learjet 35 manual

FlightSafety International, Inc.Marine Air Terminal, LaGuardia Airport

Flushing, New York 11371(718) 565-4100

www.flightsafety.com

FlightSafetyinternational

LEARJET 30 SERIES

PILOT TRAINING MANUALVOLUME 2

AIRCRAFT SYSTEMS

Courses for the Learjet 30 Series are taught at the following FlightSafety learning cen-ters:

Tucson Learning Center1071 E. Aero Park Boulevard\Tucson, AZ 85706(800) 203-5627FAX (520) 918-7111

Wichita (Learjet) Learning CenterTwo Learjet WayWichita, KS 67209(800) 491-9807FAX (316) 943-0314

Atlanta Learning Center1010 Toffie TerraceAtlanta, GA 30354(800) 889-7916FAX (678) 365-2699

Copyright © 2005 by FlightSafety International, Inc. All rightsreserved. Printed in the United States of America.

INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES

LIST OF EFFECTIVE PAGES

Dates of issue for original and changed pages are:

Original ...... 0 .......... June 1998Revision ..... .01........ Feb. 2004Revision ..... .02........ Aug. 2004Revision ..... .03........ Feb. 2006

NOTE:For printing purposes, revision numbers in footers occur at the bottom of every pagethat has changed in any way (grammatical or typographical revisions, reflow of pages,and other changes that do not necessarily affect the meaning of the manual).

THIS PUBLICATION CONSISTS OF THE FOLLOWING:

*Zero in this column indicates an original page.

Page *RevisionNo. No.

Cover .................................................. .03Copyright ................................................ 0iii–viii .................................................. .03SYL-i–SYL-iv ........................................ .02SYL-1–SYL-2 ........................................ .02SYL-3–SYL-4 .......................................... 0SYL-5–SYL-14 ...................................... .021-i–1-iv .................................................... 01-1............................................................ 01-2 ........................................................ .011-3............................................................ 01-4 ........................................................ .011-5–1-6 .................................................... 01-7–1-10................................................ .011-11–1-14 ................................................ 02-i–2-iv .................................................... 02-1............................................................ 02-2 ........................................................ .012-3............................................................ 02-4 ........................................................ .012-5............................................................ 02-6 ........................................................ .012-7–2-14 .................................................. 02-15 ...................................................... .012-16–2-26 ................................................ 0

3-i–3-iv .................................................... 03-1–3-5 .................................................... 03-6–3-7.................................................. .013-8............................................................ 03-9–3-10................................................ .013-11 .......................................................... 03-12 ...................................................... .013-13–3-14 ................................................ 04-i–4-iv .................................................... 04-1............................................................ 04-2 ........................................................ .014-3–4-6 .................................................... 05-i–5-iv .................................................... 05-1–5-6 .................................................... 05-7 ........................................................ .015-8–5-13 .................................................. 05-14 ...................................................... .015-15 ...................................................... .025-16–5-18 ................................................ 06-1–6-2 .................................................... 07-i–7-iv .................................................... 07-1–7-4 .................................................... 07-5–7-6.................................................. .017-7–7-9 .................................................... 07-10 ...................................................... .01


LEP-1

7-15–7-18 ................................................ 07-19–7-20.............................................. .017-21–7-23 ................................................ 07-24 ...................................................... .017-25–7-30 ................................................ 08-i–8-iv .................................................... 08-1–8-4 .................................................... 08-5 ........................................................ .018-6............................................................ 09-i–9-iv .................................................... 09-1–9-14 .................................................. 010-i–10-iv ................................................ 010-1–10-22 .............................................. 010-23 .................................................... .0110-24........................................................ 011-i–11-iv.................................................. 011-1–11-4 ................................................ 011-5 ...................................................... .0111-6–11-7 ................................................ 011-8 ...................................................... .0111-9–11-11 .............................................. 011-12–11-13 ......................................... .0111-14–11-20 ............................................ 012-i–12-iv ................................................ 012-1–12-6 ................................................ 012-7....................................................... .0112-8–12-10............................................ .0212-11–12-12 ............................................ 013-i–13-iv ................................................ 013-1.......................................................... 013-2 ...................................................... .0113-3–13-6 ................................................ 014-i–14-iv ................................................ 014-1–14-2 ................................................ 014-3–14-4.............................................. .0114-5–14-13 .............................................. 014-14 .................................................... .0114-15–14-16 ............................................ 0

15-i–15-iv ................................................ 015-1–15-3 ................................................ 015-4–15-5.............................................. .0115-6.......................................................... 015-7–15-8.............................................. .0115-9–15-12 .............................................. 015-13 .................................................... .0115-14........................................................ 015-15 .................................................... .0115-16–15-26 ............................................ 016-i–16-iv .............................................. .0316-1–16-3 ................................................ 016-4 ...................................................... .0116-5.......................................................... 016-6 ...................................................... .0116-7–16-20 .............................................. 016-21..................................................... .0116-22–16-42.......................................... .0317-i–17-iv ................................................ 017-1 ...................................................... .0117-2.......................................................... 017-3–17-5.............................................. .0117-6.......................................................... 017-7–17-8.............................................. .0117-9–17-12 .............................................. 0WA-1 ........................................................ 0WA-2–WA-15 ........................................ .01APP-i–APP-ii ............................................ 0APP-1–APP-6 .......................................... 0ANN-1–ANN-6 ........................................ 0



NOTICE

The material contained in this training manual is based on information obtainedfrom the aircraft manufacturer’s Airplane Flight Manual, Pilot Manual and Mainten-ance Manuals. It is to be used for familiarization and training purposes only.

At the time of printing, it contained then-current information. In the event of conflictbetween data provided herein and that in publications issued by the manufactureror the FAA, that of the manufacturer or the FAA shall take precedence.

We at FlightSafety want you to have the best training possible. We welcome anysuggestions you might have for improving this manual or any other aspect of ourtraining program.

FOR TRAINING PURPOSES ONLY


CONTENTS

SYLLABUS

Chapter 1 AIRCRAFT GENERAL

Chapter 2 ELECTRICAL POWER SYSTEMS

Chapter 3 LIGHTING

Chapter 4 MASTER WARNING SYSTEM

Chapter 5 FUEL SYSTEM

Chapter 6 AUXILIARY POWER UNIT

Chapter 7 POWERPLANT

Chapter 8 FIRE PROTECTION

Chapter 9 PNEUMATICS

Chapter 10 ICE AND RAIN PROTECTION

Chapter 11 AIR CONDITIONING

Chapter 12 PRESSURIZATION

Chapter 13 HYDRAULIC POWER SYSTEMS

Chapter 14 LANDING GEAR AND BRAKES

Chapter 15 FLIGHT CONTROLS

Chapter 16 AVIONICS

Chapter 17 MISCELLANEOUS SYSTEMS

WALKAROUND

APPENDIX

ANNUNCIATOR PANEL

INSTRUMENT PANEL POSTER

Revision .02 SYL-i

SYLLABUSCONTENTS

Page

LEARNING CENTER INFORMATION .......................................................................... SYL-1

DESCRIPTION OF TRAINING FACILITIES ................................................................. SYL-1

TYPE OF AIRCRAFT....................................................................................................... SYL-7

CATEGORY OF TRAINING ............................................................................................ SYL-7

DUTY POSITION ............................................................................................................. SYL-7

CURRICULUM TITLE..................................................................................................... SYL-7

CURRICULUM PREREQUISITES.................................................................................. SYL-7

Core Training Curriculum Prerequisites..................................................................... SYL-7

Prerequisite Experience .............................................................................................. SYL-8

FLIGHTSAFETY TRAINING POLICY........................................................................... SYL-9

DESCRIPTION OF INITIAL COURSE ........................................................................... SYL-9

Ground Training ......................................................................................................... SYL-9

Flight Training.......................................................................................................... SYL-10

Qualification Check.................................................................................................. SYL-10

DESCRIPTION OF TRANSITION COURSE (FAR 135).............................................. SYL-10

Ground Training ....................................................................................................... SYL-10

Flight Training.......................................................................................................... SYL-10

Qualification Check.................................................................................................. SYL-11

COURSE OBJECTIVES ................................................................................................. SYL-11

TRAINING SCHEDULE (TYPICAL)............................................................................ SYL-12



LEARJET 30 Series P ILOT TRA IN ING MANUAL

Revision .02 SYL-iii

ILLUSTRATIONS

Figure Title Page

SYL-1 Wichita Facility Floor Plan ................................................................................. SYL-3

SYL-2 Tucson Facility Floor Plan.................................................................................. SYL-4

SYL-3 Atlanta Facility Floor Plan.................................................................................. SYL-5






LEARNING CENTER INFORMATIONFlightSafety International is an aviation training company that provides type-specific trainingprograms for over 50 different models of aircraft, using a fleet of 100 simulators. FlightSafetyoperates over 38 Learning Centers, including Centers in Europe and Canada.

Training for the Learjet is conducted at the FlightSafety Learning Centers in Wichita, Kansas;Tucson, Arizona; and Atlanta, Georgia. The Centers are owned and operated by FlightSafetyInternational and are located at the following address:

FlightSafety International FlightSafety InternationalLearjet Learning Center Tucson International AirportTwo Learjet Way 1071 East Aero Park BoulevardP.O. Box 9320 Tucson, Arizona 85706Wichita, KS 67209

FlightSafety InternationalAtlanta Learning Center1010 Toffie TerraceAtlanta, GA 30354

DESCRIPTION OF TRAINING FACILITIESEach classroom and briefing room is adequately heated, lighted, and ventilated to conform tolocal building, sanitation, and health codes. The building construction prevents any distractionsfrom instruction conducted in other rooms or by flight operations and maintenance operationson the airport.

Classrooms are equipped for computer-based presentations, controlled from a specially designedlectern. A standard overhead projector is mounted on the lectern. Most lectern and student po-sitions are equipped with a student responder system. Cockpit panel posters are on display.

Briefing rooms are equipped with cockpit panel posters, a white liquid chalkboard, a table, andchairs for individual or small-group briefings. A floor plan of the Centers follows.

SYLLABUS

Revision .02 SYL-1FOR TRAINING PURPOSES ONLY

Wichita Learning Center

Classroom Size in StudentNumber Feet Capacity

1 20X25 182 20X25 123 20X25 184 20X20 125 20X30 246 20X20 127 20X25 18

Tucson Learning Center

1 22X19 182 22X19 183 22X20 184 24X20 185 24X20 126 16X17 107 8X12 2

Atlanta Learning Center

104 16X20 8105 18X25 12113 18X25 12

Revision .02SYL-2 FOR TRAINING PURPOSES ONLY



SYL-3FOR TRAINING PURPOSES ONLY



Learjet Learning CenterWichita, Kansas

SECOND FLOOR

MAINENTRANCE

EXIT

EXIT EXIT

COMPUTER ROOM

ST

OR

AG

ES

HO

PE

LEC

.

MA

INT

EN

AN

CE

SELFLRNGROOM

SELFLRNGROOM

BR BR BR BR BR BR BR

ENGSUPVSR

25CPT

35CPT

LVL B35

LVL A25

LVL C35

RECEPTIONAREA

MANAGER

SCHEDULING

CMPTRRM

DIR. OFTNG

DIR. OFSTNDS

STORAGE

MAINTENANCETRAINING

LAB

CLASSROOM7

CLASSROOM2

CLASSROOM1

CLASSROOM3CLASSROOM

6CLASSROOM

4

MEN

LOUNGE

WO

ME

N

CLASSROOM5

JANITOR

TE

LEP

HO

NE

GR

OU

ND

SC

HO

OL

INS

TR

UC

TO

RS

EXIT


Figure SYL-1. Wichita Facility Floor Plan

SYL-4 FOR TRAINING PURPOSES ONLY



BALCONY

SIMULATOR ROOM117

SIMULATOR ROOM132

COMPUTER ROOM124

COMPUTER ROOM120

BRIEF127

CPM128

CPM131

BRIEF129

BRIEF130

BRIEF133

BRIEF135

MAINT.RM126

BRIEF121

BRIEF118

MEN WOMEN

EXERCISEROOM

103 MARKETING104

CUST.SUPPORT

MGR.106

ASS’TMGR107

CUSTOMER SUPPORTREPRESENTATIVE

105MANAGER109

ADMIN

108

DOT102

TOILETSTORAGE

101

STORAGE

CPM119

CPM137

BRIEF136

CPM122

BRIEF123

CPM134

MAINTENANCE125

LOBBYCONFERENCEROOM

110

DOMT115

INSTRUCTORS114

INSTRUCTORS111

PROGRAMMGRS

112

VEST FIRST FLOOR

Learjet Learning CenterTucson, Arizona


LEAR 31

LEAR 35(FC350)

LEAR 60

LEAR 25 LEAR 55 LEAR 35(200)

LEAR 45

CHALLENGER 601-3R

PROGRAMMANAGER

113

DOSPROGRAMMANAGER

116

CLASSROOM 216

CLASSROOM 215

STORAGE214

CLASSROOM 213

CLASSROOM 212

CLASSROOM 211

CLASSROOM 210

CLASSROOM 209

CLASSROOM 207

CLASSROOM 205

MAINT. LAB204

CLASSROOM 17203

SELFLEARNING

202

LIBRARY201

VEND206

LOUNGE

CO

FF

EE

CLASSROOM 224

CLASSROOM 223

CLASSROOM 222

CLASSROOM 221

CLASSROOM 220

CLASSROOM 219

CLASSROOM 218

STORAGE217

MEN WOMEN

ST

OR

AG

E20

8

TELE TELE

SECOND FLOOR

Figure SYL-2. Tucson Facility Floor Plan

Revision .02 SYL-5FOR TRAINING PURPOSES ONLY



PUMPROOM

ELECT.

COMPUTERROOM

14

CRJSIMULATOR

15

CRJSIMULATOR

16

KING AIR200

SIMULATOR

17JET-STAR

ATR CABINTRAINER

LEAR 35SIMULATOR

ROLL-UP DOORROLL-UP DOOR

PUMPROOM

HEATEXCHANGE

12

LEAR 31ASIMULATOR

13

LEAR 60SIMULATOR

ROLL-UP DOOR

CLASS113

CLASS115

CLASS120

CLASS118

EXERCISE PARTS

STORAGE

PUMPROOM

ATR 42/72SIMULATOR

4

EMB 120SIMULATOR

5

LEAR 45SIMULATOR

6

CLASS107

CLASS109

CLASS111

PUMPROOM

PUMPROOM

CR

J

CAB

IN T

RAI

NER

CORP.CABINTRAINER

DOORTRAINER

KING AIR350

SIMULATOR8

DASH 8100/200/300SIMULATOR

9

CLASS116

CLASS114

CLASS112

CLASS110

PUMPROOM

CRJSIMULATOR

2

CITATION IISIMULATOR

3

RECEPTION DESKADMINISTRATIVE

CUST.SUPP.MGR.

CSMA

COOR.RECORDSSUPPORT

PHONES CLASS102

CLASS104

CLASS106

CIT

FTDCRJCPT

WOMEN MEN

DOS

VENDING

NET-WORKADMIN

SERVER

ELECT

ELEVATOR

CLASS105

CLASS108 CRJ

SIMULATOR1

JANITORJAA

RDOC

DOMT

STAIRS

CUSTOMERLOUNGE

COFFEE

CONF.ROOM

MKT.MGR.

CNTR.MGR.

LOBBY

STAIRS

DOT

ASST.MGR.

CLASS101

CLASS103

SMOKEROOM

ASA

ELECTEMPLOYEE

LOUNGECOFFEE

CRJ 700SIMULATOR

7

Atlanta Learning Center1010 Toffie Terrace

Atlanta, GeorgiaFirst Floor


Figure SYL-3. Atlanta Facility Floor Plan (1 of 2)

Revision .02SYL-6 FOR TRAINING PURPOSES ONLY



BR207

BR209

BR216

BR218

BR211

BR213

BR222

BR220

BR215

BR217

BR226

BR224

PHONE

INSTRUCTORSP.M.

P.M.

P.M.

P.M.STAIRS

P.M.

LOBBY

WOMEN MEN

CLASS203

CLASS205

RAFTDEMONSTRATION

ELEVATOR

BR219

SIMULATORBAY

ROOF

ROOF BELOW

CLASS201

P.M.P.M.

P.M.

INSTRUCTORS

COMPUTERROOM

ROOF

ROOF

ROOF

MGR.FTD MAIN

SIMULATORSHOP

TECH.LIB.

BR221

BR228

BR230

BR223

227 229 231

BR225

BR234

BR236

BR238

BR232

KIN

GA

IRC

PM

REST

RO

OM

EMBR

AERC

PM

BR306

BR305

BR304

BR303

CO

FFE

EB

RE

AK

A

RE

A

CLASS210

CLASS206

CLASS208

CLASS202

CLASS204PHONE

CLASS212

CLASS214

BR310

BR309

BR308

BR307

BR314

BR313

BR312

BR311

ENG.OFF.REST-

ROOM

BR302

BR301

SELFLEARNINGCENTER

Second Floor

SIMULATORBAY

SIMULATORBAY

ELECT. JANITOR

ELECTRICAL

COMPUTERROOM

OPENBELOW

Atlanta Learning Center1010 Toffie Terrace

Atlanta, Georgia


Figure SYL-3. Atlanta Facility Floor Plan (2 of 2)




TYPE OF AIRCRAFTLR-JET

CATEGORY OF TRAININGInitial and Transition training for a LR-JET type rating added to an existing pilot certificate orthe issuance of an Airline Transport Pilot Certificate with a LR-JET type rating.

DUTY POSITIONPilot-in-Command (PIC)

CURRICULUM TITLELR-JET Series Pilot Training Course and Advance Simulation Training Program

CURRICULUM PREREQUISITES

CORE TRAINING CURRICULUM PREREQUISITES

§61.63

A pilot may enroll in this course and complete all of the items of the practical test required fora LR-JET type rating that are authorized to be accomplished in the flight simulator, then com-plete the items not approved for flight simulator in flight in a LR-JET Series airplane, if the pilot:

1. Holds a private pilot certificate with an airplane rating.

2. Holds an instrument rating.

3. Has a minimum of 1,000 hours flight experience in airplanes as a pilot. (May be waived atthe discretion of the Center Manager).

4. Holds a MEL catagory rating without centerline thrust limitation.

§61.57

A pilot who meets the above requirements of §61.63 may concurrently apply for an Airline Trans-port Pilot certificate with a LR-JET type rating, providing the pilot:

1. Holds a commercial pilot certificate or an ICAO recognized Airline Transport Pilot or Com-mercial Pilot License without restrictions.

SYL-7

A




2. Meets the eligibility requirements of §61.153.

3. Has passed the written test required by §61.155.

4. Meets the experience requirements of §61.159.

Prerequisite ExperienceThe curriculum is designed to accommodate pilots with varied levels of experience. Depend-ing on the pilot’s experience and fight simulator approval level, the pilot may qualify for either100% flight simulator curriculum or a combination curriculum using both flight simulator andaircraft. If a 100% flight simulator practical test is not accomplished, then aircraft training andtesting will be required.

1. Pilots who meet the appropriate requirements in §61.63 (e)(4) or §61.157 (g)(3) may obtainan unlimited type rating.

2. Pilots who meet the appropriate requirements in §61.63 (e)(5) or §61.157 (g)(4) may be is-sued an added rating with pilot-in-command limitations. Fifteen hours of supervised oper-ating experience as PIC accomplished IAW §61.63 (e)(8)(ii) or §61.15(g)(6)(ii) will be re-quired to remove this limitation.

3. Pilots who do not meet the appropriate requirements in items 1 or 2 above may still beissued a type rating with pilot-in-command limitations. Twenty-five hours of supervisedoperating experience as PIC accomplished IAW §61.63 (e)(12)(ii) or §61.157 (g)(9)(ii) will remove this limitation.

4. Pilots completing training and testing who do not want an SOE limitation on their certifi-cate may complete the following tasks on a static airplane or in flight, as appropriate:

a. Preflight Inspection

b. Normal Takeoff

c. Normal ILS Approach

d. Missed Approach; and

e. Normal Landing

5. Pilots who receive training/testing in a Level A or B simulator must complete the practicaltest in the aircraft.

SYL-8




FLIGHTSAFETY TRAINING POLICYThe policy is to train to “Proficiency” based on need-to-know information.

DESCRIPTION OF INITIAL COURSEThe LR-JET Initial Course is scheduled for twelve days and consists of the following pro-grammed hours:

GROUND TRAININGGeneral Operational Subjects ........................................................................................................ 6.0

Aircraft Systems ............................................................................................................................ 31.0*

Systems Integration ........................................................................................................................ 2.0

Prebrief/Postbrief ............................................................................................................................ 9.0

Oral Exam and Pre/Postbriefings for Qualification .................................................................. 3.0

Total Ground Training .................................................................................................................. 51.0

* If a pilot requires FMS training, he/she will receive an additional 4.0 hours of training.

SYL-9




FLIGHT TRAINING

Flight Training (Simulator) .................................................................................................... 10.0Aircraft Flight Training (Typical) (If Necessary) ..................................................................... 2.0

Total Flight Training............................................................................................................... 12.0

QUALIFICATION CHECK

Flight Simulator........................................................................................................................ 2.0Aircraft ..................................................................................................................................... 0.5

Total Qualification Hours ......................................................................................................... 2.5

An applicant may choose to take the entire practical test in the aircraft rather than in the simulator.

The 10 hours of flight simulator training is for left-seat training. Normally, a pilot is trained as acrew with another pilot. A pilot training not as a crew will receive an additional 5 hours of simu-lator training.

DESCRIPTION OF TRANSITION COURSE (FAR 135)The LR-JET Initial Course is scheduled for twelve days and consists of the following pro-grammed hours:

GROUND TRAINING

General Operational Subjects ........................................................................................................ 6.0

Aircraft Systems ............................................................................................................................ 31.0*

Systems Integration ........................................................................................................................ 2.0

Prebrief/Postbrief (Simulator) ...................................................................................................... 9.0

Oral Exam and Pre/Postbriefings for Qualification .................................................................. 3.0

Total Ground Training .................................................................................................................. 51.0

* If a pilot requires FMS training, he/she will receive an additional 4.0 hours of training.

FLIGHT TRAINING

Flight Training (Simulator) .................................................................................................... 10.0

Total Flight Training............................................................................................................... 10.0

SYL-10




QUALIFICATION CHECK

Flight Simulator........................................................................................................................ 2.0Aircraft Preflight ..................................................................................................................... 0.5Loft .......................................................................................................................................... 4.0

Total Qualification Hours ......................................................................................................... 6.5

The 10 hours of flight simulator training is for left-seat training. Normally, a pilot is trained asa crew with another pilot. A pilot training not as a crew will receive the hours specified in theoperators approved training program.

COURSE OBJECTIVESUpon the completion of this course, the pilot will have the necessary knowledge and skills todemonstrate that he/she is the master of the aircraft, with the successful outcome of a proce-dure or maneuver never in doubt, and to meet or exceed the requirements/standards listed inFAA-S-8081-5 Airline Transport Pilot and Type Rating Practical Test Standards.

Successful completion of the LR-JET Pilot Training Course will satisfy the requirements forthe following:

• Second-in-Command qualifications as specified in FAR 61.55

• Pilot-in-Command recent flight experience as specified in FAR 61.57

• Pilot-in-Command Proficiency Check as specified in FAR 61.58

• Additional Aircraft Ratings as specified in FAR 61.63

• Category II Pilot Authorization as specified in FAR 61.67

• Airplane Rating Requirements as specified in FAR 61.157

• Practical Test Requirements for Airplane ATP Certifications and Associated Class and TypeRatings

Successful completion of the LR-JET Initial/Transition Pilot Training Course and the subsequentLR-JET practical test will also satisfy the requirements for the following:

• Instrument Competency Check as specified in FAR 61.57 (d)

• Initial Equipment/Transition training for FAR 135 certificate holders contracting with Flight-Safety

SYL-11




TRAINING SCHEDULE (TYPICAL)Listed below is a typical schedule for the pilot training curriculum. On occasion, the schedulemay be rearranged to meet the needs of the client or Center. In addition, the times allotted foreach lesson may vary due to pilot experience and class size. The schedule consists of 12 train-ing days.

NOTESimulator hours reflect left-seat time for one pilot, performing all pilot flying duties.In addition, 1.0 hour for briefing and .5 hour for debriefing are allocated.

Day 1 FlightSafety Administration .................................................................. 0.5Classroom.................................................................................................. 7.0

Aircraft GeneralElectricalLighting

Day 2 Classroom.................................................................................................. 7.5Master Warning SystemFuelPowerplantFire Protection

Day 3 Classroom.................................................................................................. 7.5PneumaticsIce and Rain ProtectionAir ConditioningPressurizationHydraulic Systems

Day 4 Classroom.................................................................................................. 7.5Landing Gear and BrakesFlight ControlsAvionics

Day 5 Classroom.................................................................................................. 7.0Miscellaneous SystemsWeight and BalancePerformanceExamination

CPT ............................................................................................................ 2.0

Day 6 Simulator .................................................................................................. 2.0Simulator Period No. 1

Day 7 Simulator .................................................................................................. 2.0Simulator Period No. 2

SYL-12

1-i

CHAPTER 1AIRCRAFT GENERAL

CONTENTS

Page

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

GENERAL............................................................................................................................... 1-1

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

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

Fuselage ........................................................................................................................... 1-4

Wing................................................................................................................................. 1-9

Empennage..................................................................................................................... 1-10

AIRPLANE SYSTEMS ........................................................................................................ 1-10

Electrical Power Systems............................................................................................... 1-10

Lighting.......................................................................................................................... 1-10

Fuel System.................................................................................................................... 1-11

Powerplant...................................................................................................................... 1-11

Ice and Rain Protection .................................................................................................. 1-11

Air Conditioning and Pressurization.............................................................................. 1-11

Hydraulic Power Systems .............................................................................................. 1-12

Landing Gear and Brakes .............................................................................................. 1-12

Flight Controls ............................................................................................................... 1-12

Automatic Flight Control System (AFCS) .................................................................... 1-12

Pitot-Static System......................................................................................................... 1-12

Oxygen System.............................................................................................................. 1-13




1-iii

ILLUSTRATIONS

Figure Title Page

1-1 Learjet 35/36............................................................................................................. 1-2

1-2 General Dimensions ................................................................................................. 1-2

1-3 Turning Radius ......................................................................................................... 1-3

1-4 Danger Areas............................................................................................................ 1-3

1-5 Fuselage Sections ..................................................................................................... 1-4

1-6 Radome..................................................................................................................... 1-4

1-7 Nose Compartment................................................................................................... 1-4

1-8 Passenger-Crew Door ............................................................................................... 1-5

1-9 Door Latch Inspection Port ...................................................................................... 1-6

1-10 Emergency Exit ........................................................................................................ 1-7

1-11 Windshield................................................................................................................ 1-8

1-12 Window Locations (Typical) .................................................................................... 1-8

1-13 Tailcone Door ........................................................................................................... 1-8

1-14 Learjet 35/36 Wing................................................................................................... 1-9

1-15 Empennage............................................................................................................. 1-10




INTRODUCTIONThis training manual provides a description of the major airframe and engine systemsinstalled in the Learjet 35/36.

This chapter covers the structural makeup of the airplane and gives a general descriptionof the systems. No material is meant to supersede any of the manufacturer’s system oroperating manuals.

The material presented has been prepared from the basic design data, and all subsequentchanges in airplane appearance or system operation will be covered during academictraining and subsequent revisions to this manual.

The Annunciator Panel section in this manual displays all light indicators, and page ANN-1should be folded out and referred to while studying this manual.

GENERALThe Learjet 35/36 is certificated under FAR Part25 as a two-pilot transport category airplane,

approved for all-weather operation to a max-imum altitude of 45,000 feet.

CHAPTER 1AIRCRAFT GENERAL

1-1FOR TRAINING PURPOSES ONLY



STRUCTURESGENERALFigure 1-1 shows the Learjet 35/36. Thestructure consists of the fuselage, the wing, theempennage, and flight controls. The discussionon the fuselage includes all doors and windows.Figure 1-2 shows the general dimensions of the airplane.

Figure 1-3 displays the airplane turning radius.

Figure 1-4 displays the danger areas aroundthe Learjet 35/36 presented by the weatherradar emission cone, engine intakes, andengine exhaust cones.

Revision .011-2 FOR TRAINING PURPOSES ONLY



14 FT 8 IN.447.0 cm

38 FT 1 IN.1,161.0 cm

39 FT 6 IN.1,203.0 cm

8 FT 3 IN.251.0 cm

12 FT 3 IN.373.0 cm

20 FT 2 IN.615.0 cm

48 FT 7 IN.1,480.0 cm

Figure 1-2. General Dimensions

Figure 1-1. Learjet 35/36




WEATHER RADAR OPERATING

30 FT 12 FT

40 FT

700˚ F 100˚ F

ENGINE INTAKE

VALUES FOR TAKEOFF RPM APPROXIMATELY DOUBLE

ENGINE EXHAUST

Figure 1-4. Danger Areas

42 FT 2 IN.

Figure 1-3. Turning Radius

FUSELAGEGeneralThe fuselage is constructed of stressed all-metal skin with stringers. It employs the arearule design to reduce aerodynamic drag, andhas four basic sections. (See Figure 1-5.) They are:

1. The nose section which extends fromthe radome aft to the forward pressurebulkhead.

2. The pressurized section, which in-cludes the cockpit and passenger areas,extends aft to the rear pressure bulk-head. On 36 models this bulkhead isfurther forward than on 35 models toprovide space for the larger fuselagetank.

3. In both models, the fuselage fuel sec-tion starts just aft of the rear pressurebulkhead and extends to the tailcone.

4. The tailcone section extends aft of thefuel section.

The fuselage also incorporates attachmentsfor the wings, tail group, engine supportpylons, and the nose landing gear.

In addition to the pressurized cockpit andpassenger compartments, the fuselage includesthe nose wheel well, an unpressurized nosecompartment, and a tailcone compartmentused for equipment installation.

Nose SectionThe nose of the fuselage (Figure 1-6) is formedby the radome. Aft of the radome is the nosecompartment.




Figure 1-6. Radome

FWD PRESSURE BULKHEAD

FUELSECTION35 MODEL

FUEL SECTION36 MODEL

TAILCONE SECTIONPRESSURIZED SECTION 36 MODEL

PRESSURIZED SECTION 35 MODEL

NOSESECTION

AFT PRESSURE BULKHEAD(36 MODEL ONLY)

AT FRAME 18

AFT PRESSURE BULKHEAD(35 MODEL ONLY)AT FRAME 22

Figure 1-5. Fuselage Sections

Figure 1-7. Nose Compartment

The nose compartment access panels are on topof the fuselage (Figure 1-7), forward of thewindshield. The panels must be removed foraccess to various electronic components,oxygen bottle (when installed in the nose),emergency air bottle, and the alcohol anti-icing reservoir.

Pressurized SectionThe pressurized cabin lies between the forwardpressure bu lkhead and the a f t p ressu rebulkhead , and inc ludes the cockpi t andpassenger compartment. Within the passengercompartment is a 500-pound-capacity baggagearea at the back of the cabin, a lavatory, acabinet for storage of provisions, and galleyequipment (depending on the airplane).

The passenger-crew door is located on the leftside of the fuselage, just aft of the cockpit. Oneof the windows on the right side of the cabinserves as an emergency exit.

The cockpit seats two pilots and is fitted witha large, curved, two-piece windshield.

Passenger-Crew DoorThe primary entrance and exit for passengersand crewmembers is through the clamshelldoor, located on the left side of the forward

fuselage. (See Figure 1-8.) The standardentrance door is 24 inches wide, but there isan optional 36-inch door. The upper doorserves as an emergency exit, and the lowerdoor has integral entrance steps.

The upper portion of the door has both outsideand inside locking handles connected to acommon shaft through the door. Rotating eitherof these handles to the close position drivessix locking pins into holes in the fuselageframe (three pins forward and three aft) andtwo pins through interlocking arms that securethe two door halves together.

The lower door has a single locking handle onthe inside. Rotating the lower door handle tothe closed (forward) position drives two pinsinto holes in the fuselage frame (one forwardand one aft). There are a total of 10 lockingpins on the two door sections.

To facilitate alignment of the upper doorlocking pins dur ing c losing, an e lect r icactuator motor, torque tube assembly, and oneor two hooks are installed in the lower door.The hooks engage rollers installed on the upperdoor and draw the two halves together. Theactuator motor is operated from inside theairplane by a toggle switch on the lower door,and from the outside by a key switch. Shouldthe motor fail, the hooks can still be operated




Figure 1-8. Passenger-Crew Door

manually from inside. Access is provided tothe torque-tube mechanism through a panel inthe lower door, and a ratchet handle providedin the airplane tool kit can be used to operatethe torque-tube manually.

NOTEOne hook and roller is used on 24-inch doors, while two hooks androllers are used on 36-inch doors.

When the door handles are in the closedposition, the pins all contact microswitches.If any of the switches is not actuated, a red

DOOR light illuminates on the annunciatorpanel. (See Annunciator Panel section.) If thelight illuminates while the door is closed,eight inspection ports enable the crew toconfirm the position of the door-frame latchingpins by observing the position of two whitealignment marks (Figure 1-9). The two latchpins which connect the upper and lower doorsare visible through the upholstery gap at theinterface and do not have white lines.

When closing the doors from the inside, closeand latch the lower door first. Then, close theupper door and actuate the door motor switchto the closed position. This engages the hooksover rollers in the upper door, and cinches theupper door down tight while allowing thelocking pins to line up properly and meet themicroswitches as the upper door handle isrotated to the closed position. The DOORlight will remain illuminated until the hooksare backed away from the upper door rollersby reverse operation of the door motor switch.

A secondary safety latch is installed on thelower door and is separate from the door-locking system. It consists of a notched pawlattached to the door. The pawl engages a strikerplate attached to the frame when the door isclosed.This engagement holds the lower doorclosed while the locking handle is beingpositioned to the locked position. Additionally,it prevents the door from falling open as soonas the door handle is opened. The latch isreleased by depressing the pawl.

Cables and hydraulic dampers are provided tostabilize the lower door when lowering it andwhen using it as a step. The 24-inch door hasone cable and a hydraulic damper. The 36-inch door has two cables and may have anoptional hydraulic damper. The cables areconnected to takeup reels in the lower door andare also used to pull the door closed frominside the airplane.

The key switch is used to secure the door fromthe outside. By inserting a key into the switchand turning it in one direction, the actuatormotor drives the hooks to engage the upper

1-6 FOR TRAINING PURPOSES ONLY



LOCKED

NOT LOCKED

Figure 1-9. Door Latch Inspection Port

door rollers. Turning it in the other directiondrives the hooks from the rollers to permitopening the door.

NOTEAnytime the airplane is occupiedwith the entry doors locked, thehooks must be released. This permitsopening the upper door for emer-gency egress.

The red DOOR light illuminated means:• Any one of the 10 latch pins is not

engaged with its respective microswitch.• The hook drive mechanism is not com-

pletely retracted.• The door is unsafe for takeoff.

A hollow neoprene seal surrounds the door-frame; the seal has holes to allow the entry ofpressurized cabin air, forming a positive sealaround the door.

Emergency ExitA hatch near the right rear of the cabin (Fig-ure 1-10) serves as an emergency exit for alloccupants. A latching mechanism is accessiblefrom inside and outside the cabin.

The inside latch handle, located at the topcenter of the window, is pulled inward tounlock. To open from the outside, depressinga PUSH button above the window releases ahandle which must then be turned in thedirection of the arrow stamped on the handle;then the hatch may be pushed inward.

Revision .01 1-7FOR TRAINING PURPOSES ONLY



Figure 1-10. Emergency Exit

Windows

Windshield

The windshield (Figure 1-11) is divided intotwo sections, the pilot’s and copilot’s halves,and is made up of three laminated layers ofacrylic plastic. The windshield is approxi-mately one inch thick. It is impact-resistant,heated or not, and was tested against 4-poundbird strikes at 350 knots.

Passenger Windows

The cabin windows (Figure 1-12), includingthe emergency exit window, are made up of twopanes of stretched acrylic plastic with an airspace between them. They are held apart andsealed air tight by a spacer.

Fuel SectionThe fuel section, located aft of the rear pres-sure bulkhead, contains the fuselage fuel cells.

As seen in Figure 1-5, the fuel section on 35models is different from that on 36 models. On36 models, the rear pressure bulkhead hasbeen moved forward, allowing for four blad-der cells rather than two, almost doublingfuselage fuel capacity.




Figure 1-11. Windshield

Figure 1-12. Windows Locations (Typical)

Figure 1-13. Tailcone Door

Revision .01 1-9

Tailcone SectionThe tailcone section extends aft from the fuelsection to the empennage. The tailcone entrydoor (Figure 1-13) is located at the bottom ofPage 1-8. The door is hinged at the forwardedge and drops down when released by quick-release thumb latches, allowing access to thebatteries, electrical components, fuel filters,fuel computers, refrigeration equipment,engine fire extinguishers, and hydrauliccomponents.

There is an optional light switch in the tailconeequipment compartment. If inadvertently lefton, it will be turned off by the door-clos-ing action.

There is no cockpit indicator to warn the pilotif the door is open.

WINGThe Learjet 35/36 has a swept back, can-tilevered, all metal wing (Figure 1-14) whichis mounted to the lower fuselage and joined to-gether at the fuselage. Most of the wing issealed to form an integral fuel tank.

Eight fi t t ings at taching the wings to thefu se l age a r e de s igned t o p r even t w ingdeflections from inducing secondary loads inthe pressurized fuselage. Ailerons are attachedto the rear spar at three hinge points. Thesingle-slotted Fowler flaps are attached to theinboard rear spar by tracks, rollers, and hinges.The spoilers are attached to the top of thewing surface by two hinges just forward of theflaps. The tip tanks are secured to the wing attwo attach points.

The Learjet 35/36 wing is fitted with either vor-tex generators or boundary layer energizers.Whichever is used, they function to delayairflow separation over the ailerons at highMach numbers.

Airplane Serial Nos. (SN) 35-002 through 35-278 and 36-002 th rough 36-044 ( i f no tretrofitted with AAK 79-10) employ two rowsfor vortex generators bonded to the upper wingsurface forward of both ailerons.

Subsequent serial-numbered airplanes andthose modified with AAK 79-10 incorporatea “soft-flight” wing modification, whichincludes:

• Three rows of boundary layer energiz-ers (BLEs) on each wing which performthe same function as vortex generators,but are more efficient. If any are miss-ing, MMO is reduced to 0.78 M1 (FC200)or .77 M1 (FC530).

• A full-chord stall fence on each wing,inboard of the aileron, which delaysdisruption of the airflow over the aileronat high angles of attack.




Figure 1-14. Learjet 35/36 Wing

• A stall strip, affixed to the inboard sec-tion of each wing leading edge, whichgenerates a buffet at high angle of attackto warn of an impending stall

• An aileron gap seal along the leadingedge of each aileron

EMPENNAGEThe high-T-tail empennage (Figure 1-15)includes a vertical stabilizer with an attachedrudder and a horizontal stabilizer with attachedelevators.

The swept back vertical stabilizer is formed byfive spars securely connected in the tailcone.It is the mounting point for the rudder andhorizontal stabilizer. At the lower leading edgeof the stabilizer is a dorsal fin which houses aram-air scoop. Later model airplanes have theoxygen bottle located within the dorsal fin.

The horizontal stabilizer is a swept back, fullspan unit, constructed around five spars. It isattached to the vertical stabilizer at two points:

• The center aft edge attaches to a heavy-duty hinge pin.

• The center leading edge attaches to ane lec t r ica l ly opera ted screwjack toprovide pitch axis trim.

AIRPLANE SYSTEMS

ELECTRICAL POWERSYSTEMSPrimary DC electrical power is provided by twoengine-driven generators. Secondary power issupplied by two 24-volt batteries. The airplanemay be equipped with a single or dual emer-gency battery system. The airplane also has thecapability of accepting DC power from aground power unit.

DC power is used by either two or three solid-state static inverters which, in turn, supply ACpower for equipment and instruments.

LIGHTINGInterior lighting is supplied for general cockpituse and for instrument illumination. Cabinlighting is supplied for the cabin overheadlighting, individual passenger positions, andcabin baggage compartment.

