERAL

The NAT AMS 44 Dual Channel Audio Controller System consists of a single audio control panel, mounted on top of the center instrument panel.

The unit is configured with the controls lor the pilot on lhe left side of the panel, the COntrOlS lor I the copilot and passengers on the right side.

The volume level of the pilot and copilot loudspeakers are independently controlled by rotary switches. The rotary switches are located near the pilot and copilot map lights.

t

Reporl No: 01973-001 9- 13-2

Issued: January 19,1998 2 Of 3


The pilot's hand mic and the headphones of the pibt and copilot function without power and will work in emergency or equipment failure situations.

AUDIO CONTROLLER EMERGENCY MODE OPERATION

In the event of an audio controller fault or power failure, set the red mode selector to EMER. This places the controller in emergency mode operation.

NOTE

When the red mode selector is switched to the EMER position. the pilot and the copilot are removed from the ICS bus and are connected directly to their respeclive radios.

EMERGENCY MODE EFFECTS

In the EMER mode, all functions are retained by the pilot and copilot. except ICS.

In the EMER mode, all switches work exactly as they do during NORM operation, except for the RX and ICS volume controls. These volume controls will have no effect.

PILOT ISOLATION OPERATION

When the mode switch is set to ISO, the pilot is isolated from ICS audio. This is useful to prevefit passengers interferirtg during critkal flight operatlons (landing, etc.).

PASSENGER RECEIVE AUDIO

The passengers hear the radio as selected on the right (copilot) side of the controller.

The passengers will not hear any radii audio when the red mode switch on the controller is in the EMER position.

UNMUTED AUDIO WARNINGS

The dual audio control panel receives unmuted audio warnings on two inputs. All unmuted audio warnings go to both inputs. One input is amplified and is connected to the pilot and copilot positions. The other input is connected directly to the pilot position only.

If there is an amplifier failure, the pilot will continue to receive all unmuted audio warnings.

Issued: January 19,1998 Report No: 01973-001 9-13-3





DUAl QIR TRAFFIC CONTROL TRANSPONDER SYSTEM (BENDIWKING KT70 TRANSPONDERS)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when a dual system BendixIKing KT70 Air Traffic Control Transponder System is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by: &J Federal Office for Civil Aviation (FOCA) of Switzerland

Section for Type Certification

Date of Approval:

0 5. Feb. 1998


I

p~ -

Report No: 0 1973-00 t 9-14-1


SECTION 1 - GENERAL

This supplement supplies the inlormation necessary for the operation of the airplane when the dual system BENDINKING KT70 Mode S Air Traffic Control (ATC) transponder system is installed in accordance with Modification Approval Sheet PIL 12J341040.


The BendidKing KT70 and KT71 (Panel-Mounted Transponders) Pilot's Guide (BendixlKing part no. 006-08490-0002, Dated 11192 or later revision) must be readily accessible to the flight crew when operating the dual KT70 system.


No change.


Press the XPNDRlIXPNDRP switch on the center instrument panel to change operation from one ATC transponder system to the other.


No change.



Repon No: 01973-001 9- 14-2


-



Two BENDlWKlNG KT70 Mode S ATC transponder systems are installed in the aircraft. One is used as the main system, the other is used as a standby.

A push-switch installed on the center instrument panel is used to select the 'active' transponder system. The switch has two righted captions. XPNDRl and XPNDR2. The relevant caption comes on to show which transponder system is 'active'.

The two transponder systems operate independently. Each transponder system has its own antenna and encoding altimeter. The pilot's encoding altimeter is used for the XPNDRl system. The copilot's encoding altimeter is used lor the XPNDR2 system.

The two transponder systems share a common L-band suppression circuit which is connected to the Distance Measuring Equipment and the Traffic Collision and Avoidance System (if installed).

The two transponder systems are also connected to the same weight-on-wheels switch, lo slop both systems from replying to ATCRBS interrogations when the aircraft is on the ground.

Issued: January 19,1998 Repon No: 0 1973-001 9-14-3


PILOTS OPERATING HANDBOOK AND



GROUND PROXIMITY WARNING SYSTEM (ALLIEDSIGNAL AVIONICS MK-VI GPWS)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when a MK-VI GPWS is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by: ~~ Federal Ofliie for Civil Aviation (FOCA) of Switzerland

Section for Type Certi f i i t in

Date of Approval:

2 7. M ai 1998

Issued: May 27. 1998 Report No: 01 973-001 9-15-1


SECTION 1 - GENERAL

This supplement supplies the inlormation necessary for the operation of the airplane when the MK VI GPWS is installed in accordance with Modification Approval Sheet PIL 12/34/041.


The MK VI GPWS Pilot's Guide (060-4087-000 Rev A, Dated 10196 or later applicable revision) must be readily accessible to the flight crew when operating the MK VI GPWS.


SECTION 3A - ABNORMAL PROCEDURES

3A.1 GENERAL

3A.l.l Press the GPWS FLAP OVRD switch to prevent operation of the 700 LOW FLAPS" alert when you need to land with less than full flaps.

3A.1.2 Press the GIS BELOW switch to deactivate the *GLIDESLOPE" alert when you intentionally plan to go below the glidesbpe.

3A.1.3 When the GPWS INOP amber caption illuminates to indicate a System failure, the failure may be identified by Ule following methods:

(a) When the aircraft is on the ground, press and hold the GPWS test switch to perform an Abnormal Sell Test. The test will prove i f the fault is external to the GPWS (with an aural fault message) or is internal.

(b) In flight, pull the RADAR ALT circuit breaker (the GPWS will assume it is on the ground). Press and hold the GPWS test switch to perform an lnflight Self Test.

NOTE

The aural message 'RADAR ALTIMETER FAULT" will be announced because the circuit breaker is pulled.

It the lault cannot be cleared, pull the GPWS circuit breaker.

3A.1.4 The GPWS can be deactivated by pulling the GPWS circuit breaker. located on the AVIONIC 2 circuit breaker panel, when a system failure cannot be cleared.


Perform a Self Test of the GPWS prior to flight.

Issued: May 27,1998

mPILATUSW SECTION 9 ?c XI1 SUPPLEMENT 15


No change.


Factory installed optional equipment is Included in the licensed weight and balance data in Sectii 6 of the basic Pilot's Operating Handbook and FOCA Approved Alrplane Flight Manual.


GENERAL

The MK VI GPWS consists of a GPWS test switch, lour GPWS indicator lampslswitches and. a GPWS computer.

GPWS TEST SWITCH

The GPWS test switch is installed on the AVIONIC TESTS switch panel on the left of the flight compartment.

GPWS INDICATOR LAMPSISWITCHES

There are four GPWS indicator lampslswitches installed on the left and right instrument panels:

GPWS

This indicator is red. It comes on when the GPWS computer detects an alert or warning in Modes 1 (Excessive Descent Rate), 2 (Excessive Closure Rate to Terrain). 3 (Altitude Loss After Takeoff) and 4 (Insufficient Terrain Clearance).

This indicator is amber. It comes on when the GPWS computer detects a failure of the system.

This indicator switch is amber. The BELOW caption comes on when the GPWS computer detects a Mode 5 (Descent Below Glideslope) alert situation.

The CANCLD caption comes on when the pilot or copilot switch is pressed for more than one second and released.

GPWS FLAP OVRD

This indicator switch is blue. The switch is covered by a switch guard. The caption comes on when the pilot or copilot swilch is pressed and released.

Issued: May 27,1998 Report No: 01 973-001 9-15-3


GPWS COMPUTER

The GPWS computer is installed below the passenger cabin floor. It uses the following System inputs for its computations:

- AHRS roll attitude - radar altimeter altitude AGL - air data computer vertical speed and airspeed - VHF NAVl system ILS/glideslope - autopilot decision height (OH) - flaps system flaps position - landing gear system landing gear position.

The GPWS computer sends unmuted voice messages, when necessary, to the flight compartment headphones and loudspeakers. At the same time, it sends a suppression signal to the TCAS (if installed) to inh~bit TCAS voice messages when the GPWS is generating voice messages.

The (Modes 1 thru 6) GPWS voice messages used are: "SINK RATE" 'PULL-UP" "TERRAIN-TERRAIN" 'DON'T SINK" "TOO LOW, GEAR" "TOO LOW, TERRAIN 'TOO LOW, FLAPS" 'GLIDESLOPE" "BANK ANGLE" "FIVE HUNDRED" "TWO HUNDRED" "MINIMUMS - MINIMUMS"

Report No: 01 973-001 9- 15-4

Issued: May 27,1998





PC-12 AND PC-12/45 REGISTERED IN THE

REPUBLIC OF SOUTH AFRICA (RSA)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating Ihe PC-12 and PC-12/45 in the RSA. The listing contained herein supplements the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office for Qvil Aviation (FOCA) of Switzerland Secti n r Typ CG?Ricalion &L. LL-

!ate of Approval:

Issued: March 12,1998 Report No: 0 1973-001 9-16-1


-- - - ---


If installations are made to the aircratt which are not covered by the PC-12 andlor PC-12/45 type design (i.e. STC's, Major Alterations) and require operating instructions. the installer must lisr the relevant documents below and insert them in the POHIAFM. It is Ihe installers responsibil~ty to make sure that the installations are approved by the RSA CAA.

This listing is applicable to:

Aircralt Serial No: ........................

Aircralt Registration No: ........................

Report No: 01973-001 1 9-16-2

Issued: March 12.1998

Signature of the Installer

No. Equipment Details CAA Mod No.



No.

7

Report No: 01973-001 9-16-3

Equipment Detalls C M Mod No.

Signature of the Installer

-

SECTION 9 BPILAIIISW SUPPLEMENT 16 ?C %!I

Report No: 01973-001 9- 16.4

1


No. CAA Mod No.

Equipment Detalls Signature of the Installer





MECHANICAL COPILOT INSTRUMENTATION

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the Mechanical Copilot Instrumentation is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.


Date of Approval:

1 7. F2b. 1999

Issued: January 29, 1999 Report No: 01973-001 9-17-1


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the mechanical copilot instrumentation is installed in accordance with Modification Approval Sheet PIL 12/23/074.


The Mechanical Copilot Instrumentation is limited as a supplemental display only.


No change.


No change.


No change.



Repon No: 01973-001 9-17-2

1


-- -

'3PlLATUSS ?C XII

SECTlON 9 SUPPLEMENT 17

SECTION 7 - SYSTEM DESCRlPnON

ATTITUDE INDICATOR GYRO

The Attitude Indicator Gyro is located on the copilot's instrument panel between airspeed lndicator on the left and altimeter on the right side. The attitude gyro provides a source of pitch and roll information. The attitude gyro is in operation whenever the GEN 1 Bus is powered.

AN OFF FLAG WILL APPEAR WHEN POWER IS REMOVED FROM THE GYRO.

PICTORIAL NAVIGATION SYSTEM KCS 55

The Pictorial Navigation System KCS55A consists of a Directional Gyro KG-102A, a Pictorial Navigation Indicator KI-525A, a Slaving Accessory KA-JIB, and a Magnetic Azimuth Trans- mitter KMT-112. The KCS55A System is powered by the Avionics 2 Bus CB COMPASS 2.

The Directional Gyro KG-102A is located in the avionics bay under the floor. It is a remote mounted unit that, in conjunction with the KMT-112 Magnetic Azimuth Transmitter, provides a gyro-stabilized magnetic heading to the KI-525A indicator. It also supplies the pilot's EFlS EADl and EHSI with heading information. A heading source selection switch HDG lmDG 2 is installed in the pilot's instrument panel to the right of Ule GPS select switches. Select HDG 1 for heading information from AHRS 1, and HDQ 2 b r heading lnformatlon frorn.the Compass System No. 2. The selected heading source is displayed on the EFlS EADl EHSI, and on the selector switch.

The Ki-525A is located on the copilot's instrument panel below the attitude gyro. The Pictorial Navigation lndicator KI-525A provides a pictorial display of the horizontal navigation situation. It also provides manual controls for course and heading datum selection. The Pictorial Navigation indicator KI-525A receives converted VOWLOClGS navigation data from the VHF COMMAV 2 system.

The Slaving Accessory KA-SIB is located on the copilot's instrument panel to the right of the altimeter. The Slaving Accessory KA-518 contains the slaving meter, slaving switches, and corrector circuitry which compensates for the effect of local magnetic disturbances on the Magnetic Azimuth Transmitter. The SLAVEIFREE switch selects a free gyro mode when set to free, and slave gyro mode when set to slave. The CWICCW switch moves the heading card clockwise CW and counterclockwise CCW when the slavdree switch is set to free. The Magnetic Azimuth Transmitter KMT-112 senses the direction of the

Issued: January 29.1999 Report No: 01 973-001 9-17-3


earths magnetic field and transmits this information to the Pictorial Navigation Indiitor. It is located in the outer right wing.

IN CASE OF AHRS 1 SYSTEM FAIL AND HEADING SOURCE SELECT SWITCH ON HDG 2, NO AlTlTUDE INDICATION ON PILors EFIS EADI AND THE AUTO FLIGHT IS INOPERATIVE.

COURSE DEVIATION INDICATOR

The Course Deviation lndiitor KI 206 is located on the copilot's instrument panel to the right of the vertical speed indicator. The CDI indicator KI-206 is powered by the VHF COMRJAV 1. It receives converted VOWLOC and GS navigation data from the VHF C O W V 1 system.

RADIO MAGNETIC INDICATOR

The RMI Kl-229 is located on the copilot's instrument panel below the airspeed indicator. The Kl-229 is powered lrom the by the Avionic 2 Bus CB RMI 2. It receives its excitation from the 26 VAC Bus CB RMI 2. The Kl-229 receives navigation data from the VHF NAV 2 and the ADF receiver. The heading information is received from the KCS55A compass system.

AUTOMATIC DIRECTION FINDER INDICATOR (IF INSTALLED)

The Automatic Direction Finder Indicator Kl-227 could be optional installed as a replacement of the Radio Magnetic Indicator Kl-229. It is the same location as the Kl-229. The ADF W i t o r Kl-227 is powered by the ADF. The ADF indicator Kl-227 receives navigation data from the ADF receiver, and has a synchronised compass card which is driien by the KCS55A compass system.

I CAUTION (

AFTER THE START-UP COMPARE HEADING FROM Kl-227 WITH THE ACTUAL HEADING. IF READING IS NOT EQUAL, SET Kl-227 COMPASS CARD WlTH HDG KNOB TO ACTUAL HEADING.

DISTANCE MEASUREMENT EQUIPMENT

The Distance Measurement Equipment Indicator KDI-574 is located on the copilot's instrument panel above the attitude indicator. The DME indicator KDI-574 is powered by the DME. The KDl-574 receives navigation data from the DME and displays an alphanumeric readout. Such as distance in nautical miles NM, ground speed in knots KTS, and time tofirom station in

Report NO: 01973-001 917-4

Issued: January 29,1999 4015


minutes MIN from the aircraft to the selected DME ground station. The DME tuning is accomplished by the selected NAV 1 or NAV 2 receiver frequency controls. The DM€ tuning source selector is installed on the pilot's instrument panel and the selected source is displayed on the KDI 574 to the rigM of the distance annunciation.

I CAUTION I THE DISTANCE MEASUREMENT EQUIPMENT INDICATOR DME KDI 574 MUST BE INSTALLED AND SERVICEABLE. THE KDI 574 SUPPLIES DISTANCE SPEED AND TIME TOIFROM STATION INFORMATION TO THE PILOTS EFlS EHSI.


- - - -

Reporl No: 01973-001 9-17-5




SUPPLEMENT NO. 18

PC-12 REGISTERED IN FRANCE

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when operating the PC-12 In France. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limlatlons, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:


Date of Approval:

Issued: October 29,1999 Revision 1 : December 16, 1999

Report No: 01 973-001 9-18-1


SECTION 1 - GENERAL <

This Airplane Flight Manual Supplement gives the information and lists the specifics necessary for the operation of the airplane on the French register in accordance with the requirements of the French DGAC. 4

1

Report No: 01973-001 418-2

I

Issued: October 29, 1999



Service Bulletin 04-005 lists the modifications (tasks) that must be done before registration of the airplane in France and before use of this supplement.

The three French DGAC-approved pages (Ref: Report 01973-001-FR, MANUEL DE VOL) must be anached to the original Swiss FOCA-approved Airplane Flight Manual (AFM) (Ref. Report 01 973-001).

I STEEP APPROACH I Steep approaches greater than 6 O are not approved. I KINDS OF OPERATION EQUIPMENT LIST

SYSTEM I EQUIPMENT VFR VFR IFR IFR ICING DAY NIGHT DAY NIGHT

Fllght Instruments:

Altimeter Airspeed Indicator PitoVStatic System AHRS

'If EFlS is on the co-pilot side

Emergency Equipment:

Emergency Locator Transmitter (ELT) 1 1 1 1 1 (French DGAC-approved unit)

PLACARDS

Exterlor Markings:

On the cabin door:

TIRER LA POIGN~E ET TOURNER POUR OUVRIR NE PAS OUVRIR LA POQTE LORSQUE LE MOTEUR TOURNE SAUF EN CAS D'URGENCE

OUVERT

Issued: October 29, 1999 Revision 1 : December 16, 1999

Report No: 01 973-001 9-18-3


On the cargo door:

( TlRER POUR OUVRIR (Accompanied by a downward pointing arrow)

APPUYER ICI POUR OUVRIR LA PORTE TIRER LA POIGNEE ET TlRER LA PORTE VERS L'EXTERIEUR NE PAS OUVRIR LA PORTE LORSQUE LE MOTEUR TOURNE SAUF EN CAS D'URGENCE

On each side of the rudder, on the top surface of each aileron and three places on the top surface of each flap (total ten places):

NE PAS POUSSER

On the emergency exit:

SORTIE DE SECOURS

POUSSER

POUSSER APRES DEVERROUILLAGE

On the left side of the rear fuselage:

APPUYER SUR LE BOUTON POUR ABAISSER LA PORTE

On each side of the engine air intake:

SURFACE TRES CHAUDE NE PAS TOUCHER

lnterlor Placards - Corporate Commuter Configuration:

Near eight ot the nine passenger seat$:

OXYGENE OXYGEN

On the rear of eight of the nine passenger seats:

LE MASQUE A OXVGENE SE TROUVE SOUS LE SI~GE

Issued: October 29, 1999 Revision 1: December 16,1999


On the rear of one of the nine passenger seats:

LE MASQUE A OXYG~NE SE TROUVE SOUS LE SBGE AVANT

On the rear of the nine passenger seats:

POUR LE DECOLLAGE ET L' ATTERRISSAGE - CEINTURE ATTACH~E - DOSSIER EN POSITION VERTICALE FOR TAKEOFF AND LANDING - FASTEN SEAT BELT - SEAT BACK MUST BE FULLY UPRIGHT

On the left cockpitlcabin divider wall:

NE PAS FUMER NO SMOKING

LE MASQUE A OXYG~NE SE TROUVE SOUS LE s I~GE OXYGEN MASK LOCATED UNDER SEAT

POUR LE DECOLLAGE ET L'ATTERRISSAGE - CEINTURE ATTACH~E - DOSSIER EN POSITION VERTICALE FOR TAKEOFF AND LANDING - FASTENSEATBELT - SEAT BACK MUST BE FULLY UPRIGHT

L'EXTINCTEUR EST SITU~ SUR LE COT^ DROlT DU COCKPIT DERR~RE LE SIEGE DU CO-PILOTE FIRE EXTINGUISHER LOCATED ON COCKPIT SIDE R.H. BULK- HEAD BEHIND CO-PILOT SEAT (Accompanied by an arrow pointing to the right)


TlRER I PULL (white letters, transparent background)

SORTIE I EXlT

TlRER 1 PULL (white letters, red background)

On the cabin door:

SORTIE I EXlT NE PAS OUVRIR EN VOL DO NOT OPERATE IN FLIGHT

Issued: October 29,1999 Repon No: 01973-001 9-18-5


POUR OUVRIR LA PORTE LEVER LA POIGNEE ET TOURNER TO OPEN LIFT LATCH ROTATEHANDLE

OUVERT OPEN

FERME CLOSED

NE PAS OUVRIR LA PORTE LORSQUE LE MOTEUR TOURNE SAUF EN CAS D'URGENCE

UNE SEULE PERSONNE A LA FOlS SUR L'ESCALIER

On the cargo door:

NE PAS ENLEVER LE COUVERCLE EN VOL

TlRER LE VERROU TlRER LA POIGN~E ET POUSSER LA PORTE VERS L'EXTERIEUR LlFT LOCKING LEVER AND PULL HANDLE PUSH DOOR OUT


Interior Placards - Executlve Configuration:

On the cabin door:

SORTIE 1 EXIT NE PAS OUVRIR EN VOL DO NOT OPERATE IN FLIGHT

POUR OUVRIR LA PORTE LEVER LA POIGNEE ET TOURNER TO OPEN LIFT LATCH ROTATE HANDLE

OUVERT OPEN

issued: October 29, 1999


FERMG CLOSED

NE PAS OUVRIR LA PORE LORSQUE LE MOTEUR TOURNE SAUF EN CAS D'URGENCE

UNE SEULE PERSONNE A LA FOlS SUR L'ESCALIER

On the cargo door:

NE PAS ENLEVER LE COUVERCLE EN VOL

TIRER LE VERROU TIRER LA POIGN~E ET POUSSER LA PORTE VERS L'EXTERIEUR LIFT LOCKING LEVER AND PULL HANDLE PUSH DOOR OUT



TlRER 1 PULL (white letters, transparent background)

SORTIE I EXIT

TlRER I PULL (white letters, red background)

Near each of six passenger seats:

NE PAS FUMER LORS DE L'UILISATION DE L'OXYG~NE DO NOT SMOKE WHILE OXYGEN IS IN USE

POUR LE DECOLLAGE ET L'ATTERRISSAGE - LE SIEGE DOlT ETRE COMPL~TEMENT REDRESS~ COMPL~EMENT EN ARRI~RE COMPL~EMENT A L'EXTCRIEUR - LA TABLE DOlT ETRE REPLICE

FOR TAKEOFF AND LANDING - SEAT MUST BE FULLY UPRIGHT FULLY AFT AND FULLY OUTBOARD - TABLE MUST BE STOWED

Near five of six passenger seats:

MASQUE A OXYG~NE A L'INTI~IEuR OXYGEN MASK INSIDE

Issued: October 29,1999 Report No: 01973-001 9- 18-7


Near four of six passenger seats:

TITER LA BANDE POUR LlBERER LE MASQUE A OXYGENE PULL TAPE FOR OXYGEN MASK (accompanied by a rearward pointing arrow)

Near two of six passenger seats:

TITER LA BANDE POUR LlBERER LE MASQUE A OXYGENE PULL TAPE FOR OXYGEN MASK (accompanied by a forward pointing arrow)

On the left cockpivcabin divider wall:

L'EXTINCTEUR EST SIT& SUFI LE COT& DROIT DU COCKPIT DERRIERE LE SIEGE DU CO-PILOTE FIRE EXTINGUISHER LOCATED ON COCKPIT SIDE R.H. BULK- HEAD BEHIND CO-PILOT SEAT (accompanied by an arrow pointing to the right)

SECTION 7 - AIRPLANE AND SYSTEMS DESCRIPTION

In the POH on page 7-169 the following Note is added after the HALF BANK paragraph:

NOTE: Depending on course changes, Half Bank mode may lead to wide turns exceeding RNP-5 volumes.

Report No: 01 973-001 9-18-8

Issued: October 29,1899 Revision 1: December 16, 1999



At the date of issue of this supplement, these are the FOCA-approved supplements (to the Airplane Flight Manual) that are also accepted by the DGAC for French registered airplanes:

NOTE

For French DGAC acceptance of supplements issued after the date of this supplement (No.18), please contact the French DGAC.

Issued: October 29, 1999 Revision 1 : December 16, 1999

Report No: 01 973-001 9- 18-9





OPERATION WITH GPS KLN 900 NAVIGATION SYSTEM

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the GPS KLN 900 Navigation System is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic PiloYs Operating Handbook and FOCA Approved Airplane flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consuH the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Issued: 22 December, 1999 Report No: 01 973-001 9-19-1


SECTION 1 -GENERAL

This supplement supplies the information necessary for the operation of the airplane when the AlliedSignal Aerospace GPS KLN 900 Navigation System is installed.

The GPS KLN 900 Navigation System is a panel-mounted unit that contains the GPS sensor, the navigation computer, a CRT display and all controls required to operate the unit. It also houses the data base cmridge which plugs directly into the front of the unit.

The data base cartridge is an electronic memory containing information on airports, navaids, intersections, SID's, STAR'S, instrument approaches, special use airspace, and other Hems of value to the pilot.

Every 28 days, AiliedSignal receives new data base information from Jeppesen Sanderson. This information is processed and downloaded onto the data base cartridges. AlliedSignal makes these data base cartridge updates available to GPS KLN 900 Navigation System users.

Provided that the GPS KLN 900 Navigation System receives adequate usable signals, it has been demonstrated to be capable of and has been shown to meet the accuracy specifications of VFRllFR en route oceanic and remote, en route domesti, terminal, and instrument approach operation within latitudes bounded by 74' North and 60° South using the WGS (or NAD 83) coordinate reference datum in accordance with the criteria of AC 20-138, AC 91-49, AC 120-33, and FAA Notice 81 10.60.

Provided that the GPS KLN 900 Navigation System receives adequate usable signals, that the airplane has alternative, serviceable navigation systems (8.g. VOR; DME; ADF) adequate to the route to be flown and full radar coverage, it has demonstrated that Ule system installation can perform Basic RNAV (RNP-5) operations in designated European airspace in accordance with JAA Temporary Guidance Leaflet No 2 (TGL 2) rev 1: AMJ 20x2 - JAA Guidance Material on Airworthiness Approval ;nd Operational Criteria for the use of Navigation Systems in European Airspace designated for Basic RNAV operations.

The KLN 900 GPS equipment as installed has been found to comply with the requirements for GPS primary means of navigation in oceanic and remote airspace, when used in conjunction with the KLN 900 prediction program. This does not constitute an operational approval.

Report No: 01973-001 9-19-2

Issued: 22 December, 1999


NOTE

Airplanes using GPS navigation Systems for oceanic IFR operations may use the KLN 900 if the installations are approved in accordance with FAA Notke 8110.60 or equivalent JAA or national documentation and their operation is approved in accordance with FAA HBAT 95-09 or equivalent JAA or National documentation

A single KLN 900 system installation may be used on short oceanic routes where only one Long Range Navigation System is required.

Issued: 22 December. 1999 Report No: 01973-001 9- 19-3

SECnON 9 SUPPLEMENT 19


A. The KLN 900 GPS Pilot's Guide, PM 006-08796-0000, (or later applicable revision) must be immediately available to the flight crew whenever navigation is predicated on the use of the system. The Operational Revision Status (0%) of the Pilot's Guide must match the ORS level annunciated on the Sell Test page.

B. IFR Navigation is restricted as follows:

(1) The system must utilize ORS level 04 or later approved revision.

(2) The data on the self test page must be verified prior to use. Verify valid altitude data is available to the KLN 900 prior to flight.

(3) IFR en route and terminal navigation is prohibited unless the pilot verifies the database contains the current update cycle or each selected waypoint for accuracy by reference to current approved data.

(4) Instrument approaches must be accomplished in accordance with approved instrument approach procedures that are retrieved from the KLN 900 data base. The KLN 900 data base must incorporate the current update cycle.

(a) Instrument approaches must be conducted in the approach mode and RAlM must be available at the Final Approach Fix.

(b) APR ACT (approach active) mode must be annunciated at the Final Approach Fix.

(c) Accomplishment of ILS, LOC, LOC-BC. LDA, SDF and MLS approaches are not authorized.

(d) When an alternate airpon is required by the applicable operating rules, it must be served by an approach based on other than GPS navigation.

(e) The KLN 900 can only be used for approach guidance if the reference co-ordinate datum system for the instrument approach is WGS-84 or NAD-83. (All approaches in the KLN 900 data base use the WGS-84 or the NAD-83 geodetic datum and approval of the NAA.)



(5) The use 01 the KLN 900 to perform Basic RNAV (RNP-5) operations in the designated European airspace is limited as follows:

(a) The system must have the GPS XPRESS software update we& the following display on the STA 3 page:

HOST01 51 5-0046 RCVR 01621-0001

(b) The KLN 900 data base must incorporate ihe current update cycle.

(c) Given a GPS constellation of 23 satellites or less (22 or less when the KLN 900 incorporates automatic pressure altitude aiding) is projected to be operational, the availability of RAlM must be confinned for the intended flight (route and time). Dispatch for Bask-RNAV must not be made in the event of predded continuous bss of RAlM of more than 5 minutes for any part of the intended I i h t . For RAlM prediction Preflight Version 2.0 or equivalent approved software must be used.

(d) Radar coverage must be available for the route to be flown.

(6) Traditional approved navigation equipment (e.g. VOR, DME. ADF) adequate for the route to be flown must be installed and sewiceable for use of the KLN 900 GPS navigation system in accordance with the operational approval.

Issued: 22 December, 1999 Report No: 01 973-001 9- 19-5



ABNORMAL PROCEDURES

A. If the KLN 900 GPS information is not available or invalid, utilize traditional operational t

navigation equipment as required.

6. If the alarm limit is exceeded during Basic-RNAV operation, revert to an alternative means of navigation.

C. H a "RAIM NOT AVAILABLE" message is displayed while conducting an inst~ment approach, terminate the approach. Execute a missed approach if required.

D. If a 'RAIM NOT AVAILABLE" message is displayed during Basic-RNAV operation. navigation may continue using the KLN 900 provided crosscheck wlh VOR, DME andtor NDB information shows an acceptable level of navigation performance (RNP- 5). Otherwise revert to an alternative navigation as required.

E. Refer to the KLN 900 Pilot's Guide, Appendices B and C, for appropriate pilot actions to be accomplished in response to annunciated messages.

Report No: 01973-001 9-19-6




FAMILIARITY WITH THE EN ROUTE OPERATION OF THE KLN 900 DOES NOT CONSTITUTE PROFICIENCY IN APPROACH OPERATIONS. DO NOT ATTEMPT APPROACH OPERATIONS IN IMC PRIOR TO ATTAINING PROFICIENCY IN THE USE OF THE KLN 900.

OPERATION

A. Normal operating procedures are outlined in the KLN 900 GPS Pilot's Guide, PIN 006- 08796-0000 (ORS 04 or later applicable revision).

B. During Basic-RNAV operation, traditional navigation equipment (e.g. VOR, DME and ADF) should be selected to available ground based aids so as to allow immediate cro~s-~he~king or reversion in the event of loss of GPS navigation capability.

C. Before joining and during operation on an RNAV route, correct operation of the RNAV system shall be established. This includes that:

the routing is in accordance with the clearance

the navigation performance accuracy meets RNP-5 requirements

SYSTEM ANNUNClATORSISWITCHESICONTROLS

A. GPS Status message (MSG) annunciation - In addition to the MSG annunciation on the KLN 900 GPS, the white MSG annunciation on the EFlS display will flash for five seconds (then go solid) to alert the pilot of a situation that requires attention. Press the MSG button on the KLN 900 GPS to view the message. Appendix B of the KLN 900 Pilots Guide contains a list of all of the message page messages and their meanings.

B. GPS waypoint 0 alerting - The KLN 900 GPS will provide navigation along a curved path segment to ensure a smooth transition between two adjacent legs in the night plan. This feature is called turn anticipation. Approximately 20 seconds prior to the beginning of turn anticipation, the arrow preceding the active waypoint will begin flashing on the FPL 0 page, going solid upon initialization of the turn, and extinguishing upon turn completion. At the same time, the white WPT annunciation on the EFlS display will flash for five seconds, then go solid and extinguish upon turn completion.