Exterior lighting includes the combinationlanding-taxi light on each main gear, navigationlights, anticollision lights, strobe lights, and arecognition light. A second recognition light andwing ice inspection light may be available.




Figure 1-15. Empennage

1-11

FUEL SYSTEMFuel is contained in integral wing tanks, tiptanks, and in a bladder cell fuselage tank justaft of the rear pressure bulkhead. The 36 modelhas a larger fuselage tank than the 35 model.

Fueling is accomplished through filler caps inthe top of each tip tank.

POWERPLANTThe Learjet 35/36 is powered by two GarrettTFE731 turbofan engines. The TFE731 is alightweight, two-spool, front fan-jet engine. Ithas a reverse-flow annular combustion cham-ber which reduces the overall length and resultsin more efficient combust ion and coolerexternal surfaces of the turbine section.

The low-pressure rotor consists of a four-stage,axial compressor and a three-stage, axialturbine rotating on a common shaft. The axial-flow fan assembly is located at the forwardend of the engine and is gear-driven by thelow-pressure rotor.

The high-pressure spool incorporates a single-stage, high-pressure centrifugal compressorand a single-stage axial turbine constructedas a single unit. The high-pressure spool drivesthe accessory section.

The high-pressure spool is located betweenthe low-pressure compressor and the low-pressure rotor shaft passing through its center.

The engines are mounted on external pylonsand are accessed by upper and lower nacellecovers. An access door on the outboard side of each nacelle is provided to check engine oil quantity.

Fire detectors are located in each engine nacelleand two engine fire extinguisher bottles in the tailcone.

Each engine supplies both high-pressure (HP)and low-pressure (LP) bleed air which is usedeither independently or in combination for

anti-icing, pressurization, cabin temperaturecontrol, and the Aeronca thrust reversers, if installed.

ICE AND RAIN PROTECTIONThe anti-icing systems use engine bleed air,electric heating, and alcohol.

Bleed air is used to heat the wing leading edge,the horizontal stabilizer leading edge, wind-shields, nacelle lips, and on some airplanes, theengine fan spinners. Bleed air is also used toremove rain from the windshield.

Electrically heated systems include pitot tubes,static ports, P2T2 sensors, and the stall warn-ing vanes.

An alcohol system is used for radome anti-icing and to back up the pilot’s windshieldbleed-air anti-icing.

AIR CONDITIONINGAND PRESSURIZATIONRegulated engine bleed air is diluted into thepressurized compartment through a heatexchanger where it is cooled by ram air fromthe dorsal inlet. Cabin temperature is regulatedby controlling the amount of bleed air allowedto bypass the heat exchanger.

Pressurization is regulated by controlling theamount of air that is exhausted from the cabin.Control is maintained by a pressurizationcontroller module and an outflow valve. Thecontroller module provides fully automaticcontrol of pressurization as well as manualmode. It ensures that the airplane is depres-surized on the ground, and causes automaticpressurization to occur on takeoff. Built-insafeguards prevent over-under pressurization.

A Freon refrigeration system and an optionalauxiliary cabin heater supplement the normalair conditioning system; they may be usedwhen the engines are not operating, provideda ground power unit is connected. Both systemsare completely independent of the bleed-airpressurization system.




HYDRAULIC POWER SYSTEMSThe hydraulic system supplies pressure forthe operation of the landing gear, gear doors,brakes, flaps, spoilers, and Dee Howard thrustreversers, if installed. A single reservoirsupplies fluid to the two engine-driven pumpsthrough fire shutoff valves.

An electric auxiliary pump can pressurize allsystems except the spoilers. It draws fluidfrom the same reservoir. The auxiliary supplyline is not affected by the fire shutoff valves.

LANDING GEAR AND BRAKESThe Learjet 35/36 has a retractable tricyclelanding gear which is electrically controlledand hydraulically operated.

An emergency air bottle, located in the rightside of the nose compartment, can be used toextend the landing gear or for emergencybraking, or both, in case of hydraulic orelectrical failure.

The self-centering nose gear has a single wheeland incorporates an electrical nosewheelsteering system which has variable authority,depending upon taxi speed.

Each main gear has dual wheels, each equippedwith multiple-disc brakes. Hydraulic brakingis controlled from either the pilot’s or copilot’sstation. A fully modulated antiskid systemprovide maximum braking performance whileprotecting against skids.

FLIGHT CONTROLSThe Learjet 35/36 uses manually actuatedpr imary f l ight controls . Pi lot inputs aret r ansmi t t ed v ia cab les , be l l c ranks , andpushrods to the ailerons, rudder, and eleva-tors. There are no hydraulic or electric powerboosts for these systems. Primary control trimsare electrically controlled and operated.

Secondary flight controls (spoiler/spoileronand flaps) are electrically controlled andhydraulically operated.

AUTOMATIC FLIGHT CONTROLSYSTEM (AFCS)The automatic flight control system (AFCS)includes a f l ight director, autopilot , and yaw dampers.

The flight director system generates roll andpitch commands by means of a single-cue V-bar display in the pilot’s attitude directorindicator. Programming and annunciation ofselected modes is accomplished on the AFCScontrol panel in the center glareshield.

The two-axis autopilot provides control of thero l l and p i t ch axes . When engaged , theautopilot responds to the flight director asprogrammed, or the pilot may elect to operatethe autopilot in a basic attitude-hold mode bycanceling all flight director modes, in whichcase the command bars are biased out of view.

Dual yaw dampers are installed for control ofthe yaw axis. Intended for full-time in-flightoperation, either yaw damper must be engagedafter takeoff. Functioning to dampen yaw andprovide turn coordination, the yaw damper(s)operate independently, whether or not the au-topilot is engaged.

PITOT-STATIC SYSTEMThe type of system used to supply pitot andstatic pressure to the pilot’s and copilot’sinstruments depends on whether the FC 200 orFC 530 automatic flight control system (AFCS)is installed.

FC 200 models use a conventional pitot-staticsystem consisting of one heated pitot tubemounted on each side of the nose section andtwo heated static ports flush-mounted on eachside of the nose compartment. The air data




1-13

sensor uses the copilot’s pitot line for pitot-pressure, while its static pressure is provided bytwo additional heated static ports installed onthe nose, forward of the windshield. An alternateunheated static port inside the nose compartmentis provided for the pilot’s static system.

FC 530 models use a Rosemount-designedpitot-static system which physically integratestwo static ports into each of two pitot tubes,one mounted on each side of the nose section.The air data sensor uses the copilot’s pitot andstatic lines.

An unheated static port is located on the rightside of the nose compartment to provide a staticsource for the pressurization control module.

OXYGEN SYSTEMThe oxygen system consists of the crew andpassenger distribution systems connected to ahigh-pressure oxygen storage cylinder locatedin the nose compartment on early 35 and 36models. On airplane SNs 35-492 and 36-051and subsequent, the cylinder is located in thevertical stabilizer.




2-i

CHAPTER 2ELECTRICAL POWER SYSTEMS

CONTENTS

Page

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

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

DC POWER............................................................................................................................. 2-2

Batteries ........................................................................................................................... 2-2

Generators ........................................................................................................................ 2-3

Ground Power .................................................................................................................. 2-5

Circuit Components ......................................................................................................... 2-6

Distribution ...................................................................................................................... 2-9

AC POWER........................................................................................................................... 2-15

Inverters ......................................................................................................................... 2-15

Controls.......................................................................................................................... 2-16

Indicators ....................................................................................................................... 2-16

Distribution .................................................................................................................... 2-17

EMERGENCY BATTERY.................................................................................................... 2-19

General........................................................................................................................... 2-19

Single Emergency Power System .................................................................................. 2-19

Dual Emergency Power System..................................................................................... 2-20

SCHEMATICS ...................................................................................................................... 2-20

QUESTIONS......................................................................................................................... 2-24




2-iii

ILLUSTRATIONS

Figure Title Page

2-1 Component Locations............................................................................................... 2-2

2-2 Battery Location ....................................................................................................... 2-2

2-3 Battery Switches....................................................................................................... 2-3

2-4 Battery Temperature Indicator.................................................................................. 2-4

2-5 Generator Location................................................................................................... 2-4

2-6 Generator Switches................................................................................................... 2-4

2-7 Generator Indicators ................................................................................................. 2-5

2-8 Ground Power Connector ......................................................................................... 2-6

2-9 Basic DC Distribution .............................................................................................. 2-7

2-10 Current Limiter Panel ............................................................................................... 2-6

2-11 Typical Circuit-Breaker Panels—SNs 35-002 through 35-201 and 35-205, and 36-002 through 36-040 (Not Incorporating AMK 78-13)............................... 2-10

2-12 Typical Circuit-Breaker Panels—SNs 35-202 and Subsequent, except35-205, 36-401 and Subsequent, and Airplanes Incorporating AMK 78-13 ......... 2-11

2-13 Equipment Powered by Battery Charging Bus and Generator Buses .................... 2-12

2-14 Main DC Bus Power .............................................................................................. 2-12

2-15 Essential DC Bus Power—SNs 35-002 through 35-201 and 35-205, and36-002 through 36-040 (Not Incorporating AMK 78-13)...................................... 2-14

2-16 Essential DC Bus Power—SNs 35-202 through 35-508, except 35-205,36-041 through 36-053, and Prior Airplanes Incorporating AMK 78-13.............. 2-14

2-17 Essential DC Bus Power—SNs 35-509 and Subsequent and 36-054 andSubsequent, and Prior Airplanes Incorporating AMK 85-1................................... 2-15

2-18 Inverter ................................................................................................................... 2-15

2-19 Inverter Switches.................................................................................................... 2-16

2-20 AC Bus Switch and AC Voltmeter ......................................................................... 2-17

2-21 AC Distribution ...................................................................................................... 2-18




2-22 Emergency Battery Controls and Indicators .......................................................... 2-19

2-23 Electrical System—SNs 35-002 through 35-205 and 36-002through 36-040 (Not Incorporating AMK 78-13).................................................. 2-21

2-24 Electrical System—SNs 35-202 through 35-204, 35-206 through 35-508,36-041 through 36-053, and Prior Airplanes Incorporating AMK 78-13.............. 2-22

2-25 Electrical System—SNs 35-509 and Subsequent, 36-054and Subsequent, and Prior Airplanes Incorporating AMK 85-1............................ 2-23

2-iv FOR TRAINING PURPOSES ONLY



INTRODUCTIONPrimary DC electrical power is provided by two engine-driven brushless DC generatorsrated at 30 volts, 400 amperes each. A single generator is capable of sustaining normalDC load. Secondary DC electrical power is supplied by two batteries. In the event of adouble generator failure, the airplane batteries will provide power for a limited periodof time.

A ground power unit can also provide the DC electrical power needed for systemoperation or engine starting. Electrical power for AC-powered equipment is providedby two (or an optional third) solid-state static inverters located in the tailcone. The invertersrequire DC input power for operation. An emergency battery is provided in case of totalairplane electrical failure to operate a standby attitude gyro and the landing gear andflaps. A second emergency battery may be installed at the customer’s option to poweradditional equipment such as an emergency communication radio, transponder, oremergency directional gyro.

GENERALThe electrical system incorporates a multiplebus system for power distribution intercon-nected by relays, current limiters, overload

sensors, and circuit breakers, which reactautomatically to isolate a malfunctioning bus.Manual isolation is also possible by openingthe appropriate circuit breakers.

#1 S

ERVO

SYSTEM

BATT HOT

BAT OFF

AC

GEN

#1 D

C

GEN

#1 E

NG

OIL PL

CHAPTER 2ELECTRICAL POWER SYSTEMS




The batteries are capable of operating the min-imum equipment for night instrument flightfor approximately 30 minutes in the event bothgenerators become inoperative. An emergencybattery is provided to operate an emergencyattitude gyro, and gear and flap systems in theevent of total airplane electrical system failure.

It is possible to power the entire DC and ACelectrical systems from the airplane batter-ies, an engine-driven generator, or GPU.

Figure 2-1 shows the major electrical powersystem component locations.

DC POWER

BATTERIESTwo batteries located in the tailcone (Figure2-2) provide the secondary source of DC power.

Each battery has a removable cover and a casewhich is vented and cooled by overboard con-nections. The batteries are of sufficient ca-pacity to supply normal ground electricalrequirements and may be used for engine start-ing when external power is not available.

Lead-Acid

Lead-acid batteries are enclosed in a plasticcase. Nickel-cadmium (nicad) batteries areenclosed in a stainless steel case. On airplaneSNs 35-341 and 36-050 and subsequentequipped with lead-acid batteries, a sump jarhas been added to contain any electrolytespillover. A sponge saturated in a baking sodaand water solution neutralizes the acid. AMK81-5A makes this installation available inearlier airplanes.




EMERGENCY BATTERY(IES)

CIRCUIT-BREAKER PANELS

PRIMARY, SECONDARY,AND AUX INVERTERS

GENERATOR BATTERIES CURRENT-LIMITER PANEL

Figure 2-1. Component Locations

Figure 2-2. Battery Location

Each battery is connected to its respectivebattery bus through a 20-amp current limiterfor hot-wired circuits.

Charging nicad batteries with a GPU is notrecommended.

Charging lead-acid batteries in the airplane isnot recommended because of poor GPU outputregulation.

ControlsTwo battery switches (Figure 2-3) are providedwhich connect the batteries in parallel to thebattery-charging bus when the switches are on.The switches are labeled “BAT 1” and “BAT2,” corresponding to the respective battery.Each switch is a two-position, ON–OFF switchwhich completes a ground circuit to close itsrespective battery relay in the ON position.(see Figure 2-13.)

The battery relays require approximately 16volts (minimum) from the respective battery.If either battery voltage is less than 16 volts,the respective battery relay will not close, inwhich case that battery cannot be connected

to the airplane electrical system for the purposeof operating electrical equipment (exceptequipment hot-wired to its battery bus), nor canit be charged by a GPU or the generators. Theairplane batteries are always connected inparallel (including during engine starts) whenboth batteries switches are on.

IndicatorsElectrical system indicators are grouped in acluster on the upper portion of the centerinstrument panel. A single DC VOLTS meter,connected to the battery charging bus througha 5-amp current limiter, indicates the highestvoltage input to the bus by batteries, genera-tors, or a GPU. To read individual battery volt-age , on ly one ba t t e ry a t a t ime may beconnected to the battery-charging bus withthe generators off and a GPU not connected.Airplane generators and GPUs normally putout a higher voltage than the batteries; there-fore, when either of these is powering thebat tery-charging bus, generator or GPUvoltage will be indicated (Figure 2-7).

Airplanes with nickel-cadmium batteries areequipped with battery temperature indicatorsand overheat warning light systems. Thetemperature indicators and warning lights areattached through two electrical connectors onthe face of each battery case to temperaturesensors and thermal switches on each battery.

A dual-indicating temperature gage is in-stalled on the lower portion of the copilot’sinstrument panel (Figure 2-4). Two red warn-ing lights labeled “BAT 140” and “BAT 160”are installed in the annunciator panel and il-luminate if either or both batteries reach 140to 160° F, respectively.

GENERATORSTwo engine-driven DC generators, one on eachengine (Figure 2-5), provide the primarysource of DC power. Each brushless genera-tor is rated at 30 VDC, 400 amperes.Cooling




Figure 2-3. Battery Switches

air is routed from a scoop on the engine nacelleto the associated generator. During normaloperation, both generators operate in parallelthrough the solid-state voltage regulatorslocated in the tailcone. As long as both batteryswitches are on, either generator charges bothbatteries through the associated 275-ampcurrent limiter. The generators supply DC powerto all DC-powered equipment on the airplane.

Generator voltage is regulated to 28.8 VDCfor lead-acid batteries and to 28.5 VDC fornicad batteries. On airplanes SNs 35-148 and

subsequent and 36-036 and subsequent, single-generator voltage is reduced as load increasesduring ground operation and any time a starteris engaged to limit amperage. This designfeature protects the 275-amp current limitersduring engine start. The generator control panel,located in the tailcone, contains relays for thebatteries, starters, GPU over-voltage control,and an equalizer circuit for load sharing.

Controls Two starter-generator switches are located onthe center switch panel. They are three-positionswitches labeled “GEN,” “OFF,” and “START”(Figure 2-6). In the GEN position, current isprovided to the generator field through theIGN & START circuit breaker, which auto-matically connects the generator bus, and theamber GEN caution light extinguishes.

Two generator reset buttons labeled “L GENRESET” and “R GEN RESET” on the centerswitch panel (Figure 2-6) provide for resettingthe generator in case of failure. Provided theGEN-OFF-START switch is in the GEN posi-tion, momentarily depressing the reset button




Figure 2-5. Generator Location

Figure 2-4. Battery Temperature Indicator

Figure 2-6. Generator Switches

resets the overvoltage relay, completes a powercircuit to the voltage regulator, and restores thegenerator to normal operation.

IndicatorsTwo AMPS meters (one for each generator)indicate the load, in amperes, being carried byeach gene ra to r (F igu re 2 -7 ) . The l oadindication is measured at the voltage regulator.

Generator voltage is displayed on the DCVOLTS meter.

An amber L or R GEN caution light on theglareshield annunciator panel (AnnunciatorPanel Section) illuminates if the associatedgenerator switch is turned off, if the generatorfails, or if the generator is tripped off by theovervoltage cutout relay.

GROUND POWERA ground power unit (GPU) can be connectedto the airplane through the receptacle located onthe left side of the fuselage, below the left en-gine (Figure 2-8). The receptacle connects GPUpower to the battery charging bus through apower relay controlled by an overvoltage circuit.The overvoltage circuit samples GPU voltageprovided through a control relay (Figure 2-9).At least one battery switch must be turned onto close the control relay, allowing the over-voltage circuit to sample GPU voltage, and, ifbelow 33 volts, the power relay closes to com-plete the GPU-to-battery charging bus connec-tion. The GPU should be regulated to 28 volts.Due to tower shaft torque limits, it must be lim-ited to 1,100 amps for engine starts. It shouldbe capable of producing at least 500 amps. IfGPU voltage exceeds 33 volts, the overvoltagecircuit causes the power relay to open, therebydisconnecting the GPU from the electrical sys-tem to prevent damage to voltage-sensitiveequipment.




Figure 2-7. Generator Indicators

CIRCUIT COMPONENTS

Current LimitersThroughout the electrical distribution system,various-size current limiters are placed atstrategic locations to prevent progressive totalelectrical failure. A current limiter is similarto a slow-blow fuse in that it will carry morethan its amp-rated capacity for short periodsof time. Extreme or prolonged overloadingwill cause a current limiter to fail, thus iso-lating that particular circuit and precludingprogress ive fa i lure of o ther e lec t r ica lcomponents. Current limiters are not resettable.When a current limiter has blown, it must be re-placed. It should be replaced if it shows dis-coloration or other signs of heating or over-loading. The current-limiter panel is locatedin the tailcone (Figure 2-10).

There are two current limiters (one on eachgenerator) that are not located on the current-limiter panels in the tailcone. A 10-amp currentlimiter on each generator is part of the paral-leling circuit.

Two types of current limiters are used in thesystem. The lower amperage current limiters(50 amps or less) are red and have a pin thatprotrudes if it is blown. The higher amperagecurrent limiters are made of a gray ceramicmaterial with a small window that allows visualinspection of current-limiter integrity.

There are two 275-amp current limiters in themain current-limiter panel which connect thegenerator buses with the battery-charging bus.On airplane SNs 35-002 through 35-147 and 36-002 through 36-035, testing of these current lim-iters is accomplished manually. On airplanesSNs 35-148 through 35-389, except 35-370,and 36-036 through 36-047, testing of the cur-rent limiters is accomplished using the rotarysystems test switch. For all of the aboveairplanes, AMK 80-17 provides two amberlights, one for each current limiter, which allowscontinuous monitoring. On airplane SNs 35-370, 35-390, and 36-048 and sub-sequent, asingle red CUR LIM light on the glareshield an-nunciator panel allows continuous monitoringof the 275-amp current limiters.




Figure 2-8. Ground Power Connector Figure 2-10. Current Limiter Panel

The 275-amp current-limiter annunciatorlight(s) are illuminated by 1-amp overloadsensors wired across the current- l imiterterminals. Failure of a current limiter resultsin a surge of current through the overloadsensor causing it to trip, thereby illuminatingthe light. In flight, it is important to know ifthe cur rent l imi te rs have b lown. On a l lairplanes with or without the current limiterannunciator light(s), current-limiter statusmay be determined by close observation ofvoltmeter and ammeter indications. If onlyone has failed, no difference will be noted one i ther ind ica to r s ince power f rom eachgenerator still flows to the battery charging busthrough the opposite current limiter. Failure ofboth current limiters, however, could be rec-ognized since the voltmeter will read battery

voltage (less than 25 volts). On airplanes priorto SNs 35-509 and 36-054 not modified byAMK 85-1, this failure will eventually resultin the depletion of the batteries since they arethe only source of power to the essential buses.The generators have been separated from theload of the essential buses and are now sup-plying power to only the main buses and thegenerator buses. This greatly reduced loadingwill be reflected by abnormally low ammeterreadings on both generators.

On airplane SNs 35-509 and 36-054 andsubsequent or earlier airplanes with AMK 85-1 installed, a failure of both 275 amp currentlimiters will not result in the separation of thegenerators from the essential buses. Genera-tor loads, therefore, will remain relatively




BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

DC VOLTS

010 30

50

AMPS0

100 200 300400

AMPS0

100 200 300400

LEFT ENGINESTART FUNCTIONS

RIGHT ENGINESTART FUNCTIONS**

OFFSTART START

OFF

REG REGLH

GENRH

GENL GENBUS

R GENBUS

10 A 10 A

275 A 275 A

L PWRBUS

R PWRBUS

FIELD FIELD

BAT CHGBUS

GEN

1. STARTER RELAYS2. STANDBY PUMP RELAY3. IGNITION POWER4. MOTIVE FLOW VALVE (SNs 35-002 THROUGH 35-057 AND 36-002 THROUGH 36-017)

*

OVERVOLT

CUTOUT

L S

TALL

WA

RN

EN

TRY

LTS

DO

OR

AC

TR

L BATBUS

20 AL

BAT

GNDPWRUNIT

20 ARBAT

R BATBUS

R S

TALL

WA

RN

Figure 2-9. Basic DC Distribution

normal. The generators have, however, beenseparated from the battery buses and batterycharging bus. Consequently, the batteries areno longer being charged and will be slowly de-pleted by electrical equipment which drawspower from either battery bus or the batterycharging bus. Battery condition should bemonitored using the DC voltmeter. On air-planes with the single CUR LIM annunciatorlight, if one limiter blows in flight, DC voltsand amps should be monitored closely sincethe CUR LIM light remains illuminated andwill not alert the pilot to subsequent failure ofthe other limiter.

RelaysRelays are used at numerous places throughoutthe electrical distribution system, particularlyin circuits with heavy electrical loads. Therelays function as remote switches to make orbreak power circuits. This arrangement allowsthe control circuit wiring to be a lighter gagesince less current is required to operate therelay. Relays control the power circuits for thebatteries, GPU, starters, generators, inverters,and left and right main buses. Instrument panelswitches or circuit breakers complete thecontrol circuits to operate the relays.

Overload Sensor

Overload sensors are used in the power circuitsto the left and right main buses and in thepower circuits to each inverter. These overloadsensors react thermally to electrical loads inexcess of their design capacity. In reacting,they electrically ground the relay controlcircuit causing the associated control circuitto trip, which causes the relay to open andbreak the power circuit. Once the overloadcondition has been removed, the overloadsenso r coo l s and r e se t s au toma t i ca l l y ;however, the control circuit breaker must bereset manually. The overload sensors in themain bus power circuits are rated at 70 amps,and the overload sensors for the inverter powercircuits are rated at 60 amps.

Circuit BreakersCircuit breakers are located on two circuit-breaker panels in the cockpit, one left of thepilot’s seat and one right of the copilot’s seat.On FC 200 AFCS airplanes, three additionalcircuit breakers, located under the pilot’s seaton the autopilot electric box, provide powerfor the autopilot flight director, and yawdamper annunciator lights. The DC circuitbreakers are the thermal type, and the ACcircuit breakers are the magnetic type. Am-perage ratings are stamped on the top of eachcircuit breaker. The circuit breakers are ar-ranged in rows according to the buses whichse rve them to s impl i fy the i so la t ion o findividual buses or circuits. Basically, all cir-cuit breakers in the top row (both sides) are onthe 115-VAC and 26-VAC buses; in the secondrow they are on the main DC buses (exceptthree which are power bus circuit breakers).Additionally, thrust reversers (if installed) arecontrolled by main bus circuit breakers whichare physically installed on the left and rightpanels, third and fourth rows. The third andfourth rows on airplane SNs 35-002 through35-201 and 35-205, and 36-002 through 36-040are on the DC essential bus. On airplane SNs35-202 and subsequent, except 35-205 and36-041 and subsequent, and earlier airplanesincorporating AMK 78-13, the third and fourthrows are on the essential A and B DC buses andsubsequent, and earlier airplanes incorporat-ing AMK 78-13, the third and fourth rows areon the essential A and B DC buses.

Circuit breakers located on the third and fourth rows, but not powered by the essentialbuses, are:

• L STALL WARN, DOOR ACTR andENTRY LTS (left battery bus items)

• R STALL WRN (right battery bus item)

• T/R EMER STOW and T/R POS IND(left main bus item—Aeronca)




2-9

• T/R CONT (r ight main bus i tem—Aeronca)

• T/R POWER and T/R CONT (left andright main bus items—Dee Howard)

See Figures 2-11 and 2-12 for typical repre-sentations of the circuit-breaker panels.

DISTRIBUTIONThe airplane uses a multiple bus, multipleconductor, electrical distribution system.Buses and major circuits are protected byrelays, current limiters, overload sensors, andcircuit breakers to preclude total failure. Thisarrangement also allows isolation of mal-functioning buses. All circuit breakers areaccessible to the crew during flight.

Battery BusesThe left and right battery buses are connectedto the left and right batteries, respectively,through 20-amp current limiters (Figure 2-9).The battery buses are always “hot,” providedthe battery quick-disconnects are connected.The battery buses supply power to the fol-lowing items:

• Left battery bus

• Left stall warning system

• Entry lights (step lights, baggage com-partment lights, and tailcone inspec-tion light)

• Door actuator motor

• Right battery bus

• Right stall warning system

Battery bus items must be turned off beforeleaving the airplane to prevent battery discharge.

Battery-Charging BusThe battery-charging bus enables the genera-tors or GPU to charge the batteries and is thecentral distribution point for the DC electricalsystem. It is powered by the batteries and GPUthrough their associated power relays by eithergenerator through the respective left and right275-amp current limiters (Figure 2-9).

One or both batteries can power the entireelectrical system for a limited period of time,with the exception of the Freon air conditionerand auxil iary heater. Because their highamperage requirement would quickly depletethe batteries, these items are isolated by anopen relay which will not close until a GPUor generator is on and operating.

On airplanes SNs 35-002 through 35-508 and36-002 through 36-053, when not incorpo-rating AMK 85-1, the essential buses areconnected directly to the battery charging bus(Figure 2-15 and 2-16).

On all airplanes, the following equipment isdirectly connected to the battery charging bus(Figure 2-13):

• DC VOLTS meter

• Freon air conditioner and auxiliary heater

• Recognition light(s)

• Auxiliary hydraulic pump

• Fuel flow indicating system

• Auxiliary inverter (if installed)

• Utility light (if installed)

• Primary pitch trim motor (FC-530 AFCSonly)

• Left and right engine starters







T/R EMERSTOWT/R POSNIND

PILOTPANEL

COPILOTPANEL

L ACBUS

AUXBUS

FLOODLT

PRI VM

PRIDIR GY

NOSESTEER

MACHTRIM

AIR DATA

AFCSPITCH

AFCSROLL

DMEREAD

SPARE

E.L.LTS

26 VACBUS

L OILPRESS

PRIRMI

NAV 1

ADF 1

ALTRIC

ANTISKID

TONEGEN

PRIDME

AUXCAB HT

CABLTS

R NAVCOMP

L TURBRPM

GALLEY

HT VALIND

PRIVERT GY

FLTDIR

AT TD

PRI YAWDAMP

L LDGTAXI LT

SQUATSW

WSHLDKT

ANTISKID

CAB HTAUTO

CABBLD

FREONCONT

RDNGLTS

NOSESTEER

AIRDATASEN

STROBELTS

NAVLTS

HFCOMM

DME 1

ADF 1

PRIINV

L IGN& START

L MAINBUS

L ESSBUS

AUD 1

COMM1

NAV 1

ATC 1

L FANRPM

WRNLTS

INSTRLTS

L PITHT

PRI FLTDIR

R NAVCONT

PRIAFCS

AFCSPITCH

AFCSROLL

AFCSYAW

S WRNHT FUEL

ITSN

L ITT

FUELCMPTR

FUSEVAL

OXYVAL

AIRBL

YAW

ROLL

PITCH

RAM AIRTEMP

L ICEDET

R ICEDET

OILTEMP

FUELQTY

FLAPS

GEAR

AUXCOM

FUELCMPTR

R ITT

RECOGLT

R FANRPM

FUELITSN

SPOILER

L STBYPMP

R STBYPMP

L JETPMP

R JETPMP VAL

L FWSOV

R FWSOV

L FIREEXT

R FIREEXT

L FIREDET

R FIREDET

L AIRIGN

R AIRIGN

ESSBUS TIE

R ESSBUS

AUD 2

COMM 2

NAV 2

ATC 2

UHF

WRNLTS

INSTRLTS

R PITHT

CABPRESS

VLFRCVR

SECAFCS

SECG/S

S WARNHT

AUXINV

TESTSYS

SEC PTRIM

SEC DME

R STALLWARN

TRCONT

HFCOMM

DMEREAD

R NAVSTBY

LSTALLWARN

ENTRLTS

DOORACTR

TAB FLAPPOSN

AC BUS

MAIN BUS

ESSENTIAL BUS

LEGENDBATTERY BUS

POWER BUS

FOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

SECFLTDIR

MAINBUS TIE

R MAINBUS

R IGNSTART

SECINV

EMERBAT 1

EMERBAT 2

ADF 2

R LDG &TAXI LT

BCNLTS

RADLTDM

RADAR

ALCPMP

STEREO

L NACKT

R NACHT

FUSLGPMP

FILL &XFER

HRMETER

TOILET

R TURBRPM

ALTM

CABINHT MAN

BATTEMP

SECDME

ADF2

NAV 2

SECRMI

R OILPRESS

26 VACBUS

FLTDR (MC)

FLTDR ATT

CMPTR

SECVERT GY

SECDIR GY

RADAR

SECVM

E.L.LTS

SPOILERON

R AUXBUS

R ACBUS

AC BUSTIE

SEC YAWCAMP

FLTDR HEAD

STAB &WING KIT

AIRDATASEN

Figure 2-11. Typical Circuit-Breaker Panels—SNs 35-002 through 35-201 and 35-205, and36-002 through 36-040 (Not Incorporating AMK 78-13)




PILOTPANEL

COPILOTPANEL

L ACBUS

FLOODLTS

PRI VM

NOSESTEER

MACHTRIM

AIR DATA

AFCSPITCH

AFCSROLL

RADAR

ELLTS

26 VACBUS

L OILPRESS

PRIRMI

NAV 1

ADF 1

ALTM

TONEGEN

AUXCAB HT

CABLTS

L TURBRPM

GALLEY

HT VALIND

HFCOMM

FLTDIR

ATTD

L LDG& TAXI

LTS

SQUATSW

WSHLDHT

L VACHT

CABBLOW

FREONCONT

RDNGLTS

NOSESTEER

AIRDATA

SEN

STROBELTS

NAVLTS

MODEPWR

DME 1

ADF 1

PRIINV

L IGN& ST

L MAINBUS

L ESSBUS

AUD 1

COMM1

NAV 1

ATC 1

L FANRPM

WRNLTS

AIRBLEED

VLFNAV

CLOCK

EMERLT

S WRNHT L AIR

IGN

OXYVAL

T/R EMERIND

T/R EMERSTOW

T/R POSN

IND

FUELJTSN

INSTRLTS

AFCSYAW

AFCSROLL

PRIAFCS

YAW

ROLL

PITCH

L PITOTHT

L ICEDET

R ICEDET

OILTEMP

FUELQTY

FLAPS

GEAR

AUXCOM

INSTRLTS

R AIRIGN

TESTSYSTEM

FUELJTSN

SPOILER

L STBYPMP

R STBYPMP

L JETPMP VAL

L FWSOV

R FWSOV

L FIREEXT

AUXINV

L FIREDET

R ESSB BUS

L ESSB BUS R ESS

A BUS

AUD 2

COMM 2

NAV 2

ATC 2

R FANRPM

WARNLTS

R FIREDET

R FIREEXT

CABPRESS

SECFLT DIR

SECAFCS

ADSPNEU V

S WARNHT

FUELCOMPT

SENSRHTR

UHFPHONE

T/R CONT

PASSINFO

FPA

L STALLWARN

ENTRLT

DOORACTS

AC BUS

MAIN BUS

ESSENTIAL "A" BUS

LEGEND

POWER BUS

BATTERY BUSESSENTIAL "B" BUS

FOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

MAINBUS TIE

R MAINBUS

R IGN& ST

SECINV

EMERBAT 1

EMERBAT 2

ADF 2

RECOGLT

R LDG &TAXI LT

BCNLTS

CABINTEMP

RADAR

ALCSYS

STEREO

ANTISKID

R NACHT

FUSLGPMP

FILL &XFER

BATTEMP

TOILET

R TURBRPM

ALTM

RADALTM

ADF 2

NAV 2

SECRMI

R OILPRESS

26 VACBUS

SEC FLTDIR

SAT/TAS

SECFLT DIR

SECVM

E.L.LTS

SPOIL -ERON

R AUXAC BUS

R ACBUS

AC BUSTIE

SEC F/DCMD

SEC F/DATTD

AIRDATA

SEN

L AUXAC BUS

PRIDIRGY

PRIVERT

GY

PRI YAW

DAMP

PRIFLT DIR

MACHA/S IND

PRIHDG

& CRS

LHMODVAL

L ITT

RAMAIR

TEMP

PRIFLTDIR

FUELCOMPTR

WINGINSP

LT

HF COMM

ESS BBUSTIE

R JETPMPVAL

TAB &FLAPPOSN

RPIOTOHT

ESS ABUSTIE

R ITT

CLOCK

SECPITCHTRIM

RSTALLWRN

STAB &WINGHT

RHMODVAL

SECDIRGY

SECVERTGY

SECRATEGYRO

SECYAWDAMP

SECHDG& CRS

MACHASIND

VLFHDGEXT

Figure 2-12. Typical Circuit-Breaker Panels—SNs 35-202 and Subsequent, except 35-205,36-041 and Subsequent, and Airplanes Incorporating AMK 78-13




LGEN BUS

RGEN BUSBAT CHG BUS

275 A275 A

20 A A

LLDGANDTAXI

LIGHT

LSTARTER

150 A30 A50 A10 A20 A20

FREONCOMP

MOTORANDAUX

HEATER

RRECOGLIGHT

HYDPUMP

FUELFLOW

IND

AUXINVERTER*

PRIPITCHTRIM

(FC-530)R

STARTER

RLDGANDTAXI

LIGHT

30 A 5 A 5 A

DC VOLTS

010 20

30

UTILITYLIGHT **

LRECOGLIGHT

*IF INSTALLED

BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

Figure 2-13. Equipment Powered by Battery Charging Bus and Generator Buses

Figure 2-14. Main DC Bus Power

DC VOLTS

010 20

30

LMAIN

RMAIN50 A

BUSTIEPOWER

RELAY

70 AOVERLOADSENSOR

L MAIN BUSCB

POWERRELAY

70 AOVERLOAD

SENSOR

R MAIN BUSCB

LPWR BUS

RPWR BUS

LGEN BUS

LGEN

RGEN

10 A

RGEN BUS

10 A

BAT CHG BUS275 A

CL275 A

CL

RBAT

LBAT

GPU

BATTERY POWERGENERATOR POWERGROUND POWER

LEGEND

2-13

Generator BusesThe left and right generator buses distributepower to the right and left main buses, theprimary and secondary inverters, and the leftand right power buses (Figures 2-14 and 2-21).On airplanes SNs 35-509 and subsequent, 36-054 and subsequent , and prior airplanesincorporating AMK 85-1, the generator busesalso power the respective essential A and Bbuses (Figure 2-17). On all airplanes, thelanding/ taxi l ights are connected to therespective generator bus (Figure 2-13). Thegenera to r buses can be powered by thebatteries, a GPU, or either generator.

Power BusesThe left and right power buses are poweredfrom the respective generator bus through a 10-amp current limiter. Each power bus providespower to three circuit breakers which controlthe respective engine starting and generatorfunctions, main bus power, and inverter power,as follows:

• The L or R MAIN BUS circuit breakercontrols the respective main bus powerrelay which connects the respectivegenerator bus to the main bus anytimeDC power is available (Figure 2-14).

• The L o r R IGN & START c i r cu i tbreaker: (1) controls the respectivestarter relays and standby fuel pumprelay and provides starting ignitionpower (through the thrust lever idleswitch) when the GEN-OFF-STARTswitch is in the START position; (2)provides power to the generator fieldwhen the switch is in the GEN position(Figure 2-13).

• The PRI or SEC INV circuit breakercontrols the respective inverter powerrelay which connects the respective gen-erator bus to the inverter when the in-verter switch is turned on (Figure 2-21).

The power bus circuit breakers are located atthe forward end of the respective circuit-breaker panels on what is generally referred toas the “main bus row,” those labeled “L” and“R IGN & START” and “PRI” and “SEC INV”are in no way related to, or affected by, the mainbuses; however, the L and R MAIN BUS circuitbreakers are, in that they control the relayswhich power the main buses (Figure 2-14).

Main DC BusesThe left and right main buses are poweredfrom the respective left and right generatorbuses through a 70-amp overload sensor anda power relay. The power relay is energizedc losed wheneve r t he re i s power on therespective power bus and the associated MAINBUS circuit breaker is closed. The left andright main buses are connected to each otherby a 50-amp MAIN BUS TIE circuit breakerwhich is normally closed for load equalization(Figure 2-14).

In the event of an overload on either main bus,the respective overload sensor causes theaffected MAIN BUS circuit breaker to trip.This deenergizes the power relay which opensto break the power circuits; then the MAINBUS TIE circuit breaker opens when it is forcedto accept the overload and cannot, resulting inautomatic isolation of the faulty bus.







BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

DC VOLTS

010 20

30

R ESSL ESS

REG

REG

LGEN

RGEN

LGEN BUS

275 ABAT CHG BUS

LBAT

GPU

RBAT

275 A

RGEN BUS

40 A

50 A

20A

40 A

50 A

Figure 2-15. Essential DC Bus Power—SNs 35-002 through 35-201 and 35-205, and 36-002through 36-040 (Not Incorporating AMK 78-13)

Figure 2-16. Essential DC Bus Power—SNs 35-202 through 35-508, except 35-205, 36-041through 36-053, and Prior Airplanes Incorporating AMK 78-13

BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND R ESS AL ESS A

40 A

50 A

40 A

50 A

DC VOLTS

010 20

30

BAT CHG BUS

LBAT

GPU

RBAT

REG

REG

LGEN

RGEN

LGEN BUS

275 A 275 A

RGEN BUS

40 A

50 A

40 A

50 A

20A

R ESS BL ESS B

20A

Essential DC BusesOne of three different bus configurations willapply to a given airplane, depending onp roduc t i on s e r i a l number and AMKapplicability (Figures 2-15, 2-16, and 2-17).

The left and right essential buses are poweredfrom the battery charging bus, or from therespective left or right generator buses (asapplicable) through a 50-amp current limiterand a 40-amp ESS BUS circuit breaker, and areconnected to each other by a 20-amp ESS BUSTIE circuit breaker which is normally closedfor load equalization.

In the event of an overload on one of theessential buses, the respective ESS BUS cir-cuit breaker opens, followed by the ESS BUSTIE circuit breaker which is forced to acceptthe overload and cannot, resulting in auto-matic isolation of the faulty bus. The currentlimiters provide backup for their respectiveESS BUS circuit breakers.