Issued: 22 December, 1999 7 of 21

Report No: 01973-001 9-19-7


I WARNING 1 TURN ANTICIPATION IS AUTOMATICALLY DISABLED FOR FAF WAYPOINTS AND THOSE USED EXCLUSIVELY IN SlDlSTAR WHERE OVER-FLIGHT IS REQUIRED. FOR WAYPOINTS SHARED BETWEEN SID/STAR AND PUBLISHED EN ROUTE SEGMENTS (REQUIRING OVER- FLIGHT IN THE SIDISTAR), PROPER SELECTION ON THE PRESENTED WAYPOINT PAGE IS NECESSARY TO PROVIDE ADEQUATE ROUTE PROTECTION ON THE SIDISTAR.

C. The OBS switch on the KLN 900 bezel (face) is used to select one of two basic course modes of KLN 900 operation, either:

(1) OBS, which is single waypoint with omni bearing (OBS) selection through that waypoint (like a VOR).

(2) LEG which is automatic leg sequencing between waypoints.

0 D. The course control knob on the EFIS control panel provides analog course input to the KLN 900 in OBS mode when the GPS source is selected on the EFIS. When other than GPS navigation sources are selected, GPS course selection in OBS mode is digital through the use of the controls and display of the KLN 900.

E. GPS omni bearing or leg CRS (white) OBS (whitey LEG (green) course switch/annunciator - Used to select the basic modes of KLN 900 operation, either:

1) single waypoint with omni - bearing course (08s) selection through that waypoint (like a VOR)

2) automatic leg sequencing (LEG) between waypoints.

NOTE

Either LEG or OBS wiU illuminate during system self test depending upon switch position.

Report No: 0 1973-001 9-19-8



F. A GPS approach APR (white) ARM (white)/ACT (green) remote switch/annunciator is located to the left of the EIS, below the alimeter. The APR and ARM annunciators are white and the ACT annunciator is green. The APR (white) annunciator is repeated on the EHSl display. The remote switch is wed to:

(1 ) Manually select or deselect approach ARM (or deselect appmach ACT).

and

(2) Annunciate the stage of approach operation either armed (ARM) or activated (ACT).

Sequential button pushes if in ACT would first result in approach ARM and then approach arm canceled. Subsequent button pushes will cycle between the armed state (1 an approach is in the flight plan) and approach arm canceled. Approach ACT cannot be selected manually.

PILOT'S DISPLAY

Lettlright steering information is presented on the EFlS as a function of the navigation source selection on the EFlS control panel.

AUTOPILOT COUPLED OPERATION

The KLN 900 may be coupled to the autopilot by first selecting GPS on the NAVlGPS switch. The autopilot approach mode (APR) should be used when conducting a coupled GPS appmach.

NOTE

NAV or APR coupled DME arc intercepts can result in excessive overshoots (aggravated by high ground speeds and intercepts from the arc).

Tracking the GPS in NAV autopilot mode with a 2 0.3 nm CDI scale factor selected is not recommended. The GPS CDI scale factor may be changed on the KLN 900 MOD page.

Issued: 22 December. 1999 Report No: 01973-001 9- 19-9


APPROACH MODE SEQUENCING AND RAIM PREDICTION

NOTE

The special use airspace alert will automatically be disabled prior to flying an instrument approach to reduce the potential for message congestion.

A. Prior to arrival, select a STAR if appropriate from the APT 7 page. Select an approach and an Initial Approach Fix (IAF) from the APT 8 page.

NOTE

Using the right hand outer knob, select the ACT (Active Flight Plan Waypoints) pages. Pull the right hand inner knob out and scroll to the destination airport, then push the inner knob in and select the ACT 7 or ACT 8 page.

To delete or replace a SID, STAR or approach, select FPL 0 page. Place the cursor over the name of the procedure, press ENT to change it, or CLR then ENT to delete it.

6. En route, check for RAlM availability at the destination airpon ETA on the STA 5 page.

NOTE

RAlM must be avtilable at the FAF in order to fly an instrument approach. Be prepared to terminate the approach upon loss of RAIM.

C. At 30 NM from the FAF:

(1) Verify automatic annunciation of APR ARM.

(2) Note automatic dbar scaling change from * 5.0 NM to * 1.0 NM over the next 30 seconds.

(3) Update the KLN 900 altimeter bar0 setting as required.

(4) Internally the KLN 900 will transition from en route to terminal integrity monitoring.



D. Select GPS NAV 5 page or applicable EFlS presentation, to fly the approach procedure.

(1) If receiving radar vectors, or need to fly a procedure turn or holding pattern, fly in OBS until inbound to the FAF.

NOTE

OBS navigation is TO-FROM (like a VOR) wilhout waypoint sequencing.

(2) NoPT routes including DME arcs are flown in LEO. ihe Final Awnad~ Fix WAF) to the MAP,

NOTE

for DMF arc interceots. NAV or APR coupled DME arc intercepts can result in excessive overshoots (aggravated by high ground speeds andlor intercepts from the arc).

FLYING FINAL OUTBOUND FROM AN OFF-AIRPORT VORTAC ON AN OVERLAY APPROACH; BEWARE OF THE DME DISTANCE INCREASING ON FINAL APPROACH, AND THE GPS DISTANCE-TO-WAYPOINT DECREASING. AND NOT MATCHING THE NUMBERS ON THE APPROACH PLATE.

E. At or before 2 NM from the FAF inbound:

(1) Select the FAF as the active waypoint, if not accomplished already.

(2) Select LEG operation.

F. Approaching the FAF inbound (within 2 NM.):

(1) Verify APR ACT.

(2) Note automatic dbar scaling change fmm f 1.0 NM to * 0.3 NM over the 2 NM inbound to the FAF.



(3) Internally the KLN 900 will transition from terminal to approach integrity I

monitoring.

G. Crossing the FAF and APR ACT is Mt annunciated:

(1) Do not descend.

(2) Execute the missed approach.

H. Missed Approach:

(1) m (2) Navigate to the MAP (in APR ARM if APR ACT is not available).

NOTE

There is MI automatic LEG sequencing at the MAP.

in accordance with the published missed approach procedure, verify or change the desired holding fix and press ENT.

,

GENERAL NOTES 4

The data base must be up to date for non-precision instrument I

approach operation.

Only QW approach can be in the active flight plan at a time. 1

If the destination airpon is the active waypoint at the time of the instrument approach selection, the active waypoint will shift automatically to the chosen IAF.

Checking RAIM prediction for your approach while en route using the STA 5 page is recommended. A self-check occurs automatically within (

2 NM of the FAF. APR ACT is inhibited without RAIM.

Data cannot be altered, added to or deleted from the approach procedures contained in the data base. (DME arc intercepts may be relocated along the arc through the SUPER NAV 5 or the FPL 0 pages).

Repon No: 01973-001 9-19-12



Some approach waypoints do not appear on the approach plates (Including in some instances the FAF).

Waypoint suffixes in the flight plan:

m MAP

h missed approach holding fix.

The DME arc IAF (arc intercept waypoint) will be:

(a) On your present position radial off the arc VOR when you load the IAF into the flight plan.

(b) The beginning of the arc if currently on a radial beyond the arc liml.

To adjust the arc intercept to be compatible with a current radar vector, bring up the arc IAF waypoint in the SUPER NAV 5 page scanning field, or applicable EFlS GPS presentation, or under the cursor on the FPL 0 page, press CLR, then ENT. Fly the arc in LEG. Adjust the HSI or CDI course pointer with reference to the desired track value on the SUPER NAV 5 page, or applicable EFlS GPS presentation (i will flash to remind you). LeWright DBAR information is relative to the arc. Displayed distance is M along the arc but direct to the active waypoint. If desired, select NAV 2 page for digital DME arc distance to and radial from the reference VOR.

The DME arc IAF identifier may be unfamiliar. Example: D098G where 098 stands for the 098O radial off the referenced VOR, and G is the seventh letter in the alphabet indit ing a 7 DME arc.

APR ARM to APR ACT is automatic provided:

(a) You are in APR ARM (normally automatic).

(b) You are in LEG mode

(c) The FAF is the active waypoint

Issued: 22 December. 1999 13 of 21

Reporl No: 01 973-001 9-19-13


(d) Within 2 NM of the FAF.

(e) Outside of the FAF.

(1) Inbound to the FAF.

(g) RAlM is available.

Direct-To operation between the FAF and MAP cancels APR ACT. Fly the missed approach in APR ARM.

Flagged navigation inside the FAF may usually be restored (not guaranteed) by pressing the GPS APR button changing from ACT to ARM. Fly the missed approach.

The instrument approach using the KLN 900 may be essentially automatic starting 30 NM out (with a manual barometer setting update) or it may require judicious selection of the OBS and LEG modes.

APR ARM may be cancelled at any time by pressing the GPS APR bunon. (A subsequent press will reselect it.)


No change





SECTION 7 -AIRPLANE AND SYSTEMS DESCRIPTION

DESCRIPTION (REF. FIG. 7-1)

The KLN 900 GPS Navigation Syslem processes signals from a maximum of elght satellites to give the pilot present pos%ion information and to display guidance information with respect to a flight plan defined by the pilot. The pilot uses the controls on the front panel of the GPS unit to enter the flight plan.

The KLN 900 GPS Navigation System calculates crosstrack error, distance-to-waypoint, groundspeed, track angle, time to waypoint, bearing to waypoint and vertical navigation data. The memory of the GPS holds the present position, the pilot-specified waypoints and up to 26 diferent flight plans with up to 30 waypoints each.

The GPS Navigation System comprises these components:

one KLN 900 GPS navigation unit

one KA 92 GPS antenna

one APR - ARMlACT switch

a one CRS OBSLEG switch

The Computer Aided Test (CAT) connector on the aircraft can be used as a database loader. This is an alernative method to the practice of replacing the database card in the front GPS unit.

The GPS antenna is located on the top of the aircraft, above the totward door.

A GPS navigation unit is located in the aft center pedestal.

An APR - ARMIACT switch and a CRS OBSllRG switch are located on the instrument panel.

Issued: 22 December, 1999 Report No: 01973-001 9-19-15


Report No: 01973-001 9-19-16

Figure 7-1. GPS Schematic

Issued: 22 December, 1999 16 of 21


I


Figure 7-2. GPS Controls

17 of 21 Report No: 01973-001

9-19-17


GPS DISPLAY CONTROL SWITCHES (REF. FIG. 7-2)

The GPS display is a Cathode Ray Tube (CRT). The display screen is divided into a maximum of five segments, the number of segments depending on the mode of operation selected. Most operations show a five-segment display, with two of the segments larger than the other three. These are used as the left page data display and the right page data display. The lefl and right page data controls are used to select and change data on these pages. The lefl and right page data controls have an inner knob and an outer knob.

The outer knob:

sets differen1 pages when the cursor is off

moves the cursor on the display when the cursor is on

The inner knob:

8 selects the next page of the page group when the cursor is off

8 changes the data below the cursor when the cursor is on

1. Left Page Data Control

r NAV (navigation)

FPL (flight plan)

MODE (mode)

TRIP (trip planning)

CALC (flight calculator)

STAT (status)

SETUP (set-up)

OTHER

2. Right Page Data Control:

NAV (navigation)



ACTV (active waypoint)

a REF (reference waypoint)

CTR (center waypoint)

APT (airport waypoint)

VOR (VOR waypoint)

NDB (non-directional beacon waypoint)

a INT (intersection page)

a PULL SCAN: when the right page data control inner knob is pulled out, it allows the pilot to view the data in the database

The functions of the remaining switches are:

(a) PUSH ON the GPS navigation unit is switched on when the control is pushed in and switched off when it is pulled out.

(b) BRT the light intensity of the display changes when the control is rotated. The panel Iihting is controlled from the COCKPIT LIGHTS - SIDE PNL controls.

4. Left and Right CRSR (Cursor)

Switches the cursor on or off. The cursor is a block of light and it is moved on the display with the left or right page data controls.

5. Direct-To (Arrowed D)

Shows the navigation data from the aircraft's present position to the selected waypoint on the display panel.



6. CLR (Clear)

Removes the data shown on the display panel.

7. ENT (Enter)

The data under the cursor is put in the computer memory.

8. OBS

When the OBS navigation mode is used, each push of the switch causes the mode to allernate beiween OBS and LEG.

9. ALT (Altitude)

Shows the altitude page and data on the altitude alert mode.

10. NRST (Nearest)

Shows a list of the nearest nine airports.

1 1. MSG (Message)

When there is a situation which requires attention, the MSG prompt flashes at the bottom right of the display. The MSG button is then pressed to show the message page.

12. SAVE

Used to create a user waypoint at the present position.

13. Card Ejection Button

The GPS navigation unit is turned off and this button is pressed to eject the database card from the front of the unit.

APR - ARMIACT SWITCH

The APR - ARMIACT switch is used to arm, disarm or deactivate the approach mode.

CRS - OBSILEO SWITCH

The CRS - OBSILEG switch is used to select either OBS or LEG mode.

Report NO: 01 973-001 9-19-20

Issued: 22 December, 1999 20of 21


ALTITUDE ALERT FUNCTION

The aircraft is wired tor the altitude alert function, which can be set on or off when configuring the GPS navigation system options.

When the altitude alert option is used, the GPS navigation system receives altitude information from the air data computer and, when necessary, sends an altitude alert warning signal to the audio integrating system. The alliiude alert warning signal is connected to the unmuted input of the audio control panel. The warning tone frequency is 1 kHz. The tone volume is adjusted on the SET 9 page. The altitude alerts include:

3 short tones: the aircraft is within 1000 11 of the selected alliiude

2 short tones: the aircraft has reached the selected altitude

- 2 short tones - pause - 2 short tones: the aircraft has deviated from the selected altitude by more than 300 ft

1 short tone - 1 long tone - 1 short tone: height above airport alert

DATA LOADING

The GPS will not navigate during data loading, therefore it is necessary to load the data base before the flight. The published database can be updated by either of the following methods:

using a personal computer (connected to the CAT connector)

replacing the database card in the GPS navigation unit, which is accessible through the front panel of the unit

PUBLISHED DATA BASE

The published data base is an electronic memory containing information on airports, navaids. intersections, SID's, STAR'S, instrument approaches, special use airspace, and other items of value to the pilot.

Every 28 days, AlliedSiinal receives new data base information from Jeppesen Sanderson. This information is processed and downloaded onto the data base cards. Atliedsignal makes these data base card updates available to KLN 900 GPS navigation system users.






SECOND PITOT STATIC SYSTEM AND

COPILOT FLIGHT INSTRUMENTS

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane FligM Manual when a second pitot static system and copilot flight instruments are installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in Me basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and perlormance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Right Manual.

Federal Office for Civil Aviation (FOCA) of Switzerland Section lor Type Certification

Date of Approval:

1 a Aug. 2000 L

Issued: January 17,2000 Revision 1 : April 13,2000

I

Report No: 01 973-001 9-20- 1


This supplement supplies the information necessary for the operation of the airplane when the

I second pitot static system and copilot flight instruments are installed in accordance with !

Modification Approval Sheets PIL 12/30/085 and 12/23/095.

I t

SECTION 1 -GENERAL

There are two configurations of the copilot flight instrument option. The configurations cover the requirements of the PC-12 and PC-12/45 aircraft. Confiiurations 1 and 2 are applicable I

the PC- 12/45 and Configurations l a and 2a are applicable to the PC-12. The only difference between the PC-12 and PC-12/45 instrument installation are the speed limitation zonal markings on the airspeed indicator. 1


3.19.6 PITOT AND STATIC PROBE DEICE FAILURE IN ICING CONDITIONS

Indlcatlon: P-S FAlL 1 or P-S FAlL 2 annunciator on or PROBES DP ICE CAWS ADVISORY does rxrt come on with operation of probes heating. l

A PlTOT AND STATIC DEICE FAILURE IN ICING CONDITIONS CAN CAUSE AN INCORRECT INDICATION ON THE AS1 AND/OR ALTIMETER AND VSI.