AC POWER

INVERTERSAlternat ing current to the AC electr icalinstruments and electronic equipment isprovided by two or three 1,000-VA, solid-state static inverters located in the tailcone(Figure 2-18). The third (auxiliary) inverter isoptional. During normal operation both, orall three, inverters are on and operate inparallel. It is recommended that the auxiliaryinverter, if installed be operated in conjunctionwith the primary and secondary inverters toextend inverter life.




BATTERY POWER

GENERATOR POWER

GROUND POWER

LEGEND

DC VOLTS

010 20

30

L ESS A R ESS A

20 A

L ESS B R ESS B

20 A

40 A

50 A

40 A

50 A

REG

REG

LGEN

LGEN BUS

RGEN BUSBAT CHG BUS

275A 275A

RGEN

LBAT

RBAT

GPU

Figure 2-17. Essential DC Bus Power—SNs 35-509 and Subsequent and 36-054 andSubsequent, and Prior Airplanes Incorporating AMK 85-1

Figure 2-18. Inverter

The primary and secondary inverters arepowered from the respective left and rightgenerator buses through a 60-amp overloadsensor and a power relay. The power relay isenergized closed whenever there is power onthe respective power bus, the associated PRIor SEC INV circuit breaker is closed, and theinverter switch is on (Figure 2-21).

In the event an inverter becomes overloaded,i.e., a shorted inverter, the respective overloadsensor causes the affected PRI or SEC INVcircuit breaker to trip. This energizes the powerrelay which opens to break the power circuit,resulting in automatic isolation of the faultyinverter. If installed, the auxiliary invertercircuits differ only in that they are poweredfrom the battery charging bus, and the powerrelay is controlled by the AUX INV circuitbreaker on the right essential bus (Figure 2-21).

Inverter output is 115-volt, 400-Hz, single-phase , a l te rna t ing current . Some of theinstruments and avionics require 26-VACpower. This 26-volt power is furnished by twostep-down transformers located in the cockpitjust aft of the circuit-breaker panels. These twotransformers take 115-VAC input from therespective 115-VAC buses and step it down to26-VAC output. Other components in thesystem include power relays, a parallelingbox, overload sensors, circuit breakers, andinverter lights on the glareshield for primary,secondary, and auxiliary inverters.

The paralleling box is the central control unitfor the AC electrical system. It incorporatesload equalizer and frequency synchronizer/phaser circuits through which it maintainsinverter load balance and frequency/phasesynchronization. It also causes illuminationof the associated annunciator lights for cer-tain malfunctions.

CONTROLS Two (or three) inverter switches, one for eachinverter (PRI, SEC, and optional AUX) areinstalled on the pilot’s lower instrument panel.The primary and secondary inverter switcheshave two positions, respectively, labeled“PRI–OFF” and “SEC–OFF.” The auxiliaryinverter switch, if installed, is labeled “ON-–OFF” (Figure 2-19). If the optional auxiliaryinverter is installed, an additional switchlabeled “L BUS–R BUS” is also installed.This switch is used to direct the auxiliaryinverter output to either the left or right AC busas needed. In case of an inverter failure, theauxiliary inverter will not automatically as-sume the operation of the failed inverter un-less the auxiliary inverter is turned on and theL/R BUS switch is properly positioned.

INDICATORSTwo red inverter warning lights labeled “PRIINV” and “SEC INV” a r e i n s t a l l ed onglareshield. If the optional auxiliary inverter




Figure 2-19. Inverter Switches

is installed, there is an amber AUX INV lighton the glareshield (Annunciator Panel section).

The corresponding inverter annunciator lightilluminates when inverter output is below 90volts or if bus load is less than 10 volt-amps.

The primary and secondary inverter lights alsoilluminate when the respective inverter switchis turned off. The AUX INV light, however, il-luminates only for an auxiliary inverter fail-ure with the switch turned on.

A single AC voltmeter (Figure 2-20) indicatesthe voltage on the LH or RH AC bus, de-pending on the position of the AC BUS switch.This two-position switch labeled “PRI–SEC”selects which bus the AC voltage is being mea-sured from. To check individual inverter volt-age, only the inverter to be checked should beturned on.

DISTRIBUTION

115-Volt AC Buses (L and R)Alternating current from the inverters isdistributed through the paralleling box to therespective left and right AC buses (Figure 2-21). Primary inverter output goes to the leftbus; secondary to the right bus. Auxiliaryinverter output (if installed) may be selectedto either the left or the right bus.

All circuit breakers on the left 115-VAC busare located on the top row of the left circuit-breaker panel. The right 115-VAC bus circuitbreakers are on the top row of the right circuit-breaker panel. The first circuit breaker on thetop row of the right panel is the 7 1/2-amp ACbus-tie circuit breaker. The second circuitbreaker on the top row of the right panel andthe first circuit breaker on the top row of theleft panel are the L and R AC BUS 10-amp busfeeder circuit breakers.

26-Volt AC Buses (L and R)Two step-down transformers draw 115-VACpower from the left and right 115-VAC buses,reduce the voltage output to 26 VAC, andconnect to the 26-VAC buses for equipmentrequiring 26-VAC power.

The 26-VAC BUS breakers are approximatelytwo-thirds of the way aft on the top row ofeach panel. All circuit breakers aft of the re-spective 26-VAC BUS breakers power 26-VAC equipment.




Figure 2-20. AC Bus Switch andAC Voltmeter




26 VAC

BUS

26 VAC

BUSTRANS TRANSL AC

BUSR ACBUS 2 A2 A SEC VM7.5 APRI VM

P S

010 50

AC VOLTS

30

L AUXAC BUS CB10 A 10 A10

A10A

L AC BUSCB

R AUXAC BUS CB

L BUSAUX INVPRI-SEC SW

R AC BUSCB

PARALLELING BOX

SECONDARYINVERTER

POWERRELAY

60 ASEC INVCB

SEC INVSW

R PWRBUS

L PWRBUS

10A

L GENBUS

R GENBUS

LGEN

RGEN

BAT CHGBUS

275A

275A

LBAT

RBAT

GPU

50 A

R ESS B

AUX INVCB

AUX INVON-OFF SW

60 A

POWERRELAY

10A

POWERRELAY

60 A PRI INVCB

PRI INVSW

PRIMARYINVERTER

AUXINVERTER *

R BUS

*OPTIONAL EQUIPMENT

LEGENDBATTERY POWER

GENERATOR POWER

GROUND POWER

INVERTER POWER

Figure 2-21. AC Distribution

EMERGENCY BATTERY

GENERALThe airplane may be equipped with either asingle (standard) or a dual (optional) emergencybattery system. The battery, or batteries, may beinstalled in either the nose compartment or thetailcone, and provide an emergency electricalpower source for selected equipment in the eventof total airplane electrical system failure.

Emergency batteries may be nickel-cadmium(nicad) or lead-acid. The nickel-cadmiumbattery is standard up to airplane SNs 35-462and 36-052. These airplanes are equipped withAC-powered standby attitude indicators, andthe battery packs contain a built-in inverter.On later airplanes, lead-acid batteries and aDC-powered standby attitude indicator arestandard. Lead-acid batteries may be retrofittedto earlier airplanes.

The nicad battery provides 25 VDC at 3.8 ampere-hours and contains an inverter andtransformer that provide 115 VAC and 4.6VAC. The lead-acid battery provides 24 VDCat 5.0 ampere-hours.

Both NO. 1 and NO. 2 emergency batteries receive a trickle-charge from the normal air-plane electrical system through the respec-tive EMER BAT 1 and EMER BAT 2 circuitbreakers on the right main bus anytime poweris on the bus. The trickle-charge is providedeven when the switches are off, but at a reducedrate. Controls and indicator location are illus-trated in Figure 2-22.

SINGLE EMERGENCYPOWER SYSTEMIf an airplane is equipped with a single emer-gency battery, the cockpit switch is labeled“EMER PWR.” There is an amber EMR PWRannunciator light on the pilot’s instrumentpanel that illuminates when power from theemergency battery is being used but the trickle-charge from the airplane electrical system hasbeen lost.

The EMER PWR switch has three positions—ON, STBY, and OFF. The emergency batterypowers the following equipment with theswitch in the ON or STBY position.

• ON

• Standby attitude indicator, indicatorlighting, and annunciator light

• Landing gear control circuits and gearposition lights

• Flap control circuits

• STBY

• Standby attitude indicator, indicatorlighting, and annunciator light




Figure 2-22. Emergency Battery Controlsand Indicators

With the switch in ON or STBY, the standbyattitude indicator is powered from the emer-gency battery. If power is available from theairplane electrical system, the emergency bat-tery is replenished as it provides power for thestandby attitude indicator. The other equipmenttied to the emergency battery is normally pow-ered by the airplane electrical system and ispowered by the emergency battery only whennormal electrical power is off or has failed.

Normally, the EMER PWR switch is in the ONposition. In the event of electrical system fail-ure, the EMR PWR light will illuminate whenpower from the associated emergency batteryis being used and is not receiving a trickle-charge. In the event of a total airplane electri-ca l sys t em fa i lu re , t he approved AFMrecommends that the EMER PWR switch beplaced in STBY until gear or flap operation isrequired to conserve battery life. Since only thestandby attitude indicator is powered in STBY,battery life is approximately 3 hours and 45minutes versus 30 minutes in the ON position.

DUAL EMERGENCYPOWER SYSTEMThe dual emergency battery system has twoswitches labeled “EMER PWR” (“BAT 1” and“BAT 2”). An amber EMR PWR annunciator

light for each power supply is installed on thepilot’s instrument panel. The applicable EMRPWR light will illuminate when power fromthe associated emergency battery is being usedand is not receiving a trickle-charge.

The BAT 1 switch operates the same systems asdescribed under Single Emergency Power Sys-tem. The BAT 2 switch has two positions—OFF and BAT 2. When turned on, power fromthe No. 2 emergency power supply is availableto illuminate the EMR PWR 2 light and oper-ate predetermined electrical equipment shouldthe normal electrical system fail. The auxiliarycommunication radio is the most common equip-ment powered by BAT 2; however, its installa-tion and use is optional. The pilot must turn offthe emergency battery switch(es) before leavingthe airplane. If the airplane power is turned offwith the emergency battery switch(es) in ON orSTBY, the emergency batteries will continue topower the emergency battery equipment andlose their charge.

SCHEMATICSThe following schematics (Figures 2-23, 2-24,and 2-25) are provided to show the three basicelectrical circuit configurations, differing onlywith respect to the essential buses.







26 VAC

BUS

26 VAC

BUSTRANS TRANSL AC

BUSR ACBUS 2 A7.5 A

AC VOLTS

010 30

50

P S

2 A

10 A 10 A 10 A 10 A

PARALLELING BOX

PRIMARYINVERTER

AUXINVERTER

SECONDARYINVERTER

LMAIN

RMAIN50 A

70 A

R PWRBUS AMPS

0100 200 300

400

60 A

R IG

N/S

TA

RT

/GE

N

10A

REG RHGEN

R GENBUS

275A

BAT CHGBUS

40 A

50 A

R ESS BL ESS B20 A

DC VOLTS

AC K

K

60 A

40 A

50 A

275A

REGLHGEN

L GENBUS

*OPTIONAL EQUIPMENT

10A

L PWRBUS

AMPS0

100 200 300400

60 A 70 A

L IG

N/S

TAR

T/G

EN

OVERVOLT

CUTOUT

L BATBUS

R BATBUS

LBAT

RBAT

GNDPWRUNIT

20 A20 A

*

LEGENDBATTERY POWER

GENERATOR POWER

GROUND POWER

INVERTER POWER

Figure 2-23. Electrical System—SNs 35-002 through 35-205 and 36-002 through36-040 (Not Incorporating AMK 78-13)




AC VOLTS

010 30

50

DC VOLTS

AC K

K

AMPS0

100 200 300400

26 VAC

BUS

26 VAC

BUSTRANS TRANSL AC

BUSR ACBUS2 A 7.5 A 2 A

50 A

P S

10 A10 A 10 A10 A

PARALLELING BOX

PRIMARYINVERTER

AUXINVERTER

SECONDARYINVERTER

R IG

NIS

TAR

T/G

EN

R PWRBUS

LMAIN

RMAIN

60 A

50 A 50 A

70 A60 A

L IG

N/S

TAR

T/G

EN

L PWRBUS

10A

L GENBUS

R GENBUSREG REGLH

GENRH

GEN

0

AMPS

100 200 300400

275A

275A

BAT CHGBUS

R ESS A

20 A

40 A 40 A

OVERVOLT

CUTOUT

*OPTIONAL EQUIPMENT

L BATBUS

R BATBUS

LBAT

RBAT

GNDPWRUNIT

20 A 20 A

BATTERY POWER GENERATOR POWER GROUND POWER

LEGENDINVERTER POWER

10A

70AR ESS B

L ESS A

L ESS B

20 A

60 A

*

Figure 2-24. Electrical System—SNs 35-202 through 35-204, 35-206 through 35-508, 36-041through 36-053, and Prior Airplanes Incorporating AMK 78-13




AC VOLTS

010 30

50

DC VOLTS

AC K

K

AMPS0

100 200 300400

26 VAC

BUS

26 VAC

BUSTRANS TRANSL AC

BUSR ACBUS2 A 7.5 A 2 A

50 A

P S

10A

10A10 A 10 A

PARALLELING BOX

PRIMARYINVERTER

AUXINVERTER

SECONDARYINVERTER

R IG

NIS

TAR

T/G

EN

60 A

R PWRBUS

10A

LMAIN

RMAIN

60 A

50A

60 A

L IG

N/S

TA

RT

/GE

N

L PWRBUS

10A

L GENBUS

R GENBUSREG REGLH

GENRH

GEN

AMPS0

100 200 300400

275A

275A

BAT CHGBUS

L ESS A

L ESS B R ESS B

R ESS A20 A

20 A

40 A 40 A

50 A

OVERVOLT

CUTOUT

*OPTIONAL EQUIPMENT

L BATBUS

R BATBUS

LBAT

RBAT

GNDPWRUNIT

20 A 20 A

BATTERY POWER GENERATOR POWER GROUND POWER

LEGENDINVERTER POWER

Figure 2-25. Electrical System—SNs 35-509 and Subsequent, 36-054 andSubsequent, and Prior Airplanes Incorporating AMK 85-1




QUESTIONS1. The DC voltmeter indicates:

A. Battery voltage onlyB. Generator voltage onlyC. Voltage on the battery busesC. Voltage on the battery charging bus

2. When a GPU is used for engine start, theoutput value should be:A. Regulated to 24 voltsB. Regulated to 28 volts and limited to

1,100 ampsC. Regulated to 33 ± 2 voltsD. Regulated to 28 volts and limited to

500 amps

3. The buses that the airplane batteries powerare:A. Battery buses onlyB. Battery and battery-charging buses

onlyC. All buses except the 115 VACD. All buses including AC if an inverter

is on

4. A generator failure is indicated when:A. One ammeter indicates less than 25

ampsB. The GEN switch is in the ON position

and the GEN light remains illumi-nated after activating RESET.

C. The GEN light is extinguished.D. The DC voltmeter reads less than 28

volts.

5. If airplane electrical power fails and theEMER PWR BAT 1 switch is ON, thesystems powered by the emergency bat-tery are:A. Standby attitude gyro onlyB. Flaps and gear onlyC. Flaps, gear, and spoilerD. Standby attitude indicator, gear, and

flaps

6. If both 275-amp current limiters fail inflight:A. The essential buses will remain pow-

ered by the airplane batteries.B. The essential buses will remain pow-

ered by the generators.C. The battery-charging bus will fail im-

mediately.D. Both inverters will fail.

7. Illumination of a PRI or SEC inverterlight indicates:A. The inverter is operating.B. The inverter output is less than 90

VAC, or there is less than 10 volt-amps draw on the inverter.

C. The inverter switch is off.D. B and C

8. The AC voltmeter will indicate:A. Right AC bus voltage with the AC

BUS switch in PRIB. Left AC bus voltage when the AC BUS

switch is in PRIC. The AC loadD. The voltage on the 26-VAC buses




9. If an overload sensor shuts off power toa main bus, power may be restored by:A. Resetting the control circuit breaker

after the overload sensor resetsB. Changing the overload sensorC. Automatic action after the current

limiter coolsD. Automatic action after the overload

sensor cools

10. To unlock the entrance door when thebatteries are dead:A. Plug in a GPU and use a key.B. Plug in a GPU with 33 ±2 VDC or

less on the small pin and use a key.C. Remove both batteries for charging

and reinstall.D. Enter airplane through the emergency

hatch, place the emergency batteryswitch to ON, and activate the interiordoor switch.

11. With a dual-generator failure in flight, the airplane batteries will support theminimum night IFR equipment load forapproximately:A. 60 minutesB. 2 hours 45 minutesC. 30 minutesD. 30 minutes with fully charged emer-

gency batteries and emergency BAT1 in standby position

12. Inverter output is:A. 115 VAC, 400 HzB. 115 VAC and 26 VAC, 400 HzC. 26 VAC, 400 HzD. 115 VAC and 26 VAC, 1,000 Hz

13. The approved AFM recommends that aGPU be used for engine start when the am-bient temperature is:A. 10° C or belowB. 0° F or belowC. 15° F or belowD. 32° F or below

14. When either primary or secondary in-verter light illuminates, the first step ofcorrective action is:A. Pull the AC bus-tie circuit breakerB. Turn the respective inverter switch

off.C. Check for open INV or AC BUS cir-

cuit breaker(s).D. Reduce the load on the failed AC bus.

3-i

CHAPTER 3LIGHTING

CONTENTS

Page

INTRODUCTION ................................................................................................................... 3-1

GENERAL............................................................................................................................... 3-1

INTERIOR LIGHTING........................................................................................................... 3-2

Cockpit Lighting .............................................................................................................. 3-2

Cabin Lighting ................................................................................................................. 3-4

Emergency Lighting......................................................................................................... 3-6

EXTERIOR LIGHTING ......................................................................................................... 3-8

General ............................................................................................................................. 3-8

Landing-Taxi Lights......................................................................................................... 3-9

Recognition Light ............................................................................................................ 3-9

Strobe Lights.................................................................................................................. 3-10

Navigation Lights........................................................................................................... 3-10

Anticollision Lights ....................................................................................................... 3-10

Wing Inspection Lights .................................................................................................. 3-11

Tailcone Area Inspection Light (Optional) .................................................................... 3-12

QUESTIONS......................................................................................................................... 3-13




3-iii

ILLUSTRATIONS

Figure Title Page

3-1 Interior Lighting Controls ........................................................................................ 3-2

3-2 Cockpit Map Lights.................................................................................................. 3-4

3-2A Reading Lights (Typical).......................................................................................... 3-4

3-3 Overhead Lights Control (Typical) .......................................................................... 3-4

3-4 Advisory Lights and Controls .................................................................................. 3-5

3-5 Emergency Cabin Door Light, Emergency Exit Light, and Wing Inspection/Egress Light ........................................................................... 3-6

3-6 Emergency Lights Control ....................................................................................... 3-7

3-7 Exterior Lighting Locations ..................................................................................... 3-7

3-8 Exterior Lighting Controls ....................................................................................... 3-8

3-9 Landing-Taxi Lights ................................................................................................. 3-8

3-10 Recognition Light..................................................................................................... 3-9

3-11 Strobe and Navigation Lights................................................................................... 3-9

3-12 Anticollision Lights................................................................................................ 3-10

3-13 Wing Ice Inspection Light ...................................................................................... 3-11

3-14 Wing Ice Inspection Light Control......................................................................... 3-11

3-15 Tailcone Inspection Light Switches ....................................................................... 3-12




INTRODUCTIONAirplane lighting is divided into interior, exterior, and emergency (if installed) lightingpackages. Interior lighting provides illumination of both the cockpit and cabin areas undernormal conditions. The cockpit area is provided with general illumination and specificlighting for instruments and map reading. Cabin area lighting provides illumination forthe standard warning signs and specific area illumination for passenger safety and con-venience. Exterior lighting consists of navigation, landing-taxi, anticollision, recogni-tion, and strobe lights. An optional tailcone area inspection light and two lightingpackages to illuminate the wing are available.

An emergency lighting system may be installed as optional equipment, serving to illu-minate the cabin interior and egress points in the event of airplane electrical power fail-ure. There are two basic configurations, depending on airplane serialization.

GENERALCockpit lighting consists of the instrumentlights, floodlight, electroluminescent lighting,and map lights, all adjustable for intensity with

rheostat controls. The electroluminescent light-ing illuminates the lettering on the variousswitch panels, pedestal, and circuit-breakerpanels. Optional map lights may be installed,and consist of a flexible-neck light located

EXIT

CHAPTER 3LIGHTING




on each pilot’s sidewall panel or one of twooverhead light installations, depending on air-plane serialization. Cabin lighting consists ofeight fluorescent upper center-panel lights(four on 36 models), two door entry lights, bag-gage compartment lights, individual readinglights, and the no smoking/fasten seat beltssign. The optional emergency lighting sys-tems illuminate the fluorescent upper center-panel lights, and other lights at the exits.Exterior lights include landing-taxi lights,wing and tail navigation lights, anticollisionbeacons, one (or two optional) recognitionlight, and high-intensity strobe lights. A winginspection and egress light (which may be partof the emergency lighting option) illuminatesthe right wing area to check for ice accumu-lation, and for emergency egress. An optionalwing ice inspection light is available

on late models which is not a part of the emer-gency lighting system. An optional light insidethe tailcone does not require airplane batteryswitches to be on for operation.

INTERIOR LIGHTING

COCKPIT LIGHTING

GeneralSome cockpit lighting systems use both in-candescent and fluorescent bulbs and, conse-quently, require both AC and DC power.Controls for lighting are either on the deviceor as illustrated in Figure 3-1.




Figure 3-1. Interior Lighting Controls

Instrument Panel FloodlightsA single fluorescent light tube is installedunder the glareshield to illuminate the in-strument panel. It is controlled by the FLOODrheostat switch on the pilot’s side panel (Fig-ure 3-1). Electrical power required is 115 VACsupplied through the FLOOD LT circui tbreaker on the left (primary) AC bus.

Instrument LightsIncandescent lighting is installed for pilot,copilot, and center instrument panels, pedestalindicators, and the magnetic compass. Thelights are controlled with the INSTR rheostaton the pilot’s side panel and both INSTR andPEDESTAL rheostats on the copilot’s sidepanel. DC power for the lights is suppliedthrough the respective INSTR LTS circuitbreakers on the respective left and right es-sential buses.

Pilot’s INSTR LightsThe pilot’s INSTR rheostat provides lightingcontrol for the pilot’s flight instruments, en-gine instruments, clock, electrical indicators,oil temperature indicators, altitude indicator,and the radar edge lighting.

Copilot’s INSTR LightsThe copilot’s INSTR rheostat provides light-ing control for the copilot’s flight instruments,the magnetic compass, cabin temperature in-dicator, BAT TEMP indicator (if installed),landing gear control panel, EMERGENCYAIR and HYDRAULIC PRESSURE indica-tors, and the pressurization control panel.

PEDESTAL LightsThe PEDESTAL rheostat on the copilot’s sidepanel provides lighting controls for the flightdirector panel and the pedestal.

Switch Panel LightingElectroluminescent lighting is used to illu-minate the lettering on all switch panels andboth circuit-breaker panels.

Electroluminescent (EL) lighting uses 115VAC supplied through the EL LTS circuitbreakers on the left (primary) and right (sec-ondary) AC buses, respectively. The lights arecontrol led wi th the EL PANEL rheosta tswitches on the pilot’s and copilot’s side pan-els, respectively.

EL PANEL Rheostat(Pilot’s Sidewall)The EL rheostat controls all edge lighting onthe switch panels to the left of a line runningvertically between the radar and radio panels.This control includes dimming for the audiocontrol panel, the left circuit-breaker panel andthe pilot’s microphone jack panel.

EL PANEL Rheostat (Copilot’s Sidewall)The EL PANEL rheostat controls all edgelighting on switch panels to the right of the ver-tical line established in the preceding para-graph. It also controls lighting for the copilot’smicrophone jack panel, audio panel, and theright circuit-breaker panel.

Maps Lights (Optional)When installed, the airplane may have one ormore of three different map light options: (1)flexible neck light on each pilot’s sidewallpanel, with an ON-OFF rheostat for intensitycontrol (Figure 3-1); (2) a reading light andgasper assembly, installed in the cockpitheadliner for each pilot , incorporating arheostat for light intensity adjustment and alight pattern adjustment lever (Figure 3-2);(3) a dome light assembly, mounted on eachside of the headliner just forward of the upperair outlets incorporating a rocker-operated




switch (labeled ON-REMOTE) with an unla-beled center off position (Figure 3-2) and aswivel-mounted light.

All installations are powered through theINSTR LTS circuit breakers on the left andright essential buses. In the REMOTE position,the dome lights are powered from the ENTRYLT circuit breaker on the left battery bus.

CABIN LIGHTING

GeneralPassenger compartment lighting consists ofreading lights, overhead lights, entry lights,no smok ing / fa s t en sea t be l t s i gns , andrefreshment cabinet lights.

Reading LightsThe reading lights are mounted in the uppercenter panel above the seats on each side of




Figure 3-2. Cockpit Map Lights

Figure 3-2A. Reading Lights (Typical)

Figure 3-3. Overhead Lights Control(Typical)

the cabin. There are individual switches foreach light. The lights are adjustable for posi-tion and use DC power supplied through theRDNG LTS circuit breaker on the left main bus(Figure 3-2A).

Overhead LightsThe cabin overhead light system consists offour (three on 36 models) fluorescent lightsrecessed in each side of the upper panel, acabin lights power supply, a three-positionswitch, a cabin lights relay assembly, and aCAB LTS circuit breaker on the left main bus.

Normally, the lights are controlled with thethree-position switch located on the leftservice cabinet forward of the entry door (Fig-ure 3-3).

In the event of cabin depressurization, thelights automatically illuminate full brightwhen the cabin altitude reaches 14,000 feet.On airplanes equipped with the optionalemergency lighting system, three overheadlights illuminate automatically in the eventof airplane electrical power failure.

Entry LightsThe entry light system consists of a STEPLIGHT switch and light on the left servicecabinet forward of the entry door (Figure 3-3),and another directly over the door opening.Power from the left battery bus is suppliedthrough the ENTRY LT circuit breaker on theleft battery bus; therefore, the lights are oper-able when the airplane BAT switch is in OFF.

Baggage Compartment LightsTwo lights are installed in the aft baggagecompartment and, on 36 model airplanes, onelight is installed in the forward baggage com-partment. Aft baggage compartment lights arecontrolled by a switch on the left servicecabinet forward of the entry door (Figure 3-3) and are powered through the ENTRY LTcircuit breaker on the left battery bus. Thefo rward baggage compar tmen t l i gh t i scontrolled by a switch on the forward end ofthe upper center panel.




Figure 3-4. Advisory Lights and Controls



Passenger Advisory LightsThe no smoking/fasten seat belts advisory lightsystem consists of two fixtures (one on 36 mod-els) (Figure 3-4), a switch on the center switchpanel, and the RDNG LT circuit breaker on thele f t ma in bus . The swi tch has th reepositions—NO SMOKING/FASTEN SEATBELT–OFF–FASTEN SEAT BELT. When theswitch is moved from OFF to either of the otherpositions, an audible tone sounds and the ap-propriate symbols illuminate. A RETURN TOSEAT light (if installed) in the lavatory is apart of the advisory light system. Location ofthe fixtures varies with cabin configuration.

Cabinet LightsThe cabinet light system varies with cabinconfiguration and consists of various lightswithin the refreshment cabinet, microswitchesactuated by doors or drawers, power sup-plies, and a circuit breaker on the right es-sential bus.

EMERGENCY LIGHTING

Cabin Interior and WingInspection and Egress LightsIf these lights are installed, the airplane isequipped with two nickel-cadmium (nicad)battery power supplies and a control module,which function to illuminate selected areas

automatically in the event of airplane DCpower failure. An emergency light is installedin the upper cabin door (F igure 3-5) toi l luminate the lower cabin door and theimmediate door area. A second light illumi-nates the emergency exit window area. Anexterior wing inspection/egress light option-ally installed on the right side of the airplaneis adjacent to the emergency exit window andilluminates the exterior egress area. The cabinupper-center panel (fluorescent) lights illu-minate the cabin interior. When activated, oneof the power supplies turns on the cabin upper-center panel lights, while the other powersupply turns on the upper cabin door light,the emergency exit light, and the wing in-spection/egress light. The nicad battery packsare charged through the EMER LTS circuitbreaker on the right essential bus.

Control ModuleThe EMER LIGHT TEST switch on the pilot’s(or center) switch panel (See Figure 3-6)provides the test function for the system andfor automatic illumination of the emergencylights in the event of an interruption of normalDC electrical power. The switch has threepositions: TEST, ARM and DISARM. Settingthe switch to TEST simulates a failure ofnormal DC electrical power and illuminates theupper cabin entry door light, the emergencyexit light, and the cabin overhead fluorescentlights. Setting the switch to ARM will armthe system to illuminate the emergency lights


Figure 3-5. Emergency Cabin Door Light, Emergency Exit Light, and WingInspection/Egress Light

in the event of a failure of normal DC electricalpower. Setting the switch to DISARM iso-lates the emergency lights from the emergencybatteries. The switch should be set to ARMprior to takeoff. If the switch is in the DISARMposition and at least one BAT switch is on, theamber l igh t ad jacen t to the swi tch wi l lilluminate to remind the pilot that the switch

should be set to ARM. The switch should beset to DISARM prior to set t ing the BATswitches to OFF.

The WING INSPECTION light switch (in-cluded as part of the emergency lightingsystem), located adjacent to the EMERGLIGHT TEST–ARM–DISARM switch, may beused independently of the rest of the emer-gency lighting system to visually check for iceaccumulation on the wing leading edge. Turn-ing the switch on illuminates the exterior winginspection/egress light.

The EMERGENCY LTS. switch on the leftservice cabinet near the entry door providesa means for manual i l luminat ion of theinterior emergency lights. When the switchis set to EMERGENCY LTS., the upper cabinentry door light, the emergency exit light, thecabin overhead fluorescent lights, and thewing inspection/egress light (if installed)will illuminate. For normal operation, the




Figure 3-6. Emergency Lights Control

RECOGNITION LIGHT LANDING–TAXI LIGHTS WING ICE INSPECTION LIGHTANTICOLLISION LIGHT

NAVIGATION LIGHT STROBE LIGHT NAVIGATION LIGHT ANTICOLLISION LIGHT

RECOGNITION LIGHT LANDING–TAXI LIGHTS WING ICE INSPECTION LIGHTANTICOLLISION LIGHT

NAVIGATION LIGHT STROBE LIGHT NAVIGATION LIGHT ANTICOLLISION LIGHT

Figure 3-7. Exterior Lighting Locations

switch should be set to OFF, allowing auto-matic illumination of the emergency lightsin the event of a failure of the normal electricalsystem.

EXTERIOR LIGHTING

GENERALThe exterior lighting systems consist of thelanding-taxi lights, navigation lights, anti-collision lights, recognition light(s), strobelights, and an optional wing ice inspection




Figure 3-8. Exterior Lighting Controls

Figure 3-9. Landing-Taxi Lights

light (Figure 3-7). The exterior lighting con-trols are shown in Figure 3-8.

LANDING-TAXI LIGHTSThe landing light system consists of one 450-watt lamp mounted on each main landing gearstrut (Figure 3-9), one 20-amp current limiterfor each side in the current-limiter panel,relays, dimming resistors, and the L and RLDG LT switches on the center switch panel.The L and R landing light switches have threepositions: OFF, TAXI and LDG LT. DC powerto operate the relays comes from the left andright main buses, respectively.

Setting the L or R LDG LT switch to TAXIcloses a relay which shunts DC power from therespective generator bus through a resistorwhich limits current to the lamp element. Mov-ing the switch to LDG LT closes a secondrelay, allowing current flow to bypass theresistor, thereby increasing the brightness ofthe lamp. The 20-amp current limiters protectthe power circuits between the respectivegenerator bus and lamp filament.

Regardless of switch position, the lights willnot illuminate unless the respective landinggear down-and-locked switches are closed andprovide a ground. It is recommended that thelights be operated in the L and R LDG LTmodes as sparingly as possible. Lamp servicelife is shortened in the LDG LT mode becauseof the higher current flow.

RECOGNITION LIGHTA 250-watt recognition light is installed in thenose of the right tip tank (Figure 3-10). Thelight is controlled with the RECOG LT switchon the copilot’s lighting control panel. Whenturned on, DC power, applied through theRECOG LT ci rcui t breaker on the r ightessential bus, closes a control relay andconnects power through a 30-ampere currentlimiter to the light. A second recognitionlight may be installed in the let tip tank as op-tional equipment.




Figure 3-10. Recognition Light Figure 3-11. Strobe and Navigation Lights

STROBE LIGHTSThe strobe light system consists of a strobelight mounted inside each navigation lightfixture, a power supply for each strobe (Fig-ure 3-11), a STROBE LT switch on the copi-lot’s lighting control panel, a DC STROBE LTScircuit breaker on the left main bus, and atiming circuit module that causes the strobesto flash. Each power supply is protected by aninternal 3-ampere fuse.

NAVIGATION LIGHTSThe navigation light system consists of onelamp in the outboard side of each tip tank,two lamps in the upper aft tail fairing, a NAVLT switch on the copilot’s lighting controlpanel, and a NAV LTS circuit breaker on theleft main bus.

All three navigation lights are controlled bythe NAV LT switch. Additionally, setting theNAV LT switch to ON automatically dimsmost instrument panel and pedestal “peanut”lights, and activates the landing gear positionlight dimmer rheostat.

ANTICOLLISION LIGHTSAnticollision lights are installed on top of thevertical stabilizer and on the bottom of thefuselage (Figure 3-12). The lights are con-trolled by a BCN LT switch on the copilot’slighting control panel. Each light is a dual-bulb light and each bulb oscillates 180 degreesat 45 cycles per minute. The beam is concen-trated by an integral lens, and an illusion of 90flashes per minute occurs due to the oscillation.The lights operate on DC power through theBCN LT circuit breaker on the right main bus.




Figure 3-12. Anticollision Lights

WING INSPECTION LIGHTSTwo separate installations are designed toilluminate the wing area for signs of ice (Fig-ure 3-13). Both are optional. One light is in-stalled on the right side of the fuselage adjacentto the lower forward corner of the emergencyexit window. This light is designed to illumi-nate the leading edge of the right wing and ad-ditionally serves as an illumination source foremergency egress over the wing. The light isdesignated the “wing inspection and egresslight,” and may be installed as an integral part of the earlier emergency lighting system,or as a selected option not involving theemergency lighting system. In either case, asecond option may include a second lightinstalled on the left side of the fuselage directlyopposite the right-hand light, serving as aninspection light for the left wing. The WINGINSPECTION control switch is located on theemergency lighting panel or on the instrumentpanel (Figure 3-6).




Figure 3-14. Wing Ice Inspection Light Control

Figure 3-13. Wing Ice Inspection Light

On airplane SNs 35-416 and 36-048 andsubsequent, another option provides a lightinstalled in the fuselage below the copilot’swindow. It is designed to illuminate a blackspot on the right wing leading edge. A coveringof ice obscures the spot, enabling ice detectionat night when the light is turned on. This lightis designated the “WING INSP light” (Figure3-13) and is operated by a push-button switchlocated forward of the rheostats on the copilot’sright side panel (Figure 3-14).

TAILCONE AREA INSPECTIONLIGHT (OPTIONAL)When installed, this light is located in thetailcone, directly above the entry door. An

ON—OFF switch is positioned inside the doorat the forward left side of the opening. Amicroswitch installed on the forward rightside of the opening breaks power to the lightwhen the door is closed (Figure 3-15). Powerfor operating the tailcone area inspection lightoption is provided by the left battery busthrough the ENTRY LT circuit breaker (pilot’scircuit-breaker panel), which permits operationof the light without turning airplane poweron. However, on some airplanes, the light ispowered by the battery charging bus througha 5-amp current limiter, in which case anaircraft battery must be turned on to operatethe light.




Figure 3-15. Tailcone Inspection Light Switches




1. The instrument panel flood light controlis located:A. On the lightB. Just forward of the warning panelC. On the pilot’s side panelD. On the copilot’s side panel

2. The cockpit map lights are controlled:A. With an ON-OFF switch on the copi-

lot’s side panelB. With the overhead map light rheostat

on the copilot’s side panelC. With an integral rheostat and a pattern

leverD. Automatically, relative to ambient

light

3. The cabin overhead light control switchesare located on the:A. Right forward refreshment pedestalB. The entrance door thresholdC. Left forward service cabinetD. Light assembly

4. When a cabin overhead light switch isturned on, first select:A. ONB. OFFC. DIMD. BRT

5. The lights that are illuminated by theemergency lighting system are the:A. Instrument panel floodlights and elec-

troluminescent lightsB. Cabin overhead lights, wing egress

light, and emergency exit lightC. Navigation lightsD. Strobe lights

6. The emergency lighting switch positionused during normal operation is:A. DISARMB. ARMC. TESTD. EMER LT

7. The lights that come on when cabin alti-tude reaches 14,000 feet or higher are the:A. Passenger advisory lightsB. Lavatory lightsC. Cabin overhead panel lightsD. Reading lights

8. The wing ice inspection light switch (ifinstalled) is located on the:A. Pilot’s switch panelB. Light assemblyC. Overhead panelD. Copilot’s right sidewall

9. The lights that require inverter power are the:A. Cabin overhead lightsB. FLOOD and EL lightsC. INSTR lightsD. NAV lights

10. The lights that can be operated with theairplane batteries turned off are the:A. Entry lights and baggage compart-

ment lightB. Overhead lightsC. Passenger advisory lightsD. Reading lights

QUESTIONS

4-i

CHAPTER 4MASTER WARNING SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................... 4-1

GENERAL............................................................................................................................... 4-1

GLARESHIELD ANNUNCIATOR LIGHTS......................................................................... 4-2

MASTER WARNING LIGHTS .............................................................................................. 4-2

TEST........................................................................................................................................ 4-2

INTENSITY CONTROL......................................................................................................... 4-3

BULB CHANGE ..................................................................................................................... 4-3

ILLUMINATION CAUSES .................................................................................................... 4-3

QUESTIONS ........................................................................................................................... 4-6




4-iii

ILLUSTRATION

Figure Title Page

4-1 Test Switch ............................................................................................................... 4-2

TABLE

Table Title Page

4-1 Annunciator Illumination Causes............................................................................. 4-3




INTRODUCTIONThe master warning system provides a warning for airplane equipment malfunctions,anindication of an unsafe operating condition requiring immediate attention, and an indi-cation that a system is in operation.

GENERALThe warning light system incorporates twohorizontal rows of red, amber, and green lights(see Annunciator panel section), which alertthe pilots to various conditions or switch po-sitions, and are located on the center portionof the glareshield just above the autopilot-flight director panel. These lights are referredto as glareshield annunciator lights. TwoMSTR WARN lights on the instrument panel,one in front of each pilot, flash when any redlight on the glareshield panel illuminates.These flashing lights serve to draw pilot at-tention to the glareshield lights and, thereby,to the malfunctioning system.

Provision is made to test all glareshield an-nunciator lights with two switches, one lo-cated on either end of the glareshield justbeneath the glareshield lights panel.

The intensity of the glareshield annunciatorlights is controlled automatically.

There may be other annunciator lights lo-c a t e d o n t h e i n s t r u m e n t p a n e l , c e n t e rpedestal, or thrust reverser control panel (ifinstalled). These lights function as system ad-visory annunciators.

TEST

CHAPTER 4MASTER WARNING SYSTEM




GLARESHIELDANNUNCIATOR LIGHTSThe red, amber, and green glareshield lightsreceive power from the lef t and/or r ightessential DC buses through the respectiveWRN LTS circuit breakers. The red lights areused to indicate the more critical malfunctions.Amber lights denote cautionary items, andgreen lights indicate conditions which maybe normal but need to be announced.

If a glareshield annunciator light illuminatesand the condit ion is corrected, the l ightextinguishes; should the condition recur, thelight again illuminates.

Five of the glareshield annunciator lights give a f lashing indication under the fol-lowing conditions:

1. SPOILER—If spoilers and flaps areboth extended (flaps more than 13°)

2. STALL (L or R)—If the angle-of-at-tack indicators reach shaker limits(yellow band)

3. FIRE (L or R)—If the warning sys-tem detects a fire or overtemperaturecondition in the engine nacelle

NOTEOn airplane SNs 35-002 through 35-431 and SNs 36-002 through 36-049, the MSTR WARN lights maynot cancel when any of these redglareshield lights are flashing.

MASTER WARNINGLIGHTSAnytime a red glareshield annunciator lightilluminates, the red MSTR WARN lights on thepilot’s and copilot’s instrument panels alsoilluminate and flash. Pressing either MSTRWARN light causes both MSTR WARN lightsto extinguish (except when triggered by aflashing red annunciator light on the early air-planes mentioned above. However, the red

glareshield annunciator light remains illumi-nated as long as the causative condition exists.

TESTDepressing either of the two test switchesunder the glareshield (See Figure 4-1) causesthe following lights to illuminate:

• All glareshield annunciator lights andboth MSTR WARN lights

• FIRE warning lights

• Marker beacon lights (if installed)

• Thrust reverser panel annunciator lights(if installed)

• AFCS/control panel annunciator lights(FC-530 AFCS)

• ANTISKID lights

• AIR IGN lights

• Fuel panel lights

• Copilot’s flight director annunciatorlights

• Dual PITOT HT indicator l ights (ifinstalled)

• Starter-engaged lights (if installed)

• Rotary test switch current limiter light(if installed)




Figure 4-1. Test Switch

INTENSITY CONTROLA photoelectric cell outboard of each FIREhandle (Figure 4-1) automatically adjusts theglareshield annunciator light intensity forexisting cockpit light conditions. The otherinstrument panel and pedestal annunciatorlights dim when the navigation light (NAVLT) switch is turned on.

BULB CHANGEGlareshield annunciator light lenses can beremoved for bulb replacement.

ILLUMINATION CAUSESThe tabula t ion in Table 4-1 shows eachannunciator light label, color, and cause forillumination.

NOTESome l i gh t s a r e op t i ona l , andarrangements may vary betweenairplanes.




ANNUNCIATOR CAUSE FOR ILLUMINATION ANNUNCIATOR CAUSE FOR ILLUMINATION

DH

LOWFUEL

L FUELPRESS

R FUELPRESS

SPOILER

SPOILER

DOOR

AUGAIL

PITOTHT

FUELFILTER

L ENGICE

R ENGICE

L FUELCMPTR

R FUELCMPTR

LSTALL

RSTALL

L VGMON

R VGMON

At or below altitude set on radio altimeter

Fuel is below 400-500 pounds in either wing tank

Less than 0.25-psi fuel pressure to engine (Light extinguishes at 1 psi.)

Steady–Spoilers not locked down (normal if extended)

Flashing–Spoilers deployed with 13° or more flaps extended (normal on landing roll)

One of 10 latch pins not fully engaged or hook motor not fully retacted

1. Spoilers split 6° or more

2. Spoiler and aileron split 6° or more in spoileron mode

1. One or both pitot heaters is in-operative with the switches on.

2. One or both pitot heat switches is off.

Differential pressure is 1.25 psi across one or both airframe fuel filters. Fuel is bypassing the filter.

1. Switch ON–Insufficient pressure to nacelle or fan spinner or failure of valve(s) to open

2. Switch OFF–Nacelle or fan spinner valve(s) open

1. Switch is off

2. Computer has failed with the switch on.

1. Steady–system is off or failed. (During pusher actuation it is normal.)

2. Flashing–In shaker range

One motor in the vertical gyro has failed.

(FC 200)

(FC530)

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES




CAUSE FOR ILLUMINATIONANNUNCIATORCAUSE FOR ILLUMINATIONANNUNCIATOR

MACHTRIM

PRIINV

SECINV

AUXINV

LO OILPRESS

STABOV HT

WSHLDOV HT

STEERON

BLEEDAIR L

BLEEDAIR R

LGEN

RGEN

CABALT

WINGOV HT

WSHLDHT

ALCAI

BAT140

BAT160

ENGSYNC

TOTRIM

CURLIM

ARMED

FIREPULL

MSTRWARN

LOWHYD

FUELXFLO

L LOOIL

R LOOIL

PITCHTRIM

System is inoperative with speed above 0.69 Mach and autopilot disengaged. If above 0.74 Mach, the overspeed warning horn sounds.

1. Inverter is off.

2. Inverter switch is on and output is less than 90 volts, or less than 10 volt-amps

Inverter has failed with the switch on.

Oil pressure on one or both engines is below 23 ±1 psi.

Stabilizer structural temperature is above 215° F.

Windshield heat has been shut off by a temperature limit.

GND–High or low limit

AIR–High limit only

Nosewheel steering is engaged

1. Overtemperature of pylon (250° F) or duct (590° F/645°F)

2. Both lights–Manifoldoverpressure (47 psi) on SNs35-082, 35-087 through 35-106,35-108 through 35-112, 36-023through 36-031, and AMK 76-7

Indicated generator is off or has failed.

Cabin altitude has reached 8,750 ±250 feet and controller has automatically switched to manual control.(Late ECS Only) FC 530 AFCS

Wing structural temperature is above 215° F.

The windshield anti-ice valve is open.

1. Late ECS–the alcohol tank is empty.

2. Early ECS–alcohol system pres-sure is low.

One or both batteries' temperature is 140° F or more.

One or both batteries' temperature is 160° F or more.

The engine sync switch is on with the nose gear down and locked.

Airplane is on the ground and the horizontal stabilizer is not trimmed for takeoff.

Failure of either or both 275-amp current limiter (SNs 35-370, 35-390 and subsequent and 36-048 and subsequent).

Fire-extinguishing bottles are armed.

Fire/overheat is detected in associated engine.

A red light on the master warning panel is illuminated.

LOW HYD–Hydraulic system pressure is 1,125 psi or less.

FUEL XFLO–Fuel crossflow valve is open.

L LO OIL, R LO OIL–Indicated engine oil pressure is low.

1. Primary pitch trim is running at fast rate with flaps up.

2. Primary pitch trim has a fault (potential runaway).

3. Wheel master switch is depressed

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES (Cont)




ANNUNCIATOR CAUSE FOR ILLUMINATION ANNUNCIATOR CAUSE FOR ILLUMINATION

LH ENGCHIP

RH ENGCHIP

EMERPWR 1

EMERPWR 2

START L

START R

COMPTRWARN

NAC HEATON

L NACHEAT

R NACHEAT

OR

AIR IGN L AIR IGN R

ANTI-SKIDGEN

L CURLIMITER

R CURLIMITER

R PITOTHEAT

PARKBRAKE

WSHLDDEFOG

L R

Ferrous metal particles are detected in indicated engines oil.

Indicated emergency battery is powering the connected systems.

Indicated starter is engaged.

HSI headings are not within 7°.

L or R NAC HEAT switches are ON.

Indicated NAC HEAT switch is ON.

Ignition system is activated.

Indicated antiskid generator is inoperative.

Indicated 275-ampere current limiter has failed.

1. Indicated pitot heat switch is off.

2. Switch is on and indicated pitot heat has failed.

1. Parking brake is set.

2. Parking brake handle is not fully in after releasing parking brake.

1. Illuminates momentarily when WSHLD DEFOG is set ON.

2. Indicates an over heat/underheat conditiion when ON.

(AMK 80-17)

L R

START L L PITOTHEAT

Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES (Cont)




1. All glareshield annunciator lights andsystem advisory annunciator lights can betested by:A. The rotary test switchB. Depressing each individual lightC. Depressing either glareshield TEST

switchD. Shutting the represented system off

2. When a red glareshield annunciator lightilluminates, another annunciation thatoccurs is:A. Only the pilot’s MSTR WARN light

flashes.B. Both MSTR WARN lights illuminate

steady.C. Only the copilot’s MSTR WARN light

illuminates.D. Both MSTR WARN lights flash.

3. An illuminated glareshield annunciatorlight suddenly extinguishes, indicating:A. Five minutes have passed.B. The malfunction no longer exists.C. Three minutes have passedD. The MSTR WARN lights have been

reset.

4. The glareshield annunciator light inten-sity is adjusted:A. Automatically by photoelectric cellsB. By depressing the TEST buttonC. By depressing each individual capsuleD. By depressing the DIM button

5. The flashing MSTR WARN lights can bereset by depressing either MSTR WARNlight:A. Unless a red glareshield annunciator

is flashingB. AnytimeC. Unless a red glareshield annunciator

is illuminated steadyD. Unless an engine FIRE PULL light

illuminated steady

QUESTIONS

5-i

CHAPTER 5FUEL SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................... 5-1

GENERAL............................................................................................................................... 5-1

FUEL TANKS AND TANK VENTING SYSTEM ................................................................ 5-3

General ............................................................................................................................. 5-3

Tip Tanks.......................................................................................................................... 5-3

Wing Tanks ...................................................................................................................... 5-3

Fuselage Tank................................................................................................................... 5-3

Ram-Air Vent System ...................................................................................................... 5-4

FUEL INDICATING SYSTEMS ............................................................................................ 5-4

Fuel Quantity Indicating System/Low Fuel Warning ...................................................... 5-4

Fuel Flow Indicating System ........................................................................................... 5-7

FUEL DISTRIBUTION .......................................................................................................... 5-7

General ............................................................................................................................ 5-7

Boost Pumps .................................................................................................................... 5-7

Motive-Flow Fuel and Jet Pumps..................................................................................... 5-8

Filters ............................................................................................................................... 5-9

Main Fuel Shutoff Valves (Firewall)................................................................................ 5-9

Low Fuel Pressure Warning Lights.................................................................................. 5-9

Pressure-Relief Valves ..................................................................................................... 5-9

Fuel Drain Valves............................................................................................................. 5-9




FUEL TRANSFER SYSTEMS............................................................................................... 5-9

Crossflow System............................................................................................................. 5-9

Normal Transfer System ................................................................................................ 5-10

Gravity-Flow Transfer System....................................................................................... 5-11

Float and Pressure Switches........................................................................................... 5-11

Pressure-Relief Valves ................................................................................................... 5-12

Fuselage Fuel Fill-Transfer Operations ......................................................................... 5-12

TIP-TANK FUEL JETTISON SYSTEM .............................................................................. 5-13

FUEL SERVICING ............................................................................................................... 5-13

General........................................................................................................................... 5-13

Safety Precautions.......................................................................................................... 5-13

Aviation Gasoline........................................................................................................... 5-15

Anti-icing Additive ........................................................................................................ 5-15

Refueling........................................................................................................................ 5-15

QUESTIONS......................................................................................................................... 5-16

5-ii FOR TRAINING PURPOSES ONLY



5-iii

ILLUSTRATIONS

Figure Title Page

5-1 Fuel System .............................................................................................................. 5-2

5-2 Ram-Air Scoop and Overboard Drain...................................................................... 5-4

5-3 Fuel Vent System...................................................................................................... 5-5

5-4 Fuel Control Panels .................................................................................................. 5-6

5-5 Fuel Flow Indicator .................................................................................................. 5-7

5-6 Jet Pump Schematic.................................................................................................. 5-8

5-7 Fuel Drain Locations.............................................................................................. 5-10

5-8 Airplane Grounding Points .................................................................................... 5-14

5-9 Prist Blending Apparatus ....................................................................................... 5-14

5-10 Refueling Filler Cap............................................................................................... 5-15




INTRODUCTIONThe Learjet 35/36 series fuel system consists of the fuel tanks, tank venting, indicating,distribution, transfer, and jettison systems.

This chapter covers the operation of the fuel system up to the engine-driven fuel pumps.At that point, fuel system operation becomes a function of the engine. Refer to Chapter7, “Powerplant,” for additional information.

GENERALThe fuel storage system consists of tip tanks,integral tanks in each wing, and a fuselagetank. A crossflow valve permits fuel transferbetween the wings for fuel balancing.

Each wing tank contains a jet pump and anelectric standby pump to supply fuel to the en-gine on the same side. Tip tank and fuselage tankfuel must be transferred into the wing tanks byjet pumps and an electric pump, respectively.

A ram-air system is used to vent all tanks.Drain valves are provided to remove conden-sation and contaminants from the low pointsin the fuel tanks, and to drain the contents ofthe vent system sump.

Tip tank fuel can be jettisoned, if required.

Figure 5-1 depicts the Learjet 35/36 seriesfuel systems.

0

2

4 6

8

10

MAINFUEL

LBS X 100

CHAPTER 5FUEL SYSTEM







FILLER CAP

QUANTITY PROBE

BOOST PUMP

JET PUMP

FILTER

ENGINE PUMP

PRESSURE RELIEF VALVE

CHECK VALVE

FLAPPER VALVE

FLOAT SWITCH

SUPPLY

LOW PRESSURE

HIGH PRESSURE

GRAVITY (TRANSFER)LOW PRESSURE (FILL)

LEGEND

F

P

TOSUMP

FUEL JETTISONSHUTOFF VALVE

LOW FUELPRESSURESWITCH

DIFFERENTIALPRESSURESWITCHFUEL

SHUTOFFVALVE

75-PSIRELIEFVALVE

CROSSFLOWVALVE

WINGPRESS SW

TRANSFERVALVE

FUSELAGETANK

TRANSFERLINE

EMPTY LIGHTPRESSURESWITCH

MOTIVEFLOWVALVE

MOTIVEFLOWFUEL

MODEL 35WITHOUTGRAVITY-FLOW LINE

F

P

PP

P

PP

TOSUMP

FUEL JETTISONSHUTOFF VALVE

LOW FUELPRESSURESWITCH

DIFFERENTIALPRESSURESWITCH

FUELSHUTOFF

VALVE

75-PSIRELIEFVALVE

CROSSFLOWVALVE

WINGPRESS

SW

TRANSFERVALVE TRANSFER

LINE

FUSELAGEVALVE

GRAVITY-FLOWLINE

EMPTY LIGHTPRESSURESWITCHFUSELAGE

TANK(MODEL 36

TANK SHOWN)MOTIVEFLOWVALVE

MOTIVEFLOWFUEL

MODEL 36 AND MODEL 35WITHGRAVITY-FLOW LINE

MODEL 36 AND MODEL 35WITHGRAVITY-FLOW LINE

F

P

PP PP

P

Figure 5-1. Fuel System

FUEL TANKS AND TANKVENTING SYSTEM

GENERALThe total usable fuel capacity is approximately6,238 pounds for the 35 model and approxi-mately 7,440 pounds for the 36 model. Unus-able (trapped) fuel is included in the airplanebasic weight and is not reflected in any fuelquantity indications.

TIP TANKSEach tip tank capacity is 1,215 pounds of us-able fuel; capacity is reduced to 1,175 poundswith installation of a recognition light. Thetanks are permanently attached to the wingsand are positioned at 2 degrees nosedownrelative to the airplane centerline. Baffles areinstalled to minimize slosh and prevent adverseeffects on the airplane center of gravity duringextreme pitch attitudes.

A jet pump installed in each tip tank transfersfuel into the wing tank. Approximately one-half of the fuel will gravity-flow through twoflapper valves into the wing tank; however, anyfuel at a level lower than one-half full must betransferred using the jet pump. A standpipe isinstalled in each jet pump transfer line to pre-vent fuel from being siphoned from the wingtank to the tip tank when the applicable engineis shut down.

The tip tank is vented through two vent floatvalves located in the forward and aft ends ofthe tank.

A fuel probe in each tip tank provides infor-mation to the fuel quantity indicating system.

All tip-tank fuel can be jettisoned through avalve in the tank tailcone, if required.

A filler cap on each tip tank is used to servicethe entire airplane fuel system.

WING TANKSEach wing tank extends from the airplane cen-terline to the tip tank and holds 1,254 poundsof usable fuel. Areas which are not part of thewing fuel cell are the main landing gear wheelwell, the leading edge forward of spar 1 (wingheat area), and the trailing edge between spars7 and 8 (flap, spoiler, and aileron areas).

The 2.5-degree wing dihedral makes the in-board portions of the wing tanks the lowestareas. In each wing tank, a jet pump and anelectric standby pump are located within theseareas and will remain submerged in fuel untilthe tanks are nearly empty.

Wing tank ribs and spars act as baffles to mini-mize fuel shifting. Flapper valves located inthe wing ribs allow unrestricted inboard flowof fuel and limit outboard flow. Two pressure-relief valves at the centerline rib equalize in-ternal pressures between the two wing tanks.The wing tanks begin to fill through the twotip tank flapper valves as tip tank fuel in-creases beyond one-half full.

Three fuel probes in each wing tank provide in-formation to the fuel quantity indicating system.

FUSELAGE TANKThe fuselage tank consists of rubber bladderfuel cells located between the aft pressurebulkhead and tailcone section. The 35 modelsare equipped with two fuel cells with a capacityof 1,340 pounds of usable fuel, while the 36models are equipped with four fuel cells witha capacity of 2,542 pounds of usable fuel. De-pending on the airplane, either one or two fuellines connect the fuselage tank to the wingtanks for filling and transfer. This is explainedin the Fuel Transfer Systems section.

One fuel probe provides information to thefuel quantity indicating system.




RAM-AIR VENT SYSTEMA ram-air scoop located on the underside ofeach wing (Figure 5-2) supplies positive airpressure in flight to a manifold which directlyvents the fuselage tank and both tip tanks.Each wing tank is indirectly vented to its owntip tank through a length of tubing, the ends ofwhich extend to the uppermost area of each tank (Figure 5-3). The ram-air scoops, by design,do not require heating to remain ice-free. Twovent float valves are located in each tip tank,and one in the fuselage tank on 35 models.The float valves close when the fuel level reachesthe vent ports, preventing fuel from entering thevent lines. A vacuum relief valve in each tiptank and the fuselage tank opens to allow airto enter the tanks should vacuum conditionsoccur. Each tip tank has two pressure reliefvalves which protect the tanks from excessivepressure. The pressure relief valves are set at1.0 and 1.5 psi, the second valve providing abackup in case the first valve fails.

Thermal expansion of fuselage fuel in 35 mod-els is accounted for by an open-ended ventline which bypasses the vent float valve (36models use three open-ended vent lines) tovent pressures overboard through the ram-airscoops. A sump, installed in the vent manifold,located at the bottom center fuselage just aftof the main landing gear, collects any fuel

which might enter the vent lines. A ventdrain valve permits draining of the sump toensure that the vent line to the fuselage tankis unobstructed.

FUEL INDICATINGSYSTEMS

FUEL QUANTITY INDICATINGSYSTEM/LOW FUEL WARNINGThe fuel quantity indicating system includesan indicator and tank selector switch locatedon the fuel control panel (Figure 5-4). A redLOW FUEL warning light (Annunciator Sec-tion) illuminates when either wing tank fuellevel is low.

The fuel quantity indicating system uses DCpower from the right essential bus throughthe FUEL QTY circuit breaker. The six-posi-tion rotary selector switch enables the pilot tocheck the fuel quantity in each of the fivetanks and the airplane total fuel quantity. Thefuel quantity for the position selected is readon the fuel quantity indicator. The quantitiesprinted beside each selector switch positionindicate usable fuel capacities in pounds.




Figure 5-2. Ram-Air Scoop and Overboard Drain




WINGVENT

OPENVENTTUBE**

FLOATVALVE*

VACUUMRELIEF

FUELVENT

DRAIN

OVERBOARDDRAIN

FLAMEARRESTER

RAM-AIRSCOOP

*35 MODELS ONLY**THREE VENTS ON 36 MODELS

FLOATVALVE

(TYPICAL)

PRESSURE RELIEF

VACUUMRELIEF

TO AMBIENT

VACUUMRELIEFVALVE

1.5-PSIRELIEFVALVE

1.0-PSIRELIEFVALVE

Figure 5-3. Fuel Vent System




Figure 5-4. Fuel Control Panels

*OPTIONAL ON SNs 35-299 THROUGH 35-596. STANDARD ON SNs 35-597 AND SUBSEQUENT.

L ON R

L ON R

FUEL JTSN

OPEN

CLOSECROSS FLOW

EMPTYXFER

OFFFILL

FULL

}OPEN

CLOSEFUS VALVE

FUSTANK

JET PUMP

L WING1254

L TIP1215

TOTAL6238LBS

R WING1254

R TIP1175

FUS1340

STANDBY PUMPS

L ON R

MODEL 35

0 0 0 0

OPEN

0

1

2LBS x 1000

3FUEL

QUANTITY

45

6

7

MODEL 36

L ON R

L ON R

FUEL JTSN

OPEN

CLOSECROSS FLOW

EMPTYXFER

OFFFILL

FULL

OPEN

CLOSEFUS VALVE

FUSTANK

JET PUMP

L WING1254

L TIP1215

TOTAL7440LBS

R WING1254

R TIP1175

FUS2542

STANDBY PUMPS

L ON R

0

1

2LBS x 1000

3FUEL

QUANTITY

45

6

78

0 0 0 0

There are nine capacitance fuel probes. Onefuel probe is located in each tip tank and in thefuselage tank. Each wing tank has three probeswired in parallel. The inboard probe in the leftwing contains a temperature compensator whichadjusts quantity readings for all switch selec-tions for fuel density change due to temperature.

I f the compensator probe i s uncovered ,erroneous fuel quantity indications could beencountered at all switch positions.

Each probe uses an electrical capacitancemeasuring system to sense the fuel level. It thentransmits an electrical signal to the cockpit in-dicator where it is read as pounds x 1,000 onthe dial.

Each wing tank has fuel low-level float switch.When either wing tank fuel level reaches 400to 500 pounds remaining, the respective floatswitch actuates the red LOW FUEL light on theannunciator panel to indicate low wing fuelquantity (Annunciator Section). When flyingwith the LOW FUEL light on, limit pitch atti-tude and thrust to the minimum required.

FUEL FLOW INDICATINGSYSTEMA single fuel flow indicator, with two pointers(L and R) provides a readout of pounds of fuelflow per hour (Figure 5-5). A fuel counter(Figure 5-4) located on the fuel control panelprovides a four-digit readout in pounds of fuelconsumed by both engines. It should be resetto zero using the reset button adjacent to thecounter before starting the first engine. Bothindicators are powered from the battery charg-ing bus through a 10-amp current limiter.

FUEL DISTRIBUTION

GENERALEach engine is supplied with fuel from itsrespective wing fuel system; there is nocrossfeed capability. Either the wing standbypumps or the wing jet pumps supply fuel under

pressure to the engine-driven pumps. Duringengine start, the respective wing standby pumpis automatically energized when the GEN-–START switch is p laced in the STARTposition. When turbine speed (N2) reaches45% or 50%, or when the START switch ismoved to OFF or GEN (computer off starts),the wing standby pump is deenergized and thewing jet pump then provides fuel to the engine.The wing jet pumps and standby pumps havecheck valves on the output side to preventreverse flow when they are inactive.

BOOST PUMPSSubmerged DC-powered boost pumps areinstalled at three different locations—onestandby pump in each wing adjacent to thejet pump, and one transfer pump in the fuse-lage tank.




Figure 5-5. Fuel Flow Indicator

The standby pumps are used:

• For engine start (automatically ener-gized with starter switch activation)

• As a backup for the wing jet pumps

• For wing-to-wing crossflow

• For filling the fuselage tank (automati-cally energized with the XFER–FILLswitch in the FILL position)

Both standby pumps are deactivated when theXFER–FILL switch is in the XFER position.

The transfer pump is used to transfer fuse-lage tank fuel to the wing tanks.

The s tandby pumps a re powered by therespective L or R STBY PMP circuit breakerson left and right essential buses; the fuselagepump, from the FUSLG PMP circuit breakeron the right main bus.

MOTIVE-FLOW FUELAND JET PUMPSHigh-pressure fuel from the engine-drivenfuel pumps is the source of motive-flow fuelto operate the jet pumps. The fuel is routed

through the motive-flow valves to the jetpumps, where it passes through a small orificeinto a venturi. The low pressure created in theventuri draws fuel from the respective tank, re-sulting in a low-pressure, high-volume outputfrom the jet pump (Figure 5-6).

Motive-flow pressure varies with engine rpmand is regulated to 300 psi maximum. Conse-quently, jet pump discharge pressure also varieswith engine rpm. At idle, discharge pressureis approximately 10 psi, while at full-power set-tings, discharge pressure is approximately 12psi.

There are four jet pumps—one in each wingtank adjacent to the standby pump, and one ineach tip tank. The wing tank jet pumps drawfuel from the wing tanks and supply low-pres-sure fuel to the engine-driven, high-pressurefuel pumps. Wing jet pump output can besupplemented by the wing standby pump toensure positive pressure to an engine. The tiptank jet pumps draw fuel from the tip tanks anddeliver it directly to the cavities where thestandby pumps and jet pumps are located.

Jet pumps require no electrical power andhave no moving parts. They are controlled bytwo jet pump switches (Figure 5-4) whichelectrically open and close the motive-flow




FUEL SUPPLY HIGH PRESSURE LOW PRESSURE

OUTPUT

WING TANKSTRUCTURE

INPUT

LEGEND

Figure 5-6. Jet Pump Schematic

valves. Power is provided by the respective Lor R JET PUMP VAL circuit breaker on the leftand right essential buses. The amber indicatorlights next to the switches illuminate when themotive-flow valves are in transit or are not inthe position selected on the switch. Each jetpump switch (and motive-flow valve) controlsboth jet pumps (wing and tip) on that side.

FILTERSA fuel filter is installed in each engine feed lineto filter the fuel before it enters the engine-driven fuel pump. Should the filters becomeclogged, the fuel is allowed to bypass them.A differential pressure switch installed in eachfilter assembly illuminates the one amberFUEL FILTER annunciator light if either orboth filters are bypassing fuel (AnnunciatorPanel section).

MAIN FUEL SHUTOFF VALVES(FIREWALL)The fuel shutoff valves are powered from theessential buses through the L and R FW SOVcircuit breakers and are controlled by the FIREhandles on the glareshield. Pulling either FIREhandle closes the associated valve, and push-ing the FIRE handle in opens the valve. Thevalves will remain in their last positions shouldDC power fail.

LOW FUEL PRESSUREWARNING LIGHTSA low fuel pressure switch is located betweenthe fuel shutoff valve and the engine-drivenfuel pump in each engine feed l ine. Theswitches cause illumination of the appropri-ate red L or R FUEL PRESS annunciator lightwhen fuel pressure drops below 0.25 psi. Thelight extinguishes when pressure increasesabove 1.0 psi. Illumination of a FUEL PRESSwarning light is an indication of loss of fuelpressure to the engine. The probable cause isfailure of the affected wing jet pump.

The engine-driven pump is capable of suc-tion-feeding enough fuel to sustain engine op-eration without either the wing standby pumpor jet pump. However, 25,000 feet is the high-est altitude at which continuous operationshould be attempted in this event.

PRESSURE-RELIEF VALVESA 75-psi relief valve is installed in each mainfuel line on the engine side of the main shut-off valve. The valves relieve pressure buildupcaused by thermal expansion of trapped fuelwhen the engines are shut down by venting fueloverboard.

FUEL DRAIN VALVESDrain valves are located at low points through-out the fuel system for draining condensationor sediment. A small amount of fuel should bedrained from each valve during the exterior pre-flight inspection. The valves, spring-loaded tothe closed position, are located as follows: onefor each tip tank sump, one for the crossflow line,one for each wing sump, one for each engineline, one for each fuel filter, one (or two) for thefuselage tank line(s), and one (or two) for thefuselage tank sump(s) (Figure 5-7).

There is one drain valve located at the fuel ventsump. This valve must be completely drainedduring the exterior preflight inspection toprevent possible blockage of the fuselage ram-air vent line.

FUEL TRANSFERSYSTEMS

CROSSFLOW SYSTEMA DC motor-driven valve is installed in thecrossflow manifold connecting the two wingtanks (Figure 5-1). It is opened during fuse-lage fuel transfer and filling operations, andfor wing-to-wing fuel balancing. The valve is




controlled by the CROSS FLOW switch or theXFER–FILL switch on the fuel control panel(Figure 5-4) and is powered through the rightmain bus FILL & XFER circuit breaker. Ad-ditionally, on those airplanes equipped with thegravity-flow transfer line, the valve is con-trolled by the FUS VALVE switch which ispowered from the left essential bus FUSVALVE (or FUSE VAL) circuit breaker.

The amber light (Figure 5-4) adjacent to theCROSS FLOW switch illuminates when thevalve is in transit or is not in the positionselected. A green or amber FUEL XFLOannunciator light (Annunciator Panel section)on the glareshield illuminates continuouslywhenever the crossflow valve is fully open.

If wing fuel imbalance occurs, as in single-engine operation, crossflow is accomplishedby opening the crossflow valve and turning onthe standby pump in the heavy wing, while en-suring that the opposite standby pump is off.The transfer rate is approximately 50 poundsof fuel per minute.

With both engines operating, opening thecrossflow valve to balance fuel should not beattempted when a red FUEL PRESS light isilluminated unless it can be accomplishedbelow 25,000 feet. To do so would divert pres-sure from the affected engine-driven pump tothe crossflow line. Instead, asymmetric powersettings may be used to balance fuel, if nec-essary. The above considerations do not applyto single-engine operations, and normal cross-flow operations may be performed as usual.

NORMAL TRANSFER SYSTEMThe Learjet models 35/36 each have a fueltransfer line connecting the fuselage tanktransfer pump with the crossflow manifold(Figure 5-1). A DC motor-driven transfer valveinstalled in the line controls fuel movementbetween the fuselage and wing tanks. Thevalve is controlled by the XFER-FILL switchlocated on the fuel control panel. When theswitch is positioned from OFF to XFER, thetransfer and crossflow valves are sequencedopen and the transfer pump is energized




RH ENGINE FUELLINE DRAIN

FLUSH SUMP DRAIN

WING SUMP DRAIN

FUEL VENT DRAIN

FUEL FILTER DRAIN

CROSSFLOW DRAIN

LH ENGINE FUELLINE DRAIN

FLUSH SUMP DRAIN

WING SUMP DRAIN

FUSELAGE LINE DRAIN*

FUSELAGE TANK SUMP DRAIN*

FUEL FILTER DRAIN

* THE 35 MODELS WITH OPTIONAL FUEL LINE AND THE 36 MODELS HAVE TWO FUSELAGE LINE DRAINS AND TWO FUSELAGE TANK SUMP DRAINS.

RIGHT WING LEFT WING

Figure 5-7. Fuel Drain Locations

automatically, while both standby pumps aredeactivated. When the switch is positionedfrom OFF to FILL, the transfer and crossflowvalves are sequenced open, and both standbypumps are energized automatically. When theswitch is positioned from either XFER or FILLto the OFF position, the transfer pump orstandby pumps (whichever the case may be) aredeenergized and the transfer and crossflowvalves are sequenced closed. The amber lightadjacent to the XFER–FILL switch illuminateswhen the valve is in transit or is not in theposition selected (Figure 5-4). The valve ispowered through the right main bus FILL &XFER circuit breaker.

On 35 models without the optional gravity-flow line, the transfer line is connected to theright side of the crossflow valve. On all 36models, and 35 models with the optional gravity-flow line, the transfer line is connected to theleft side of the crossflow valve.

GRAVITY-FLOWTRANSFER SYSTEMAs an option of SNs 35-299 through 35-596,and as standard equipment on 35-597 andsubsequent, and on all 36 models, a DC motor-driven fuselage valve is installed in a secondfuel line, connecting the fuselage tank with thecrossflow manifold on the right side of thecrossflow valve (Figure 5-1). The valve iscontrolled by the FUS VALVE switch on thefuel control panel. When the FUS VALVEswitch is positioned to OPEN, both the fuselagevalve and the crossflow valve simultaneouslyopen, allowing fuel to gravity-flow from thefuselage tank to both wings. When fuselagefuel is transferred in this manner, 162 poundsof fuel remain in the fuselage tank. The fuse-lage valve is also controlled by the XFER–FILL switch. When placed to FILL, the trans-fer valve, fuselage valve, and crossflow valveare sequenced open, and the standby pumps areenergized to pump wing tank fuel throughboth fuel lines into the fuselage tank. Thefuselage valve remains c losed when theXFER–FILL switch is positioned to XFER.The amber light adjacent to the FUS VALVE

switch will illuminate when the fuselage valveis in transit or is not in the position selected(Figure 5-4). If either standby pump switchis on, the FUS VALVE switch is renderedinoperative, and neither the fuselage valvenor the crossflow valve will open if the FUSVALVE switch is moved to OPEN. Conversely,if the FUS VALVE switch is already in theOPEN position (fuselage valve and crossflowvalve open), turning either standby pumpswitch on will automatically cause the fuse-lage valve and crossflow valves to sequenceclosed. The fuselage valve is powered throughthe left essential bus FUSE VAL (or FUSVALVE) circuit breaker.

FLOAT AND PRESSURESWITCHES

Fuselage Fuel Tank Float SwitchWhen filling the fuselage tank, a float switchmounted inside the tank actuates when thetank is full. The switch:

• Illuminates the green FULL light on thefuel control panel

• Deenergizes the standby pumps

• Closes the transfer and crossflow valves

• Closes the fuselage valve (all airplanesequipped with the gravity-flow transferline)

The green FULL light on the fuel control panel(Figure 5-4) remains illuminated until theXFER–FILL switch is turned off.

Fuselage Tank Low-PressureSwitchThe fuselage tank low-pressure switch is in-stalled in the fuselage transfer line to alertthe pilot when fuselage fuel is depleted. Withthe XFER–FILL switch in the XFER position,the switch senses low pressure in the line andilluminates the white EMPTY light on the fuel




control panel (Figure 5-4) when either of twoconditions exists:

• The tank is empty.

• The fuselage transfer pump fails.

The switch actuates when pressure dropsbelow 2.75 psi and resets at 3.75 psi as pres-sure increases.

Wing Fuel Pressure SwitchA wing fuel pressure switch is installed toprevent internal overpressurization of thewings during transfer of fuselage tank fuel. Theswitch, located in the right main wheel well,deenergizes the fuselage transfer pump whenwing fuel pressure reaches 5 psi; the switchresets and energizes the pump again when thepressure drops below 2.5 psi.

PRESSURE-RELIEF VALVESTwo one-way pressure relief valves are locatedat wing rib 0.0, which separates the left andright wing fuel tanks. Each valve, relieving inthe opposite direction, opens at 1 PSID toequalize fuel pressure between the wing tankswhen crossflowing or transferring fuel.

FUSELAGE FUELFILL-TRANSFER OPERATIONS

Fill OperationFuel may be pumped from the wings to thefuselage tank using the FILL position on theXFER–FILL switch. The FILL position maybe used for CG considerations in flight; however,it is normally used only during fuel servicing.

When the XFER–FILL switch is placed to theFILL position:

• The transfer valve opens, then

• The crossflow valve opens, then

• The standby pumps are energized, and

• The fuselage tank float switch is enabled.

If the tank is to be filled to capacity, the floatswitch actuation automatically:

• Deactivates the standby pumps

• Closes all valves

• Illuminates the green FULL light, whichwill go out when the XFER–FILL switchis turned off

The filling process may be terminated at anypoint by turning the XFER–Fill switch off.

Transfer OperationsThe normal method of transferring fuselagefuel in flight is accomplished by using the XFERposition on the XFER–FILL switch. When theswitch is placed in the XFER position:

1. The transfer valve opens.

2. The crossflow valve opens.

3. The fuselage transfer pump is energized.

4. The standby pumps are disabled.

5. The white EMPTY light (pressureswitch) is enabled.

Gravity-flow is also possible on all airplanesthrough the normal transfer line, should thetransfer pump fail. The amount of fuel trapped(unusable) is approximately 162 pounds. Therate of gravity transfer will, however, be slowerthan when using the fuselage valve, if installed.

When the XFER–FILL switch is placed in theOFF position:

• The transfer pump is deenergized, and

• Operation of the standby pumps is onceagain possible.

• The transfer valve closes, then

• The crossflow valve closes.




The alternate method of transferring fuselagefuel in flight is only possible on airplanesequipped with the gravity-flow line by usingthe OPEN position on the FUS VALVE switch.However, prior to doing so, it is essential tofirst assure that the XFER–FILL switch is off,and that both standby pump switches are off.Then, when the FUS VALVE switch is placedin the OPEN position, the fuselage valve andcrossflow valve open simultaneously. Thevalves are not sequenced as they are when usingthe XFER–FILL switch.

When the amount of fuel in the wing tanksbegins to decrease, the FUS VALVE switchmay be turned off, and the transfer process maybe completed using the normal t ransferprocedure. On airplanes with the gravity-flowline, approximately 162 pounds of fuel will betrapped (unusable) if the gravity-flow lineonly is used to transfer fuselage fuel.

TIP-TANK FUELJETTISON SYSTEMA DC motor-driven valve in the tailcone ofeach tip tank provides the capability of jetti-soning tip-tank fuel. One FUEL JTSN switchon the fuel control panel (Figure 5-4) controlsboth tip-tank jettison valves. When the FUELJTSN switch is placed to the ON position, thejettison valves open and two amber lightsilluminate continuously on the fuel controlpanel, indicating that the valves are fully open.The jettison tubes are scarfed, which createsa low-pressure area that helps pull the fuelout of the tank(s). This, in combination withthe force of gravity, enables the entire contentsof both tanks to be jettisoned. Fuel jettison isfaster while the airplane is in a noseup attitude.It takes approximately five minutes to jettisonfuel from the tip tanks. The left- and right-handjettison valves are protected independentlyby the FUEL JTSN circuit breakers located onthe left and right essential buses, respectively.

FUEL SERVICING

GENERALFuel servicing includes those proceduresnecessary for fueling and adding anti-icingadditives.

Fueling is accomplished through a filler capin the top of each tip tank. Fuel then begins toflow by gravity from the tip tanks into thewing tanks as the tip tanks reach one-half full.The wing standby pumps pump fuel to thefuselage tank when the XFER–FILL switch isset to FILL.

At normal temperatures some water is alwaysin solution (dissolved) with fuel. At highalt i tudes, fuel undergoes a cold-soakingprocess and small amounts of water come outof solution and subsequently freeze. The anti-icing additives specified for use in the Lear-jet 35/36 are Hi-Flo Prist and QUELL. Eitheradditive prevents the growth of microbiologicalorganisms in the fuel. Fuel containing anti-icing additive conforming to MIL-I-27686requires no further treatment.

SAFETY PRECAUTIONSRefueling and defueling should be accomplishedonly in areas which permit free movement offire equipment.

Figure 5-8 shows the airplane grounding points.

When adding anti-icing additives (Figure 5-9), follow the manufacturer’s instructionsfor blending.







Figure 5-8. Airplane Grounding Points

FUEL ADDITIVE BLENDER HOSE

FUEL NOZZLE

TRIGGER

RING

HANDLE

HI-FLO

PRIST

(OR)

MIL-I-

27686

Figure 5-9. Prist Blending Apparatus

AVIATION GASOLINEAviation gasoline (MIL-D-5572D, Grades80/87, 100/130, and 115/145) may be used asan emergency fuel and mixed in any propor-tion with the various approved jet kerosene-base fuels.

ANTI-ICING ADDITIVEAll fuels used must have an approved blendedanti-icing additive. Depending upon fueling lo-cation and type of fuel, the additive may or maynot be blended at the refinery. If not blendedat the refinery, the additive must be blendedat the time of fueling. Refer to the AFM for theapproved MIL-Specs. Compare the MIL-Specof the anti-icing additive to be blended withthe referenced MIL-Specs in the AFM to de-termine the correct blending amounts.

REFUELINGRefueling is accomplished through the tip-tank filler caps (Figure 5-10). The fuel beginsto flow by gravity into the wing tanks as thetip tanks reach one-half full. The standbypumps are used to fill the fuselage tank. (SeeFuel Transfer Systems, this chapter.) A groundpower unit should be used, if possible, becauseof the requirement to operate the standbypumps. Refer to the approved AFM for detailedrefueling procedures.