1. PROBES switch Cycle from OFF to ON

2. PlTOT DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

3. STATIC DE ICE Check. Do not reset unless tripped circuit breaker (Generator 2 busbar)

4. COPILOT PITOT DE ICE Check. Do not reset unless tripped circuit breaker (Generator 1 busbar)

5. COPILOT STATIC DE ICE Check. Do not reset unless tripped I circuit breaker (Generator 1 busbar)

6. PROBES circuit breaker Check. Do not reset unless tripped (Battery busbar)

Report No: 01973-001 9-20-2

Issued: January 17.2000 Revision 1 : April 13.2000




IF CAPTION P-S FAlL 1 OR PROBES DE ICE STAYS IN FAILURE STATUS:

8. Autopilot Disconnect

9. Aircraft DEPART ICING CONDITIONS to positive lOAT atmosphere, H possible



12. Landing approach Center EFlS AOA fast sbw pointer with PUSHER ICE MODE and flaps 15". Maintain speed above shaker activation. The total landing distance will be bnger by up to 55%.

3.20.5 PITOTISTATIC SYSTEM FAILURE

Indkatlon: P-S FAlL 1 annunciator on.

1 1. Airspeed indicator, altimeter Utilize cross panel instrument and vertical speed indicator monitoring to view the copilots panel

air data instruments


4.6 BEFORE TAXllNG

10. PROBES switch

4.8 BEFORE TAKEOFF

I 6A. Copilot flight instruments

4.12 CRUISE I

2A. Copilot directional gyro (if installed)

L

L


I

Set to ON and check green liiht PROBES DE ICE on amber light AOA DE ICE off P-S FAlL light off

CHECK and SET

SET HEADING Heading to be set to AHRS or magnetic compass every 15 minutes

Report No: 01973-001 9-20-3



No change.


Factory installed optional equipment is included in the licensed weight and balance data in '

Section 6 of the basic Pilot's Operating Handbook and FOCA Approved Airplane Fliht Manual. t

Report No: 01 973-001 9-20-4

Issued: January 17.2000 Revision 1 : April 13,2000



DUAL PlTOT STATIC SYSTEM

GENERAL

An optional second pitot static system can be installed (Ref. Figure 1).

DESCRIPTION

A second heated pitot head is installed under the left wing and a heated dual passage static port is installed on each side of the rear fuselage in place of the single system static ports. The existing pitot static system comprising the pitot head on the right wing and one passage from the left and right static ports supply pitot and static pressure to the pilot's instruments and is designated the No. 1 system. The second system designated No. 2 supplies static pressure to the coptiit's air speed indicator, altimeter and vertical speed indicator and pitot pressure to the copliot's ah speed indicator.

Drain traps for the No. 2 system are installed in the left bottom wing root for the pitot system and the rear fuselage behind the pressure bulkhead lor the static system.

The heaters for the pitot heads and static ports for the No. 1 and 2 systems are controlled by the DE ICING PROBES switch installed on the pilot's bwer right switch panel. Electrical power for the No. 1 system heating is supplied through the PlTOT DE ICE and STATIC DE ICE circuit breakers on the generator 2 bus. Electrical power for the No. 2 system heating is supplied through the COPILOT PlTOT DE ICE and COPILOT STATIC DE ICE circuit breakers on the generator 1 bus.

A P-S FAlL 1 and 2 annunciator is installed on the center instrument panel and is connected to the pitot and static heater current load sensors of No.1 and 2 systems. The individual lamps in the annunciator can be checked by pressing the LAMP switch on the ELECTRIC TEST panel installed on the rear left cockpit sidewall.

When electrical power is applied to the aircraft with the DE ICE PROBES switch OFF the amber P-S FAlL 1 and 2 annunciator will come on. When the DE ICE PROBES switch is set to ON the PROBES DE ICE advisory caption will come on and the P-S FAlL 1 and 2 annunciators will go off. If any of the pitot head heaters or static port heaters fail, the amber P- S FAlL and a 1 or a 2 annunciator will come. The '1" caption will come on ii the right side pilot pitot heater or static port heater has failed. The "2" caption will come on if the left side copilot pitot heater or the static port heater has failed.

issued: January 17,2000 Revision 1 : April 13, 2000


2OPILOT INSTRUMENT INSTALLATION

The optional copilot instrument installation configuration 1 (Ref. Figure 2) or confiration 2 Ref. Figure 3) can be installed with the second pitot static system. Configurations 1 and 2 are ipplicable the PC- 12/45 and Configurations l a and 2a are applicable to the PC-12. They only iitfer in the zonal markings for operating speed limitations on the airspeed indicator.

:onfiguration 1 and l a comprise:

1 Second pitot static system Airspeed indicator Vertical speed indicator

1 Altimeter Master Warning/Caution switches

Configuration 2 and 2a comprise:

r Second pitot static system Airspeed indicator Vertical speed indicator Altimeter Directional gyro Attitude indicator Master Warning/Caution switches

All the copilot instruments are illuminated with 28 v or 5 v power supply from the lighting control unit. The lighting intensity is controlled by a copilot dim control switch on the Center console.

DESCRIPTION

AIRSPEED INDICATOR

The copilots Airspeed Indicator (ASI) is the same type as installed in the pilots instrument panel but without the V, and M, switches. The copilots AS1 is connected to the second pitot static system.

The PC-12 and PC-12/45 are equipped with the same basic AS1 type but ditfer in their zonal markings for operating speed limitations.

VERTICAL SPEED INDICATOR

Report No: 01973-001 9-20-6



The copilots Vertical Speed Indicator (VSI) is the same type as installed in the pilots instrument panel and is connected to the second static system.

ALTIMETER

The copilots altimeter Is a non-encoding type and k connected to the second static system.

DIRECTIONAL GYRO

The Directional Gyro (DO) provides the copilot with an unstabilized source of aircraft heading information. The DG is a self-contained vertical axis gyroscope, mounted to show aircraft heading. The rotor is erected when the knob is depressed, as this is an unstabilized instrument the displayed heading will drift over time and needs to be set in reference to the AHRS or magnetic compass every 15 minutes.

The DG is powered via a 2 Amp COPILOT DG circuit breaker connected to the Avionic 1 Bus. The DG operation is continuous whenever the Avionic 1 Bus is powered. An OFF flag will appear when power is removed from the DG.

ATTITUDE INDICATOR

The Attitude Indicator (Al) provides the copilot with a source of pitch and roll information. The Al is self contained, vertical axis gyroscope, mounted in a pitch gimbal, which is mounted in a roll gimbal. The Al is capable of operation through 360" of aircraft pitch and roll displacement. The Al incorporates stops on the inner (pitch) axis to limit pitch axis freedom to +I- 85" relative to the roll gimbal. The pitch range of 360" is accomplished using controlled procession when the pitch stops are contacted. There is an inclinometer mounted on the face of the Al to present aircraft slip and skld.

The Al is powered via a 2 Amp COPILOT ATT IND circuit breaker connected to the Avionic 2 Bus. The Al operation k continuous whenever the Avionic 2 Bus is powered. An OFF flag will appear when power is removed from the Al.

MASTER WARNINWCAUTION SWITCHES

Master waming and caution switches are installed at the top of the copilots instrument panel. They are connected in parallel to the pilots master waming and caution switches and alert the crew to any changes on the CAWS indication panel. Pushing the applicable master warning or caution switch on either the pilots or copilots instrument panel will cancel the indication. The CAWS warning or caution annunciator that triggered the master warning or caution indication will remain illuminated. Additionally an audible tone will sound through the cockpit overhead speaker and or crew headsets anytime a master waming or caution indicator comes on. The individual lamps in the of the master warning and caution indicators can be checked by pressing the Lamp Test switch on the system test panel located on the left cockpit sidewall.

~ssued: January 17,2000 Revision 1 : April 13, 2000

Report No: 01973-001 9-20-7

SECTIO

N 9

SUPPLEM

ENT 20


9-20-8

Figure 1 D

ual Pitot Static System

Issued: January 17,2000 R

evision 1 : April 13.2000


Figure 2 Copilot Instrument Panel Configurations 1 and l a


Report No: 01973-001 9-20-9


PlLOrS LOWER RIGHT PANEL r. l

I DE ICINQ PROBES SWITCH

ATTITUDE MCATOR

AIRSPEED WDICATOR

Reporl No: 01973-001 9-20- 10

Figure 3 Copilot Instrument Panel Configurations 2 and 2a



PILOTS OPERATING HANDBOOK AND



MSN 294 SPECIFIC EQUIPMENT

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when a VNHF Com System, Mobile Phone system and a second copilot Audio Control panel is installed in accordance with Pilatus Project PIL 12/OQIO88. The information contained herein supplements or supersedes the Information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information no1 contained in this supplemenl, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.


Date of Approval:

0 3, Hfn 2000

I Issued: February 28,2000 Report No: 01 973-001

9-21 - 1


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when a VIUHF Com System. Mobile Phone system and a second copilot Audio Control panel is installed in accordance wilh Piatus Project PIL 12/00/088.


Transmitters COM 1, COM 2 and COM 3 may be used by both the pilots and copilot's mike selector. A coordination procedure including COM 4 operation must be established.

The C-5000 Communication Management Controller Operatofs Manual No. 150-1352-000, Rev C or later applicable revision must be readily accessible to the flighl crew when the Wulfsberg Flexcorn II VlUHF transceiver system is being used.

The Wullsberg Flexcom II VIUHF system must not be used in the frequency range 118.000 - 137.000 MHz during IFR operations (emergency: loss of VHF COM, is exempted)

Placard adjacent to,COM 4

TRANSMITTING ON COM 4 WITHIN THE VHF COM FREQUENCY BAND

MAY CAUSE INTERFERENCE IN THE


Dual COM failure on COM 1 and COM 2: COM 4, Wulfsberg Flexcom II VIUHF may be used.


Second Audio system: No change to the basic POH

For operations of VIUHF Wulfsberg Flexcom II: Refer to C-5000 Communication Management Controller Operator's Manual

Mobile Phone systemt: Refer to ACM 2000 Operatoh Manual


No change.

Repon No: 01973-001 9-21-2

Issued: February 28.2000

r PILATUS W PC12



Factory inslalled optional equipment is included in the licensed'weight and balance dala in Section 6 of the basic Pilot's Operating Handbook and FOCA Approved Airplane Flighl Manual.

SECTION 7 - SYSTEM DESCRlPnON

GENERAL AVIONICS INSTALLATION

Customer specific communication equipment consists of the following systems:

Second Audio system KMA 24H-70 at copilot's slation COM 4 Wulfsberg Flexcom I1 VJUHF at copilot's station MobUe Communicallon system ACM 2000 in the cockpit and cabin

SECOND AUDIO CONTROL PANEL

A second copilot Audio Control Panel Bendix King KMA 24K70 is installed in the copilot's inslrument panel. The operaling mode of the copilot's audio panel is the same as that of the pilol's. Each audio panel has access to the three intercom stations in the cabin and the cockpil inlercom line tor the pilot and copilot. A summing amplifier is installed to match certain warning audio signals lor the lwo audio panels. Power supplies to the copilot audio syslem are through two AUDIO COPILOT circuit breakers, one on the BATTERY BUS (pilol's side) and one on the AVIONIC 2 BUS (copilot's side) circuit breaker panels.

VIUHF COMMUNICATION SYSTEM

DESCRIPTION

The V/UHF system COM 4 Global Wulfsberg C-5000 Flexcom 2 is a stand alone multichannel AMIFM voice communicalion syslem, power is supplied through two circuit breakers:

MISSION RADIO TWRX on the AVIONIC 1 BUS (pilot's side) circuit breaker panel MISSION RADIWCTRL on the AVIONIC 1 BUS (pilot's side) circuil breaker panel

An automatic self lest of the system is done each lime the system is switched on. The syslem is self monitored for faults duripg operation. The system can be operated in frequency mode or preset channel mode.

The system comprises a control panel installed in the bwer right copilot's inslrument panel, a transceiver inslalled below the passenger compartment floor and an anlenna inslalled in the center of the bottom fuselage.

Issued: February 28,2000 3 of 4

Report No: 01 973-001 9-2 1-3


I CAUTION I Transmilling on COM 4 within Ihe VHF COM Frequency band may cause inlerference in the VHF COM 2 transceiver.

MOBILE PHONE SYSTEM

DESCRIPTION

The AirCell ACM 2000 AirlGround Phone System comprises handsels, a transceiver, and an anlenna. One handset is installed behind the pilot on the lelt cockpit divider. It is in reach ol Ihe copilot. The olher handset is inslalled in the center ol Ihe left cabin sidewall. The handsets provide voice communications, control and programming of the system. The handsel receives information to and lrom the transceiver. Handset cradJes provide a secure storage lor h e handsels when not in use. The transceiver is installed below the passenger compartment floor. The receiver uses two separate internal transceivers. One is used for operation on the ground and the olher is used lor operation while in the air. The antenna is lnslalled on the bottom of Ihe luselage in line with the crewlpassenger door.

Power supply to the system is through the MOBILE PHONE circuit breaker on the AVIONIC 2 BUS (copilot's side) circuit breaker panel.

1

Report No: 01873-001 9-21-4

Issued: February 28,2000


PlLors OPERATING HANDBOOK AND



EMERGENCY POWER SYSTEM

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Alrplane Flight Manual when the emergency power system is Installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the bask Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office for Civil Aviation (FOCA) of Switzerland S

Issued: September 1,2000 Revision 1: May 9, 2001

Report No: 01973-001 9-22-1


SECTION 1 - GENERAL

This supplement supplies the information necessaly for the operation of the airplane when the emergency power system is installed in accordance with Modification Approval Sheet PIL

I 12/24/093 or at aircraft build on MSN 321 and 401 & UP.


No change.


3.0 EMERGENCY PROCEDURES

The following step is added to all the Emergency Procedures where the electrical System

I gangbar (MSN 101-320 and 322-400) or MASTER POWER switch (MSN 321 and 4Q18 UP) is operated:

EMER PWR SYSTEM switch OFF


3.15.1 The standby attitude indicator is operative.


4.4 BEFORE STARTING ENGINE

4.4.1 PROCEDURE

I 1. EPS switch TEST (minimum 5 seconds) I

2. BAT TEST indicator Check green LED comes on and stays on during test

I 3. EPS switch

4. EPS ON indicator

I 5. BAT switch BAT 1 and 2 switches

6. EPS ON indicator

4.18 SHUTDOWN

I 1. EPS switch

Report No: 01973-001 9-22-2

ARMED

Check amber LED comes on

(if 2"d battery installed)

Check amber LED goes off

OFF

Issued: September 1,2000 Revision 1: May 9, 2001



No change.




ELECTRICAL

GENERAL

Emergency Power system option: a lead-acid battery 24 V 5 Ah

DESCRIPTION

POWER SUPPLIES (MSN 101-320 and 322-400) I The Emergency Power System (EPS) uses a 24 V 5 Ah lead-acid battery to supply power to the EPS bus. The COMMMAV 1 and the second Attitude Indicator are connected to the EPS bus. During normal operation and when the EPS is armed, a relay connects the EPS bus to the battery bus through the EPS. If the battery bus fails or there is a total power loss, the EPS battery will continue to provide power through the EPS bus to the COMMlNAV 1 and the second Attitude Indicator. Power is also supplied to the lighting circuits of the CDI and the second Attitude Indicator.

POWER SUPPLIES (MSN 321 and 401 & UP) I The Emergency Power System (EPS) uses a 24 V 5 Ah lead-acid battery to supply power to the second Attitude Indicator. During normal operation and when the EPS is armed, a relay connects the SECOND ATT IND circuit breaker to the battery bus through the EPS. If the battery bus fails or there is a total power loss, the EPS battery will continue to provide power lo the second Attitude Indicator. Power is also supplied to the lighting circuit of the second Attitude Indicator.

I CONTROLS AND INDICATORS

The Emergency Power System (EPS) is controlled from a switch on the Overhead Electrical Power Management Panel (MSN 101-320 and 322-400) or the pilot's lower right switch panel (MSN 321 and 401 & UP). The switch EMER PWR SYSTEM has the positions ARMED, OFF and TEST. There are two indicators EPS ON (amber) and BAT TEST (green) adjacent to the

I Issued: September 1, 2000 Revision 1 : May 9, 2001 3 of 4

Report No: 01 973-001 9-22-3


switch. With the switch in the TEST position (minimum 5 seconds) the BAT TEST indicator will come on to show that the system is serviceable. The switch is put in the ARMED position before flight and if there is a total power loss, the EPS ON indicator will come on to show that the system is supplying power.

Issued: September 1,2000 Revision 1 : May 9,2001

- --

Report No: 01 973-001 9-22-4





MuLn FUNCTION DISPLAY (KMD 850)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when the KMD 850 Multi-function display is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:

Federal Office for Civil Aviation (FOCA) f Switzerland ~ e c i i o w c @ i f i i R

Date Cr of App vat:

17. AFR. 2001

Issued: November 14,2000 Revision 1 : April 2,2001

Reporl No: 01973-001 9-23-1


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the multi function display (KMD 850) is installed in accordance with Modification Approval Sheet PIL 12/34/109.


No change.


No change


No change


No change.

SECTION 6 - WEIQHT AND BALANCE


, Report NO: 01973-001 1 9-23-2

Issued: November 14,2000 Revision: 1 : April 2, 2001



MULTI-FUNCTION DISPLAY (KMD 850)

Ref. Fig 7.1

The KMD 850 Multi-function display (MFD) is an independent Liquid crystal display (LCD) which shows:

A moving map display

Weather radar data (if installed)

Terrain data (il EGPWS or TAWS is installed)

Traffic data (if TCAS is installed). I The MFD is installed in the center instrument panel and contains the controls and indicators necessary to operate the MFD in each of its operating modes. The MFD has the following controls and display:

1. Display area

The display area operates in the mode set by the function select keys. The functions that can be set on the display area are:

Map function - When the start up sequence is complete, the MFD shows a "TOP0 ON" MAP page, if the MFD has a valid GPS input the map corresponds to the position of the aircraft

Weather Radar function - When the WX tunction key is pushed the weather radar page is shown

Auxiliary function -When the AUX function key is pushed, the MFD shows an auxiliary code cover page. This provides access to the MFD set up pages.

Terrain function - (if EGPWS is installed) When the TERR function key is pushed, the MFD shows the terrain page for the EGPWS I Traffic function - (if TCAS is installed) When the TRFC function key is pushed, the MFD shows the traffic page for TCAS.

Issued: November 14,2000 Revision 1 : April 2,2001

Report No: 01 973-001 9-23-3


2. Joystick

The joystick moves a pointer around the LCD display. The joystick is used to point at items on the map page for further information and for measuring range and bearing to specific points.

When the WX radar function is selected, the joystick controls the I angle and track line of the antenna. On the AUX setup pages the joystick is used to select and change the settings for a given set up field.

3. Control keys

The control keys are used to manipulate the page that is currently displayed. Either a soft label on the LCD on the left side of the key or the control key annunciator on the right side of the control key indicates the functionality of each key.

The control key functions are MODE. RNG UP, RNG DOWN, VlEW and OVLY.

MODE -This key makes the display sequence through all available modes associated with the displayed page

RNG UP - This key increases the range scale up one level on the displayed page. The default range on start up is 80 nm.

RNG DOWN -This key decreases the range scale down one level on the displayed page.

VlEW - This key makes the display sequence through the available views associated with the displayed page.

OVLY -This key allows data from more than one source to be displayed simultaneously on the display screen.

The MFD has a fault indicator located between the RNG UP and RNG DOWN keys. When there is a hardware fault detected, the letter "F" in a circle appears.

4. Inner and Outer control knobs

The inner and outer control knobs have various functions as indicated by the relevant soft label when active. If the weather radar page is selected the inner knob controls the gain of the weather radar in the ground mapping mode. The outer knob acts as the weather radar function selector for the standby, test and on funCti0ns.

5. Selected function indicators

Report NO: 01973-001 1 9-23-4

Issued: November 14.2000 4 of 6 Revision: 1 : April 2. 2001


The selected function indicators come on when the applicable function key is pushed to set a particular function of the display.

6. Function select keys

The function select keys select the available data sources to be displayed on the display area. The function select keys have the following captions:

MAP

WX (Weather Radar H installed)

TRFC (TCAS il installed)

TERR (Terratn if EGPWS or TAWS is installed)

AUX (Set-up pages)

When a function select key is pushed, the indicator above the key is illuminated to show that the function has been selected for display. If the key is pressed multiple times in sequence, the available pages associated with the particular function are shown on the display area.

7. OFFION control

The OFFION control sets the MFD to on or off.

8. Regional map data card

The regional map data card is a front panel loading PCMCIA flashdisk. The card contains application software and a customized Jeppesen aeronautical database. Three regions are covered on different databases, the Atlantic, Americas and Pacific. The appropriate database card for your area must be installed before switching on the MFD. The database contains data VORs, NDBs, intersections and special use airspace. The data card also has a topographical database which includes rivers, roads, lakes, coastlines, cities, rail tracks, and towers.

9. Brightness control

The brightness control changes the intensity of the display.

Detailed operating instructions or technical information for the KMD 850 Multi-Function Display may be obtained by consulting the latest available revision of the pilots guide Pt No 006-18222- 0000.

Issued: November 14,2000 Revision 1 : April 2.2001

Report No: 01973-001 9-23-5 (


Report No: 01973-001 I 9-23-6

Figure 7-1 Multifunction Display (KMD 850)

Issued: November 14,2000 6 of 6 Revision: 1 : April 2,2001





ENHANCED GROUND PROXIMITY WARNING SYSTEM (HONEYWELL MK-VI EGPWS)

This supplement must be attached to the Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual when a MK-VI EGPWS is installed in accordance with Pilatus Drawings. The information contained herein supplements or supersedes the information in the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual only in those areas listed. For limitations, procedures and performance information not contained in this supplement, consult the basic Pilot's Operating Handbook and FOCA Approved Airplane Flight Manual.

Approved by:


Date of ~ ~ ~ r o v i l :

Issued: June, 22 2001

- -

Report NO: 01973-001 9-24-1


SECTION 1 - GENERAL

This supplement supplies the information necessary for the operation of the airplane when the MK VI EGPWS is installed in accordance with Modification Approval Sheet PI1 121341103.


- SWITCH LEGEND

GND PROX WARN

GPWS INOP

TERR INOP

G/S P/CNCL

BELOW WS

GPWS PiTEST

GIs CNCLD

The MK VI EGPWS Pilot's Guide (060-4314-000 Initial Issue dated April 2000 or later applicable revision) must be readily accessible to the flight crew when operating the MK V1 EGPWS. Navigation using the terrain database is not permitted.

FUNCTION

Ground Proximity Warning

Ground Proximity Warning System inoperative

Terrain Function Inoperative

Glideslope press to cancel

Below Glideslope

Ground Proximity Warning System press to test

Glideslope cancelled

I


SECTION 3A - ABNORMAL PROCEDURES

3A.1 GENERAL

3A.l.l Press the GPWS FLAP OVRD switch to prevent operation of the "TOO LOW FLAPS" alert when you need to land with less than full flaps. Flap override is automatically deselected below 50 feet AGL.

3A.1.2 Press the GIS PlCNCL switch to deactivate the "BELOW GIS" alert when you intentionally plan to go below the glideslope. A G/S CNCLD amber caption will illuminate, it will automatically reset after landing or if the aircraft climbs above 2000 feet, or a non ILS frequency is selected.

Report No: 01973-001 9-24-2


-- - -

IP ILATUSF SECTION 9

PC12 SUPPLEMENT 24

3A.1.3 Press the TERR INHIBIT switch to prevent nuisance or unwanted warnings when operating at an airport not in the terrain database. Terrain inhibit requires manual deactivation.

3A.1.4 Press the GPWS STEEP APP switch if steep approaches greater than 4" and within the operational limitations of the aircraft are performed to eliminate unwanted 'SINKRATE' and 'PULL UP' aural warnings. GPWS steep approach requires manual deactivation

3A.1.5 Flight outside the installed database region will initiate an amber TERR INOP caution, all enhanced functions dependant on the database are inoperative. The terrain display will display Magenta, depicting unknown terrain. All GPWS functions remain active.

3A.1.6 When the GPWS INOP amber caption illuminates to indicate a system failure, the failure may be identified by the following method:

(a) On ground, press and hold the GPWS test switch to perform a System Self Test. The test will prove if the fault is external to the EGPWS (with an aural fault message) or is internal.

If the fault cannot be cleared, pull the EGPWS circuit breaker.

The alrcraft must not be flown until the pilot has verified that operational requirements are met with the systems available.

(b) In flight, if the fault cannot be cleared, pull the EGPWS circuit breaker.

Continue flight - Report on landing.

3A.1.7 The EGPWS can be deactivated by pulling the EGPWS circuit breaker, located on the AVIONIC 2 circuit breaker panel, when a system failure cannot be cleared.


Perform a Self Test of the EGPWS prior to flight.


No change.



Issued: June, 22 2001 Report No: 01 973-001 9-24-3



GENERAL

The Enhanced GPWS Mk VI provides an enhanced capability of reducing accidents caused by controlled flight into terrain. The system achieves this by receiving a variety of aircraft parameters as inputs, then applying alerting algorithms to provide the flight crew with aural messages and visual annunciation and display. The EGPWS provides the flight crew with enhanced terrain awareness while tollowing an ATC flight plan clearance. The terrain and obstacle display is shown on the KMD 850 Multi-function display.

The EGPWS has an integral GPS receiver for accurate position determination and a regional terrain database of geographical topography and physical obstructions. There are three regional terrain databases available which cover the Americas, the Atlantic and the Pacific regions. Only one regional terrain database can be installed at a given time and this is indicated by the version number and letter designation during a system self test.

The installed regional database limits the availability of the Enhanced functionality to operations within the installed region, operation outside of the installed regional database does not affect the basic GPWS functionality. If the incorrect region is installed lor the current location of the aircraft, the Enhanced functions are inoperative, this is indicated by a TEAR INOP annunciation and Outside database region indicated during a self test.

The Mk VI EGPWS consists of two GPWS test switches, nine GPWS indicator lampslswitches, a GPS antenna and. an EGPWS computer.

EGPWS TEST SWITCH

A test switch for the EGPWS system is installed on the left and the right instrument panel. It has the legend 'GPWS PITEST' in illuminated white letters on a black background.

EGPWS INDICATOR LAMPSISWITCHES

There are four indicator lampslswitches for the EGPWS installed on the left and the right instrument panels and three indicator lamps/switches are installed on the left crossbar panel:

GND PROX WARN (LEFT AND RIGHT INSTRUMENT PANELS)

The GND PROX WARN indicator consists of black letters on a black background when not illuminated. It comes on with red illuminated letters on a black background when the EGPWS computer detects an alert or warning in Modes: 1 (Excessive Descent Rate), 2 (Excessive Closure Rate to Terrain), 3 (Altitude Loss After Takeoff) and 4 (Insufficient Terrain Clearance).

Report No: 01973-001 1 9-24-4


SPILATUSC SECTION 9

PC12 SUPPLEMENT 24

GPWS INOPKERR INOP (LEFT AND RIGHT INSTRUMENT PANELS)

This indicator is a split screen annunciator, The upper GPWS INOP consists of black letters on an black background. The text comes on in amber when the EGPWS detects a failure of the system. The lower TERR INOP consists of black letters on a black background. The text comes on in amber when the EGPWS Terrain mode has a disabled function.

GIS PICNCL I BELOW GIS (LEFT AND RIGHT INSTRUMENT PANELS)

This momentary indicator switch has a split screen annunciator. The upper GIs PICNCL consists of illuminated white letters on a black background. The lower BELOW GIS caption consists of black letters on a black background. The text comes on in amber on a black background when the EGPWS computer detects a Mode 5 (Descent Below Glideslope) alert situation.

GPWS PfrEST I GIs CNCLD (LEFT AND RIGHT INSTRUMENT PANELS)

This is a momentary action switch with a split screen annunciator. The upper GPWS PlTEST legend consists of illuminated white letters on a black background. The lower GIs CNCLD consists of black letters on a black background. The text comes on in amber on a black background when a BELOW GIS caution has been cancelled using the GIS PCNCL switch. This shows that the EGPWS mode 5 function has been inhibited. The legend will stay illuminated until the EGPWS reverts to ground mode, or the aircraft climbs above 2000 feet, or a non ILS frequency is selected.

TERR INHIBIT (LEFT CROSSBAR PANEL)

This indicator switch consists of illuminated white letters on a black background when inoperative. When operated the switch comes on with blue letters on a black background. When operated the switch will inhibit the aural and terrain warnings associated with the enhanced features of the EGPWS, allowing the aircraft to operate without nuisance or unwanted warnings. This function is particularly useful when operating at airfields that are not in the terrain database. The terrain awareness display remains fully operational.

GPWS STEEP APR (LEFT CROSSBAR PANEL)

This indicator switch consists of illuminated white letters on a black background when inoperative. The switch is covered by a switch guard. The caption comes on with blue letters on a black background when the switch is selected to enable a steep approach greater than 4" and within the operational limitations of the aircraft without nuisance aural alerts. GPWS steep approach requires manual deactivation.

GPWS FLAP OVRD (LEFT CROSSBAR PANEL)

This momentary indicator switch consists of illuminated white letters on a black background when inoperative. The switch is covered by a switch guard. The caption comes on with blue letters on a black background when the switch is pressed and released. The switch is selected to enable late deployment of landing flaps without

Issued: June. 22 2001 Report NO: 01 973-001 9-24-5


receiving nuisance alerts (Too Low Flaps). GPWS flap overide is automatically deselected below 50 feet AGL. 1

EGPWS COMPUTER I

The EGPWS computer is installed below the passenger cabin floor. It uses the following system inputs for its computations:

- AHRS roll attitude I

- radar altimeter altitude AGL - air data computer vertical speed and airspeed - VHF NAVl system ILSfglideslope - EFlS 1 symbol generator decision height (DH) - flaps system flaps position - landing gear system landing gear position - KM0850 multi-function display range data - Autopilot autopilot engaged

4

The EGPWS computer sends unmuted voice messages, when necessary, to the flight compartment headphones and loudspeakers. At the same time, it sends a suppression signal to the CAWS and TCAS (if installed) to inhibit voice messages from the CAWS and TCAS when the EGPWS is generating voice messages.

The (Modes 1 thru 6) EGPWS voice messages used in order of priority are:

"PULL-UP PULL U P "TERRAIN-TERRAIN" "OBSTACLE OBSTACLE" "CAUTION TERRAIN" "CAUTION OBSTACLE" "TERRAIN "MINIMUMS - MINIMUMS" "TOO LOW, GEAR" "TOO LOW, TERRAIN "Altitude Callouts" (see NOTE 1) "TOO LOW, FLAPS" "SINK RATE SlNK RATE" "DON'T SlNK DON'T SINK" "GLIDESLOPE "BANK ANGLE BANK ANGLE"

Report No: 01973-001 9-24-6


- -


NOTE

The following altitude callouts are enabled:

MINIMUMS MINIMUMS FIVE HUNDRED TWO HUNDRED ONE HUNDRED FIFTY FORTY THIRTY TWENTY TEN

The enhanced feature of the EGPWS is the ability to alert the crew to and provide a display of potential conflict with terrain. Terrain conflict alerts will initiate a specific aural message and illuminate the GND PROX WARN annunciator. The EGPWS keeps a synthetic image of local terrain in front of the aircraft for display on Ihe KMD 850 multi-function display.

Terrain is displayed as a variable density dot pattern in green, yellow or red. The pattern density and color being a function of how close the terrain or obstacle is, relative to the altitude of the aircraft.

There are two different background terrain awareness display modes Standard and Peaks.