Figure 5-10. Refueling Filler Cap




1. Trapped fuel weight:A. Must be added to the weight of fuel

taken on board when servicing theairplane

B. Is included in the airplane basicweight for a i rp lanes cer t i fied inthe U.S.

C. Must be accounted for in the fuselagetank for CG purposes

D. May be disregarded since it is lessthan 200 pounds

2. With the exception of the FUEL JTSNlights, all other amber lights on the fuelcontrol panel, when illuminated steady,indicate that the respective:A. Valves are cycling or the pumps are

properly operating.B. Valves are in the correct position; the

pumps are inoperative.C. Switch position agrees with the valve

position or pump operation.D. Valve position disagrees with the

switch position.

3. The red LOW FUEL light illuminateswhen:A. 350 pounds total fuel remains.B. 250 to 350 pounds remain in either

wing, depending on the airplane SN.C. 400 to 500 pounds total fuel remainsD. 400 to 500 pounds remain in either

wing

4. The standby pumps are used for all thefollowing functions except:A. Engine startB. As a backup for the main jet pumpsC. Wing-to-wing crossflow with a wing

tank jet pump inoperativeD. Wing-to-fuselage transfer of fuel

5. The crossflow valve opens:A. Only when the CROSS FLOW switch

is set to OPENB. Only when the CROSS FLOW switch

is set to OPEN or the XFER–FILLswitch is set to XFER

C. Anytime electrical power is lostD. Whenever the CROSS FLOW, XFER–

FILL, or FUS VALVE switches aremoved f rom the OFF or CLOSEposition

6. Steady illumination of an amber transfervalve light indicates:A. The valve has failed to close.B. The valve has failed open.C. The valve has operated correctly.D. The valve has failed to move to the

position commanded by the XFER–FILL switch.

7. Illumination of the red L or R FUELPRESS light indicates:A. Fuel pressure to the respective engine-

driven fuel pump is low.B. Fuel pressure to the respective engine

is too high for safe operation.C. A fuel filter is bypassing.D. Fuel pressure to the respective engine

is optimum for engine start.

8. When the XFER–FILL switch is placedto the FILL position, the:A. Float switch is disabled.B. Wing standby pumps are disabled.C. Fuselage valve closes.D. Crossflow valve opens.

QUESTIONS

9. Motive-flow fuel for the jet pumps issupplied by the:A. Engine-driven fuel pumpsB. Wing standby pumpsC. Fuselage transfer pumpD. Motive-flow control unit

10. The amber FUEL FILTER light indicates:A. Low fuel pressure to the engine-driven

pump; the standby pumps should beturned on

B. That both fuel filters are being by-passed; the light does not illuminateif only one filter is bypassed

C. That one or both fuel filters are beingbypassed

D. That only the secondary fuel filters arebeing bypassed

11. The amount of fuel trapped in the fuselagetank after completion of gravity transfervia the fuselage valve is approximately:A. 62 poundsB. 162 poundsC. 262 poundsD. None

12. The wing fuel pressure switch:A. Turns off the fuselage transfer pump

when wing fuel pressure reaches 5 psi

B. Turns on the fuselage transfer pumpwhen wing fuel pressure is below 5 psi

C. Turns off the wing standby pumpswhen wing fuel pressure reaches 5 psi

D. Turns on the wing standby pumpswhen wing fuel pressure is below 5 psi

13. When using any mixture of aviationgasoline:A. Do not take off with fuel temperature

lower than –54° C (–65° F).B. Restrict flights to below 15,000 feet.C. Both jet pumps and both standby

pumps must be on and the pumps mustbe operating.

D. All of the above answers are correct.

14. The Learjet 35/36 requires anti-icingadditive:A. At all timesB. Only when temperatures of –37° C

and below are forecastC. Only for flights above 15,000 feetD. Only for flights above FL 290




The information normally contained in this chapter is

not applicable to this particular aircraft.

Revision .01

7-i

CHAPTER 7POWERPLANT

CONTENTS

Page

INTRODUCTION ................................................................................................................... 7-1

GENERAL............................................................................................................................... 7-1

MAJOR SECTIONS................................................................................................................ 7-2

Air Inlet Section ............................................................................................................... 7-2

Fan Section....................................................................................................................... 7-2

Compressor Section ......................................................................................................... 7-3

Combustor Section ........................................................................................................... 7-3

Turbine Section ................................................................................................................ 7-3

Exhaust Section................................................................................................................ 7-3

Accessory Section ............................................................................................................ 7-3

OPERATING PRINCIPLES.................................................................................................... 7-3

OIL SYSTEM.......................................................................................................................... 7-5

General ............................................................................................................................. 7-5

Indication ......................................................................................................................... 7-6

Operation.......................................................................................................................... 7-7

FUEL SYSTEM....................................................................................................................... 7-8

General ............................................................................................................................. 7-8

Fuel Pressure .................................................................................................................... 7-8

Motive-Flow Lockout Valve and Pressure Regulator ...................................................... 7-8

Fuel Control Unit (FCU).................................................................................................. 7-8




Electronic Fuel Computer ................................................................................................ 7-8

Start Pressure Regulator (SPR) ...................................................................................... 7-11

Surge Bleed Valve .......................................................................................................... 7-11

Fuel Flow ....................................................................................................................... 7-11

Flow Divider .................................................................................................................. 7-11

Fuel Spray Nozzles ........................................................................................................ 7-12

Operation ....................................................................................................................... 7-12

IGNITION SYSTEM ............................................................................................................ 7-12

General........................................................................................................................... 7-12

Automatic Mode ............................................................................................................ 7-12

Selective Mode............................................................................................................... 7-13

Indication ....................................................................................................................... 7-13

ENGINE CONTROLS .......................................................................................................... 7-13

STARTERS............................................................................................................................ 7-14

General........................................................................................................................... 7-14

Operation ....................................................................................................................... 7-16

Other Start Functions ..................................................................................................... 7-17

ENGINE INSTRUMENTATION.......................................................................................... 7-18

General........................................................................................................................... 7-18

Turbine Speed (N2) ........................................................................................................ 7-18

Turbine Temperature (ITT)............................................................................................ 7-18

Fan Speed (N1) .............................................................................................................. 7-18

ENGINE SYNCHRONIZER SYSTEM................................................................................ 7-19

General........................................................................................................................... 7-19

Control ........................................................................................................................... 7-19




7-iii

Indication ....................................................................................................................... 7-19

Operation ....................................................................................................................... 7-19

THRUST REVERSERS (OPTIONAL EQUIPMENT) ........................................................ 7-20

General........................................................................................................................... 7-20

Aeronca Thrust Reversers.............................................................................................. 7-20

Dee Howard TR 4000 Thrust Reversers ........................................................................ 7-24

QUESTIONS......................................................................................................................... 7-29




7-v

ILLUSTRATIONS

Figure Title Page

7-1 Major Sections.......................................................................................................... 7-2

7-2 Airflow Diagram .................................................................................................... 7-4

7-3 Oil Servicing Access .............................................................................................. 7-5

7-4 ∆ P Indicator............................................................................................................. 7-5

7-5 Engine Instruments................................................................................................... 7-6

7-6 Oil System Schematic .............................................................................................. 7-7

7-7 Electronic Fuel Computer ........................................................................................ 7-9

7-8 Computer Inputs and Outputs................................................................................... 7-9

7-9 Fuel Computer and SPR Switches ......................................................................... 7-10

7-10 Fuel Counter ........................................................................................................... 7-11

7-11 Engine Fuel System................................................................................................ 7-12

7-12 Center Switch Panel ............................................................................................... 7-13

7-13 Throttle Quadrant ................................................................................................... 7-14

7-14 Left Start Circuit—SNs 35-002 through 35-147 and36-002 through 36-035........................................................................................... 7-15

7-15 Left Start Circuit—SNs 35-148 through 35-389, except 35-370and 36-036 through 36-047.................................................................................... 7-15

7-16 Installation Of AAK 81-1....................................................................................... 7-16

7-17 Left Start Circuit—SNs 35-370, 35-390 and Subsequentand 36-048 and Subsequent ................................................................................... 7-17

7-18 ENG SYNC Indicator ............................................................................................ 7-19

7-19 ENG SYNC Control Switches ............................................................................... 7-19

7-20 Thrust Reverser (Aeronca) ..................................................................................... 7-20

7-21 Thrust Reverser Control Panel (Aeronca) .............................................................. 7-21

7-22 Thrust Reverser (Dee Howard TR 4000) ............................................................... 7-24

7-23 Thrust Reverser Control Panel (Dee Howard) ....................................................... 7-25




INTRODUCTIONThis chapter describes the powerplants installed on Learjet 35/36 series airplanes. Inaddition to the powerplant, the chapter describes such engine-related systems as oil, fuel,ignition, engine controls and instrumentation, engine synchronization, the Aeronca andthe Dee Howard thrust reversers, and all pertinent powerplant limitations.

GENERALAll 35/36 series airplanes are powered by twoaft fuselage-mounted TFE731-2-2B turbofanengines. Optional thrust reversers are availableeither as a factory installation or as a retrofit.

The TFE731 series engine is manufacturedby the Garrett Turbine Engine Company atPhoenix, Arizona. The engine is a lightweight,

twin-spool turbofan. The fan is front mountedand gear driven.

Each engine develops 3,500 pounds of thrust,static at sea level, up to 72˚ F (+22˚C).

The modular design concept of the engine fa-cilitates maintenance.

#1 DCGEN

CHAPTER 7POWERPLANT




MAJOR SECTIONSFor descriptive purposes, the engine (Figure7-1) is divided into seven major sections asfollows:

1. Air inlet

2. Fan

3. Compressor

4. Combustor

5. Turbine

6. Exhaust

7. Accessory

AIR INLET SECTIONThe air inlet section is a specially designed,sound-reducing structure enclosing the fan

and its associated planetary gear-drive. The fanshroud is armored for blade containment.

FAN SECTIONThe fan section includes the single-stage axialfan, an integral spinner, and the fan planetarygear assembly which is driven by the low-pressure rotor. The rpm of the LP rotor isdesignated “N1” (commonly referred to as“fan speed”).

The planetary gear provides the requiredgear reduction for the fan. The rpm of the LProtor (N1 ) is read on the FAN SPEED indi-cator (Figure 7-5). Engine thrust is set usingthis instrument.




Figure 7-1. Major Sections

The fan performs two functions: (1) its outerdiameter accelerates a large air mass at arelatively low velocity into the full-lengthbypass duct, and (2) the inner diameter of thefan accelerates a smaller air mass to the four-stage axial-flow compressor.

COMPRESSOR SECTIONThe compressor sect ion includes a low-pressure (LP) compressor and a high-pres-sure (HP) compressor.

The LP compressor incorporates four axialstages. Stall-surge protection is provided forthe LP compressor by an au tomat ica l lycontrolled surge bleed valve.

The HP compressor consists of a single-stagecentrifugal impeller driven by the HP turbine.

COMBUSTOR SECTIONThe combustor section includes an annular re-verse-flow combustion chamber enclosed in aplenum. (Two 180˚ directional changes in air-flow take place through the combustor section.)

Twelve duplex fuel atomizers (spray nozzles)and two igni ter plugs are located in thecombustion chamber.

TURBINE SECTIONThe turbine section, consisting of a single-stage axial HP turbine and a three-stage axialLP turbine, is located in the path of theexhausting combustion air.

The single-stage HP turbine, rigidly joinedwith the HP compressor, forms the HP spoolwhich rotates independently about the LP rotorshaft. The rpm of the HP spool is designatedN2 (commonly referred to as turbine speed).The rpm of the turbine (N2 rpm) is read on theTURBINE SPEED indicator (Figure 7-5). Thisis a supporting engine operation instrument.

The three-stage LP turbine assembly is rigidlyconnected to the LP compressor assembly bya common shaft, forming the LP rotor. The for-ward end of the rotor shaft is geared to the plan-etary gear assembly which drives the fan.

EXHAUST SECTIONThe exhaust section consists of the primary andbypass air exhaust ducts. The primary exhaustsection directs the combustion gases to theatmosphere. The bypass air exhaust directsthe fan bypass air to the atmosphere.

ACCESSORY SECTIONThe accessory section consists of a transfergearbox and an accessory drive gearboxlocated on the lower forward side of the engine.The transfer gearbox is driven by a tower shaftand bevel gear from the HP spool. A horizontaldrive shaft interconnects the transfer gearboxto the accessory drive gearbox to drive thefollowing accessories:

• Oil pump

• Fuel pump and mechanical governorwithin the fuel control unit (FCU).

• Hydraulic pump

• DC generator

In addition to these accessories, a DC startermotor is mounted on the accessory drive gear-box to turn the HP spool for engine starting.

OPERATINGPRINCIPLESThe fan (Figure 7-2) draws air through theengine nacelle air inlet. The outer diameter of the fan accelerates a moderately large a i r mass through the fan bypass duct to provide direct thrust. The inner diameter of the fan accelerates a smaller air mass into theLP compressor.




Air is progressively compressed as it passesthrough the LP compressor, then to the HPcompressor where a substantial increase inpressure results. Air leaving the HP com-pressor is forced through a transition duct intoa plenum chamber surrounding the combustor.The compressed air is allowed to enter thecombustor through holes and louvers designedto direct the flow of combustion air and tokeep the flame pattern centered within thecombustor. Each of the duplex fuel nozzlessprays fuel in two distinct patterns, resultingin efficient , control led combustion. Themixture is initially ignited by the two igniterplugs. The expanding combustion gases,generating extremely high pressures, aredirected to the HP turbine which extractsenergy to drive the integral HP compressor

and the accessory section through the towershaft. The combustion gases continue to ex-pand through the three-stage LP turbine whichextracts energy to drive the LP compressorthrough the LP rotor shaft and the fan throughthe planetary gear. The combustion gases arethen exhausted through the exhaust duct. Theresulting thrust created by the combustionair adds to the thrust generated by the fanthrough the bypass air duct to produce thetotal propulsion force. At sea level, the fancontributes 60% of the total rated thrust,diminishing as altitude increases. At 40,000feet, the fan contributes approximately 40%of the total thrust . Engine core rotat ion(looking forward) is c lockwise, and fanrotation is counterclockwise.




Figure 7-2. Airflow Diagram

OIL SYSTEM

GENERALThe oil system provides cooling and lubrica-tion of the engine main bearings, the planetarygear, and the accessory drive gear.

Oil is contained in a tank on the right side ofthe engine. Access for servicing and levelchecking (Figure 7-3) is located on the out-board side of each nacelle.

The engine-driven oil pump incorporates onepressure element, four scavenge elements, anda pressure regulator (Figure 7-6).

The pressure element draws oil from the tankand provides pressure lubrication for all bear-ings and gears. The scavenge elements returnoil to the tank.

A bypass oil filter removes solids from the oil.A red pop-out ∆P indicator provides visual in-dication of a clogged filter. It can be checkedthrough a spring port (Figure 7-4) on the rightside of each engine nacelle. The indicatorbutton should be flush with the housing; if itis not, maintenance is required before flight.




LEFT ENGINE ACCESS RIGHT ENGINE ACCESS

Figure 7-4. ∆P Indicator

Figure 7-3. Oil Servicing Access

Oil cooling is fully automatic and is achievedby a combination of sectional air-oil coolersin the fan bypass duct and a fuel-oil coolermounted on the engine. Temperature andpressure bypass protection is provided forthe oil coolers.

Oil venting is provided and controlled by analtitude compensating breather-pressuriz-ing valve.

INDICATIONOil pressure is displayed on a single indicatorwith dual (L-R) needles (Figure 7-5) on the

engine instrument panel which require 26-VACelectrical power from the L and R OIL PRESScircuit breakers located on their respective Land R 26-VAC bus.

A single red LO OIL PRESS light on theannunciator panel provides warning of low oilpressure (Annunciator Panel). An optionalinstallation provides for dual lights labeled“L LO OIL” and “R LO OIL,” usually locatedoutboard of either engine FIRE handle. Thelight(s) is illuminated by a pressure switch oneach engine when pressure drops to 23 psi.




Figure 7-5. Engine Instruments

With the single LO OIL PRESS light installa-tion, the light is wired in parallel from the pres-sure switch on each engine. When this lightilluminates, the affected engine must be de-termined by checking the oil pressure indicator.

Oil temperature is displayed on individualgages (Figure 7-5) on the upper right side ofthe engine instrument panel. Power for thesegages is supplied through the OIL TEMP circuitbreaker located on the right essential bus.

A chip detector is installed in the scavengereturn line. It is used by maintenance to checkfor the presence of ferrous particles in the oil.As optional equipment, the detectors may beconnected to amber LH and RH ENG CHIPlights installed on the glareshield just to theright of the right-hand engine FIRE handles(Annunciator Panel).

OPERATIONFigure 7-6 illustrates operation of the engineoil system.




VENT

NOS. 4 AND 5BEARING

NO. 6BEARING

TRANSFERGEARBOX

BREATHERPRESSVALVE

CHIPDETECTOROIL TANK

SCAVENGE RETURN

REGULATOR

OIL PUMPS

FUEL IN

TEMPCONTROLBYPASSVALVE

PLANETARYGEARS

NOS. 1, 2 AND3 BEARINGS

OIL COOLER

FUEL OUT

FILTER∆PBYPASS

ACCESSORYGEARBOX

P S S S S

T P

AIR-OILCOOLER

SUPPLY

PRESSURE

SCAVENGE

LEGENDAIR

FUEL

ELECTRIC

Figure 7-6. Oil System Schematic

AIR-OILCOOLER

AIR-OILCOOLER

Figure 7-6. Oil System Schematic

FUEL SYSTEM

GENERALThe engine fuel system provides for fuelscheduling during engine starting and accel-eration to idle, operational acceleration anddecelerat ion, and steady-state operat ionthroughout the entire operating envelope ofthe airplane.

FUEL PRESSUREEngine fuel pressure is generated by a two-stage engine-driven pump. The centrifugal LPstage increases inlet fuel pressure from theairplane fuel system and directs fuel througha bypassable fuel filter with a ∆P indicatorbutton to the HP stage. The HP pump increasesfuel pressure to the valve required for efficientoperation of the fuel control unit (FCU). Inaddition, the HP fuel pump supplies the motive-flow fuel for operation of the fuel tank jetpumps. (See Chapter 5, “Fuel System.”)

MOTIVE-FLOW LOCKOUTVALVE AND PRESSUREREGULATORThe lockout valve remains closed initially duringengine start to ensure sufficient pressure to theFCU. The valve gradually opens fully as fuelpressure increases during the start. On earlierairplanes the motive-flow shutoff valve is alsoclosed when the START-GEN switch is moved toSTART. A pressure regulator maintains motive-flow line pressure for efficient jet pump operation.

FUEL CONTROL UNIT (FCU)The FCU schedules fuel flow to the fuel nozzles.Its primary mode of operation is the automaticmode (fuel computer on). In automatic the FCUresponds to electrical signals from the fuelcomputer. The secondary mode of operation isthe manual mode (fuel computer off or failed).In manual the FCU responds mechanically tothrust lever movement. The FCU includes (Fig-ure 7-11):

• A mechanical fuel shutoff valve, oper-ated by thrust lever movement betweenCUT-OFF and IDLE

• A DC potentiometer, mechanically po-si t ioned by thrust lever movement,which electrically transmits this aspower lever angle (PLA) to the com-puter for automatic operation

• A manual mode solenoid valve which isnormally energized open by the fuelcomputer for automatic mode opera-tion. It is deenergized closed for man-ual mode operation.

• A DC torque motor which schedulesfuel flow in automatic mode in responseto electrical signals from the computer.

• A mechanical flyweight governor, drivenby the engine fuel pump to (1) limit en-gine overspeed to 105% N2 in the auto-matic mode and (2) govern engine rpmrelative to thrust lever position in themanual mode.

• A pneumatically controlled meteringvalve, which (1) restricts fuel flow in theevent of engine overspeed and (2) sched-ules fuel flow in manual mode.

• Pneumatic circuits to channel and con-trol P3 bleed-air pressure to pneumati-cally position the metering valve.

• An ultimate overspeed solenoid valveenergized by the fuel computer at 109%N1 or 110% N2 to shut off fuel.

ELECTRONIC FUELCOMPUTER

GeneralTwo electronic fuel computers are located inthe tailcone area (Figure 7-7). They operate onDC power from the L and R FUEL CMPTRcircuit breakers on the left and right essentialbuses, respectively.




Automatic Mode OperationThe computer controls the fuel flow based onthrust lever position (PLA) and atmosphereconditions, while automatically maintainingN1, N2, and ITT within prescribed limits topermit optimum engine acceleration rates.The computer provides engine overspeedprotection and also controls the surge bleedvalve to prevent compressor stalls and surges.During engine start, the computer providesau tomat ic fue l enr ichment , s t a r t e r d i s -engagement, and termination of ignition andstandby fuel pump operation.

The computer receives input signals repre-senting the following engine parameters (Fig-ure 7-8):

• N1 (fan speed)

• N2 (turbine speed)

• PLA (power level angle)

• PT2 (inlet pressure)

• TT2 (inlet temperature)

• ITT (interstage turbine temperature)




Figure 7-8. Computer Inputs and Outputs

Figure 7-7. Electronic Fuel Computer

The computer analyzes these signals andproduces output signals which are sent to thetorque motor (to control fuel flow) and to thesurge bleed valve (to control compressor air-flow). Thrust lever movement mechanicallymoves a power lever angle potentiometer,which furnishes a variable electrical signal(PLA) to the computer. This is the commandinput for a specific thrust setting. Fuel flow ismetered by the torque motor to produce andmaintain the desired thrust. Inlet temperatureand pressure (PT2/TT2), N1, N2, and ITTsignals are used to optimize engine accelera-tion rates and limit thrust and temperaturewithin normal limits. By powering one or theother of the two surge bleed valve controlsolenoids, the computer opens or closes thesurge bleed valve during engine accelerationand deceleration to prevent compressor stallsand engine surges.

In automat ic opera t ion , the mechanica lflyweight governor section limits engineoverspeed to 105% N2 rpm. Should the 105%governing function fail, the computer energizesthe ultimate overspeed solenoid valve closedat 109% N1 or 110% N2 to shut off fuel flowto the engine.

IndicationThe computer constantly monitors input andoutput signals and, with the exception of ITTinput loss, automatically reverts to manualmode if these signals are lost. In this case, orif computer power is lost, the amber “L” or “RFUEL CMPTR” annunciator light illuminates.In some cases it may be possible to regainnormal operation. Refer to Section IV, Ab-normal Procedures, of the AFM.

Manual Mode OperationWhen the computer fails or is turned off, the fuelcontrol unit assumes manual control of the fuelmetering to the engine. The torque motor valveis deenergized and opens fully. The fuel flowis controlled by the mechanical flyweightgovernor section, functioning as an onspeedgovernor, utilizing the metering valve. Thesurge bleed valve automatically goes to the1/3-open position and remains there.




Figure 7-9. Fuel Computer andSPR Switches

START PRESSUREREGULATOR (SPR)Fuel enrichment is automatically controlled bythe fuel computer during starting up to 200˚C.It may be extended manually to assist engineacceleration during starting in cold ambienttemperatures (below 0˚ F) or during airstart atlow altitude/high airspeed if light-off doesnot occur withing 5 seconds after moving thethrust lever to IDLE. This additional fuel iscontrolled by a three-position switch (Figure7-9) labeled “SPR L” and “R.” The switch isspring-loaded to the center (off) position.When additional start fuel is required, theswitch must be held in the L or R position andreleased when ITT indicates between 300˚ Cand 400˚ C. SPR is a computer function andis available only in the computer-on mode.Manual SPR overrides the automatic temper-ature-limiting feature of the computer. There-fore, ITT monitoring during SPR operationis extremely important. It should be used onlyduring starting and discontinued when ITT isin the 300˚ C to 400˚ C range.

SURGE BLEED VALVEThe surge bleed valve functions to maintain asafe surge margin in the LP compressor byspilling some LP compressor air into thebypass duct, thus preventing LP compressorstalls and surges during acceleration anddeceleration when large LP-HP rpm mismatchoccurs. The surge bleed valve has three posi-tions: FULL OPEN, FULL CLOSED, and 1/3OPEN. Surge valve position is controlled bytwo fuel computer-operated solenoid valveswhich route P3 bleed air to a respective porton the surge valve. By energizing one solenoidvalve, the computer opens the surge valve,while energizing the other solenoid valvecloses it. By deenergizing both solenoid valves,the surge valve assumes the 1/3 OPEN posi-tion which automatically occurs if the com-puter fails or is switched off, thereby providingsome surge margin continuously while oper-ating in manual mode. In addition, the surgebleed valve will assume the FULL OPEN po-sition in the computer-on mode whenever the

PLA is 26˚ or less (42˚ on early model com-puters). During acceleration, the computerfirst signals the surge bleed valve to assumethe 1/3 OPEN position; if the surge margin can-not be maintained in this position, the com-puter will command the FULL OPEN position.The opposite is true during deceleration. Insummary, surge bleed valve position is a func-tion of the fuel computer, relative to N1, N2,and thrust lever angle.

FUEL FLOWFuel flow is sensed downstream of the FCU andis displayed on a dual-needle gage on the centerinstrument panel (Figure 7-5). The needles arelabeled “L” and “R,” and the gage is calibratedin pounds per hour times 1,000. Electrical poweris supplied directly from the battery-chargingbus through a 10-amp current limiter.

A resettable digital fuel counter (Figure 7-10)is located on the fuel control panel on the centerpedestal. The indicator is operated by the fuelflow indicating system and displays pounds offuel consumed. The indicator should be resetprior to engine starting.

FLOW DIVIDERThe flow divider splits fuel flow between theprimary and secondary manifolds to which thefuel nozzles are connected. During engine




Figure 7-10. Fuel Counter

starts, the flow divider blocks the secondarymanifold until fuel flow reaches 150 poundsper hour.

FUEL SPRAY NOZZLESThe twelve duplex fuel spray nozzles in thecombustion chamber consist of concentric pri-mary and secondary orifices that function toatomize the fuel delivered by the primary andsecondary fuel manifolds.

OPERATIONFigure 7-11 illustrates the operation of theengine fuel system in simplified format.

IGNITION SYSTEM

GENERALA solid-state, high-energy ignition systemconsists of a dual ignition exciter (mounted onthe engine) and two igniter plugs in the com-bustion chamber. Two ignition modes are avail-able: (1) automatic and (2) selective.

AUTOMATIC MODEAutomatic ignition occurs during enginestarting when the START–GEN switch on thecenter switch panel (Figure 7-12) is positionedto START and the thrust lever is moved from




LOW PRESSURE FUEL

ENGINE BLEED AIR

LEGEND

ELECTRICAL

MECHANICAL

FUEL FROMWING TANK

MOTIVE FLOWLOCKOUT/

REG VALVE

TO JETPUMPS

OVERBOARDPORT

SURGEVALVE

N1

N2

PT2TT2

ITT

POWER LEVER ANGLE

MANUALSHUTOFF VALVE

ULTIMATE OVERSPEEDSOLENOID(109% N1, 110% N2)

FUEL TOSPRAY NOZZLES

BLEED-AIRPRESSURE P

THRUSTLEVER

METERINGVALVE

FLYWEIGHTGOVERNOR

(105% N2)

MANUAL MODEADJUSTMENT

POTENTIOMETER

MANUAL MODE SOLENOID(POWERED OPEN,COMPUTER ON)

DC TORQUEMOTOR

HIGHPRESSPUMP

LOWPRESSPUMP

BYPASSINDICATOR

FILTER

FUELCOMPUTER

3

HIGH PRESSURE FUEL

Figure 7-11. Engine Fuel System

CUT–OFF to IDLE position. Ignition is au-tomatically terminated (in a computer-onmode) by an electronic speed switch in thecomputer at 45% or 50% N2 as determined bythe computer model installed. Power for au-tomatic ignition is provided by the L and R IGN& START circuit breakers on the left and rightpower buses respectively.

If the computer switch is off during a starter-assisted start or if the computer reverts tomanual mode during start, ignition will con-tinue until the START–GEN switch is movedout of the START position. Ignition will alsoterminate (computer on or off) if the thrustlever is moved forward to a position repre-senting approximately 70% N2.

SELECTIVE MODESelective ignition is controlled by two-posi-tion switches labeled “ AIR IGN L” and “ AIRIGN R”(Figure 7-12) located on the centerswitch panel. When the switch is positionedto AIR IGN, the igniters will operate contin-uously. Ignition power is supplied by the L orR AIR IGN circuit breakers on the left and rightessential buses, respectively.

Selective use of air ignition is required for alltakeoffs and landings, and also for windmillingairstarts. It may by used continuously whenflying in heavy precipitation, icing condi-tions, or turbulent air.

INDICATIONAn amber light (Figure 7-12 and AnnunciatorPanel section) located above each AIR IGNswitch will be on whenever power is beingsupplied to the associated ignition exciter. Theignition lights (if on) will dim if the NAV LTSswitch, located on the right switch panel, is on.

ENGINE CONTROLSEngine control is achieved by thrust levers(Figure 7-13) mounted on a quadrant on thecenter pedestal. The levers can be moved fromthe fully aft or CUT–OFF position throughthe IDLE position to the fully forward, max-imum power position. A stop is provided at theIDLE position which requires raising a re-lease trigger on the outboard side of each leverbefore the lever can be moved to CUT–OFF.

The thrust lever is connected to the FCU by acable. In the automatic mode, thrust lever po-sition is relayed to the computer as an electricalsignal from a potentiometer inside the FCUthat represents thrust lever angle. In manualmode, thrust lever movement changes P3,which operates the metering valve. The thrustlever also mechanically operates a rotary fuelshutoff valve.




Figure 7-12. Center Switch Panel

Optional thrust reverser levers are piggy-backmounted on the thrust lever. (See Thrust Re-versers, this chapter).

STARTERS

GENERALEach engine starter is powered through relayscontrolled by the GEN–OFF–START switchand the fuel computer (during computer-onstarts). A “soft start” feature incorporates aresistor to minimize the effect of the initialtorque on the mechanical drive components.After a 1.5-second delay, a relay operates toallow the starting current to bypass the resistorso that full electrical potential is available tocomple t e t he s t a r t . Au toma t i c s t a r t e rdisengagement occurs at 50% N2 (45% forSNs 35-245 and subsequent, 36-045 andsubsequent, and earlier airplanes equippedwith 1142 fuel computers). On SNs 35-370, 35-390, and 36-048 and subsequent, illumination

o f a r ed l i gh t unde r t he app rop r i a t eGEN–OFF–START switch indicates that thestarter is engaged. On earlier airplanes mod-ified by AMK 80-17, the red lights may beinstalled elsewhere on the instrument panel.

The GEN–OFF–START switches are locking-lever switches. They must be pulled out tomove to the START position. It is not neces-sary to pull out for movement to any other po-sition.

When either GEN–OFF–START switch ispositioned to START for a normal computer-on start, the start sequence is initiated for thatengine. The start sequence and circuitry for theleft engine are presented herein; they are iden-tical with those for the right engine.

There are three different designs for the relaycircuits which route power to the starter.

For SNs 35-002 through 35-147 and 36-002through 36-035, the relays are wired in parallel.

One relay is connected to the opposite generatorbus and the other to the battery-charging bus.This arrangement was designed to protect the275-amp current limiters during initiation ofeach engine start sequence (Figure 7-14).

For SNs 35-148 through 35-389, except 35-370,and 36-036 through 36-047, the relays are againwired in parallel, but both are connected to thebattery-charging bus (Figure 7-15). This designchange includes automatic single-generatorvoltage reduction on the ground and duringairstarts, resulting in 275-amp current limiterprotection when the first generator is switchedon and during initiation of the start sequence onthe second engine.

For SNs 35-002 through 35-389, except 35-370,and 36-002 through 36-047, two separatemodifications have been introduced to the start-ing circuits:

• AMK 80-17 provides a red starter-en-gaged light for each starter to provide in-dication of starter engagement (Figures7-14 and 7-15). Location of the lights isleft to customer specification.




Figure 7-13. Throttle Quadrant




Figure 7-14. Left Start Circuit—SNs 35-002 through 35-147 and 36-002 through 36-035

GPU

RBAT

RGENBUS

RGEN

275AMPCL

LBAT

BATTERY-CHARGING

BUS

275 AMPCL

NO. 1RELAY

LEFT STARTCIRCIUT

BOTH RELAYS: • ENERGIZED WITH START SWITCH IN START • DEENERGIZED BY FUEL COMPUTER (45% OR 50% N2) • IF FUEL COMPUTER IS OFF , RELAYS REMAIN ENERGIZED UNTIL START SWITCH IS MOVED FROM START POSITION.

RIGHT STARTCIRCIUT- SAMEAS LEFT

STARTERENGAGED

LIGHT

NO. 2RELAY

*WHEN INSTALLED BY AMK 80-17

LSTARTER*

GPU

BOTH RELAYS: • ENERGIZED WITH START SWITCH IN START • DEENERGIZED BY FUEL COMPUTER (45% OR 50% N2) • IF FUEL COMPUTER IS OFF , RELAYS REMAIN ENERGIZED UNTIL START SWITCH IS MOVED FROM START POSITION.

*WHEN INSTALLED BY AMK 80-17

STARTERENGAGED

LIGHT

NO. 2RELAY

LSTARTER

NO. 1RELAY

275 AMPCL

BATTERY-CHARGING

BUS

275AMPCL R

GENBUS

RGEN

LBAT

RBAT

RIGHT STARTCIRCIUT- SAMEAS LEFT

LEFT STARTCIRCIUT

*

Figure 7-15. Left Start Circuit—SNs 35-148 through 35-389, except 35-370 and 36-036 through 36-047

• AAK 81-1 installs a third starter relay inseries between the two existing relays andthe starter motor; the circuits which ener-gize the relays are redesigned. AMK 80-17is a prerequisite or concurrent requirementfor this modification (Figure 7-16).

For SNs 35-370, 35-390, and subsequent, and36-048 and subsequent, two starter relays arewired in series to the battery-charging bus,and the red starter-engaged lights are stan-dard. (Figure 7-17).

OPERATION

SNs 35-002 through 35-389,except 35-370, and 36-002through 36-047 with or withoutAMK 80-17With the airplane battery switches on, movingthe GEN–OFF–START switch to STARTconnects DC power through the IGN & STARTcircuit breaker to energize the starter relays.S t a r t e r engagemen t occu r s a l ong w i thillumination of the starter-engaged light ifAMK 80-17 i s i n s t a l l ed . Wi th t he fue lcomputer on, starter disengagement occursautomatically when power is removed from thestarter relay circuit. At this time the starter-engaged light (if installed) extinguishes.

SNs 35-002 through 35-389,except 370, and 36-002 through 36-047 whenincorporating AMK 80-17 andAAK 81-1All three starter relays must be energized topower the starter and illuminate the starter-en-gaged light. With the airplane battery switcheson, the two parallel relays are energized closedthrough the IGN & START circuit breakeranytime the GEN–OFF–START switch is in theOFF or START position. The third relay isalso energized from the IGN & START circuitbreaker, but only when the start switch is inthe START position. If the fuel computer is onfor the start, it will automatically deenergizethe third relay when N2 reaches 45% (or 50%,depending on which computer is installed).The starter is then disengaged and the starter-engaged l ight ext inguishes . Moving theGEN–OFF–START switch to GEN deener-gizes the two parallel relays to back up therelease of the third relay. If either of the twoparallel relays, plus the third relay, remain inthe closed position, the starter-engaged lightremains in the closed position, the starter-engaged light remains illuminated and thestarter remains powered. The only way to dis-engage the starter in this event is to removeelectrical power from the battery-charging




STARTER-ENGAGEDLIGHT

LSTARTER

THIRD RELAY:

• ENERGIZED WITH GEN/START SWITCH IN START • DEENERGIZED BY FUEL COMPUTER (45% OR 50%) AND GEN/START SWITCH IN OFF OR GEN

BATTERY-CHARGING

BUS

ORIGINAL RELAYS:

• ENERGIZED WITH GEN/START SWITCH IN OFF OR START• DEENERGIZED WITH GEN/START SWITCH IN GEN

Figure 7-16. Installation of AAK 81-1

bus by turning off both batteries and both gen-erators. If the starter-engaged light remainsilluminated after start, consult Section IV, Ab-normal Procedures, of the approved AFM.

SNs 35-370, 35-390 andSubsequent, and 36-048 andSubsequentThere are two relays in series between thebattery-charging bus and the starter (Figure 7-17). Both must be energized to power thestarter and illuminate the starter-engaged light.With the airplane battery switches on, the No.1 relay is energized through the IGN & STARTcircuit breaker anytime the GEN–OFF–STARTswitch is in the OFF or START position. TheNo. 2 relay is also energized from the IGN &START circuit breaker, but only when theGEN–OFF–START switch is in the STARTposition. If the fuel computer is on for thestart, it will automatically deenergize the No.2 relay when N2 reach 45%. The starter-en-gaged l i gh t ex t i ngu i she s . Mov ing t he

GEN–OFF–START switch to GEN deener-gizes the No. 1 relay to back up the release ofthe No. 2 relay. If both relays fail in theenergized position, the starter-engaged lightremains illuminated, and the starter remainspowered. The only way to disengage the starterin this event is to remove electrical powerfrom the battery-charging bus by turning offboth batteries and both generators. If thestarter-engaged light remains illuminated afterstart, consult Section IV, Abnormal Proce-dures, of the approved AFM.

OTHER START FUNCTIONSIn addition to the starter, a number of othercircuits are affected when the GEN–OFF–STARTswitch is placed in START. The standby fuelpump in the associated wing is energized, the ig-nition is armed, and the Freon air-conditioningand auxiliary heating systems are disabled.

Additionally, on SNs 35-002 through 35-057and 36-002 through 36-017, the motive-flow




L STARTER-ENGAGED

LIGHT

LBAT

RBAT

GPU

BATTERY- CHARGING

BUS

275 AMPCL

275AMPCL

RGEN

RGENBUS

NO. 2RELAY

NO. 1RELAY

LEFT STARTERCIRCUIT

RIGHT START CIRCUIT— SAMEAS LEFT

• ENERGIZED WITH START SWITCH IN OFF OR START• DEENERGIZED IN GEN

• ENERGIZED WITH START SWITCH IN START• DEENERGIZED BY COMPUTER ABOVE 45% N2• IF FUEL COMPUTER IS OFF , RELAY REMAINS ENERGIZED UNTIL START SWITCH IS MOVED FROM START POSITION

LSTARTER

Figure 7-17. Left Start Circuit—SNs 35-370, 35-390, and Subsequent,and 36-048 and Subsequent

control valve must automatically cycle closed,or the starter relays will not energize. Whenthe associated thrust lever is moved from CUT-OFF to IDLE, a switch in the throttle quadrantcloses and activates the ignition system,causing the ignition light to illuminate. Whenturbine speed reaches 45% or 50% (depend-ing on computer model), the fuel computerremoves power from the start relay(s). Thiscauses the starter to disengage and terminatesignition and standby pump operation. Thestart sequence can be aborted at any point byplacing the thrust lever to CUT-OFF and theGEN–OFF–START switch to OFF. If enginestart is accomplished with the fuel computeroff, the starter is not automatically disengagedafter starting. The pilot must position theGEN–OFF–START switch to OFF to terminatestarter engagement and ignition.

Af t e r t he eng ine r eaches i d l e rpm, t heGEN–OFF–START switch may be placed tothe GEN position. The generator may be turnedon when a GPU is connected; however, it ispreferable to place the GEN–OFF–STARTswitch to OFF after starting engines until theGPU is disconnected.

On SNs 35-002 through 35-147 and 36-002through 36-035 (during battery start), afterthe first engine is started, one battery switchmust be turned off prior to selecting GEN onthe GEN–OFF–START switch. This actionreduces the initial load on the generator andprotects the 275-amp current-limiter. On laterairplanes this procedure is not required, andthe GEN position may be selected immediatelyafter start.

When the GEN–OFF–START switch is movedfrom START, those systems which weredisabled during the start can now be operated.

ENGINEINSTRUMENTATION

GENERALThe primary engine instruments are mountedin two vertical rows on the center instrument

panel (Figure 7-5). From top to bottom theseinstruments are:

• Turbine speed (N2 rpm)

• Turbine temperature (ITT)

• Fan speed (N1 rpm)

TURBINE SPEED (N2)Turbine speed (N2 rpm) is remotely sensedby a dual monopole transducer installed in thetransfer gearcase. One output signal is sent tothe turbine speed (N2) indicator, and another tothe fuel computer. This indicator includes ananalog scale and pointer calibrated in percent-age of maximum design rpm, and a digitalcounter, calibrated in tenths of percent. A redOFF flag will appear on the face of the indica-tor to indicate loss of DC power to the indica-tor. The indicators are powered through the LR TURB RPM circuit breakers located on theleft and right main buses, respectively.