In the Standard Mode, terrain data is shown using colors and shading patterns which correspond to the vertical displacement between the elevation of the terrain and the current altitude of the aircraft. Red and Yellow dot patterns indicate terrain near or above the current altitude of the aircraft. Solid Red and Yellow colors indicate warning and caution areas relative to the flight path of the aircraft when an alert is active. Medium and low density green display patterns indicate terrain which is below the aircraft and within 2000 feet of the aircraft altitude. Terrain which Is more than 2000 feet below the aircraft is not displayed.

In the Peaks mode, additional density patterns and level thresholds are added to the Standard mode display levels and patterns. These additional levels are based on absolute terrain elevations relative to the range and distribution of the terrain in the display area. The Peaks mode display is a merged display applicable to all flight phases. At altitudes safely above all terrain for the chosen display range, the terrain is shown independent of aircraft altitude emphasizing the highest and lowest elevations. This gives greater situational awareness. The Peaks mode display includes a solid green level to indicate the highest non threatening terrain. The standard lower density green display patterns indicate mid and upper terrain in the display area as well as terrain that is within 2000 feet of the aircraft. Terrain identified as water (0 feet MSL) is displayed as a cyan color dot pattern. Additionally with the Peaks mode display, two elevation numbers indicating the highest and lowest terrain currently being displayed are overlaid on the display. The elevation numbers indicate terrain in hundreds of feet above sea level (MSL). The terrain elevation numbers are displayed with the highest terrain number on top and the lowest terrain number beneath it. The elevation numbers are unique to the Peaks mode.

The terrain alerting algorithms continuously compute the terrain clearance envelopes ahead of the aircraft. If the boundaries of these envelopes conflict with terrain elevation data in the terrain database, then alerts are issued. Two envelopes are computed, one corresponding to a terrain caution alert and the other to a terrain warning alert. Terrain awareness caution and warning

Issued: June, 22 2001 Report No: 01973-001 9-24-7

SECTION 9 EPILATUSE SUPPLEMENT 24 PC12

alerts are inhibited below 30 feet of radio altitude, within lnm of the runway or below 60 knots groundspeed.

When the required conditions have been met to generate a terrain or obstacle caution alert, the terrain image on the KMD 850 multi-function display is enhanced to highlight the threatening terrain as solid yellow for caution threats and the appropriate aural alert is given. When the required conditions have been met to generate a terrain or obstacle warning alert, the display image on the KMD 850 multi-function display is enhanced to highlight the terrain as solid red and the appropriate aural alerl is given.

Report No: 01 973-001 9-24-8


- - -

I

I PILOT'S OPERATING HANDBOOK

Safety & Operational Tips r

SECTION 10 SAFETY AND OPERATIONAL TlPS

Subject

GENERAL

SAFETY TlPS

SECTION 10

SAFETY AND OPERATIONAL TlPS

TABLE OF CONTENTS

OPERATIONAL TlPS

ANTI-COLLISION LIGHTS CROSSWIND OPERATION FLAMMABLE MATERIALS, PRESSURE VESSELS AND EQUIPMENT LOCATIONS REMOVAL OF SNOW, ICE AND FROST FROM THE AIRCRAFT OPERATIONS FROM PREPARED UNPAVED SURFACES PASSENGER BRIEFINGS

I Issued: February 14, 1994 Revision 10: September 1,2000

t

Page

Report No: 01 973-001 104

SECTION 10 SAFETY AND OPERATIONAL TIPS

Report No: 01 973-001 I 10-Y


Issued: February 14, 1994 Revision 1: June 10,1994

FPILATUSS SECTlON 10 ?c XI1 SAFE1 Y AND OPERATIONAL TIPS

GENERAL

This section provides information for the operation of the airplane

SAFETY TIPS

Pilots who fly above 10,000 feet should be aware of the need for physiological training. It is recommended that this training be taken before flying above 10.000 feet and receive refresher training every two or three years.

Inlormation on the location of flammable materials, pressure vessels and equipment locations for crash-firerescue purposes is given in Figure 1.

OPERATIONAL TIPS

ANTI-COLLISION LIGHTS I Anti-collision strobe lights should not be operating when flying through cloud, fog, or haze Reflected light can produce spatial disorientation.

CROSSWIND OPERATION

Takeoff

It is possible, if required, to hold the aircraft stationary with the brakes while the engine is at max takeoff power. When the brakes are released rapid and aggressive use of the rudder and possibly some small application of brake is necessary to establish and maintain the centerline but, once rolling, directional control is easy with rudder only. Holding the elevator neutral will keep the nosewheel on the ground and assist in maintaining directional control.

In strong crosswinds the aircraft establishes a drift angle of up to 10" while accelerating to rotation speed. I In gusty conditions it is recommended to rotate at V, + 10 Kls. On rotation the aircraft yaws considerably further into wind and automatically establishes the heading necessary to track the runway centerline.

Landing I It is recommended to use the wing down technique. At approximately 100 to 200 ft on approach to the runway. apply rudder to align the longitudinal axis of the aircrafl to the runway I Issued: February 14, 1994 Revision 6: Dec 6,1996

Report No: 01973-001 10-1


and put on bank In the opposite direction to maintain the runway centerline. The aircraft is then llown in a sidesl~p to touch down inillally on one wheel. As soon as one wheel touches, lower the other two to the runway and immediately select either the condition lever to ground idle or the PCL to beta or reverse. Once the aircraft is eslablished on the runway it can be stopped as normal with brakes or reverse power without dilliculty. Do not attempt heavy braking in a strong crosswind as the into wind wheel will tend to lock more easily.

In conditions of strong turbulence it is recommended, if runway length permits, to fly the approach with reduced llap dellection to increase IAS and aileron efficiency. It is also recommended to increase the approach speed for the chosen flap setting by 50% o l the difference between the wind mean speed and max gust speed, to give a greater speed margin over the stall.

FLAMMABLE MATERIALS, PRESSURE VESSELS AND EQUIPMENT LOCATIONS

Refer to Figure 10-1 for the location of these items

Report No: 01973-001 10-2

Issued: February 14, 1994 Revision 6: Dec 6.1996


DIRECT VIEW WINDOW PASSENGEWREW DOOR CARGO DOOR (OPENS INWARDS)

PROPELLER. DANGERAREA

ENGINE EXHAUST ENGINE OIL VERY HOT AREA 14 5 ORT GAS STRUT STRUT 1668 PSI GAS STRUT

SHUT-OFF LEVER a FUEUAIR SEP TANK FUEL FILTER

MIL-H.5606

TRANSMITTER

-=zz2 R134 REFRIGERANT EMERGENCY EXIT HAND FIRE PROPELLER COOLING UNIT

834 PSI - STATIC PORT MAIN WHEEL OXYGEN CYLINDER S! VERY HOT AREA 3 FUSIBLE PLUGS VERY HOT AREA I850 PSI i?

Figure 10-1. Flammable Materials. Pressure Vessels and Equipment Locations

Issued: February 14, 1994 Revision 6: Dec 6.1996

Report No: 0 1 9 7 3 - 0 0 1

10-3


REMOVAL OF SNOW, ICE AND FROST FROM THE AIRCRAFT

GENERAL

The aircraft must be clear of all deposits of snow, ice and frost adhering to the lifling and control surfaces immediately prior to takeoff. The clean aircraft concept is essential to safe ll~ght operations. The pilot in command of the aircraft has the ultimate responsibility to determine if the aircraft is clean and in a condition for safe llight. This may require the use of ground deicing techniques.

Manual methods of deicing provide a capability in clear weather to clean the aircraft to allow a safe takeoff and flight. Deicing lluids can be used to quickly remove frost and to assist in melting and removal of snow. In inclement, cold weather conditions, the only alternative is sometimes limited to placing the aircraft in a hangar to perform the cleaning process.

For a complete guide to deicing reference should be made to the Advisory Circular AC 135-17 "Pilot Guide. Small Aircraft Ground Deicing".

DEICING METHODS AND EQUIPMENT

Manual methods of deicing such as brooms, brushes, ropes, squeegees etc. can be used to remove dry snow accumulations and to remove the bulk ol large wet snow deposits. These manual methods require that caution be exercised to prevent damage to the aircraft skin or components.

Deicing fluids can t)e used to quickly remove frost and to prevent or retard ice formation overnight. They can also be used to assist in melting and removal of snow or other ice formations that may develop as a result of treezing rain and assisting in the removal of ice or lrost formations accumulated during a previous flight. Portable spray equipment such as pressurized containers with spray wands, hand pumps attached to a supply tank can be used lo apply the fluids. Fluids can also be applied by a mop or brush from a bucket to melt the ice to the extent that it can be removed by manual means. Mobile ground support equipment's ex~st with the capability of heating deicing tluid to high temperatures and dispensing large quantities of fluid at high pressures. This type of equipment with the fluid temperature limited to 160" F (70" C) can be used to deice the aircraft.

Report No: 01973-001 1 10-4

Issued: February 14, 1994 Revision 7: July 1,1997


DEICING FLUIDS I Various deicing fluids are commercially available. The pilot should understand the criticality of effective quality control in order to ascertain that the fluid supply conforms to the user need. The following deicing fluids are approved for use on the PC-12.

Common Primary Active Name I Ingredients

Traditional North

American

Ethylene. propy lene, diethy lene

glycols andlor isopropyl alcohol

AEA Type 1 SAE Type 1 (AMS 1424) IS0 Type 1

Propylene. andlor

diethylene glycol

Viscosity

Low

Low

Primary I Notes Use

Deicing

DEICING AND ANTI-ICING THE AIRCRAFT

Includes SAE AMS 1425. SAE AMS 1427, AF 3609. Mil-A-4823. other pre- 1993 Mil-Spec Fluids and other commercially available fluids

Deicing

The wings are the main lifting surfaces of the aircraft and must be free of snow and Ice to operate efficiently. Deicing of the wings should begin at the leading-edge wing tip. sweeping in the aft and inboard direction. If ice accumulation is present in the control surface cavities, i t may be necessary to spray lrom the trailing edge forward. Tail surfaces should be deiced in a similar manner to the wing. Move the horizontal stabilizer to nose down for a better visual check. The area adjacent to the elevator balance horns and the horizontal stabilizer should be thoroughly inspected. The propeller should be thoroughly deiced in the static mode makirlg sure all blades are uniformly clean. The fuselage should be deiced from the top down. Any deicing fluid must be removed from the cockpit windows to maintain optimal visibility. Passenger and cargo doors must also be deiced in order to ensure proper operation. All door hinges, locks and seals should be inspected to make sure that they are free of contamination. Under no circumstances should deicing fluid be applied to static ports, pilot head. AOA transmitters, cockpil windows, air intakes and the engine.

Propylene Glycol based fluids not lo be used undiluted at OAT less

than 14' F (-10" C)

DEICING THE ENGINE AREA

Minimal amounts of de~cing fluid should be used to deice the engine external area. Avoid the engine air inlet. Fluid res~due on engine compressor blades can reduce englne performance or cause stall or surge. In addition, this will reduce the possibility of glycol vapors entering the aircraft through the engine air bleed system Engine intake areas should be inspected for the presence of ice immediately after shutdown. Any accumulation should be removed while [he

Issued. February 14. 1994 Revision 7: July 1.1997

Report No: 01973-001 10-5


engine is still warm and before the installation of intake covers. A light coating of deicing fluid appl~ed to the intake covers w~l l assist in preventing the covers freezing to the nacelle.

HEALTH EFFECTS \

Pilots must be aware of the potential health elfecls of deicing and anli-icing fluids in order to ensure proper precautions are taken during deicing operations and to better ensure the well- being of passengers and crew.

POST DEICING/ANTI-ICING CHECKS

The follow~ng areas should be checked alter delcing operations have been carried out:

wing leading edges, upper and lower surfaces, aileron surfaces including the wing seals horizontal stabilizer leadlng edges, upper and lower surfaces elevator surfaces particularly at the balance horns vertical stabilizer and rudder side surfaces flaps propeller engine, 011 cooler and ECS air intakes. Inertial separator and screen fuselage static ports, pitot head. AOA transm~tters and temperature probes fuel tanks and vents landing gear.

A thorough walk-around prellight inspection is more important in temperature extremes because they may effect the alrcrall or its performance. Do not be tempted to hurry the prellight. This is the time to do a most thorough prellight inspection and to make sure that all protective covers been removed. ,

A pre-takeotl contamination walk round check should be made 5 minutes prior to beginning takeoff to make sure the wings and control surlaces are free of frost, snow and ice.

TakeoH should not be attempted unless the pilot ;n command has ascertained that all crltical surfaces of the aircraft are free of adherlng ice, snow or frost formations.

Report No: 0 1973-00 1 10-6

Issued: February 14. 1994 Revlsion 7: July 1.1997

=PILATUSZ SECTION 10 PC Xll SAFETY AND OPERATIONAL ~ P S

DEICINO WORKSHEET

Refer to the POH Section 10 for complete informatbn

-

b the temperature at or below 50' F (10" C ) where visible mdsture, or idng conditions encountered during descent, approech. or taxi-in of the previous flight?

Deice the akcreft It any frost. ice, slush or srow is presently adhering (or will adhere prior to takeoff) to the No wings. control surfaces. engine inkt t

Yes

Review airport advlsory page, and fluid required: - W i n g fluid available - deicing location and distance from runway *

*I Deice the aircraft I I

Will takeoff be accomplished within 5 minutes after the completion of the Post Deicing Check?

Do a Pre Takeoff Contamination Check

, X G + i i y e s contamination? J

Issued: February 14. 1994 Revision 7: July 1.1997

I

TAKEOFF I

Report No: 01 973-001 10-7

SECTION 10 mPILATUSIB6 SAFETY AND OPERATIONAL TIPS K12

OPERATIONS FROM PREPARED UNPAVED SURFACES

The aircraft is constructed for operations from prepared unpaved sudaces.

Prepared unpaved surfaces are taxi-ways and runways that are prepared and approved for aircraft operations with a surface other than tarmac or concrete.

PREPARED UNPAVED SURFACES SUITABLE FOR AIRCRAFT OPERATIONS VARY GREATLY AND SOME MAY NOT BE SUITABLE FOR OPERATIONS. IT IS THE RESPONSIBILITY OF THE PILOT IN COMMAND TO MAKE SURE THAT EACH TAXI-WAY AND RUNWAY SURFACE IS FIT FOR USE AT THE INTENDED AIRCRAFT WEIGHT BEFORE COMMENCING OPERATIONS ON IT.

The following factors should be considered when deciding if a surface is fit for operation or when operating from prepared unpaved surfaces:

SURFACEHARDNESS

A prepared unpaved surface may be hard after a period of dry weather but atter rain can become soft. The wheels of a heavy aircraft can sink into soft surfaces causing a large increase in drag. This can make taxiing difficult or impossible and increase the takeoff ground roll distance considerably, sometimes to the point where V, cannot be achieved. How deep the wheels sink in, varies with aircraft weight and surface condition. It may be possible to operate a light weight aircraft when it is not possible to operate it at maximum take off weight.

SURFACE ROUGHNESS

The taxi-way and runway surface should be smooth. Undulations, depression or bumps can cause longitudinal pitching of the aircraft which may cause a significant reduction in propeller ground clearance. Particular care should be exercised in lono orass which can conceal hard objects and depressions and also at the borders between gr& and concrete suifaces.

SURFACE TYPE

Loose stones or gravel can cause propeller or airframe damage. The propeller creates turbulence which lilts stones into the air which then are struck by following blades or are accelerated rearwards to hit the airframe. The risk of damage is reduced if the aircralt is allowed to accelerate forwards before high power is selected and if reverse thrust is not used below 30 kts forward speed.

Wet or fresh grass on a hard surface is slippery and has a lower coellicient of friction than short dry grass. Takeon and stopping distances may increase. On a son surface landing ground roll may decrease but takeoff ground roll may increase.

Report No: 01 973-001 10-8



On sandy or dusty surfaces, or where loose grass is present, reverse thrust can cause a loss of forward visibility and particles ingested into the air intake can cause increased engine wear.

INERTIAL SEPARATOR

When operating from any surface where there is a risk of dust, sand or other material entering the engine intake, it is recommended to open the inertial separator.

On takeoff from hot and high airfields with the inertial separator open it may not be possible to obtain maximum takeoff power (44 psi) and the takeoff performance will consequently deteriorate.

AIRCRAFT INSPECTION

When operating from prepared unpaved surfaces where there are loose stones, gravel, grit, sand, dust or cut grass etc. there is always a risk of propeller or airframe damage or blockage of air inlets. After operations from prepared unpaved surfaces, where a risk of damage or contamination exists, the aircraft should be thoroughly inspected.

BEFORE STARTING ENGINE

Make sure the area under and adjacent to the propeller is clear of loose stones or other objects which could damage the propeller or enter the engine or oil cooler air inlets.

TAXllNQ

1. Use minimum power to prevent stone damage particularly when moving away from rest and when turning.

2. Be alert for surface unevenness or obstructions which could cause propeller damage.

3. To turn the aircraft on soft or slippery surfaces using nosewheel steering assisted by brake will help to keep the power low. (Reducing the risk of damage to the propeller or runway surface). If possible avoid making small radius turns.

TAKEOFF

When aligned for takeoff set a low power before brake release. After brake release, as the aircraft begins to accelerate, move the power lever steadily fonvards to achieve Takeoff power. This procedure will reduce the risk of damaging the propeller by loose stones on the ground.

LANDING

BEFORE LANDING ON A PREPARED UNPAVED RUNWAY CHECK THAT THE SURFACE IS FIT FOR OPERATION AT THE INTENDED WEIGHT.


Reporl No: 01973-001 10-9


PASSENGER BRIEFINGS

GENERAL

In Sections 3 and 4 there are procedural actions that call for the pilot to brief the passengers. They fall into two categories those forming part of an emergency procedure and the more regular type ones for taxiing prior to takeoff and before landing. Tips for passenger briefings during an emergency cannot be specified as each situation will place a different demand on the pilot. However, much of the content in the Taxiing briefing tips can be used to brief the passengers, if time permits. Tips for the recommended subjects that should be covered for the regular passenger briefings are given in the following lists:

TAXIING (Section 4, para 4-7)

For aircraft with a standard cabin interior:

Stow hand baggage under the seats Put the seat back in the upright position Switch ofl electronic equipment Fasten seat belts and tighten lap strap Mention how to locate, remove and put on the passenger oxygen masks Mention the location and usage of the emergency exits Mention to remain buckled up during cruise in case of unexpected turbulence Mention the safety on board cards for more detailed information about the safety features (if available)

For aircraft with an executive cabin interior:

Stow hand baggage in the seat or cabinet drawers Move the seat to the required position for takeoff (as per the placard adjacent to each seat) Position the seat headrest to support the head Stow the tables, cabinet drawers, seat drawers and legrests Switch off electronic equipment Fasten seat belts and tighten lap strap Mention how to locate, remove and put on the passenger oxygen masks Mention the location and usage of the emergency exits Mention to remain buckled up during cruise in case of unexpected turbulence, but that the shoulder strap may be released once the fasten seat belt sign has been switched off Mention the safety on board cards for more detailed information about the safety features (if available)

Report No: 01 973-001 10-10



BEFORE LANDING (Section 4, para 4-14)

For aircraft with a standard cabin interior:

Stow hand baggage under the seats Put the seat back in the upright position Switch off electronic equipment Fasten seat belts and tighten lap strap

a Remain seated and buckled until the aircraft has come to a standstill and the engine is turned off

For aircraft with an executive cabin interior:

Stow hand baggage in the seat or cabinet drawers Move the seat to the required position for landing (as per Ule placard adjacent to each seat) Position the seat headrest to support the head Stow the tables, cabinet drawers, seat drawers and legrests Switch off electronic equipment Fasten seat belts and tighten lap strap Remain seated and buckled until the aircraft has come to a standstill and the engine is turned off


Report No: 01 973-001 10-11


Report No: 01 973-001 10-12



I PILOT'S OPERATING I

t HANDBOOK

Equipment List

I AIRPLANE EQUIPMENT LlST

PC- 12


Report No. 02047

SERIAL No.: REGISTRATION No.:

DATE:

PllATUS AIRCRAFT LTD. CH-6370 STANS SWITZERLAND

Within this report contains the Airplane Equipment List for the above indicated airplane at the time of license at the factory. This list itemizes the equipment installed lor the Basic Empty Weight condition and is grouped according to the ATA 100 System.

Items marked 'X' are included in the Basic Empty Weight as recorded in Figure 6-2. Airplane Basic Empty Weight, in the Pilot's Operating Handbook and Airplane Flight Manual Report 01973. Unmarked items are either optional or alternative items not delivered with the airplane.

1

Issued: June 30,1994 Revision 3: Aug 4,1995

b

Report No: 02047 1

AIRPLANE EQUIPMENT LIST ePkATusF PC XI1


Report No: 02047 2

~ssued: June 30,1994 Revision 3: Aug 4,1095

XPILATUS w AIRPLANE EQUIPMENT LIST P

PC XI1

r LOG OF REVISIONS


Revision Number

1 2 3 4

5

Report No: 02047 3

Page Number

1-1 6 1 - 16 1-1 6 1 1 0 1-1 1 1-12 1-1 2 1-6 i -9 1-10 1-1 1 1-12

Description

Complete revision Complete revision Complete revision Clock Part No. changed Altimeter, encoding option added EFIS, Argus moving map display option added Stormscope, processor and antenna option added AOA transmitters changed and temp. controller added Stabilizer trim actuator and warning system changed CAWS control unit changed Encoding altimeter mass and arm added Stormscope mass and arm added

AIRPLANE EQUIPMENT LIST SPLATUS3

PC XI1


Report No: 02047 4

Issued: June 30,1994 Revision 3: AUQ 4.1995

WPLATUSF AIRPLANE EQUIPMENT LIST PC XI1

Issued: June 30.1994 Revision 3: Aug 4,1995

Report No: 02047 5

Mark 'X' if

installed

X

X X X X X X X X X

X X X X X X . X X X X X X X

X X

X X X X X X X

ITEM

AIR CONDmONlNG

Valve. High Pressure Bleed Air Shut-off Valve. Low Pressure Non-Return Valve, Primary Shut-off Assembly. Refrigeration Package Valve. Temperature Control Injector, Water Assembly. Water Separator Switch, Overtemperature Sensor, Duct Temperature Switch, High Pressure Bleed Air Overpressure Controller, Temperature Valve, ECS Temperature Control Valve, Non-Return (Distribution) Assembly. Shut-off Valve Pump. Jet Valve, Check Valve, Solenoid Air NC Valve, Solenoid Air NO Controller, Outflow Valve Tank. Auxiliary Volume Valve. Cabin Safety Valve. Cabin Outflow Controller. Manual Pressure Compressor, Vapor System Evaporator Module. Cabin Fan LH Venturi, 2 in. Venturi, 1 % in. Evaporator Module, Cabin Fan RH Evaporator Module, Cabin Fan Flood

Heater. Cabin Valve, Shut-off, Firewall AlVPX Indicator Cab Rate Indicator Heater. Cockpit Switch. Cabin Pressure Differential Switch, Maximum Cabin Altitude Recirculation Blower

ATA Code

21

PllATUS PARTNO.

959.90.20.131

963.30.11.101 959.90.20.132 959.90.20.1 11 959.90.20.134 959.90.20.141 521.41.12.060 973.81.32.121 959.90.20.121 973.81.15.101

959.90.20.123 959.90.20. 1 34 963.30.1 1 .I02 521.22.12.071 959.90.91 .I61 959.90.91.151 963.14.1 1.001 963.14.1 1.002 959.90.91.141 959.90.91.1 1 1 959.90.91 .I31 959.90.91 .I21 959.90.91 .I01 959.90.22.101 959.90.22.1 12 963.41.12.073 963.41.12.074 959.90.22.1 13 959.90.22.1 13

969.67.81.501 963.30.1 1.001 965.23.23.201 965.16.21.001 969.67.81.501 965.23.21.532 965.23.21.531 959.90.22.151

MASS

kg

0.710

0.071 1.210 8.225 1.725 0.024 1.013 0.120 0.043 0.113

0.431 0.077 0.045 0.800 0.075 0.032 0.227 0.227 0.790 0.127 0.807 0.807 0.136 23.131 3.040 0.326 0.270 3.040 3.040

4.000 0.633 0.380 0.371 4.000 0.170 0.170 1.900

ARM m

2.40

2.15 2.94 2.55 2.87 2.34 2.62 2.76 2.83 2.12

3.05 4.60 4.25 3.02 4.30 4.32 4.27 4.27 3.45 3.06 4.29 4.29 3.47 10.17 9.40 2.94 2.31 9.40 9.40

5.22 3.13 3.52 3.52 5.22 4 41 4.41 9.40


Report No: 02047 6


ARM

m

9.40

8.40 3.25 3.50 3.50 8.40 11.50 11.50 6.38 6.38 11.50 11.50 4.89 5.10 5.10 5.97 3.13 5.50

3.50 3.50 3.50 4.50 4.50

7.00 3.50 3.50 3.50 8.40 12.70

Mark 'X ' il

tnstalled

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

X X

X

X X

X X

ATA

Code

22

23

MASS

kg

1.900

1.630 0.440 0.400 0.515 0.720 2.290 0.565 2.290 0.565 3.200 0.565 1.300 0.700 0.700 0.590 0.650 0.205

0.640 0.275 0.275 0.280 0.280

1.600 0.729 2.310 2.310 1.319 0.754

ITEM

Recirculation Blower

AUTO-FLIGHT

AUTOPILOT

Autopilot Control Computer Computer. Air Data KDC 222 Panel. KMC 321 Control Selector, KAS 297C Altitude Adapter, KTA 336 Pitch Trim Actuator. KSA 372 Pitch Servo Mount. KSM 375 Pitch Servo Actuator, KSA 372 Roll Servo Mount, KSM 375 Roll Servo Actuator. KSA 372 Yaw Servo Mount. KSM 375 Yaw Servo Stick Push ContMlT Gen Transmitter. AOA, Wing LH Transmitter, AOA. Wing RH Stick Pusher Actuator Stick Shaker Actuator Temperature Controller (2) (ea)

COMMUNICATIONS

AUDIO SYSTEM

Panel, Audio Selector Hand Microphone Pilot Hand Microphone Co-Pilot Headset. Pilot Headset, Co-Pilot

VHF COMINAV 1 SYSTEM

Converter. NAV. Dual Indicator VOWLOCIGSICV Transceiver, KX 155 Transceiver, KX 165 Antenna, COM 1 Antenna. NAV System

PILATUS PARTNO.

959.90.22.152

985.92.03.104 985.92.03.1 11 985.92.03.121 985.92.03.131 985.92.03.181 985.92.03.161 985.92.03.173 985.92.03.161 985.92.03.171 985.92.03.163 985.92.03.1 71 975.44.23.103 975.44.21.415 975.44.21.416 978.61.1 1.102 978.61.11.143 975.44.21.417

985.99.1 1.033 984.22.1 1.301 984.22.1 1.301 984.32.16.001 984.32.16.001

975.96.32.482 985.99.1 1.302 985.99.11.256 985.99.1 1.254 984.81.10.203 984.81.10.280


Issued: June 30.1 994 Revision 5: Dec 7.1995

Report No: 02047 7

Mark 'X' if

installed

X

X

X X X X X

X X X X X X X

X

X

X X

MASS

kg

0.095

2.310 2.310 1.319

0.675 2.940 6.500 0.634

36.400 12.970 7.960 4.267 4.267 4.500 1.369 1.369 0.340 0.100 0.085 0.085 0.085

2.500

0.105

16.470 16.470 2.415 1.880

ATA Code

24

25

ITEM

Coupler Antenna NAV llNAV 2

VHF COMINAV 2 SYSTEM

Transceiver. KX 155 Transceiver. KX 165 Antenna. COM 2

HF SYSTEM

Control Panel. KCU 951 HF Transceiver, KTR 953 Coupler, KAC 952 HF Antenna

ELECTRICAL POWER

Battery, Ni-Cad StarterIGenerator 2nd Generator Inverter. Battery Inverter, Generator Electr Cargo Door Actuation System Cross Tie, BAT GEN Cross Tie. GEN 2 RF Filter Interference (2) (ea.) Device. Pmteclon, Overvoltage Sensor, Battery Current Sensor. Generator 1 Sensor. Generator 2

EQUIPMENT FURNISHINGS

Emergency Locator Transmitter Narco ELT-910 Antenna. ELT

COCKPIT

Pilot Seat LH Co-Pilot Seat RH LH Bulkhead RH Bulkhead (Standard Cabin only)

ARM m

3.25

3.50 3.50 7.77

3.50 9.04 10.00 11.27

10.72 2.67 2.60 4.64 4.64 9.30 4.64 4.64 4.64 10.95 10.60 4.64 4.64

10.50

11.27

4.20 4.20 4.56 4.56

PILATUS PARTNO.

984.81.17.413

985.99.1 1.256 985.99.1 1.254 984.81.10.203

985.99.1 1.901 985.99.1 1.906 985.99.1 1.903 523.1 1.12.005

976.17.31.301 978.91.23.431 978.87.24.121 975.04.21.103 975.04.21.103 552.71.12.020 972.55.37.003 972.55.37.004 980.61.1 1.003 988.21.16.101 975.50.85.104 975.50.85.104 975.50.85.104

985.13.80.841

985.13.80.807

959.30.01.1 11 959.30.01.1 12 950.10.00.047 950.10.00.050


Report No: 02047 8

Issued: June 30,1994 Revision 3: Aug 4,1905

Mark 'X'

11 ~nstalled

X X

MASS

kg

11.340 11.340 1 1.340 11.340 11.340 11.340 11.340 11.340 1 1.340 1.650 1.650 1.650 1.697

17.822 17.822 17.538 17.530 17.538 17.538 5.215 5.215 5.215

20.100

12.271 14.204 1.950 1.950 1.900 1.697

0.300 2.260

ARM

m

5.88 5.88 6.40 6.40 7.21 7.21 8.02 8.02 8.80 6.65 7.37 9.20 9.15

5.75 5.75 7.08 7.08 8.15 8.56 6.43 6.43 7.92 4.85

5.39 5.30 6.65 7.37 9.20 9.15

1.79 4.46

ATA

Code

26

ITEM

CORPORTE COMMUTER CABIN (9 Seat Conftguration)

Seat PAX 1 LH FF Seat PAX 2 RH FF Seat PAX 3 LH FF Seat PAX 4 RH FF Seat PAX 5 LH FF Seat PAX 6 RH FF Seat PAX 7 LH FF Seat PAX 8 RH FF Seat PAX 9 RH FF Partition Net at Frame 24

or at Frame 27 Baggage Bay Net (Frame 34) Bulkhead (Frame 34)

EXECUTIVE CABIN (6 Seat Configurat~on)

Seat PAX 1 LH AF Seat PAX 2 RH AF Seat PAX 3 LH FF Seat PAX 4 RH FF Seat PAX 5 LH FF Seat PAX 6 RH FF Sidewall table A. LH, stowable Sidewall table 0. RH, stowable Sidewall table C, RH, stowable WC Cabinet RH, fwd, with curtain Stowage Cabinet, AH, fwd, with CDIRadio Audio System Refreshment Cabinet. LH, fwd Partition Net at Frame 24

or at Frame 27 Baggage Bay Net (Frame 34) Bulkhead (Frame 34)

FIRE PROTECTION

Fire Detection Sensor Fire Extinguisher

PllATUS

PART NO.

959.30.01.501 959.30.01.502 959.30.01.501 959.30.01.502 959.30.01.501 959.30.01.502 959.30.01.501 959.30.01 502 959.30.01.502 950.10.00.073 950.10.00.073 950.10.00.071 950.10.00.053

959.30.10.601 959.30.10.602 959.30.10.603 959.30.10.604 959.30.30.603 959.30.30.604 950.10.00.306 950.10.00.306 950.10.00.306 950.10.00.307

950.10.00.316 950.10.00.309 950.10.00.315 950.10.00.315 950.10.00.314 950.10.00.341

975.12.10.012 959.08.06.201

~PILATUSF AIRPLANE EQUIPMENT LIST

Report No: 02047 9

I. Issued: June 30,1994

MASS

kg

0.290 0.290 5.000 2.948 0.1 13 0.150 0.324 0.155 1.618

0.200 0.310 1.432 0.080 0.059 0.059 0.840 0.862 0.320 1.960 0.165 0.318 0.156 0.012 0.023 0.059 0.024 0.1 00

0.100 0.080 0.080

9 850 1.750 0.055 0.540

PILATUS PART NO.