TURBINE TEMPERATURE (ITT)Turbine temperature is sensed by ten parallel-wired thermocouples located between the HPand LP turbines. An averager circuit providestwo output signals—one to the turbine tem-perature indicator, and the other to the fuelcomputer. The indicator includes an analogscale and pointer, calibrated in degrees Cel-sius, and a digital counter, calibrated to thenearest whole degree. A red OFF flag will ap-pear on the face of the indicator. The indica-tors are powered through the L and R ITTcircuit breakers located on the left and rightessential buses, respectively.

FAN SPEED (N1)Rotation of the LP rotor is sensed by a dualmonopole transducer installed under a coverplate at the aft end of the LP rotor shaft. Oneoutput signal is sent to the fan speed (N1) in-dicator, and the other to the fuel computer. Allother operational aspects of the indicator areidentical with the turbine speed indicator ex-cept that the indicators are powered through




the L and R FAN RPM circuit breakers locatedon the left and right essential buses, respec-tively.

NOTEThe fan speed (N1) indicators arethe primary power indicators.

ENGINE SYNCHRONIZERSYSTEM

GENERALThe engine synchronizer system is installed onairplane SNs 35-067 and subsequent, and 36-018 and subsequent as standard equipment. Itincorporates a synchronizer control box whichuses N1 or N2 inputs from both engine fuelcomputers to enable automatic or manual syn-chronization of the engines.

CONTROLThe system incorporates a single R ENGSYNC indicator located on the pilot’s lowerinstrument panel (Figure 7-18), and two ENGSYNC switches located immediately belowthe thrust levers (labeled “SYNC–OFF” and“TURB–FAN,” respectively (Figure 7-19).The system operates manually (with theSYNC–OFF switch in the OFF position) or au-tomatically (with the SYNC–OFF switch in theSYNC position) to maintain the right enginefan or turbine in sync with the left engine fanor turbine as determined by the TURB-FANswitch.

INDICATIONAn amber ENG SYNC light (Annunciator Panelsection) on the glareshield annunciator panelwill be illuminated anytime the nose gear isdown and locked with the SYNC–OFF switchin the SYNC position. The R ENG SYNC(SLOW/FAST) indicator indicates right enginerpm deviation from that of the left engine.

OPERATIONManual synchronization is accomplished byselecting OFF on the SYNC–OFF switch. TheR ENG SYNC indicator shows SLOW or FASTout-of-sync condition of the right-hand engine(slave engine) relative to the left-hand engine(master engine). The pilot has the option ofselecting either N2 or N1 as the rpm referenceby using the TURB–FAN switch.




Figure 7-18. ENG SYNC Indicator

Figure 7-19. ENG SYNC Control Switches

Automatic synchronization is accomplishedby selecting SYNC on the SYNC–OFF switch.If the engines are within approximately 2.5%rpm of each other, the right engine will auto-matically synchronize to the left engine. It isnecessary, therefore, to manually sync to within2.5% initially. As in manual sync, either the N2OR N1 may be selected as the rpm reference.

DC electrical power is supplied to the systemfrom the left essential bus through the leftFUEL CMPTR circuit breaker to the L FUELCMPTR switch.

The amber ENG SYNC annunciator lightserves as a reminder that the system should beturned off.

The engine sync system is inoperative if eitherfuel computer is off or failed.

THRUST REVERSERS(OPTIONAL EQUIPMENT)

GENERALThe Learjet 35/36 series airplanes may beequipped with either a cascade thrust reversersystem, manufactured by Aeronca, Inc., or a tar-get reverser system (TR 4000), manufactured

by the Dee Howard Co. Effective with airplaneSNs 35-507 and 36-054, either system is avail-able for retrofit, but only the target system isavailable during production.

AERONCA THRUSTREVERSERS

GeneralThe Aeronca thrust reverser system incorpo-rates a translating structure (Figure 7-20)which forms the after body of the engine na-celle. When deployed, it exposes cascadevanes, while simultaneously operating twoblocker doors that block engine exhaust ducts,thereby deflecting all exhaust in a forwarddirection through the cascade vanes.

The translating structure is deployed and stowedby an air motor using HP bleed air from theassociated engine and sequenced electrically bymicroswitches operated by the reverser levers(Figure 7-13) . The sys tem incorpora tesautomatic stow and stow-prevention features tominimize the possibility of inadvertent de-ployment on the ground and in flight, andinadvertent stow at high reverse thrust settings.The system is self-arming on the ground throughcontrol circuits operating through the landinggear squat switch relay box.




Figure 7-20. Thrust Reverser (Aeronca)

ControlThe reverser levers control the deploy andstow cycles and engine power when thereversers are deployed. The thrust reversercontrol panel (Figure 7-21) is located in thecenter of the glareshield above the annunciatorpanel. It incorporates a rocker selector switchfor normal and emergency operations, sevenannunciator l ights which provide visualevidence of normal sequencing and certainabnormal conditions, and a test switch forperforming system test functions.

Thrust Reverser Control Panel

TEST Button

The TEST button provides a means of checkingoperation of the bleed valve and, on someairplanes, also checks the blocker door positionindicating circuits. When depressed, the whiteBLEED VALVE lights should illuminate, and,on a i rp lanes incorpora t ing AMK 81-6(installation of blocker door position indicator[DPI] switches), the white UNLOCK lights will

flash to indicate that the blocker doors arecorrectly stowed.

BLEED VALVE Lights

In addition to the test function above, thewhite BLEED VALVE lights illuminate asreverse thrust is increased to indicate that HPbleed air to the air motors is shut off. Thisprevents inadvertent stow commands.

DEPLOY Lights

The two white DEPLOY lights illuminate whenthe corresponding thrust reverser is fully deployed.Both DEPLOY lights must be illuminated;otherwise, the reverser lever solenoid interlockswill not release to permit thrust increase.

UNLOCK Lights

In addition to the test function above, the twowhite or amber UNLOCK lights illuminatesteady while the translating assembly is intransit during the deploy and stow cycles; thatis, the reversers are not fully deployed orlocked in the stowed position.




Figure 7-21. Thrust Reverser Control Panel (Aeronca)

NORM-EMER STOW Switch

In the NORM position, the red rocker switchprovides the electrical circuitry for all normaland automatic functions. In the EMER STOWposition, all normal electrical circuits arebypassed, and a separate circuit applies stowcommands to the reversers.

EMER STOW Light

The amber EMER STOW light illuminateswhenever the NORM–EMER STOW switch isin the EMER STOW position and the emer-gency stow circuits have been activated, thusrendering the normal system inoperative.

System Operation

Arming

The reversers are automatically armed fornormal operation when the following con-ditions exist:

• The T/R circuit breakers are closed.

NOTEThe T/R POS and T/R EMER STOWcircuit breakers are located on the leftmain bus, and the T/R CONT circuitbreaker is located on the right main bus.

• The airplane is on the ground (squatswitch relay box is in the ground mode).

• The NORM-EMER STOW switch is inthe NORM position.

• Both thrust levers are a t the IDLEposition.

Electrical power for deployment will not beavailable unless both thrust levers are at IDLEand both reverser levers are raised to thedeploy position.

Deploy

When the reverser levers are moved to thedeploy position (the first “hard stop”), themain-thrust levers are locked in the IDLEposition, and N1 rpm increases to approxi-mately 55%-60%. Switches are operated byeach reverser lever to complete circuits thatenergize pneumatic latch releases (two perreverser) to unlock the translating assembly.Switches on each latch function to:

1. Illuminate the UNLOCK lights.

2. Shut off bleed air to the windshieldheat, nacelle heat, and wing/stabi-lizer heat systems (for approximately3 seconds).

3. Energize the air motor directionalcontrol solenoid valve which routesHP bleed air through an air inletva lve , in to the a i r motor on therespective reverser.

The air motor transmits torque to drive the trans-lating structure aft, exposing the cascade vanes.As the assembly approaches its aft limit, theblocker doors are closed, the DEPLOY lightilluminates, the UNLOCK light extinguishes,and the reverser lever solenoid-operatedinterlock releases.

The reverser lever solenoid-operated inter-lock prevents movement of the reverser leversaft of the idle-deploy position until bothDEPLOY lights are illuminated. If the pilot isapplying excessive aft pressure on the reverserlevers when the DEPLOY lights illuminate, thesolenoid-operated interlock will not release,and reverse thrus t above approximate ly55%–60% N1 will not be possible. The inter-lock will release when aft pressure is relaxed.

For single-engine reversing, both thrust leversmust be at IDLE, and both reverser leversmust be raised to the deploy position in orderto deploy the reverser on the operating en-gine. Since the reverser on the inoperative en-gine will not deploy, the solenoid interlock willnot release; therefore, reverse thrust on the op-erating engine is limited to reverse idle (55%to 60% N1).




Reverse ThrustAfter both DEPLOY lights illuminate (two-en-gine operation) and the solenoid-operatedinterlocks release, the reverser levers can bepulled further aft to increase engine power.There is no limitation on engine thrust whenusing reverse except that the normal forwardthrust limitations still apply.

Stow PreventionAs reverse thrust (N1) is increased, a pressureswitch in each reverser system causes the bleedvalve on the corresponding system to openand illuminate the BLEED VALVE light. Thisisolates the bleed-air system from the airmotors (by closing the air inlet valve) untilstow is commanded by the reverser levers orwith the EMER STOW switch, thus prevent-ing inadvertent stow on either engine whichcould cause significant thrust asymmetry.

At 60 KIAS the reverser levers should besmoothly returned to the idle-deploy position.

When the engines have reached the reverse-idlerpm (approximately 55%–60% N1), the pilotmay stow the reversers by moving the reverserlevers to the full forward position.

Normal StowWhen the reverser levers are moved from theidle-deploy position to the full forward anddown position (stow), they operate switchesthat send a stow signal to the directionalcontrol solenoid of the air motor. The bleedvalve closes, admitting bleed air into the airmotor, causing it to drive the translatingst ructure toward the s tow posi t ion. TheDEPLOY lights extinguish and, simultane-ously, the UNLOCK lights illuminate. Whenthe thrust reversers are fully stowed and thepneumatic latches engage the translatingstructure, the UNLOCK lights will extinguish.As in the deploy cycle, bleed air is shut off tothe windshield, nacelle, and wing/stabilizerheat systems for approximately 3 secondswhen the stow cycle is initiated.

Abnormal Indications

UNLOCK Light (Steady)

If either thrust reverser fails to completely stow,or if any of the pneumatic latches fails to engageafter stowing, the corresponding UNLOCKlight will remain illuminated. Also, if a pneu-mat ic la tch d isengages a t any t ime, thecorresponding UNLOCK light will illuminate.

The automatic stow circuit is activated anytimean UNLOCK light illuminates with the reverserlevers in the stowed position. Stow pressure willbe applied until the UNLOCK light extinguishes.

UNLOCK Light (Flashing)

A flashing UNLOCK light is a function ofmodificat ion AMK 81-6 ( instal la t ion ofb locke r doo r pos i t i on i nd i ca to r [DPI ]switches). Proper stowing of the blocker doorsis essential for continued operation. An -undetected jammed blocker door could resultin inadvertent deployment of the affectedthrust reverser. Each blocker door (upper andlower) actuates a DPI switch when in the prop-erly stowed position. If the stow cycle is com-plete (latches engaged) and one of the DPIswitches is not actuated, the correspondingUNLOCK light will flash to indicate thejammed blocker door. Since damage to thesystem has occurred, repairs are required priorto the next takeoff.

A flashing UNLOCK light at any other timeindicates a malfunctioning DPI switch, butthe blocker doors are still properly stowed.This does not preclude operating the reverserson landing.

BLEED VALVE Light

When the reversers are stowed, illuminationof a BLEED VALVE light means that the bleedvalve is open. This isolates the bleed-air-sys-tem from the air motor, and deployment ofthe affected reverser will not be possible.




Automatic StowThe thrust reversers incorporate an auto-stowprovision. If any of the pneumatic latches release(UNLOCK light illuminates) when the reverserlevers are stowed, electrical power from theT/R CONT circuit breaker is applied to open thebleed-air valve and to the directional solenoid,causing the air motor to stow the translatingstructure. Stow pressure will be maintaineduntil the UNLOCK light extinguishes.

Emergency StowThe NORM–EMER STOW switch is normallyleft in the NORM position. The EMER positionis designed for inadvertent UNLOCK orDEPLOY conditions when the reverser leversare stowed. Power is provided by the TR EMERSTOW circuit breaker on the left main bus.

In the case of the UNLOCK or DEPLOY con-dition in flight, the EMER position on theswitch in not functional with the thrust leversset at any power setting above approximately70% N1. It is therefore imperative that if theEMER selection is made for any reason due toa reverser malfunction, the amber EMER STOWindicator light be monitored. If the powersetting is sufficiently high to prelude the emer-gency stow circuits from functioning, the amberlight will not illuminate, and the appropriate

thrust lever must be retarded until the lightilluminates. Illumination of the EMER STOWlight gives visual indication that the emergencystow circuits have, in fact, been activated.

In the event of a system malfunction whileintentionally operating in the reversing range,there is nothing to preclude use of the EMERSTOW selection at any time, and doing so willimmediately command all components to stow,and illuminate the amber EMER STOW light.

All thrust reverser normal, abnormal, andemergency procedures are contained in thesupplement section of the approved AFM.

DEE HOWARD TR 4000THRUST REVERSERS

GeneralThe Dee Howard thrust reversers incorporatea hydraulically operated system (Figure 7-22) consisting of a pair of clamshell doorsforming the afterbody of the engine nacelle.When deployed, the doors deflect all exhaustin a forward direction. The reverser hydraulicsystem is integral with the airplane’s hydraulicsystem for normal operation. It is equippedwith a separate accumulator and a one-waycheck valve which enable one deploy and stow




Figure 7-22. Thrust Reverser (Dee Howard TR 4000

cycle in the event of airplane hydraulic sys-tem failure. The accumulator preload pres-sure is 900–1,000 psi.

An automatic emergency stow system, whichincludes an automatic throttle-retard feature,is incorporated to provide protection againstinadvertent deployments.

Two pairs of spring-loaded latches (one paireach side) secure the doors when stowed. Hy-draulic actuators operate each pair of latches,the doors, and a throttle-retard mechanism. Hy-draulic pressure is supplied by a selector valvewhich incorporates four separate solenoid valvesthat are electrically sequenced by microswitches.One of the solenoid valves (the isolation valve)blocks hydraulic pressure at the selector valveinlet until the system is fully armed. The otherthree solenoid valves are for latch release, doorstow, and door deploy.

ControlThe reverser levers control the deploy andstow cycles and engine power when thereversers are deployed. The thrust reversercontrol panel (Figure 7-23) is located in thecenter of the glareshield above the annuncia-t o r pane l . I t i n co rpo ra t e s two ARM–OFF–TEST switches (one for each reverser)which provides system arming, disarming,

and testing. Four annunciator lights (two foreach reverser) provide visual indicat ion of normal sequencing and certain abnormalconditions.

Thrust Reverser Control Panel

ARM–OFF–TEST Switches

Arming, disarming, and testing are accom-plished for each reverser by use of the re-spective ARM–OFF–TEST switch. The ARMposition is wired in series with the groundmode of the squat switch relay box, as well asan IDLE switch on the respective thrust lever.The system, therefore, will only ARM whenthe airplane is on the ground and the thrustlevers are at IDLE.

The TEST posi t ion provides a means ofchecking operation of the hydraulic isolationvalve. When TEST is selected, the isolationvalve is energized open, and hydraulic pressureis applied to a pressure switch that illumi-nates the ARM light.

The Arm position enables all sequencing mi-croswitches and energizes the isolation valveopen. Illumination of the ARM light indicatesthat the isolation valve has opened and hydraulicpressure is available to the other three solenoidvalves for normal sequencing. The OFF position




Figure 7-23. Thrust Reverser Control Panel (Dee Howard)

completely disarms the deploy circuits withoutdisarming the automatic emergency stow system.

ARM Lights

The green ARM lights illuminate in conjunc-tion with the TEST and ARM functions as de-scribed above. However, should the ARM lightilluminate at any other time (i.e., in flightwith the ARM–TEST switch in the OFF po-sition), it indicates that two inboard (or out-boa rd ) doo r l a t che s a r e un locked , andautomatic activation of the emergency stow cir-cuit has occurred. This will be annunciated bya flashing DEPLOY light.

DEPLOY Lights

The amber DEPLOY lights flash during allstow/overstow cycles and illuminate steadywhen the respective reverser is in the fully de-ployed position during a normal deployment.A flashing DEPLOY light at any other timeindicates that one or more of the door latchesare unlocked (Automatic Emergency Stow thischapter).

System Operation

Arming

The reversers are armed for normal operationas follows:

• The T/R circuit breakers (two for eachreverser) are closed.

• The airplane is on the ground (squatswitch relay box is in the ground mode).

• The respective ARM–TEST switch is inthe ARM position.

• The respective thrust lever is at the IDLEposition.

• The respect ive green ARM light isilluminated.

Deploy

Raising the respective reverser lever to theidle deploy position (the first hard stop) locksthe main thrust lever at IDLE and contacts adeploy switch that energizes the latch andstow solenoid valves open. This di rectshydraul ic pressure to both la tch releaseactuators, the stow side of the door actuator,and the throttle-retard actuator. The resultingdoor “overstow” condition unloads the spring-loaded latches so that the latch release actu-ators can release them, and simultaneously, thethrottle-retard actuator is operated by the stowpressure. When the latch release actuatorsengage their respective unlock switches, thestow solenoid valve is deenergized closed, thelatch solenoid valve remains energized open,and the deploy solenoid valve is energizedopen. This directs hydraulic pressure to deploythe doors, while pressure is maintained on thelatch release actuators. When fully deployed,the doors contact a switch that illuminates theDEPLOY light steadily, deenergizes the latchsolenoid valve closed, and energizes the re-verser lever solenoid-operated lock, which re-leases to allow the reverser lever to be pulledfurther aft to increase reverse thrust.

Reverse Thrust

When the DEPLOY light(s) illuminates and thereverser lever solenoid-operated lock(s)releases, the reverser lever(s) can be pulled fur-ther aft to increase N1 to achieve the desiredresults. A second hard stop limits N1 rpm toapproximately 75%, which constitutes maxi-mum reverse thrust.

At 60 KIAS the reverser levers should besmoothly started toward the idle deploy position.

Use of maximum reverse power below 50KIAS could cause reingestion of exhaust gasesor possible foreign object damage.

After the engines have reached reverse-idlerpm (approximately 30% N1), the pilot canstow the reverser levers by returning them tothe full forward position.




Normal Stow

Returning the respective reverser lever to thefull forward and down position unlocks themain thrust lever and contacts a stow switch.This deenergizes the deploy solenoid valveclosed and energizes the stow solenoid valveopen, directing hydraulic pressure to stow thedoors and operate the throttle-retard actua-tor. The “overstow” condition allows the fourspring-loaded latches to lock into place andbreak contact with their respective latchswitches. This deenergizes the stow solenoidvalve closed, which shuts off hydraulic pres-sure to the door actuator and the throttle-retardactuator. Exhaust gas pressure and springsreturn the doors to their normal positionagainst the latching hooks.

Abnormal Indications

ARM Light Fails to Illuminate during Test

If the ARM light fails to illuminate when TESTis selected on the ARM–TEST switch, theisolation valve has failed to respond correctly,hydraulic pressure is not available, or the pres-sure switch is faulty; also, the affected reverserwill be inoperative.

ARM Light Fails to Illuminate during Normal Arming (On theGround at Idle)

If the ARM light fails to illuminate when ARMis selected on the ARM–TEST switch (on theground with thrust levers at IDLE), possiblemalfunctions are:

• Isolation valve failure

• No hydraulic pressure available

• Pressure switch failure

• Thrust lever IDLE switch failure

• Faulty squat switch relay circuitry

Steady ARM Light (ARM–TESTSwitch Off)

Steady illumination of the ARM light withthe ARM–TEST switch off indicates that twodoor latches on the same side (inboard oroutboard) are unlocked. Illumination of theARM light indicates activation of the auto-matic emergency stow circuit. This will beaccompanied by a flashing DEPLOY light.

Flashing DEPLOY Light

A flashing DEPLOY light indicates that oneor more of the door latches are unlocked.

Automatic Emergency StowThe automatic emergency stow system isdesigned to prevent inadvertent deployment atany time (ARM–TEST switch off or on). If twolatch position switches on the same side(inboard or outboard) indicate an unlatchedcondition for the doors, the result is as follows:

• The i so l a t i on va lve opens , wh ichilluminates the ARM light.

• The DEPLOY light begins to flash.

• The stow solenoid valve is energizedopen, which applies stow pressure tothe door actuator and to the throttleretard actuator, which retards the thrustlever to the idle position.

The steady ARM light and flashing DEPLOYlight remain on until the latches return to thelatched position or until power is removedfrom the control circuits.




Automatic Throttle RetardAutomatic throttle retard is designed primarilyto minimize severe thrust asymmetry whichmay occur as a result of inadvertent deploymentof a reverser during high thrust settings. Thisis accomplished by use of the overstow cyclehydraulic pressure to operate a throttle retardactuator, resulting in mechanical repositioningof the thrust lever to the IDLE position.

This feature can be checked on the ground bydeploying the reversers, pulling the reverserlevers toward a higher power position, thenquickly returning the reverser levers to the

stow position and pushing forward on the thrustlevers. Resistance to thrust lever movement will befelt until completion of the stow cycle.

All thrust reverser normal, abnormal, andemergency procedures are contained in thesupplement section of the approved AFM.

Do not use thrust reversers to back up the air-plane, and do not deploy the drag chute andthrust reversers simultaneously.

Adequate airplane control has been demon-strated with a 20-knot crosswind component, butthis value is not considered to be limiting.







l. The TFE731-2-2B engine provides 3,500pounds of thrust at:A. Sea level up to 72˚ F (22˚ C)B. All altitudes and temperaturesC. Sea level at any temperatureD. All altitudes up to 72˚ F (22˚ C)

2. The engine LP rotor (N1) consists of:A. A four-stage, axial-flow compres-

sor and a single-stage certrifugal compressor

B. A single-stage fan and a three-stage,axial-flow compressor

C. A single-stage fan, a four-stage, axial-flow compressor, and a three-stage,axial-flow turbine

D. A four-stage, axial-flow compressorand a four-stage, axial-flow turbine.

3. During a normal ground start, the ignitionlight should come on when:A. N2 reaches 10%B. The START–GEN switch is moved to

STARTC. The thrust lever is moved to idleD. N1 reaches 10%

4. The engine HP spool (N2) consists of a:A. Three-stage axial compessor and a

four-stage radial turbineB. Single-stage centrifugal compressor

and a two-stage axial turbineC. Two-stage axial compressor and a

single-stage axial turbineD. Single-stage centrifugal compressor

and a single-stage axial turbine

5. The engine instruments (N1, N2 and ITT)are powered by:A. Self-generating tachometersB. The 26-VAC busesC. The essential busesD. The DC main and essential buses

6. Electrical power for engine oil pressureindication is provided by the:A. Left and right essential busesB. Inverters through the 26-VAC busC. Battery charging busD. Pilot’s and copilot’s 115-VAC buses

7. The primary engine thrust indicatinginstrument is the:A. Turbine (N2)B. ITTC. Fan (N1)D. Fuel flow

8. The maximum ITT during engine start is:A. 832˚ CB. 870˚ C for ten secondsC. 795˚ CD. 860˚ C

9. The maximum transient ITT during take-off is:A. 860˚ C for five minutesB. 870˚ C for ten secondsC. 880˚ C for five secondsD. 865˚ C for five minutes

10. What is the maximum acceptable engineoil temperature?A. 140˚ CB. 70˚ CC. 130˚ CD. 127˚ C

11. During computer-on operation, the surgebleed valve:A. Is controlled by the fuel computerB. Remains closedC. Remains at 1/3 OPEN positionD. Has no function

QUESTIONS

12. During computer-on operation, what en-gine overspeed protection is provided?A. Only 109% N1 and 110% N2 ultimate

overspeed shutoffB. Only 105% N2 mechanical governorC. Only 109% N1 ultimate overspeed

shutoffD. Only 105% N2 mechanical governor

and 109% N1/110% N2 ultimate over-speed shutoff

13. Which o f the fo l lowing s t a t ements regarding fuel control is true in the eventof airplane electrical failure?:A. Fuel control remains in the NORMAL

mode, but overspeed protection is lost.B. Fuel control reverts to the MANUAL

mode , and u l t ima t e ove r speedprotections is lost.

C. Fuel control reverts to the MANUALmode , bu t 109% N 1 ove r speedprotection is still available if thecomputer switch is on.

D. Fuel control remains in the NORMALmode wi th no loss of overspeedprotection.

14. If the SPR switch is used during enginestart, it should be released to OFF when:A. ITT begins to rise.B. ITT reaches 200˚ CC. ITT reaches 300˚ to 400˚ CD. Engine idle rpm stabilizes.

15. The ENG SYNC light indicates:A. Engine sync i s no t tu rned on o r

has failed.B. Engine sync is operating properly.C. Engine sync is turned on, and the nose

landing gear is locked in the DOWNposition.

D. The engines are synchronized.

16. When performing a fuel control governorcheck, N2 rpm increases rapidly. The pilot must:A. Turn on the fuel computer switch

immediately, allow rpm to stabilize atidle, shut down the engine, and havethe system checked.

B. Pull the associated fire T-handle, setthe fuel computer switch to manual,and restart the engine.

C. Wait until N2 rpm stabilizes at 105%,then turn on the fuel computer switch,and, when N2 drops to idle, shut theengine down.

D. Turn on the fuel computer switch,and, i f the rpm drops to idle , no further action is necessary.

17. The major portion of total thrust at lowaltitudes is developed by the:A. FanB. LP turbineC. Core engineD. HP turbine

18. The maximum allowable N1 under all operating conditions is:A. 101% to 103% maximum continuousB. 105% for one minuteC. 103% to 105% for five secondsD. 101.5% for five minutes




8-i

CHAPTER 8FIRE PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................... 8-1

GENERAL............................................................................................................................... 8-1

ENGINE FIRE DETECTION AND INDICATORS............................................................... 8-2

Sensing Elements and Control Units ............................................................................... 8-2

FIRE or ENG FIRE Lights............................................................................................... 8-3

Fire Detection System Test .............................................................................................. 8-3

ENGINE FIRE-EXTINGUISHING ........................................................................................ 8-3

Extinguisher Containers................................................................................................... 8-3

FIRE or ENG FIRE T-Handles and ARMED Lights....................................................... 8-4

Exterior Extinguisher Discharge Indicators ..................................................................... 8-4

PORTABLE FIRE-EXTINGUISHERS................................................................................... 8-4

QUESTIONS ........................................................................................................................... 8-6




8-iii

ILLUSTRATIONS

Figure Title Page

8-1 Engine Fire Detection System.................................................................................. 8-2

8-2 Engine Fire Warning Lights and Controls (LH)....................................................... 8-3

8-3 System Test Switch .................................................................................................. 8-3

8-4 Engine Fire-Extinguishing System........................................................................... 8-5

8-5 Fire-Extinguisher Discharge Indicators.................................................................... 8-4

8-6 Portable Fire-Extinguisher ....................................................................................... 8-5




INTRODUCTIONThe Learjet 35/36 series airplanes are equipped with engine fire detection and fire-extinguishing systems as standard equipment. The systems include detection circuits whichgive visual warning in the cockpit and controls to activate one or both fire-extinguisherbottles. There is a test function for the fire detection system. One or two portable fire-extinguishers are provided.

GENERALThe engine fire protection system is composedof three sensing elements, two control units(one for each engine) located in the tailcone,one warning indicator light for each engine,two fi re -ext inguisher bot t les which areactivated from the cockpit, and a fire detectioncircuit test switch. The fire-extinguishingsystem is a two-shot system; if an engine fire

is not extinguished with actuation of the firstbott le, the second bott le is available fordischarge into the same engine. The fire bottlesare located in the tailcone of the airplane.Exterior discharge indicators provide a visualindica t ion i f e i ther fi re bot t le has beendischarged manually or by thermal expansion.

FIRE PULL

FIREWARN

CHAPTER 8FIRE PROTECTION




ENGINE FIREDETECTIONAND INDICATORS

SENSING ELEMENTSAND CONTROL UNITSWithin each engine cowling are three heat-sensing elements—one mounted on the enginepylon firewall, one mounted around the lowerengine accessory section, and one surround-ing the engine combust ion sect ion. Theelements are connected to a control unit whichmonitors the electr ical resis tance of thesensing elements. The sensing elements aremade of Inconel metal tubing filled with a pli-able, heat-sensitive ceramic material which,in turn, encloses a conductor wire at its center

that carries the DC power through the detectioncircuit. The electrical resistance of the ceramicmaterial is relatively high at normal temper-atures; consequently, there is little currentflow from the conductor wire through theceramic material to ground (outer tubing). Athigh temperatures, however, the electricalresistance decreases and allows increasedcurrent flow.

The control unit detects the increased cur-rent flow and illuminates the red FIRE orENG FIRE light in the T-handle when currentflow equates to 890˚ F at the hot section sen-sor, or 410˚ F at the engine accessory and/orfirewall sensors (Figure 8-1). DC essentialbus electrical power for the system is sup-plied through the L and R FIRE DET circuitbreakers on the pilot’s and copilot’s circuit-breaker panels.




Figure 8-1. Engine Fire Detection System

ENG FIREPULL

ELEMENTSUPPORTFRAME

410 F SENSING ELEMENTAND SUPPORT FRAME

890 F SENSING ELEMENT

410 F SENSING ELEMENT(PYLON FIREWALL)

L FIREDET

L ESSBUS

COMBUSTIONSECTION

CONTROLUNIT

FIRE AND ENG FIRE LIGHTSThe red FIRE PULL or ENG FIRE PULL warn-ing lights are part of the T-handles, one locatedat each of the glareshield annunciator panel(Figure 8-2). In the event of an engine fire, thewarning light in the T-handle will flash untilthe fire or overheat condition ceases to exist.Operation of the T-handles is explained underEngine Fire-Extinguishing.

FIRE DETECTION SYSTEM TESTThe rotary system test switch (Figure 8-3) onthe center switch panel is used to test the firedetection system. Rotating the switch to FIREDET and depressing the switch test buttontests the continuity of the sensing elements andcontrol units. A satisfactory test is indicatedby both FIRE or ENG FIRE lights flashinguntil the test button is released.

ENGINE FIRE-EXTINGUISHING

EXTINGUISHER CONTAINERSTwo spherical extinguishing agent containers arelocated in the tailcone area. Both containersuse common plumbing to both engine cowlingsvia shuttle valves, providing the airplane witha two-shot system. The agent used in the fire-extinguishing system is variously known asmonobromotrifluoromethane, bromotrifluo-romethane, or by the more common trade nameof Halon 1301. It is noncorrosive, so no cleanupis necessary after use. The agent is stored underpressure, and a pressure gage is installed oneach container. The pressure gages indicateapproximately 600 psi at 70˚ F when thecontainers are properly serviced.




Figure 8-2. Engine Fire Warning Lightsand Controls (LH)

Figure 8-3. System Test Switch

A thermal relief valve one each container isplumbed to a common discharge port (red disc)on the outside of the fuselage below the left en-gine pylon. The thermal relief valves will re-lease bottle pressure at approximately 220˚ F.

FIRE OR ENG FIRE T-HANDLESAND ARMED LIGHTSWhen a FIRE PULL or ENG FIRE PULL lightbegins to flash, it indicates a fire or overheatcondition in the respective engine cowling.Following AFM procedures, the pilot shouldfirst place the affected engine thrust lever toCUT-OFF and then pull the corresponding T-handle. Pulling out on the T-handle closes themain fuel, hydraulic, and bleed-air shutoffvalves for that engine. DC essential buselectrical power to close these valves is providedthrough the L and R FW SOV (firewall shutoffvalve) circuit breakers on the pilot’s and copi-lot’s circuit-breaker panels, respectively.

There are two ARMED lights above each T-handle. Pulling either T-handle arms the fireextinguisher system, which is indicated byillumination of the two ARMED lights abovethe handle which was pulled. Depressing ani l lumina ted ARMED l igh t momentar i lysupplies DC power to the explosive cartridgewhich discharges the contents of one fire-extinguisher bottle and allows it to flow intothe affected engine nacelle. When the ARMEDlight is depressed, a holding relay is alsoengaged which extinguishes the ARMED light,indicating that the associated bottle has beendischarged. Either ARMED light may bedepressed to extinguish the fire. Should onecontainer control the fire, the other container isstill available to either engine. (See Figure 8-4.)

NOTEIf the red warning light goes out, thecontinuity of the detection circuitshould be tested using the rotarysystem test switch.

EXTERIOR EXTINGUISHERDISCHARGE INDICATORSTwo colored disc indicators are flush-mountedin the side of the fuselage below the left enginepylon (Figure 8-5). The red disc covers thethermal discharge port. It will be ruptured ifone or both thermal relief valves have releasedbott le pressure. The yel low disc wil l beruptured if either bottle is discharged bydepressing an illuminated ARMED light. Theintegrity of the two discs is checked during theexternal preflight inspection.

PORTABLE FIRE-EXTINGUISHERSOne (standard) or two (optional) hand-heldfire-extinguishers provide for interior fire pro-tection. Location of the extinguisher(s) varieswith airplane configuration.




Figure 8-5. Fire-Extinguisher DischargeIndicators




ARMED ARMED

LHCONTAINER

RELIEF VALVE

FIRE PULL

ARMED ARMED

FIRE PULL

RH NACELLE

LEGEND

MANUAL DISCHARGE

THERMAL DISCHARGE

LH NACELLE

PRESSURE GAGE

RELIEF VALVE

BLEED-AIRSHUTOFF

VALVE

HYDRAULICSHUTOFF VALVE

FUEL SHUTOFFVALVE

BLEED-AIRSHUTOFF

VALVE

HYDRAULICSHUTOFF VALVE

FUEL SHUTOFFVALVE

MANUALDISCHARGEINDICATOR

THERMALDISCHARGEINDICATOR

RHCONTAINER

TWO-WAYCHECKVALVES

ENGINE EXTINGUISHING

PRESSURE GAGE

Figure 8-4. Engine Fire-Extinguishing System

Figure 8-6. Portable Fire-Extinguisher

QUESTIONS




l. Engine fi re-ext inguisher bot t les arelocated in:A. The nacellesB. The engine pylonsC. The tailconeD. The baggage compartment

2. The power-off preflight check of the enginefire-extinguishers includes:A. Checking the condition of one yellow

and one red blowout discB. Checking the condition of two yellow

and two red blowout discsC. Checking blowout discs and extin-

guisher charge gages, all on the leftside of the fuselage

D. Activating the system TEST switch toFIRE DET

3. When the left FIRE or ENG FIRE T-handleis pulled:A. It discharges one extinguisher into the

left nacelle.B. It closes the main fuel, hydraulic, and

bleed-air shutoff valves for the leftengine and arms both extinguishers.

C. It discharges one extinguisher andarms the second.

D. It ruptures the yellow discharge indi-cator disc.

4. When an engine fire occurs, the controlunit:A. Arms the fire-extinguishing systemB. Illuminates the MSTR WARN light

and sounds the warning hornC. Au toma t i ca l l y d i s cha rges t he

respective fire-extinguisher systemD. Causes the respective FIRE or ENG

FIRE light in the T-handle and bothMSTR WARN lights to flash

5. The fire-extinguisher agent is dischargedby:A. A temperature switchB. A mechanically fired pin at the base

of the supply cylinderC. The FIRE T-handle electrical circuitsD. Pushing an illuminated ARMED light

6. If fire persists after activating a fire bottle:A. The second fire bottle can be dis-

charged into the affected area.B. The second fire bottle can only be

used on an opposite-side fire.C. The first fire bottle can be discharged

a second time.D. No further activation of the system is

poss ib le ; bo th bot t les d i schargesimultaneously when either ARMEDbutton is pressed.

9-i

CHAPTER 9PNEUMATICS

CONTENTS

Page

INTRODUCTION ................................................................................................................... 9-1

GENERAL............................................................................................................................... 9-1

DESCRIPTION AND OPERATION ...................................................................................... 9-2

Bleed-Air Shutoff and Regulator Valves.......................................................................... 9-2

BLEED AIR Switches ..................................................................................................... 9-2

Bleed-Air Check Valves................................................................................................... 9-5

Bleed-Air Manifold.......................................................................................................... 9-5

BLEED AIR Warning Lights ........................................................................................... 9-5

HP Servo Air .................................................................................................................... 9-5

QUESTIONS ........................................................................................................................... 9-6




9-iii

ILLUSTRATIONS

Figure Title Page

9-1 BLEED AIR Switches.............................................................................................. 9-2

9-2 Pneumatic System—SNs 35-002 through 35-112 and 36-002 through 36-031....... 9-3

9-3 Pneumatic System—SNs 35-113 and Subsequent and 36-032 and Subsequent...... 9-4




INTRODUCTIONThe airplane pneumatic system uses bleed air extracted from the engine compressor sec-tions. It includes controls for regulation and distribution of low-pressure (LP) air fromthe fourth-stage axial compressor and high-pressure (HP) air from the centrifugal com-pressor. Pneumatic air is used for cabin pressurization and heating, anti-icing systems,hydraulic reservoir pressurization, and Aeronca thrust operation (if installed).

There are two basic pneumatic system configurations—airplane SNs 35-002 through 112,and 36-002 through 031, referred to in the text as “early” airplanes; and airplanes SNs 35-113 and subsequent, and 36-032 and subsequent which incorporate a major design change,including installation of the emergency valves, referred to as “current” airplanes.

GENERALBleed air from both the LP and HP engine com-pressors is provided to a shutoff and regulatorvalve on each engine. When open, these valvesregulate air pressure by selecting either LP orHP air, which is ducted to a common manifoldwhich supplies most of the pneumatic systems.Some systems use only HP air which is tappedfrom the high-pressure compressor prior to theshutoff and regulator valve. Regulated bleed-air

pressure is used for cabin pressurization and heat-ing, windshield anti-icing, engine nacelle anti-icing, wing and stabilizer anti-icing, and topressurize the hydraulic reservoir. HP air is used forfan spinner anti-icing and Aeronca thrustreversers, if they are installed. On current airplanes,HP air is used for the alcohol anti-icing system,operation of the emergency pressurization valves, asservo pressure for the cabin pressurization andtemperature control systems, and for control ofmodulating valves on airplanes with AAK 85-6.

VALVE

L R

COBLEED AIR

515

20

AIR

CHAPTER 9PNEUMATICS




Control of pneumatic bleed air is accomplishedwith the L and R BLEED AIR switches on thecopilot’s lower right switch panel and by theengine FIRE T-handles. Provision is made fordetection of overheat conditions within theengine pylon, the pylon bleed-air duct itself,and, on some airplanes, manifold overpres-sure. Visual indication is given by illumina-tion of warning lights on the glareshieldannunciator panel.

DESCRIPTIONAND OPERATION

BLEED-AIR SHUTOFFAND REGULATOR VALVESThe bleed-air shutoff and regulator valves,one on each engine, are often called “modvalves” because they modulate air pressure(Figures 9-2 and 9-3). The valves are electri-cally controlled by the BLEED AIR switchesand may also be closed by pulling the respec-tive engine FIRE T-handle. When open, thevalves operate pneumatically to maintain down-stream pressure in the manifold of 27-35 psi.

Both HP and LP bleed air are available to thevalves. As long as enough LP air is available tomeet system demands, the valves will use only LPair. If there is not enough LP air available to meetsystem demands, the valves will automatically useHP air to maintain the required pressure.

The shutoff function of each shutoff andregulator valve is provided by a solenoid-operated shutoff valve which is spring-loadedopen; DC power is required to close it. Withloss of electrical power, the shutoff and reg-ulator valves fall open. However, on airplaneSNs 35-113 and subsequent and 36-032 andsubsequent, an HP solenoid valve, which isspringloaded closed, is installed (Figure 9-3). On these airplanes, if electrical power islost, the shutoff and regulator valve fails open,but the HP solenoid valve fails closed so thatonly LP air will be available.

BLEED AIR SWITCHESOn airplanes SNs 35-002 through 112 and 36-002 through 031, the L and R BLEED AIRswitches are located on the copilot’s lower rightswitch panel (Figure 9-1). They are two-posi-tion, ON-OFF, switches, powered by the AIRBL circuit breaker on the left essential bus. Inthe ON position, the bleed-air shutoff valve(Figure 9-2) is open. In the OFF position, thevalve is closed.




EFFECTIVITY: SNs 35-113 AND SUBSEQUENTAND 36-032 AND SUBSEQUENT

EFFECTIVITY: SNs 35-002 THROUGH112 AND 36-002 THROUGH 031

Figure 9-1. BLEED AIR Switches

On airplane SNs 35-113 and subsequent and36-032 and subsequent, the L and R BLEEDAIR switches are located on the copilot’s lowerright switch panel (Figure 9-1). They are three-position, OFF–ON–EMER, switches whichcontrol their respective bleed-air shutoff andregu l a to r va lve s and t he i r r e spec t iveemergency pressurization valves.

In the OFF position, the bleed-air shutoff andregulator valve is closed, and the emergency

pressurization valve is in its normal position(Figure 9-3). In the ON position, the bleed-airshutoff and regulator valve is open, and theemergency pressurization valve remains in itsnormal position. In the EMER position, thebleed-air shutoff and regulator valve is open,and the emergency pressurization valve ismoved to the emergency position. At the sametime, the HP solenoid valve is closed, re-stricting the shutoff and regulator valve outputto LP air.




Figure 9-2. Pneumatic System—SNs 35-002 through 35-112 and 36-002 through 36-031

LEGENDHP BLEED AIR

LP BLEED AIR

REGULATED BLEED AIR

NOT APPLICABLE ON AIRPLANES EQUIPPEDWITH CONICAL SPINNERS (AAK 79-4)

AERONCA THRUST REVERSERS

SNs 35-082, 087 THROUGH 112, AND 36-023 THROUGH031 AND EARLIER SNs INCORPORATING AMK 76-7

BLEED-AIR SHUTOFFAND REGULATOR VALVE

PYLON TEMPSENSOR

HP BLEED AIR

LPBLEED

AIR

FAN SPINNER ANTI-ICE*

T/R AIR MOTOR**

NACELLE ANTI-ICE

DUCT TEMPSENSOR

STABILIZER ANDWIND ANTI-ICE

BLEEDAIR R

BLEED-AIR MANIFOLD

FLOW CONTROLVALVE

WINDSHIELDANTI-ICE

PRESSURIZATIONJET PUMP

HYDRAULICRESERVOIR

TO CABIN

47-PSI PRESSURE SWITCH***

FROMLEFT

ENGINE

*

**

***

On airplanes SNs 35-113 through 658 and 36-032 through 063, not modified by AMK 90-3,the L and R BLEED AIR switches use DCelectrical power from the L and R MOD VALcircuit breakers on the left and right main DCbuses. These circuit breakers provide power forcontrol of the bleed-air shutoff and regulatorvalves and the emergency pressurization valves.

On airplanes SNs 35-659 and subsequent and36-064 and subsequent, and earlier airplanesmodified by AMK 90-3, the L and R BLEED

AIR switches use DC electrical power from theL and R BLEED AIR and L and R EMERPRESS circuit breakers on the left and rightmain DC buses .The BLEED AIR circui tbreakers provide power for control of the bleed-air shutoff and regulator valves. The EMERPRESS circuit breakers provide power forcontrol of the emergency pressurization valves.

See Chapter 12, “Pressurization,” for addi-tional information on the emergency pressur-ization valves.




HP BLEED AIR

LP BLEED AIR

REGULATED BLEED AIR

LEGEND

REGULATED SERVO AIR

NOT APPLICABLE ON AIRPLANESEQUIPPED WITH CONICAL SPINNERS(AAK 79-4)

AERONCA THRUST REVERSERS

AND BLEED-AIR SHUTOFF ANDREGULATOR VALVE (AAK 85-6)

FLOW CONTROL VALVE

FAN SPINNER ANTI-ICE *

T/R AIR MOTOR**

HP BLEED AIR

HPSOLENOID

VALVELP

BLEEDAIR

BLEED-AIR SHUTOFFAND REGULATOR VALVE

NACELLE ANTI-ICE

BLEEDAIR R

PYLON TEMP SENSOR

REGULATEDBLEED AIR

DUCT TEMPSENSOR

CABINDUCT

EMER PRESSVALVE

BLEED-AIRMANIFOLD

HP SERVOAIR MANIFOLD

STABILIZER ANDWING ANTI-ICE

WINDSHIELDANTI-ICE

FLOW CONTROLVALVE

TOCABIN

HYDRAULICRESERVOIR

CABINDUCT

EMER PRESSVALVE

FROMLEFT

ENGINE

EMER PRESSVALVES

PRESSURIZATIONJET PUMP

REGULATOR

REGULATORBLEED-AIRTEMP CONTROLVALVE (H-VALVE)***

ALCOHOLANTI-ICE

HP SERVOAIR

*

**

***

Figure 9-3. Pneumatic System—SNs 35-113 and Subsequent and 36-032 and Subsequent

BLEED-AIR CHECK VALVESA check valve is installed in the bleed-airducting from each engine. Each check valveallows airflow in one direction and blocksairflow applied in the opposite direction. Thecheck valves prevent loss of bleed air duringsingle-engine operation.

BLEED-AIR MANIFOLDThe bleed-air manifold serves as a collectionpoint for regulated air pressure from either orboth engines. From the manifold, bleed air isdistributed to the flow control valve for cabinpressurization and heating, the pressurizationjet pump (on early airplanes), the windshieldanti-ice (defog) valve, the wing and horizontalstabilizer anti-ice pressure regulator valve,and the hydraulic reservoir regulator.

BLEED-AIR WARNING LIGHTSThe red BLEED AIR L and R warning lightson the glareshield annunciator panel illuminatewhen an associated pylon senor or pylon ducttemperature sensor detects excessive tem-peratures. On some airplanes, a pressure sensorin the manifold causes both lights to illuminatefor an overpressure condition.

A temperature sensor in each engine pylonoperates when pylon structure temperatureexceeds 250˚ F and illuminates the respectivered L or R BLEED AIR light on the glareshield(Annunciator Panel).

A temperature sensor installed in each enginepylon bleed-air duct causes the respective redL or R BLEED AIR light to illuminate if ducttemperature is excessive. Airplane SNs 35-002 through 35-064, and 36-002 through 36-017 use 590˚ F sensors. Later productionairplanes use 645˚ F sensors.

On airplanes SNs 35-082, 087 through 112, and36-023 through 031, a pressure sensor in theregulated bleed-air manifold causes bothBLEED AIR warning lights to illuminate if pres-sure in the manifold exceeds 47 psi. This also ap-plies to earlier airplanes incorporating AMK76-7 (relocation of cabin air distribution flowcontrol valve). (Figure 11-2 in Chapter 11.)

HP SERVO AIROn airplane SNs 35-113 and subsequent, and36-032 and subsequent (Figure 9-3), HP bleedair is tapped off the HP centrifugal compressor.The air from this tap flows through a checkvalve to the HP servo air manifold. From themanifold, air is ducted directly to the alcoholanti-icing system and through two regulators.The air from one regulator is used to controlthe position of the hot-air bypass valve (H-valve) and the bleed-air shutoff and regulatorvalve on airplanes modified per AAK-85-6.The other regulator provides air to (1) mod-ulate the flow control valve, (2) control posi-tion of the emergency valves, and (3) operatethe pressurization jet pump.







l. Pneumatic air is extracted from:A. The LP compressorB. The HP compressorC. Ram airD. Both A and B

2. With loss of DC electrical power, theshutoff and regulator valves:A. Fail closedB. Fair openC. Remain in their last positionD. Can be c losed only by pul l ing a

FIRE/ENG FIRE T-handle

3. The L and R BLEED AIR ON–OFFswitches are located:A. On the copilot’s lower right switch

panelB. On the left side panelC. On the pilot’s lower left switch panelD. On the overhead panel

4. The temperature of the bleed air in theduct between the engine and the bleed- airmanifold is monitored by the:A. Pylon overheat thermostatB. Aft fuselage equipment section ther-

mostatC. Duct temperature sensorD. Duct overheat thermostat

5. The BLEED AIR L annunciator illumi-nates:A. When the temperature in the left pylon

or the left bleed-air duct is too highB. When the pressure in the left pylon is

below the system’s operational limit C. When the left half of the bleed-air

system is operatingD. When the left half of the bleed-air

system has failed

QUESTIONS

10-i

CHAPTER 10ICE AND RAIN PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................. 10-1

GENERAL ............................................................................................................................ 10-1

ICE DETECTION ................................................................................................................. 10-2

Windshield Ice Detection............................................................................................... 10-2

Wing Ice Detection ........................................................................................................ 10-2

ANTI-ICE SYSTEMS........................................................................................................... 10-2

Engine Anti-ice System (Nacelle Heat)......................................................................... 10-2

Exterior Windshield Defog, Anti-ice, and Rain Removal System................................. 10-5

Internal Windshield Defog .......................................................................................... 10-16

Windshield/Radome Alcohol Anti-ice System............................................................ 10-18

Wing and Horizontal Stabilizer Anti-ice System ........................................................ 10-21

Pitot, Static, and Angle-of-Attack Vane Anti-ice System ........................................... 10-22

QUESTIONS....................................................................................................................... 10-24




10-iii

ILLUSTRATIONS

Figure Title Page

10-1 Anti-ice Control Panel............................................................................................ 10-4

10-2 Nacelle and Fan Spinner Anti-ice Flow ................................................................. 10-3

10-3 Defog Control Knob............................................................................................... 10-5

10-4 Windshield Anti-ice System (Airplane SNs 35-002 through 086,except 082, and 36-002 through 022,without AAK 76-7A or AMK 91-2) ........... 10-6

10-5 Windshield Anti-ice System (Airplane SNs 35-082, 087 through112,36-023 through 031, and Earlier SNs Incorporating AAK 76-7A......................... 10-8

10-6 Windshield Anti-ice System (Airplane SNs 35-113 through 662,and 36-032 through 063, without AMK 91-2).................................................... 10-10

10-7 Windshield Anti-ice System (Airplane SNs 35-663 and Subsequent,36-064 and Subsequent, SNs 35-113 through 662, and 36-032through 063 Incorporating AMK 91-2) ............................................................... 10-12

10-8 Windshield Anti-ice System (Airplane SNs 35-002 through 112, and36-002 through 031 Incorporating AAK 76-7A and AMK 91-2) ....................... 10-14

10-9 Electric Windshield Defog System—Models 35-671 andSubsequent and 36-064 and Subsequent.............................................................. 10-17

10-10 Alcohol Anti-ice System (Airplane SNs 35-002 through 112and 36-002 through 031 ....................................................................................... 10-19

10-11 Alcohol Anti-ice System (Airplane SNs 35-113 andSubsequent and 36-032 and Subsequent) ............................................................ 10-20

10-12 Wing and Horizontal Stabilizer Anti-ice System................................................. 10-21

10-13 WING TEMP and STAB TEMP Indicators......................................................... 10-23




INTRODUCTIONAnti-icing equipment on the Learjet 35/36 is designed to prevent buildup of ice on:

• The engine nacelle lip, early model fan spinner, and the inlet pressure-temperature probe

• The windshield and the radome

• The leading edge of the wings and the horizontal stabilizer

• Pitot probes, static ports, AOA vanes, shoulder static ports (if installed), and thetotal temperature (Rosemount) probe (if installed)

This system is certified for flight into known icing conditions.

GENERALAirplane anti-icing is accomplished throughthe use of electrically heated anti-ice systems,engine bleed-air heated anti-ice systems, andan alcohol anti-ice system.

Electrically heated components include pitottubes, static ports, shoulder static ports (FC-

200), the engine inlet air pressure-tempera-ture (PT2 TT2) sensors, stall warning vanes,and a total temperature (Rosemount) probe,if installed.

Engine bleed air is used to heat the wind-shields, wing and horizontal stabilizer lead-ing edges, nacelle inlets, and the engine fanspinners on earlier airplanes with the ellipti-cal (dome-shaped) spinners.

CHAPTER 10ICE AND RAIN PROTECTION




An alcohol system is installed for radome anti-icing and as a backup to the pilot’s windshieldbleed-air anti-icing.

On airplane SNs 35-643 and subsequent and36-058 and subsequent, an auxiliary wind-shield defog heat system is installed.

All anti-icing equipment must be turned onbefore icing conditions are encountered. Todelay until ice buildup is visually detected onairplane surfaces constitutes an unacceptablehazard to safety of flight.

If anti-ice systems are required during take-off, they should be turned on prior to settingtakeoff power. Appropriate takeoff power andperformance charts must be used.

Icing conditions exist when there is visiblemoisture and the indicated ram-air temperature(RAT) is + 10˚ C or below. Takeoff into icingconditions is permitted with all bleed-air anti-icing systems on. The air temperature gage(RAT) should be checked frequently when fly-ing in or entering areas of visible moisture.

During descents, the cabin altitude may increaseunless sufficient engine rpm is maintained tocompensate for the additional bleed-air use.

Anti-ice system switches are located on theanti-ice control panel (Figure 10-1).

ICE DETECTIONDuring daylight operation, ice accumulationcan be visually detected on the windshield, thewing leading edges, and tip tanks.

WINDSHIELD ICE DETECTIONDuring night operations, the windshield ice de-t e c t i on l i gh t s i nd i ca t e i c e o r mo i s tu r eformation on the windshield. Two probes, oneon the pilot’s side of the glareshield and oneon the copilot’s side, contain red lights whichcont inuous ly sh ine on the ins ide of thewindshield surface. The ice detection lights

normally shine through unseen. However, theywill reflect red spots approximately 1 1/2inches in diameter if ice or moisture has formedon the windshield.

The ice detection light on the pilot’s side isinside the anti-ice airstream; the light on thecopilot’s side is located outside the anti-iceairstream. For this reason, the copilot’s lightshould be monitored when flying in icingconditions (anti-icing equipment on). The icedetection lights are illuminated wheneverairplane electrical power is on. The lights arepowered through the L and R ICE DET circuitbreakers on the pilot’s and copilot’s essentialbuses respectively.

WING ICE DETECTIONDuring daylight conditions, ice formation onthe wing leading edges and tip tanks may beobserved visually.

During darkness, the recognition light can beused to check for ice buildup.

On airplanes with the emergency light system,the wing inspection/egress light below theemergency exit, ice buildup may be detectedon t he i nboa rd l e ad ing edge . On someairplanes, an optional wing ice inspection lightis installed on the forward right side of thefuselage and is focused on a three-inch blackdot on the wing leading edge next to the tiptank. The light is operated by a switch locatedon the copilot’s sidewall panel (Figures 3-13and 3-14 in Chapter 3.)

ANTI-ICE SYSTEMS

ENGINE ANTI-ICE SYSTEM(NACELLE HEAT)The engine anti-ice system provides anti-icingfor the engine nacelle inlet lips, the ellipticalfan spinners, and the PT2TT2 probes. Thenacelle lips are heated with regulated bleed air.The PT2TT2 probe is heated electrically. Onairplane SNs 35-002 through 244 and 36-002




through 044, not incorporating AAK 79-4,the elliptical spinner is anti-iced by high pres-sure bleed air. On aircraft SNs 35-245 andsubsequent, and 36-045 and subsequent, andearlier airplanes incorporating AAK 79-4, aconical spinner replaces the elliptical spin-ner and no anti-icing is required.

Nacelle Heat SwitchesEach engine anti-ice system is independentlycontrolled by the L and R NAC HEAT switcheslocated on the anti-ice control panel (Figure10-1).

When the NAC HEAT switch is turned on (Lor R position), electrical power is supplied toheat the PT2TT2 probe . The swi tch a lsoenergizes the fan spinner shutoff valve open(if applicable) and deenergizes the nacelle lipshutoff valve open. Selecting the OFF positiondeenergizes the fan spinner shutoff valveclosed and energizes the nacelle shutoff valveclosed. Figure 10-2 is a schematic portrayalof the engine anti-ice systems.

DC electrical power to operate the systems isprovided through the L and R NAC HT circuitbreakers on the L and R main buses.

Bleed air for nacelle lip anti-icing is takenf rom the regu la t ed b l eed -a i r l i ne ju s tdownstream from the bleed-air shutoff andregulator valve (Figure 10-2). It is ductedthrough the nacelle heat shutoff valve to a dif-fuser tube which distributes it around the innersurface of the nacelle lip and then exhausts itoverboard through a hole at the bottom of thenacelle lip.

The source of fan-spinner heat is high-pressure(HP) bleed air.

Engine Ice LightsThe amber L and R ENG ICE lights on theglareshield annunciator panel (AnnunciatorPanel section) provide a visual indication offan spinner or nacelle lip anti-ice systemmalfunction. The lights are operated by pres-sure switches in the associated fan spinner




Figure 10-2. Nacelle and Fan Spinner Anti-ice Flow

HIGH-PRESSURE BLEED AIR

REGULATED (MANIFOLD) AIR

LOW-PRESSURE BLEED AIR

LEGEND

ELECTRICAL CIRCUITS

* NOT APPLICABLE TO AIRPLANES EQUIPPED WITH CONICAL FAN SPINNERS

** SNs 35-634 AND 36-058 AND SUBSEQUENT

PT2TT2PROBE

FAN SPINNERPRESS SWITCH*

L ENG ICE

DCTO OPEN

NAC HEATON

BLEED-AIRINPUT

NAC HEAT L R

OFF OFF

DC TO CLOSE

NACELLE HEATSHUTOFF VALVE

NACELLEPRESS SWITCH

**

FAN SPINNER SHUTOFF

VALVE

SHUTOFF

AND REG

VALVE

LP

HP

NACHT

MAINBUS

*




Figure 10-1. Anti-ice Control Panel

35-002 THROUGH 35-11236-002 THROUGH 36-030

WSHLD HEATON AUTO

WSHLD &RADOME

STABWINGHEAT

PITOT HEATL R

OFF OFF

NAC HEATL R

OFF OFF

OFF MAN RADOME OFF

OFF

HOLD

RAD

35-113 THROUGH 35-64236-031 THROUGH 36-063

35-643 THROUGH 35-670 35-671 AND SUBSEQUENT36-064 AND SUBSEQUENT

WSHLD WSHLD/ HT ON RADOME

STAB WINGHEAT

PITOT HEATL R

OFF OFF

NAC HEATL R

OFF OFF

OFF OFF OFF

HOLD

Off

RAD

Off

CKPT

W/S AUXDEFOGHEAT

HOLD

RAD

WSHLDDEFOG

STAB WINGHEAT

WSHLD WSHLD HT ON RADOME

OFF OFF OFF OFF

PITOT HEAT L R

OFF OFF

NAC HEAT L R

OFF OFF

WSHLD WSHLD HT ON RADOME

STAB WINGHEAT

OFF OFF

PITOT HEAT L R

OFF OFF

NAC HEAT L R

OFF OFF

and nacelle lip bleed-air plumbing. Illumina-tion of an ENG ICE light with the associatedNAC HEAT switch on indicates that bleed-air pressure to either the fan spinner or to thenacelle lip is not sufficient to provide satis-factory anti-ice protection.

When a NAC HEAT switch is turned on or off,the respective ENG ICE light illuminates mo-mentarily until bleed-air pressure at the pres-sure switch agrees with the switch command.Under some conditions, bleed-air pressure maynot be sufficient at idle rpm to keep the pres-sure switches from illuminating the ENG ICElight. In this event, advance the thrust leversto check proper nacelle heating operation.

Illumination of either ENG ICE light NACHEAT switches in the OFF position indicatesthe presence of bleed-air pressure in the nacellelip or fan spinner plumbing due to a malfunc-tion of the nacelle lip or fan spinner anti-iceshutoff valve. Cycling the NAC HEAT switchon and back to OFF may close the open valve.

GREEN NAC HT ON Light(Airplane SNs 35-634 andSubsequent, and SNs 36-058and Subsequent)A single green NAC HT ON annunciator lightis installed on the glareshield annunciatorpanel. The light illuminates when either NACHEAT switch is on as a reminder that the na-celle heat system is operating.

EXTERIOR WINDSHIELDDEFOG, ANTI-ICE, AND RAINREMOVAL SYSTEM

GeneralThere are five different systems used in theLearjet 35/36 to provide exterior windshieldanti-icing, defogging, and rain removal. Theywill be covered individually. All systems op-erate on DC power from the WSHLD HT cir-cuit breaker on the left main bus.

Airplane SNs 35-002 through086, except 082, and 36-002through 022, without AAK 76-7A or AMK 91-2The exterior windshield heat/defog systemcan be controlled either automatically ormanually (Figure 10-4). It is also used to sup-plement cockpit heating through the pilots’footwarmers, and to provide an alternate bleed-air source for emergency pressurization.

An IN–NORMAL/OUT–DEFOG knob, lo-cated below the instrument panel to the left ofthe pedestal (Figure 10-3), manually controlsa valve which directs bleed-air either to thewindshield or to the cockpit footwarmers.

When the knob is pushed into the IN–NORMALposition, with the windshield anti-ice on, bleedair is directed into the cockpit through the pilot’sand copilot’s footwarmers. This provides




Figure 10-3. Defog Control Knob




Figure 10-4. Windshield Anti-ice System (Airplane SNs 35-002 through 086, except 082,and 36-002 through 022, without AAK 76-7A or AMK 91-2)

MANIFOLD BLEED AIR

REGULATED FLOW

LOW-LIMIT THERMOSWITCH

LEGEND

HIGH-LIMIT THERMOSWITCH

ELECTRICAL CIRCUITS

WSHLDHT

CONTROLUNIT

DEFOG PRESSUREREGULATOR VALVE

(NC)

TO WING/STABHEAT

CHECKVALVE

CHECKVALVE

TOCABIN

BLEED-AIRMANIFOLD

DEFOG SHUTOFFVALVE

BLEED AIR BLEED AIR

FOOTWARMERS

OVERBOARDDRAIN

WINDSHIELDIN

NORMALOUT

DEFOG

WSHLDOV HT

WSHLD HEATON OFF

OFF MANL MAIN

BUS

WSHLDHT

additional heat in the cockpit and an alternatesource of bleed-air flow into the cabin foremergency cabin pressurization. The knob isnormally left in the IN–NORMAL position.

When the knob is pulled out to the OUT–DE-FOG position, the bleed air is directed to theexternal windshield duct outlets for wind-shield defog, anti-ice, and rain removal.

Two windshield heat switches are located onthe anti-ice panel. One is a three-positionswitch labeled “ON” and “OFF,” and is spring-loaded to the center (neutral) position. Theother switch has two positions labeled “AUTO”and “MAN.”

Bleed air from the regulated bleed-air mani-fold is routed through two valves: the shutoffvalve and the pressure-regulator valve.

The shutoff valve is motor-driven and con-trolled by either of the two switches on the anti-ice control panel. It takes four to five secondsto cycle fully. Selecting AUTO will open theshutoff valve and illuminate the green WSHLDHT light. The light will be on whenever theshutoff valve is not fully closed. If MAN is se-lected, the shutoff valve may be opened orclosed with the ON–OFF switch. Since thisswitch is spring-loaded to neutral, it must beheld in the ON position while the valve drivestoward the fully open position. The switchmay be released before the valve reaches fullopen. The shutoff valve will then stop and re-main in an intermediate position. The shutoffvalve can be closed only by holding theON–OFF switch to OFF (with MAN selected)for at least four seconds.

The pressure-regulator valve is solenoid-operated and is deenergized closed. Its functionis to regulate the engine bleed air from the man-ifold to 16 psi. It is energized open when DCelectrical power is applied to the airplane andwill be deenergized and closed to shut off wind-shield anti-ice in case of windshield overheat.

Automatic Operation

The flow of bleed air to the windshields is con-trolled in the auto mode by the high (250˚ F)

and low (215˚F) temperature thermoswitchesinstalled in each windshield outlet nozzle.

NOTEAAK 77-6 provides for changing thehigh- and low-limit thermoswitchesto 290˚ F and 250˚ F, respectively.

For ground operation, when the low-limit ther-moswitch senses 215˚ F, it will close theshutoff valve, which extinguishes the greenWSHLD HT light. It will also illuminate thered WSHLD OV HT light. If the low-limitswitch fails, or the shutoff valve fails to close,the temperature may rise sufficiently to trig-ger the high-limit thermoswitch which re-moves power from the pressure-regulatorvalve. The red WSHLD OV HT light will illu-minate, and the green WSHLD HT light will re-main illuminated because the shutoff valve is notfully closed.

During flight, through the squat switch relaybox, the low-limit switch will close the shutoffvalve which extinguishes the WSHLD HT light.However, the red WSHLD OV HT light will notilluminate because the system is designed tocycle on the low-limit switches. If the high-temperature limit is reached in flight due tofailure of the low-limit switches, the pressure-regulator valve will close, the red WSHLD OVHT light will illuminate, and the green WSHLDHT light will remain illuminated.

Manual Operation

Selecting MAN enables the spring-loadedON–OFF switch to control the shutoff valveand, therefore, the amount of bleed air suppliedto the windshields.

On the ground, in manual mode, a low-limitthermoswitch will illuminate the red WSHLDOV HT light, but will not close either theregulator valve or the shutoff valve. However,the high-limit thermoswitch does close the pres-sure-regulator valve. Therefore, an overheatcondition is indicated by illumination of boththe green and red lights, regardless of which limitis exceeded. In flight, the low-limit ther-moswitch is disabled.




Airplane SNs 35-082, 087through 112, and 36-023through 031, and EarlierAirplanes Incorporating AAK76-7AThe exterior windshield heat/defog systemcan be controlled either automatically or man-ually (Figure 10-5). It is also used to supple-ment cockpit heating through the pilot’sfootwarmers and to provide an alternate bleed-air source for emergency pressurization.

An IN–NORMAL/OUT–DEFOG knob, lo-cated below the instrument panel to the left ofthe pedestal (Figure 10-3), manually controlsa valve which directs bleed air either to thewindshield or to the cockpit footwarmers.

When the knob is pushed in to the IN–NOR-MAL position, with the windshield anti-ice on,bleed air is directed into the cockpit throughthe pilot’s and copilot’s footwarmers. Thisprovides additional heat in the cockpit and analternate source of bleed-air flow into the




Figure 10-5. Windshield Anti-ice System (Airplane SNs 35-082, 087 through 112, 36-023through 031, and Earlier SNs Incorporating AAK 76-7A)

WSHLDOV HT

WSHLDHT

HEATEXCHANGER

DEFOG HEATEXCHANGER

WSHLDANTI-ICE

CONTROLVALVE

WSHLDANTI-ICESHUTOFF

(NC)

RAM AIR OUT RAM AIRIN

RAM-AIRMODULATINGVALVE

TEMPERATURESENSOR

SERVOPRESSURELINE

CHECKVALVE

FOOTWARMERSTO WING/STAB

HEAT

TOCABIN

OVERBOARDDRAIN

INNORMAL

OUTDEFOG

WINDSHELD

CONTROLUNIT

WSHLDHT WSHLD HEAT

ON AUTO

OFF MAN L MAIN

BUSLEGENDRAM AIR

MANIFOLD BLEED AIR

CONDITIONED AIR



ELETRICAL CIRCUIT

cabin for emergency cabin pressurization. Theknob is normally left in the IN–NORMAL position.


Two windshield heat switches are located onthe anti-icing panel. One is a three-positionswitch labeled “ON” and “OFF,” and is spring-loaded to the center (neutral) position. Theother switch has two positions labeled “AUTO”and “MAN.”

Bleed air from the regulated bleed-air mani-fold is routed through two valves: the anti-iceshutoff valve and the anti-ice control valve.

The shutoff valve is solenoid-operated and isdeenergized closed. Its function is to regulatethe engine bleed air from the manifold to 16psi. It is energized open when DC electricalpower is applied to the airplane and will bedeenergized and closed to shut off windshieldanti-ice in case of windshield overheat.

The control valve is motor-driven and con-trolled by either of the two switches on the anti-ice control panel. It takes four to five secondsto cycle fully. Selecting AUTO will open thecontrol valve and illuminate the green WSHLDHT light. If MAN is selected, the control valvemay be opened or closed with the ON–OFFswitch. Since this switch is spring-loaded toneutral, it must be held in the ON positionwhile the valve drives toward the fully openposition. The switch may be released beforethe valve reaches full open. The control valvewill then stop and remain in an intermediateposition. The control valve can be closed onlyby holding the ON–OFF switch to OFF (withMAN selected) for at least four seconds.

Operation

With windshield anti-ice on, bleed air flowsthrough the open shutoff valve and anti-icecontrol valve, and through a heat exchangerfrom which it is ducted to the outlet nozzlesat the base of each windshield. The anti-ice

heat exchanger cools the bleed air with ram air.A ram-air modulating valve operates to main-tain a 300˚ F duct temperature downstream ofthe heat exchanger by using a duct tempera-ture sensor and a regulated bleed-air servoline. The subsequent heat loss occurring inthe duct as the bleed air reaches the outletnozzles keeps the outlet airflow temperaturewithin the limits of windshield beat opera-tion. During ground operation, ram air is notavailable to cool the bleed air.

Under normal conditions, the windshield heatbleed-air temperature is automatically con-trolled. However, an overheat warning sys-tem alerts the pilot and automatically shuts offwindshield heat in the event of an overheat con-dition. A low-limit (approximately 250˚ F)and a high-limit (approximately 290˚ F) ther-moswitch is installed in each windshield out-let nozzle. The low-limit switches functiononly on the ground and are cut out by the squatswitch relay box when airborne. The high-limit switches are installed primarily to limittemperature during airborne operation, butwill also function on the ground as a backupto the low-limit switches.

If either outlet nozzle temperature reaches the250˚ F limit (ground) or 290˚ F limit (air-borne), the thermoswitch will illuminate thered WSHLD OV HT light on the glareshieldannunciator panel and cause the solenoid shut-off valve to close. The anti-ice control valvewill remain in the position it was in, but thegreen WSHLD HT light will be extinguishedwhile the solenoid shutoff valve is closed. TheWSHLD OV HT light will extinguish and theshutoff valve will open again when the tem-perature at the thermoswitch cools. If thewindshield heat has not been turned off, air-flow will resume to the windshield, the greenWSHLD HT light will illuminate, and the redWSHLD OV HT light will extinguish.

Through the squat switch relay box, the low-limit thermoswitches are disabled for 10 sec-onds after touchdown. This prevents auto-matic shutoff of bleed air at the moment oftouchdown, which could restrict the pilot’svision due to loss of rain-removal capability.




Airplane SNs 35-113 through662 and 36-032 through 063,without AMK 91-2)The WSHLD HT switch controls flow ofeng ine b l eed a i r t o t he ex t e r i o r o f t hewindshield for anti-icing, defogging, and rainremoval (Figure 10-6). This three-position

switch is labeled “ON,” “HOLD,” and “OFF,”and is located on the anti-ice control panel.

Engine bleed air from the regulated bleed-air manifold is routed through two valves:the anti-ice shutoff valve and the anti-icecontrol valve. The shutoff valve is solenoid-operated and is deenergized closed. It is




Figure 10-6. Windshield Anti-ice System (Airplane SNs 35-113 through 662, and 36-032through 063, without AMK 91-2)

WSHLDHT

HEATEXCHANGER

WSHLDANTI-ICE

CONTROLVALVE

WSHLDANTI-ICE

SHUTOFF(NC)

TO WING/STABHEAT

WSHLD HT

L MAINBUS

WSHLDHT ON

OFF

RAM AIRIN


TEMPERATURESENSOR

DEFOG HEATEXCHANGER

RAM AIR OUT

TOCABIN

CONTROLUNIT

WINDSHIELD

RAM AIR

MANIFOLD BLEED AIR

CONDITIONED AIR



ELECTRICAL CIRCUIT

LEGEND

SERVOPRESSURELINE

WSHLDHT

HOLD

energized open whenever DC electrical poweris applied to the airplane. The control valve ismotor-driven and is controlled by the WSHLDHT switch.

When the WSHLD HT switch is positioned toON, the anti-ice control valve begins to openand the g reen WSHLD HT l igh t on theglareshield annunciator panel illuminates. Thecontrol valve drives to the fully open positionwithin five to eight seconds after the WSHLDHT switch is turned to ON.

For reduced airflow to the windshield, thecontrol valve may be stopped at any interme-diate position by positioning the WSHLD HTswitch to HOLD.

With both valves open, bleed air flows througha heat exchanger from which it is ducted to theoutlets at the base of each windshield. Theanti-ice heat exchanger cools the bleed airwith ram air. A ram-air modulating valve op-erates to maintain a 300˚ F duct temperaturedownstream of the heat exchanger by using aduct temperature sensor and a regulated bleedair servo line. The subsequent heat loss oc-curring in the duct as the bleed air reachesthe outlet nozzles keeps the outlet airflowtemperature within the limits of windshieldheat operation. During ground operation, ramair is not available to cool the bleed air.

Under normal conditions, the windshield heatbleed-air temperature is automatically con-trolled. However, an overheat warning sys-tem alerts the pilot and automatically shuts offwindshield heat in the event of an overheat con-dition. A low-limit (approximately 250˚ F)and a high-limit (approximately 290˚ F) ther-moswitch is installed in each windshield out-let nozzle. The low-limit switches functiononly on the ground and are cut out by the squatswitch relay box when airborne. The high-limit switches are installed primarily to limittemperature during airborne operation, butwill also function on the ground as a backupto the low-limit switches.

If either outlet nozzle temperature reaches the250˚ F limit (ground) or 290˚ F limit (air-

borne), the thermoswitch will illuminate thered WSHLD OV HT light on the glareshieldannunciator panel and cause the solenoid shut-off valve to close. The anti-ice control valvewill remain in the position it was in, but thegreen WSHLD HT light will be extinguishedwhile the solenoid shutoff valve is closed. TheWSHLD OV HT light will extinguish and theshutoff valve will open again when the tem-perature at the thermoswitch cools. If theWSHLD HT switch has not been turned off,airflow will resume to the windshield, thegreen WSHLD HT light will illuminate, andthe red WSHLD OV HT light will extinguish.

Through the squat switch relay box, the low -limit thermoswitches are disabled for 10seconds af ter touchdown. This preventsautomatic shutoff of bleed air at the momentof touchdown, which could restrict the pilot’svision due to loss of rain-removal capability.

Bleed air is not available for windshield anti-icing with both the left and right emergencypressurization valves in the emergency position.

Airplane SNs 35-663 andSubsequent, 36-063 andSubsequent, SNs 35-113through 662, and 36-032 through062 Incorporating AMK 91-2The exterior windshield defog, anti-ice, andrain removal system is shown in Figure 10-7.

With the engines running and the BLEED AIRswitches ON, engine bleed air from the regu-lated bleed-air manifold is available to twowindshield anti-ice system valves: the anti-iceshutoff valve and the anti-ice control valve.The shutoff valve is solenoid-operated and isnormally energized open whenever electricalpower is applied to the airplane. The controlvalve is motor-driven and is controlled by theWSHLD HT switch.

The t h r ee -pos i t i on (OFF–HOLD–ON)WSHLD HT switch is located on the anti-icecontrol panel. When the WSHLD HT switchis positioned to ON, the anti-ice control valve







Figure 10-7. Windshield Anti-ice System (Airplane SNs 35-663 and Subsequent, 36-064and Subsequent, SNs 35-113 through 662, and 36-032 through 063Incorporating AMK 91-2)

HIGHLIMITS

LOWLIMITS

WSHLDHT

HEATEXCHANGER

WSHLDANTI-ICE

CONTROLVALVE

WSHLDANTI-ICE

SHUTOFF(NC)

TO WING/STABHEAT

L MAINBUS

WSHLDHT ON

OFFRAM AIR

MANIFOLD BLEED AIR

CONDITIONED AIR

IN-FLIGHT THERMOSWITCH

GROUND THERMOSWITCH

ELECTRICAL CIRCUIT

SERVOPRESSURE

LINE

RAM AIRIN


TEMPERATURESENSOR

DEFOG HEATEXCHANGER

RAM AIR OUT

TOCABIN

CONTROLUNIT

WINDSHIELD

WSHLDHT

WSHLDOV HT

HOLD

LEGEND

begins to open, and the green WSHLD HTlight on the glareshield annunciator panel il-luminates. If the WSHLD HT switch is left inthe ON position, the control valve will drivefull open in approximately five to eight sec-onds. For reduced airflow to the windshield,the WSHLD HT switch may be positioned toHOLD before the control valve reaches fullopen. The control valve will then stop and re-main in an intermediate position.

With both valves open, regulated engine bleedair flows through a heat exchanger in whichit is cooled by ram-air. The ram-air flow iscont ro l led by a pneumat ica l ly ac tua tedmodulating valve. The modulating valve sensesbleed-air temperature, downstream of the heatexchanger, through a temperature sensor, andpositions itself automatically to maintain anair temperature of approximately 300˚ F. Fromthe heat exchanger, the temperature controlledbleed air is directed forward and dispensedover the outside of both the pilot’s and copi-lot’s windshields through outlets at the baseof each windshield.

Normally, the windshield anti-ice bleed-airtemperature is maintained at a safe level by theram air modulating valve. However, an automaticshutdown and warning system has been providedto prevent windshield damage from an overheatcondition. The system uses signals from fourthermoswitches, two under the windshield heatair outlets at the base of each windshield.

One thermoswitch on each side operates onlyon the ground while the other operates on theground and in the a i r. H igh- l imi t the r -moswitches are located on the left side andlow- limit thermoswitches are on the right.

If the bleed-air temperature at the windshieldreaches a low limit (250˚ F in flight or 215˚ Fon the ground), the anti-ice shutoff valve isdeenergized closed and the green WSHLD HTlight is extinguished. When the overheat cools,the thermoswitches will reset and the anti-iceshutoff valve will reopen. If the anti-ice con-trol valve is still open, the green WSHLD HT

light will illuminate and windshield anti-iceairflow will be restored.

If the bleed-air temperature at the windshieldreaches a high limit (270˚ F in flight or 250˚ Fon the ground; 215˚ F on the ground on airplaneswith electrically heated windshields), the anti-ice shutoff valve is deenergized closed, thegreen WSHLD HT light is extinguished, and thered WSHLD OV HT light illuminates. When theoverheat cools, the thermoswitches will reset,the red WSHLD OV HT light extinguishes, andthe anti-ice shutoff valve will reopen. If theanti-ice control valve is still open, the greenWSHLD HT light will illuminate and wind-shield anti-ice airflow will be restored.

The ground limit thermoswitches are disabledfor approximately 10 seconds after landing.This prevents automatic shutoff of bleed air,which could restrict the pilot’s visibility dueto loss of rain-removal, if the outlet temper-ature is between the in-flight and ground limitsat the moment of touchdown.

With loss of electrical power, the windshieldanti-icing system will be inoperative since theanti-ice shutoff valve will be deenergized andwill close. The control valve will remain in itslast position.

Bleed air is not available for windshield anti-icing with both the emergency pressurizationvalves in the emergency position.

Airplane SNs 35-002 through112 and 36-002 through 031Incorporating AAK 76-7Aand AMK 91-2The exterior windshield heat/defog systemcan be controlled either automatically ormanually (Figure 10-8). It is also used tosupplement cockpit heating through the pilot’sfootwarmers, and to provide an alternate bleed-air source for emergency pressurization.




An IN–NORMAL/OUT–DEFOG knob, lo-cated below the instrument panel to the left ofthe pedestal (Figure 10-3), manually controlsa valve which directs bleed air either to thewindshield or to the cockpit footwarmers.

When the knob is pushed in to the IN–NOR-MAL position, with the windshield anti-iceon, bleed air is directed into the cockpitthrough pilot’s and copilot’s footwarmers.This provides additional heat in the cockpitand an alternate source of bleed-air flow intothe cabin for emergency cabin pressurization.The knob is normally left in the IN–NOR-MAL position.