978.73.18.131 978.73.18.131 978.73.14.202 978.73.20.001 978.73.20.601 978.73.20.021 975.44.16.204 975.44.16.203 978.73.20.012

949.86.76.106 963.87.22.301 968.35.21.001 963.04.26.722 963.04.26.720 963.04.26.720 528.24.12.1 10 968.84.51.106 975.35.12.601 968.84.1 1.403 968.84.71.104 968.84.71.103 963.04.26.713 115.55.07.065 963.04.26.803 963.04.26.703 1 15.55.07.209 975.37.31.234

975.37.31.233 975.37.31.232 975.37.31.231

960.30.01.151 960.30.01.291 965.61.88.101 960.30.01.271

Revision 5: Dec 7.1995

ARM m

5.97 12.60 12.60 6.77 6.00 6.00 3.88 . 3.50 4.89

5.60 3.70 3.87 3.03 3.03 3.62 3.71 2.71 2.59 5.52 6.32 5.67 5.52 5.52 5.52 5.52 5.60 5.89

5.89 5.89 5.89

6.25 6.25 6.25 3.75

ITEM

FLIGHT CONTROL

Actuator, Aileron Trim Linear Actuator, Rudder Trim Linear Actuator, Stabilizer Trim Linear Flaps Power Drive Unit Transmitter, Rotation (2) (ea.) Transmitter (3) (ea.) Indicator. Trim, Triple Indicator. Flaps Warn-System

FUEL

Filler Cap (2) (ea.) Valve, Shut-off Filter Check Valve. Firewall Check Valve, Firewall Check Valve Separator, Air Pump. Engine Driven Transmitter, Fuel Flow Pump, Centrifugal Boost (2) (ea.) Pump, Transfer Ejector (2) (ea.) Pump. Main Ejector (2) (ea.) Check Valve (2) (ea.) Check Valve (4) (ea.) Check Valve (2) (ea.) Check Valve (2) (ea.) Valve. Drain (4) (ea.) Sensor Fuel Qty. LWRH (ea) Collector Sensor Fuel Qty. LHIRH. Inner (ea) Sensor Fuel Qty, LHmH. Centre (ea) Sensor Fuel Qty, LWRH, Outer, (ea)

HYDRAULIC

Power Pack. Hydraulic Accumulator Indicator. Pressure (Mechanical) Valve. Landing Gear Selector

Mark 'X' if

installed

X X X X X X X X X

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

X X X

X X X X

ATA Code

27

28

29


Report No: 02047 10

Issued: June 30.1 94 Revision 5: Dec 7,199!

Mark 'X'

it

installed

X X X X X X

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

X X X X X X

X X X X X X

ATA

Code

30

31

32

ITEM

Handpump, Emergency Landing Gear Valve. Service Selector Actuator. Nose Gear Actuator, Main Gear (2) (ea.) Switch, Hydraulic (N2) Pressure Switch, Low Pressure

ICE AND RAIN PROTECTION

RegulatorIReliever Separator, Water Valve, Ejector Flow Control (5) (ea.) Timer. Airfoil Deice Switch. Pressure (5) (ea.) Boot Installation, RH Wing Inboard Boot Installation, RH Wing Outboard Boot Installation. LH Wing Inboard Boot Installation, LH Wing Outboard Boot Installation. RH Stabilizer Boot Installation, LH Stabilizer Brush Block, Prop Deice Timer. Prop Deice Controller. Pilot Windshield Deice Tube. Pitot Static Port, Heated (2) (ea.) Actuator, Inertial Separator

lNDlCATlNG/RECORDlNG SYSTEM

Clock CAWS Annunciator Panel Control Unit, CAWS Indicator. Cabin Temperature Kit, Inclinometer Pilot Kit, Inclinometer Co-Pilot

LANDING GEAR

Reservoir Master Cylinder Valve, Shuttle Valve, Parking Brake Hub. Main Wheel (2) (ea.) Tire. Main Wheel (2) (ea.)

PILATUS

PART NO.

968.85.82.1 1 1 960.30.01.261 960.30.01.1 1 1 960.30.01.101 973.81.14.303 973.81.14.304

959.89.01.012 959.89.01.01 0 959.89.01.008 959.89.01 ,011 973.81.14.31 t 959.89.01.032 959.89.01.034 959.89.01.031 959.89.01.033 959.89.01.036 959.89.01.035 968.29.13.222 968.29.13.221 972.81.32.201 965.11.12.302 965.11.22.102 978.73.15.301

999.61.1 1.21 2 972.81.32.012 972.81 32.003 975.15.16.101 975.96.32.431 975.96.32.431

959.47.91 .lo3 959.47.51.131 963.23.11.101 959.47.51 .I20 959.56.01.501 959.56.01.303

MASS

kg

0.790 0.500 0.940 6.660 0.072 0.072

0.410 0.590 0.285 0.245 0.057 2.449 3.225 2.449 3.225 1.594 1.594 0.1 00 2.1 15 0.235 0.411 0.411 0.960

0.755 0.830 1.055 0.120 0.020 0.020

0.090 0.250 0.171 0.123 6.680 11.680

ARM

m

4.00 6.00 3.26 5.92 6.00 6.00

3.50 5.57 6.00 6.25 6.00 5.41 5.41 5.41 5.41 13.20 13.20 0.90 2.00 ,4.51 5.50 10.20 2.65

3.50 3.50 4.95 3.70 3.50 3.50

4.45 3.10 3.10 3.30 6.25 6.25

EPRATUSF AIRPLANE EQUIPMENT LIST I PC XI1

Issued: June 30,1994 Revish 5: Dec 7,1995

Report No: 02047. 11

1

ARM m

6 25 3.88 3.88

4.13 4.13 6.00 6.00 6.00 6.00 6.00

6.00 14.38 4.38 4.38 3.00 3.50

3.50 3.50 3.50 3.50 3.50 3.50 3.50 3.50 3.50 3.50 3.80

3.50 9.00

Mark 'X' if

installed

X X X

X X X X X X X

X X X X X X

X X

X

X

X X

X X

ATA Code

33

34

ITEM

Assembly. Main Wheel Brake (2) (ea.) Hub. Nose Wheel Tire, Nose Wheel

LIGHTS

Light, Utility (2) (ea.) Assembly, Map Light (2) (ea.) Assembly. Landing Light (2) (ea.) Power Supply, Strobe (2) (ea.) Assembly. Strobe Light RH Assembly, Strobe Light LH Assembly, Wingtip Position Light RH with Radar Pod Assembly, W~ngtip Position Light LH Assembly. Tail Light Assembly, Wing Inspection Light Unit. Dim Taxi Light Instrument Light (8)

NAVIGATION

INSTRUMENTS

Altimeter, Encoding FtlMWin. hg. Allmeter, Encoding Altimeter Indicator. Airspeed, Pilot Indicator. Airspeed, &-Pilot Indicator, RMI, Pilot Indicator. RMI. Co-Pilot Ind'itor, Vertical Speed. Pilot Indicator, Vertical Speed. Co-Pilot Indicator. Attitude (2nd) Compass, MagneMc

ADF SYSTEM

Receiver. ADF. KR 87 Antenna, ADF

PILATUS PARTNO.

959.56.01.51 1 959.56.02.301 959.56.02.31 1

972.87.31.601 972.87.27.106 972.87.69.401 972.87.87.161 972.87.87.321 972.87.87.321 972.87.87.151

972.87.87.152 972.87.87.201 972.87.67.1 01 975.02.15.201 972.87.68.101 972.89.80.301

985.99.1 1.203 985.99.1 1.635 965.23.22.101 965.12.13.332 965.12.13.333 985.99.11.771 985.99.1 1.771 965.16.15.101 965.16.15.101 975.96.1 1.241 999.98.80.103

985.99.1 1 .I64 985.99.11.167

MASS

kg

7.500 2.850 4.820

0.200 0.400 0.339 0.550 0.700 0.700 0.800

0.800 0.965 0.245 0.500 0.341 0.300

1.000 1.000 1.000 1.200 0.900 1.210 1.210 0.360 0.360 0.680 0.290

1.420 1.300


Repori No: 02047 12


Mark 'X'

il

installed

X X

X X

X

X X X

X

X

ARM m

5.35 4.60

6.90 5.60

3.50 3.50 4.60

7.15 3.50 3.50 7.65 3.50 3.50 7.40 3.50 3.50 5.35 5.00

5.27 3.50 3.50 5.27 7.90 8.50

A

ATA

Code

PILATUS PART NO.

985.99.1 1.101 984.81 .?0.240

985.99.1 1.822 985.99.1 1.064

985.99.11.146 985.99.1 1.141 984.81.10.240

975.96.32.402 975.96.32.421 975.96.32.443 975.96.32.402 975.96.32.421 975.96.32.445 975.96.32.402 975.96.32.421 975.96.32.443 985.99.11.561 984.81.29.101 975.96.32.601

985.99.1 1.500 985.99.1 1.523 985.99.1 1.516 985.99.1 1.51 1 985.99.18.004 985.99.1 8.002

ITEM

DME SYSTEM

DME, KN 63 Antenna, DME/IFF

AlllTUDE HEADING REFERENCE SYSTEM

Sensor, LCR-92 Valve, KMT 112 Flux

TRANSPONDER SYSTEM

Transceiver, KT 71 Transceiver, KT 70 Mode S Antenna, ATC

EFlS

Symbol Generator. Pilot EADIEHSI Display, Pilot (2) (ea) Panel, Control, Pilot Symbol Generator. MFD EADllEHSl Display, MFD (2) (ea) Panel, CP467 Control, MFD Symbol Generator, Co-Pilot EADIIEHSI Display, Co-Pilot (2) (ea) Panel. Control. Co-Pilot Radar Altimeter KRA 405 Radar Altimeter Antenna (2) (ea) Moving Map Display. Argus

WEATHER RADAR

Weather Radar RWTX Weather Radar Indicator Weather Radar Control Panel Weather Radar Antenna Stormscope. Processor Stormscope. Antenna

MASS

kg

1.270 0.138

2.400 0.140

1.760 1.760 0.138

6.1 00 2.450 0.440 6.100 2.450 0.450 6.100 2.450 0.440 2.920 0.200

4.800 4.000 0.780 0.340 1 .880 1.750

WPILAIUSF AIRPLANE EQUIPMENT LIST SO XI1

Issued: June 30.1994 Revislon 4: Nw 14,1995

Mark 'X' if

installed

X X

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

Repwt No: 02047 13

ATA Code

35

ITEM

GPS SYSTEM

GPS, KLN 906 Database Antenna, GPS (W 918)

MARKER BEACON

Receiver, Marker Antenna. Marker

OXYGEN

Cylinder. Composite (with regulator) Gauge. Oxygen Pressure Indicator, Overboard Discharge Valve, Charging (external) Outlet. Crew Oxygen (2) (ea.) System, Pilot Breathing Crew Mask Stowage Box System, &Pilot Breathing Crew Mask Stowage Box Valve, bPosition Control Switch. Cabin Altitude Pressure Switch, Flowffransmit Pressure Gauge, Pressure

CORPORTE COMMUTER CABlN OXYGEN

Outlet. Passenger Oxygen (9) (ea.)

ARM m

3.50 3.50 4.90

3.50 11.40

3.71 4.09 4.02 4.03 4.29 4.50 4.50 4.50 4.50 4.20 4.20 4.20 4.20

1=5.32 2=6.07 33.05 4=7.01 5d.97 6=5.22 74.07 8=6.99 9=7.70

-

PILATUS PART NO.

985.99.1 1 .I92 985.99.1 1.177 985.99.1 1.193

985.99.11.001 984.81.10.260

957.12.12.21 1 957.12.12.1 80 957.12.16.901 963.02.81.101 957.12.22.212 957.12.01.206 957.10.49.921 957.12.01.206 957.10.49.921 963.32.1 1.101 973.81 S2.102 973.81.14.401 957.12.12.180

957.12.22.21 1 . . . . . . ... ... ... . . . . . . . . .

- -

MASS

kg

2.700 0.050 0.204

0.020 0.570

3.250 0.085 0.015 0.085 0.180 0.580 0.130 0.580 0.130 0.250 0.100 0.150 0.085

0.080 ... ... . . . ... ... . . . ... .. .

AIRPLANE EQUIPMENT LIST SPILATUSW PC XI1

Report No: 02047 14

Mark 'X'

11

tnstalled

X

X X

X X X

X

Issued: June 30,1994 Revision 4: Nov 14.1995

ATA

Code

36

61

71

72

ITEM

Mask. Passenger Oxygen (9) (ea.)

EXECUTIVE CABIN OXYGEN

Outlet. Passenger Oxygen (6) (ea.)

Mask. Passenger Oxygen (6) (ea.)

PNEUMATIC SYSTEM

Valve, Non-Return

PROPELLER

Assembly, HubIBladelSpinner Low Pitch Warning Switch

POWERPLANT INSTALLATION

EPA Tank Check Valves (2) (ea.) Shock Mounts (5) (ea)

ENGINE

PWC PT6A-678 Turboprop, dry

ARM

rn

1=5.43 2=6.24 3=7.06 4=7.87 5=8.68 6-5.43 7=6.24 8=7.06 9=7.87

1=5.57 2=5.57 3=7.20 4=7.24 5=7.24 6-8.61

1~5.57 2=5.57 3=7.20 417.25 517.25 6=8.61

4.50

0.83 1.03

1.81 1.81 1.94

1.80

PILATUS

PART NO.

957.12.10.21 1 ... . . . . . .

957.12.10.21 1 . . . . . . . . . ...

957.12.22.21 1 . . . . . . ... . . . ...

957.12.10.21 1 ... . . . . . . . . . . . .

536.11.12.011

968.29.13.002 973.10.11.101

571.70.12.090 963.04.23.105 944.81.21.409

968.20.13.001

MASS

kg

0.0120 ... ... . . .

0.0120 .. . ... ... . . .

0.080 ... ... ... . . . ... . . .

0.0120 ... ... . .. ... ...

1.491

80.196 0.135

0.300 0.090 1.434

244.123

XpILATUSS AIRPLANE EQUIPMENT LIST

PC XI1

Issued: June 30,1994 Revision 4: Nov 14,1995

Report No: 02047 15

- ARM

m

3.16 3.16 2.80

1.21

3.50

1.50 1.50

2.29 2 33 2.43

PlLATUS PART NO.

included in engine

941.94.11.413 941.94.1 1.412 941 94.1 1.41 1

975.21.15.206

975.29.02.013

578.10.12.039 578.10.12.040

968.32.14.102 978.73.01 .M)1 963.82.21.241

ITEM

IGNlTION

Box, Ignition Exciter

ENGINE CONTROL

Cable, PCL Cable, Condition Lever Cable, MOR

ENGINE INDICATING

Transducer, Torque (Engine Indication) Display Unit

EXHAUST

Assembly. Exhaust Stub LH Assembly, Exhaust Stub RH

OIL

Assembly, Oil Cooler Actuator, Thermal Valve, Drain

Mark 'X' if

installed

X

X X X

X

X

X X

X X X

MASS

kg

0.721 0.992 0.734

0.230

1.500

3.651 3.851

7.700 0.250 0.052

ATA Code

74

76

n

78

79


Report No: 02047 16

Mark 'X'

11

lnslalled

Issued: June 30.1994 ~ e v i s i h 3: Aug 4.1 995

ATA

Code ITEM

ADDITIONAL ITEMS (to be inserted as required)

PILATUS PART NO.

MASS

kg

ARM m

Documents

Pilatus PC-12 Pilot s Information Manual