Two windshield heat switches are located onthe anti-icing panel. One is a three-positionswitch labeled “ON” and “OFF,” and is spring-loaded to the center (neutral) position. Theo the r swi t ch has two pos i t ions labe led“AUTO” and “MAN.”

Bleed air from the regulated bleed-air mani-fold is routed through two valves: the anti-iceshutoff valve and the anti-ice control valve.




Figure 10-8. Windshield Anti-ice System (Airplane SNs 35-002 through 112 and 36-002through 031 Incorporating AAK 76-7A and AMK 91-2)

HIGHLIMITS

LOWLIMITS

WSHLDOV HT

HEAT

EXCHANGER

RAM AIR

MANIFOLD BLEED AIR

CONDITIONED AIR

IN-FLIGHT THERMOSWITCH

GROUND THERMOSWITCH

ELECTRICAL CIRCUIT

WSHLD HEAT

ON AUTO

OFF MAN

RAM AIR OUT

DEFOG HEAT

EXCHANGER

WSHLDANTI-ICE

CONTROLVALVE

WSHLDHT

L MAINBUS

CONTROLUNIT

WINDSHIELD

INNORMAL

OUTDEFOG

OVERBOARDDRAIN

CHECKVALVE

FOOTWARMERS

TO WING/STABHEAT SERVO

PRESSURELINE

TOCABIN

WSHLDANTI-ICE

SHUTOFF(NC)

TEMPERATURESENSOR

RAM AIR

INRAM-AIR

MODULATING

VALVE

LEGENDWSHLD

HT

The shutoff valve is solenoid-operated and isdeenergized closed. Its function is to regulatethe engine bleed air from the manifold to 16psi. It is energized open when DC electricalpower is applied to the airplane and will bedeenergized and closed to shut off windshieldanti-ice in case of windshield overheat.

The con t ro l va lve i s mo to r-d r iven andcontrolled by either of the two switches on theanti-ice control panel. It takes four to five sec-onds to cycle fully. Selecting AUTO will openthe control valve and illuminate the greenWSHLD HT light. If MAN is selected, thecontrol valve may be opened or closed with theON–OFF switch. Since this switch is spring-loaded to neutral, it must be held in the ONposition while the valve drives toward the fullyopen position. The switch may be releasedbefore the valve reaches full open. The controlvalve will then stop and remain in an interme-diate position. The control valve can be closedonly by holding the ON–OFF switch to OFF(with MAN selected) for at least four seconds.

Operation

With both valves open, regulated engine bleedair flows through a heat exchanger in whichit is cooled by ram air. The ram-air flow iscont ro l led by a pneumat ica l ly ac tua tedmodulating valve. The modulating valve sensesbleed-air temperature downstream of the heatexchanger through a temperature sensor andpositions itself automatically to maintain anair temperature of approximately 300˚ F. Fromthe heat exchanger, the temperature-controlledbleed air is directed forward and dispensedover the outside of both the pilot’s and copi-lot’s windshield through outlets at the base ofeach windshield.

Normally, the windshield anti-ice bleed-airtemperature is maintained at a safe level by theram-a i r modula t ing va lve . However, anautomatic shutdown and warning system hasbeen provided to prevent windshield damagefrom an overheat condition. The system uses

signals from four thermoswitches, two underthe windshield heat air outlets at the base ofeach windshield.

One thermoswitch on each side operates onlyon the ground while the other operates on theground and in the a i r. H igh- l imi t the r -moswitches are located on the left side andlow-limit thermoswitches are on the right.

If the bleed-air temperature at the windshieldreaches a low limit (250˚ F in flight or 215˚ Fon the ground), the anti-ice shutoff valve isdeenergized closed and the green WSHLD HTlight is extinguished. When the overheat cools,the thermoswitches will reset and the anti-iceshutoff valve will reopen. If the anti-ice controlvalve is still open, the green WSHLD HT lightwill illuminate and windshield anti-ice airflowwill be restored.

If the bleed-air temperature at the windshieldreaches a high limit (270˚ F in flight or 250˚ Fon the ground; 215˚ F on the ground on airplaneswith electrically heated windshields), the anti-ice shutoff valve is deenergized closed, thegreen WSHLD HT light is extinguished, and thered WSHLD OV HT light illuminates. When theoverheat cools, the thermoswitches will reset,the red WSHLD OV HT light extinguishes, andthe anti-ice shut-off valve will reopen. If theanti-ice control valve is still open, the greenWSHLD HT light will illuminate and wind-shield anti-ice airflow will be restored.

The ground limit thermoswitches are disabledfor approximately 10 seconds after landing.This prevents automatic shutoff of bleed air,which could restrict the pilot’s visibility dueto loss of rain-removal, if the outlet temper-ature is between the in-flight and ground lim-its at the moment of touchdown.

With loss of electrical power, the windshieldanti-icing system will be inoperative since theanti-ice shutoff valve will be deenergized andwill close. The control valve will remain in itslast position.




INTERNAL WINDSHIELDDEFOG

GeneralAll airplanes use conditioned engine bleedair for internal windshield defog (See Chap-ter 11, “Air Conditioning,” for additional in-f o r m a t i o n ) . O n l a t e m o d e l a i r p l a n e s ,auxiliary internal windshield defog systemshave been provided.

Internal Windshield Defog(Airplane SNs 35-643through 670)The internal windshield defog system on theseairplanes uses an electrically heated coil, in thebleed-air duct leading into the cockpit, andthe Freon air-conditioning system. It is con-trolled by a three-position (OFF–CKPT–W/SAUX DEFOG HEAT) switch on the anti- icecontrol panel.

To avoid damage to the electrically heatedcoil, the crew should ensure that adequatebleed-air flow is available in the duct to coolthe coil before using the auxiliary windshielddefog system.

Positioning the switch to CKPT applies DCpower to the coil, heating all the air cominginto the cockpit.

Positioning the switch to W/S AUX DEFOGHEAT again applies DC power to the coil,heating all the air coming into the cockpit. Italso arms the Freon air-conditioning systemso it will turn on automatically as the air-plane descends through 18,000 feet. Whenthe Freon air-conditioning system turns on,electrically actuated diverter doors on thecabin blower assembly will open and directthe cold air into the space between the cabinheadliner and the fuselage skin. This dehu-midifies the cabin air without lowering thecabin temperature excessively. (See Chapter11 for additional information on the Freonair-conditioning system.)

DC electrical power to heat the auxiliarywindshield defog coil is provided by thebattery charging bus through two, 20-ampcurrent limiters. DC control power for theauxiliary windshield defog system is providedby the AUX DEFOG circuit breaker on theleft essential A bus.

Internal Windshield Defog(Airplane SNs 35-671 andSubsequent, and 36-064and Subsequent)The internal windshield defog system on theseairplanes is shown in Figure 10-9. It uses 163-VAC power from the auxiliary and secondaryinverters and is controlled by a two-position(OFF–WSHLD–DEFOG) switch located onthe anti-ice control panel (Figure 10-1).

When the switch is positioned to WSHLDDEFOG, DC control power is applied to awindshield defog relay box. The relay boxreceives 163-VAC power, through 5-ampcurrent l imiters , f rom the auxi l iary andsecondary inverters and directs it to the heatingelements in the windshield. Each heatingelement is a thin, gold film laminated in thewindshield. The auxiliary inverter powers theelement on the left side and secondary inverterpowers the element on the right side.

Both heating elements are turned on and offtogether, but, once operating, the two elementsare controlled separately by the relay box.

Two temperature sensors are located on eachside of the windshield. One sensor is set to lookfor a windshield temperature of approximately110˚ F. When the windshield reaches 110˚ F,the sensor will signal the relay box, whichremoves electrical power from the heatingelement on that side. As the temperature cools,the relay box will reapply power to maintaina cons t an t w indsh i e ld t empe ra tu r e o fapproximately 110˚ F.







Figure 10-9. Electric Windshield Defog System—Models 35-671 and Subsequentand 36-064 and Subsequent)

WSHLDDEFOG

L R

HEATING ELEMENT(GOLD FILM)

HEATING ELEMENT(GOLD FILM)

L ESS B

L WSHLDDEFOG

R WSHLDDEFOG

R ESS B

SECINVERTER

AUXINVERTER

163 VAC IN

163 VAC OUT

110° FLEFT RIGHT

WINDSHIELD DEFOG RELAY BOX

110° F

163 VAC OUT

163 VAC IN

BELOW 90/ABOVE 150° F

BELOW 90/ABOVE 150° F

WSHLD DEFOG SWITCH

The second sensor will signal the relay boxin the event of an underheat or an overheatcondition. If the windshield temperature is ap-proximately 90˚ F or below, or approximately150˚ F or above, the sensor will signal therelay box. In either situation, the relay box willi l l u m i n a t e a n a m b e r W S H L D D E F O Gannunciator light. If an overheat conditionexists, the relay box will also remove electricalpower from the heating element in the af-fected windshield.

The d i ff e rence be tween an ove rhea t o runderheat temperature condition may bedetermined by touching the windshield. If anoverheat temperature condition is suspected,and the windshield does not cool off, the relaybox has not removed electrical power fromthe heating element and the system should beturned off.

A windshield temperature of 90˚ F or belowis common when the defog system is firstturned on, and the annunciator light willilluminate. However, the light should soonextinguish as the windshield warms up.

The WSHLD DEFOG annunciator l ight ,located to the left of the left ENG FIRE PULLT-handle, consists of three separate lights andis controlled by the windshield defog relaybox. The upper WSHLD DEFOG light willilluminate when either of the lower lightsilluminate. The lower L and R lights willi l luminate to indicate which s ide of thewindshield has malfunctioned.

The WSHLD DEFOG annunciator light willilluminate in the event of an underheat or over-heat condition, as explained above. It will alsoilluminate with loss of DC or AC electricalpower if the defog system switch is in theWSHLD DEFOG position.

The electric windshield defog system uses163-VAC power as explained previously. DCcontrol power for the system is provided by theL and R WSHLD DEFOG circuit breakers onthe L and R essential B buses.

WINDSHIELD/RADOMEALCOHOL ANTI-ICE SYSTEM

GeneralMethyl alcohol from a reservoir located in theleft side of the nose compartment is providedto prevent ice formation on the radome and,if necessary, the pilot’s windshield as a backupfor the windshield anti-ice (defog) system.The systems are operated by DC power fromthe right main bus.

There are two different systems in use.

Airplane SNs 35-002 through112 and 36-002 through 031A DC motor-driven pump supplies filteredalcohol from a 2 1/4-gallon reservoir to theradome only, or to the radome and pilot’swindshield, depending on the position selectedon the WSHLD/RADOME switch on thepilot’s anti-icing control panel.

When the switch is positioned to RAD, thepump is energized and alcohol is delivered tothe radome only due to a normally-closedsolenoid valve in the windshield supply line.In this case, a fully serviced reservoir shoulddispense alcohol for approximately 1 hourand 30 minutes.

When the switch is positioned to WSHLD/RADOME, the pump is energized and thesolenoid valve in the windshield supply line isenergized open so that alcohol is delivered toboth surfaces. Flow to the windshield isdispensed through an orifice assembly integratedwith the pilot’s defog outlet. In this case, du-ration is reduced to approximately 43 minutes.

A pressure switch installed in the radomesupply line actuates the amber ALC AI lightwhen the reservoir is empty or if the pump fails.The light will extinguish when the controlswitch is turned off (Figure 10-10).




The reservoir is vented through an open venttube located in the same area as the pitot-staticdrains on the left-hand side of the nose com-partment. A pressure relief valve operates torelieve excessive supply line pressure byreturning it to the reservoir. Some airplanes areequipped with a siphon-break valve to preventthe siphoning of fluid from the tank after thesystem has been turned off (Figure 10-10).




Figure 10-10. Alcohol Anti-ice System (Airplane SNs 35-002 through 112 and 36-002through 031)

PRESSURERELIEF

PILOT'S EXTERNALDEFOG OUTLET

ORIFICE ASSEMBLY

ANTI-ICEVALVE (NC)

WSHLD &RADOME

OFF

RADOME

RADOME

LOW-PRESSURESWITCH

ALC AI

ALCPMP

R MAINBUS

MOTOR-DRIVENPUMP

SIPHON-BREAKVALVE *

OVERBOARDVENT FILTER

LEGEND

SUPPLY

PRESSURE

RETURN

AMBIENT

ELECTRICAL CIRCUIT

EFFECTIVE WITH35-076, 36-021

*

Airplane SNs 35-113 andSubsequent and 36-032and SubsequentMethyl alcohol is stored in a 1.75-gallonreservoir. When the cockpit control switch ispositioned to WSHLD/RADOME or to RAD,circuits are completed to position a 3-wayvalve in the fluid supply line (Figure 10-11)and also to open the shutoff valve and pressureregulator in the servo bleed-air supply line.

Servo bleed air tapped from the high-pressurebleed-air manifold passes through the shutoffvalve and pressure regulator where i t isregulated to 2.3 psi and sent to pressurize thealcohol reservoir.

The alcohol is forced through a filter to thethree-way valve which is positioned accordingto the selected switch position.

The pressure relief valve, set at 2.6 psi, relievesany overpressure in the reservoir should thepressure regulator fail, and bleeds off residualpressure when the control switch is turned off.

The float switch in the reservoir illuminatesthe ALC AI annunciator when the tank isempty. The light stays on even if the switch isoff as a reminder to service the reservoir.

If the RAD position is selected, a fully servicedreservoir supplies only the radome withapproximately 2 hours and 9 minutes of al-coho l . When s e l ec t ed t o t he WSHLD/RADOME position, alcohol is also dispensedto the pilot’s defog outlet via the three-wayvalve, and duration of the supply is reducedto approximately 45 minutes. This system isstill operational if both emergency pressur-ization valves are in emergency (provided DCpower is available).




Figure 10-11. Alcohol Anti-ice System (Airplane SNs 35-113 and Subsequent and 36-032and Subsequent)

FILTER

ALCOHOLRESERVOIR

ALCSYS

R MAINBUS

WSHLD/RADOME

RAD

OFF

PRESSUREREGULATOR ANDSHUTOFF VALVE

(NC)

PRESSURERELIEFVALVE

ALCAI

WRNLTS

L & RESSBUS

HPSERVO

BLEED AIR

CHECKVALVE

TO OTHER SERVOSYSTEMS

HP SERVO BLEED AIR

REGULATED BLEED AIR

ALCOHOL SUPPLY

OVERBOARD (AMBIENT)

ELECTRICAL

LEGEND

BLEED AIR

WING AND HORIZONTALSTABILIZER ANTI-ICE SYSTEM

GeneralBleed air is used to prevent ice formation on thewing and horizontal stabilizer leading edges.The bleed air is directed from the regulated en-gine bleed-air manifold through a solenoid-operated pressure regulator valve (Figure 10-12)to the respective leading-edge surfaces.

Controls and IndicationsThe STAB WING HEAT switch located on thepilot’s anti-icing control panel controls thevalve. When the switch is moved up (on), thevalve is energized open. With the switch off, orif DC power fails, the valve deenergizes closed.

With the valve open, manifold bleed air isthen routed through the wing-s tabi l izerpressure-regulator valve where it is regulated




Figure 10-12. Wing and Horizontal Stabilizer Anti-ice System

WING

TEMP

35°

215°

WINGOV HT215°

SCUPPERSCUPPER

STAB &WING

STABWINGHEAT

STAB/WINGPRESSURE

REGULATORVALVE

(NC)CHECK VALVE

MANIFOLD

LEFT-HANDENGINE

CHECK VALVERIGHT-HANDENGINE

STAB

TEMP

35°

215° STABOV HT215°

TO FLOWCONTROL

VALVE LEGENDMANIFOLD BLEED AIR

REGULATED FLOW

ELECTRICAL CIRCUITS

R MAINBUS

to 16 psi, and finally to piccolo tubes in theleading edges of the wing and the horizontalstabilizer. After the bleed air has heated itsrespective leading edge, it continues outboardwhere it is vented overboard; each wing has ascupper vent on the underside of the leadingedge, while the horizontal stabilizer has holesat each tip.

On the glareshield annunciator panel, redWING OV HT and STAB OV HT lights areilluminated should their respective sensors(Figure 10-12) detect 215˚ F.

Separate WING TEMP and STAB TEMPindicators, located on the center instrumentpanel (Figure 10-13), indicate leading-edgeskin temperature and are color-coded asfollows:

Red—Temperature below 35˚ F (dangerof icing in visible moisture)

Green—Temperature between 35–215˚ F(normal operation)

Yellow—Temperature above 215˚ F (pos-sible overheat)

When either overheat light comes on, and thesystem is turned off, the light will remain onuntil the temperature drops to within limits.The STAB WING HEAT switch may be turnedback on, but the pilot must visually monitorthe applicable skin temperature indicator andcycle the system on and off to maintain tem-perature in the green arc.

Stabilizer heat and wing heat are not availablewhen both emergency pressurization valves (ifinstalled) are in EMERGENCY. This is coveredin Chapter 12, “Pressurization.” DC power forsystem operation is through the STAB & WINGHT circuit breaker on the right main bus.

PITOT, STATIC, ANDANGLE-OF-ATTACK VANEANTI-ICE SYSTEM

Pitot and Angle-of-AttackVane Anti-icingThe left and right pitot tubes and angle-of-attack (AOA) vanes contain electrical heatingelements. The L and R PITOT HEAT switcheslocated on the pilot’s anti-icing control panel(Figure 10-1) each supply essential bus powerto both respective heating elements. Eventhough each se t o f hea t ing e l emen t s i scontrolled by the same switch, separate circuitprotection for the AOA vane heater is provided;the L and R PITOT HT circuit breakers (forpitot heaters) and S WRN HT circuit breakers(for the AOA vane heaters) are all located onthe left and right essential buses, respectively.

On FC 530 airplanes, one heating element ineach pitot-static probe heats all of the pitot andstatic ports.

Dual amber L and R PITOT HEAT monitorlights are available as an optional feature andare located on either outboard side of theglareshield panel or on the instrument panel.On airplane SNs 35-271 and 36-045 and sub-sequent, a single amber PITOT HT light isstandard equipment and is located on the an-nunciator warning light panel (AnnunciatorPanel section). In either case, the light(s) il-luminates when the pitot heat switches areturned off or to indicate failure of power to apitot tube element (the AOA vanes are notmonitored).

Static Port Heating(FC 200 Only)There are five static ports—two on the leftside fuselage and three on the right side. Pilotinstruments are supplied static pressure bythe interconnected left front and right centerports which are heated. The interconnectedleft rear and right front static ports supply thecopilot’s static pressure and are also heated.




The right rear port, interconnected with analternate port inside the nose compartment, isused by the altitude controller and does notrequire heat.

Two additional shoulder-static ports, locatedforward of the windshield, are also heated.These ports are used by the air data sensor.

All static port heating elements are connecteddirectly to their respective L or R PITOT HTcircuit breakers. Consequently, they are heatedwhenever airplane DC power is available,provided the circuit breakers are closed (in).




Figure 10-13. WING TEMP and STABTEMP Indicators




1. Bleed air is used for anti-icing on:A. Pitot tubes and static portsB. PT2 TT2 sensorsC. Wing and ho r i zon t a l s t ab i l i z e r

leading edgesD. Conical fan spinners

2. The L or R PITOT HEAT switches alsosupply heating element power for:A. The angle-of-attack vanesB. The shoulder static ports C. The instrument static portsD. PT2 TT2 probe heater

3. The crew action required when the redWING OV HT light illuminates is:A. No action is required; the system is

automatic.B. Posi t ion the STAB WING HEAT

switch to STAB.C. Turn the STAB WING HEAT switch

to OFF or reduce power.D. Turn one BLEED AIR switch to OFF

until the light goes out.

4. The internal windshield defog systemuses:A. 230-VAC power B. 163-VAC powerC. An electrically heated coil and the

Freon air-conditioning systemD. Engine bleed-air pressure

5. Anti-icing equipment must be turned on:A. When in icing conditionsB. Before entering icing conditionsC. Before takeoffD. During climbout

6. With the loss of airplane electrical power,anti-icing will be lost on:A. All systemsB. Pitot, static, and PT2 TT2 probes onlyC. All systems except the nacelle inlet

lipsD. All systems except the windshield

and radome alcohol system

7. The L NAC HEAT switch in the up (on)position provides anti-icing to all of thefollowing except the:A. Nacelle lipB. Dome spinner (early models)C. PT2TT2 probeD. Conical spinner (late models)

8. The alcohol anti-ice system may be usedto anti-ice the:A. RadomeB. Copilot’s windshieldC. Pilot’s windshieldD. Both A and C

QUESTIONS

11-i

CHAPTER 11AIR CONDITIONING

CONTENTS

Page

INTRODUCTION ................................................................................................................. 11-1

GENERAL............................................................................................................................. 11-1

ENGINE BLEED-AIR CONDITIONING AND DISTRIBUTION ..................................... 11-2

General ........................................................................................................................... 11-2

Flow Control Valve ........................................................................................................ 11-5

Hot Air Bypass Valve (H-Valve).................................................................................... 11-5

Ram-Air Heat Exchanger............................................................................................... 11-6

Ram-Air Ventilation....................................................................................................... 11-6

Cabin and Cockpit Air Distribution ............................................................................... 11-6

Temperature Control ...................................................................................................... 11-8

AUXILIARY AIR-CONDITIONING SYSTEMS.............................................................. 11-10

General......................................................................................................................... 11-10

Distribution System ..................................................................................................... 11-10

Auxiliary Cooling System ........................................................................................... 11-14

Auxiliary Heat System (Optional) ............................................................................... 11-16

QUESTIONS....................................................................................................................... 11-19




11-iii

ILLUSTRATIONS

Figure Title Page

11-1 Engine Bleed-Air Conditioning System (SNs 35-002 through 35-086,except 35-082, and 36-002 through 36-022) ........................................................ 11-2

11-2 Engine Bleed-Air Conditioning System (SNs 35-082, 35-087 through 35-112,36-023 through 36-031, and Earlier Airplanes Incorporating AMK 76-7) ............. 11-3

11-3 Engine Bleed-Air Conditioning System (SNs 35-113 andSubsequent and 36-032 and Subsequent) .............................................................. 11-4

11-4 CABIN AIR Switch .............................................................................................. 11-5

11-5 Temperature Control Indicator ............................................................................... 11-5

11-6 Conditioned Bleed-Air Distribution ....................................................................... 11-7

11-7 CABIN CLIMATE CONTROL Panel......................................................................11-8

11-8 CABIN TEMP Indicator ........................................................................................ 11-9

11-9 Evaporator and Blower Assembly........................................................................ 11-11

11-10 Cabin Blower Grille Outlet .................................................................................. 11-10

11-11 COCKPIT AIR and CABIN BLOWER Rheostats .............................................. 11-12

11-12 Cockpit Upper Air Outlets ................................................................................... 11-13

11-13 Passenger Overhead Outlets (WEMACS) .......................................................... 11-13

11-14 Freon Refrigeration System Schematic ............................................................... 11-15

11-15 Auxiliary Heating System Components............................................................... 11-17




INTRODUCTIONAir conditioning in the Learjet 35/36 is furnished primarily by regulated engine bleedair, which is temperature controlled and distributed throughout the cabin and cockpitareas. This is the same bleed air that is used for cabin pressurization.

Additional cooling and heating is provided by a Freon refrigeration system and an op-tional auxiliary electrical heating system. These systems share a separate distributionnetwork through which cabin air is recirculated by a cabin blower and a cockpit fan.

GENERALPrimary heating and cooling is accomplishedby controlling the temperature of the bleedair entering the cabin by circulating it throughan air-to-air heat exchanger. The cabin andcockpit distribution systems differ slightly,based on airplane serial number.

Additional refrigeration cooling by the Freonsystem is available for ground operations andin flight at altitude up to a maximum of 18,000feet or 35,000 feet, depending on compressormotor part number.

Additional heating by the auxiliary electricalheating system (if installed) can be obtained forground operations and at any altitude in flight.

CHAPTER 11AIR CONDITIONING




ENGINE BLEED-AIRCONDITIONINGAND DISTRIBUTION

GENERALThis sect ion addresses the condi t ioningprocess that the engine bleed air is subjected

to before it enters the cabin area, beginninga t t h e f l ow c o n t r o l va l v e . C h a p t e r 9 ,“Pneumatics,” describes the bleed-air supplysystem. Chapter 12, “Pressurization,” de-scribes how conditioned bleed air is usedfor cabin pressurization.

Regulated engine bleed air, supplied to amanifold located in the tailcone section, isducted to the flow control valve. From the flow




CABIN AIR

AUTO

HOT

MANMANHOT COOL

FANCOLDCABIN CLIMATE CONTROL

HT VALIND

LESSBUS

LMAINBUS

CABHT

AUTO

CABIN AIR DIFFUSERS(TYPICAL)

TO SENSORBLOWER MOTOR

BLEED AIR(LEFT ENGINE)

DUCT TEMPLIMITER

DUCT TEMPSENSOR

RAM-AIRCHECK VALVE

RAM AIR IN

HEAT EXCHANGER

HOT AIR BYPASS(H-VALVE)

RAM AIR OUT

AIR DISTRIBUTIONTO LOWER CABIN DOOR

CREW OUTLETS

INTERNAL DEFOGOUTLETS

FOOTWARMER OUTLET

CABINTEMP

SENSOR

AIR DISTRIBUTIONCHECK VALVES

BLEED AIR(RIGHT ENGINE)

FLOW CONTROL VALVE

VENTURI

AIR BLEED

BLEED AIR

LEGEND

RAM AIR

CONDITIONED BLEED AIR

LMAINBUS

MAXNORMOFF

OFF

COLD

HOT TEMP

CONT

CABHT

MNL

RMAINBUS

Figure 11-1. Engine Bleed-Air Conditioning System (SNs 35-002 through 35-086, except35-082, and 36-002 through 36-022)

control valve there are three slightly differentcabin and cockpit distribution configurations,each performing the same basic functions, but

differing in component arrangement. Figures11-1, 11-2, and 11-3 depict the three basicconfigurations by airplane serial number.




CABIN AIR

AUTO

HOT

MANMANHOT COOL

FANCOLDCABIN CLIMATE CONTROL

HT VALIND

LESSBUS

LMAINBUS

CABHT

AUTO




DUCT TEMPLIMITER

DUCT TEMPSENSOR

RAM-AIRCHECK VALVE

RAM AIR IN

HEAT EXCHANGER


RAM AIR OUT

AIR DISTRIBUTIONTO LOWER CABIN DOOR

CREW OUTLETS


FOOTWARMER OUTLET

CABIN TEMPSENSOR

BAGGAGE COMPARTMENTAIR DIFFUSER(35A AIRPLANE ONLY)



FLOW CONTROL VALVE

VENTURI

PRESSURE SWITCH(47 PSI)

AIR BLEED

BLEED AIR

LEGEND

RAM AIR


LMAINBUS

MAXNORMOFF

OFF

COLD

HOT TEMP

CONT

CABHT

MNL

RMAINBUS

Figure 11-2. Engine Bleed-Air Conditioning System (SNs 35-082, 35-087 through 35-112,36-023 through 36-031, and Earlier Airplanes Incorporating AMK 76-7)

All three configurations use the flow controlvalve to control the flow of bleed air througha hot air bypass valve and an air-to-air heat

exchanger before it enters the cabin and cock-pit distribution ducting.




HEAT EXCHANGER

RAM AIR OUT

LESSBUS

AIR BLEED

BLEED AIR

LEGEND

RAM AIR


REGULATED SERVO AIR

FLOW CONTROL VALVE


DUCT TEMPSENSOR

RAM AIR CHECKVALVE

RAM AIR IN

VENTURI

HT VALIND

LMAINBUS



EMERGENCYPRESSURIZATIONVALVE

SERVOBLEED

AIR

ON

OFF

CABINAIR

AUXDEFOG/CREWHEATER *



CABINTEMP

SENSOR


BAGGAGE COMPARTMENTAIR DIFFUSER

(35A AIRPLANE ONLYL)


CHECK VALVES

DUCT TEMPLIMITER

SERVOBLEED AIR

AUTO MAN

COLD HOT

COOL

FAN

CREW OUTLETSFOOTWARMER OUTLET

CABIN

CLIMATE

OFF

* SNs 35-643 AND SUBSEQUENT AND SNs 36-064 AND SUBSEQUENT

COLD

HOTTEMP

CONT

Figure 11-3. Engine Bleed-Air Conditioning System (SNs 35-113 and Subsequentand 36-032 and Subsequent)

FLOW CONTROL VALVEThe flow control valve is a solenoid-operatedvalve which controls and regulates the flow ofbleed air into the cabin. The position of thevalve is determined by the CABIN AIR switch(Figure 11-4). The most current airplanes(Figure 11-3) use a two-position OFF–ONswitch. Earlier airplanes (Figures 11-1 and11-2) use a three-position OFF–NORM–MAXswitch. When the CABIN AIR switch is inOFF, the valve is energized and closes. Whenthe switch is in ON or NORM, the valve isdeenergized and opens. In the MAX position,the valve opens fully to provide an increase inairflow to the cabin. DC power for valve op-eration is provided through the AIR BLEEDcircuit breaker on the left essential bus.

A venturi, located downstream of the flowcontrol valve adjusts the valve to smooth outthe flow of bleed air as it enters the cabin.Airflow through the venturi is measured bypneumatic sensing lines connected to a mod-ulating mechanism in the flow control valvewhich ensures that airflow remains constantwhen engine power changes occur.

HOT AIR BYPASS VALVE(H-VALVE)A butterfly bypass valve, more commonlyreferred to as the “H-valve,” is located in the

bleed-air duct upstream of the heat exchanger.Its function is to split the flow of bleed air,directing some to the heat exchanger for cool-ing, and some to bypass the heat exchanger.The result is a mixture of the two airflows,thereby conditioning the bleed air before it en-ters the cabin area. The position of the H-va lve i s i nd ica ted on the TEMP CONTindicator located in the lower center instrumentpanel (Figure 11-5).




Figure 11-5. Temperature ControlIndicator

SNs 35-002 THROUGH 35-112AND 36-002 THROUGH 36-031

SNs 35-113 AND SUBSEQUENTAND 36-032 AND SUBSEQUENT

Figure 11-4. CABIN AIR Switch

On SNs 35-002 through 35-112 and 36-002through 36-031, the H-valve butterfly ispositioned by a DC electric motor operated bythe climate control system. Approximately25 seconds is required for the valve to travelfrom one extreme to the other. The valve willremain in its existing position in the event DCpower is lost.

On SNs 35-113 and subsequent and 36-032 andsubsequent, the H-valve butterfly is positionedpneumatically by servo bleed air (Chapter 9,“Pneumatics”) from the climate control sys-tem. No electrical circuits are involved ex-cept that the TEMP CONT indicator requiresDC power. Approximately 8 seconds is re-quired for the valve to travel from one ex-treme to the other. The valve is spring-loadedto the full cold position anytime servo airpressure is not available.

RAM-AIR HEAT EXCHANGERThe heat exchanger is located inside theta i lcone . I t cons is t s of a b leed-a i r coresurrounded by a ram-air plenum. Cool airenters the ram-air inlet in the dorsal fin andflows through the plenum, across the bleed-aircore, thus cooling the bleed air. The ram airthen exhausts overboard through a port in thelower left side of the fuselage.

The cooled bleed air flowing out of the heatexchanger core is ducted back to the bypassside of the H-valve where it mixes with hot by-passed bleed air. The resulting conditionedair is then directed into the cabin and cockpitdistribution system.

When the airplane is on the ground, do notperform extended engine operation above idlewi th t he CABIN AIR and BLEED AIRswitches positioned to ON. Since there is noram air for cooling of the bleed air, possibledamage to the air-conditioning componentscould result. Damage might also occur to in-terior cabin furnishings, as well as overheatingthe tailcone area.

On SNs 35-082, 35-087 through 35-112, and36-023 through 36-031, and earlier airplanesincorporating AMK 76-7, the flow controlvalve is located downstream of the heatexchanger. Engine bleed air is available to thehea t exchange r wheneve r an eng ine i soperating and the BLEED AIR switches areone. Because of this, a pressure switch isinstalled in the tailcone ducting prior to theheat exchanger. Should this pressure switchactuate (at approximately 47 psi), both redBLEED AIR L and R annunciator l ightsilluminate to indicate the overpressure condition.

RAM-AIR VENTILATIONIn the event that the airplane is unpressurizedin flight, air for circulation and ventilation ofthe cabin and cockpit areas is provided byram air, which is ducted into the conditionedbleed-air distribution system.

During normal operation, a one-way checkvalve in the connecting ram-air duct preventsloss of conditioned pressurization bleed airthrough the ram-air plenum exhaust port.

CABIN AND COCKPITAIR DISTRIBUTIONConditioned airflow distribution to the cabinand cockpit areas is essentially the same for allairplanes (Figure 11-6). The conditioned air isrouted from the tailcone into the cabin areathrough two ducts, one on each side of thecabin. The left duct ends at the entry door, andthe right duct continues forward to the cockpit.




Cabin Air DistributionCabin air distribution is furnished by diffusersinstalled at intervals along the two ducts, andthey direct airflow toward the floor.

A one-way distribution check valve is locatedat the aft end of each cabin duct. These valvesare functionally related to the pressurizationsys t em, a s de sc r i bed i n Chap t e r 12 ,“Pressurization.”

On SNs 35-113 and subsequent and 36-032 andsubsequent (Figure 11-3), distribution of airchanges when either (or both) emergency pres-surization valves are positioned to emergency.

If only one emergency valve is positioned toemergency, all bleed air from that engine isrouted directly into only that side’s cabindistribution duct, and temperature control of

that air is lost. However, bleed air from the op-posite engine is still subject to the normalconditioning process. One-way check valvesin the normal distribution ducting prevent theemergency airflow from being lost throughthe normal distribution system.

If both emergency valves are positioned toemergency, all bleed air from both engines isrouted directly into the respective left andright distribution ducts. Temperature controlis then sacrificed for pressurization.

Cockpit Air DistributionCockpit air distribution is provided by theducting connected to the forward end of theright hand cabin duct. Four WEMAC outlets(two on each side of the cockpit), located onthe sidewall panels and adjacent to the outboardrudder pedals, enable the pilots to control and




AFT CABIN DIFFUSER

DIFFUSER (TYPICAL)

COPILOT’S CONDITIONEDAIR OUTLETS

CHECK VALVE

LOWER CABIN DOORAIR OUTLET

PILOT’S CONDITIONED AIR OUTLETS

FOOTWARMEROUTLET

Figure 11-6. Conditioned Bleed-Air Distribution

direct the airflow as desired. A footwarmerdiffuser, located below the instrument paneljust forward of the center pedestal, directscontinuous condition air along the center floor.Two piccolo tubes installed vertically on eachside of the windshield center support structuredirect a continuous flow of conditioned airacross the forward section of each pilot’s wind-shield for the interior windshield defogging.

On SNs 35-328 and subsequent, and 36-050 andsubsequent increased continuous interior wind-shield defogging capability has been provided.Two additional piccolo tubes are installed, onefor each windshield. They are positioned hor-izontally along the lower edge and extend for-ward from the aft corner of the windshield.This position results in improved interior de-fogging for the sides of the windshield.

Interior windshield defogging can be maximizedby closing the four WEMAC outlets to divertthe maximum amount of conditioned air tothe windshield piccolo tubes.

On SNs 35-002 through 35-112 and 36-002through 36-031, additional heat is available tothe cockpit via separate footwarmers that operatefrom the windshield heat/defog system dis-cussed in chapter 10, “Ice and Rain Protection.”

TEMPERATURE CONTROLTemperature control of the engine bleed-airentering the cabin area is accomplished byvarying the position of the H-valve butterfly.As the valve opens, less bleed air is directedto the heat exchanger for cooling, while morebleed air is bypassed, and mixed with the cooledair. Manual and automatic operation of the H-valve is achieved by controls on the CABINCLIMATE swi tch pane l , loca ted on thecopilot’s lower instrument panel (Figure 11-7).

On SNs 35-002 though 35-112 and 36-002through 36-031, the climate control system isoperated e lect r ica l ly. System control i saccomplished with a rheostat and a HOT–COLD toggle switch (spring load to center).Other system components include a tempera-ture sensor located behind the copilot’s seat,a duct temperature sensor and duct temperaturelimiter ( both located in the air duct downstreamof the H-valve) ( Figure 11-1 or 11-2, as ap-plicable), and a control unit.

If the rheostat is turned fully counterclock-wise to the MAN detent, the cabin temperaturesensor and duct temperature sensor are both off.The H-valve is then controlled manually by ac-tuating the spring-loaded HOT-COLD switch.The TEMP CONT indicator (Figure 11-5) dis-plays the position of H-valve. DC power




Figure 11-7. CABIN CLIMATE CONTROL Panel


SNs 35-113 AND SUBSEQUENTAND 36-032 AND SUBSEQUENT

for manual operat ion is provided by theCABIN HT MAN circuit breaker on the rightmain bus. The TEMP CONT indicator ispowered from the HT VAL IND circuit breakeron the left main bus.

If the rheostat is out of the MAN detent, theH-valve position is determined automaticallyby the control unit, which evaluates inputsfrom the rheostat , the cabin temperaturesensor, and the duct temperature sensor. Thecontrol system then responds by continuouslymodulat ing the H-valve to mainta in thedesired temperature. DC power for automaticoperation is provided by the CAB HT AUTOcircuit breaker on the left main bus.

Whether the system is being operated manuallyor automatically, the duct temperature limitersignals the control unit if the duct temperatureincreases to approximately 350˚ F. The controlunit’s response is to drive the H-valve to thefull cold position, and direct all bleed airthrough the heat exchanger.

On SNs 35-107, 35-113 and subsequent, and 36-032 and subsequent, the H-valve is positionedpneumatically by servo bleed air (Chapter 9,“Pneumatics”), and no electrical circuits are involved. The CLIMATE CONTROL panel(Figure 11-7) incorporates two control knobs.The AUTO–MAN knob is actually a servo bleed-air selector valve. The COLD–HOT knob is aneedle valve that controls the servo air pressureapplied to the H-valve butterfly (spring-loadedto the full cold position). Other system com-ponents include a temperature sensor located inthe upper forward cabin, a duct temperaturesensor , and a duct temperature limiter locatedin the air duct downstream of the H-valve (Figure11-3). The control system consists of an inter-connected servo bleed-air network.

With the AUTO–MAN knob in the MANposition, the selector valve isolates the controlsystem from the influences of the cabin tem-perature sensor and the duct temperature sensor.Servo air pressure is routed directly throughthe needle valve (controlled by the COLD–HOT

knob ) to the H-valve butterfly, which is spring-loaded to the full cold position. Changing theCOLD–HOT knob position simply changes theservo air pressure on the H-valve, butterfly.The TEMP CONT indicator (Figure 11-5) dis-plays the relative position of the H-valve, whichis the only component in the system that re-quires DC electrical power. DC power is pro-vided through the HT VAL IND circuit breakeron the left main bus.

With the AUTO–MAN knob (selector valve)in the AUTO position, the servo pressure controlnetwork samples the needle valve setting(COLD/HOT knob position), the cabin tem-perature sensor (existing cabin temperature),and the duct temperature sensor (actual tem-perature of the bleed air inside the duct).Servo air pressures are modulated by the con-trol system, which causes the H-valve butterflyto modulate, thereby keeping the cabin tem-perature constant.

For manual or automatic operation, in the eventof a duct overheat, the duct temperature lim-iter causes the control system to shut off servoair pressure to the H-valve butterfly. This al-lows it to spring to the full cold position, di-recting all bleed air through the heat exchanger.

A CABIN TEMP indicator may be installedon the center pedestal or instrument panel toindicate the temperature in the cabin from aremote sensor (Figure 11-8).




Figure 11-8. CABIN TEMP Indicator

AUXILIARY AIR-CONDITIONINGSYSTEMS

GENERALAdditional air circulation is provided by acabin blower and a cockpit fan, ducted thoughdistribution networks that are also used bythe Freon refrigeration system (auxiliarycooler) and the optional electrical heating sys-tem (auxiliary heater).

The cabin blower and cockpit fan may be usedsimply to recirculate air within the cabin andcockpit areas, or by using the auxiliary cooleror auxiliary heater, to cool or to heat the re-circulated air.

For operational requirements on the ground(subject to certain limitations), it is possible toprecool or preheat the cabin prior to engine start.

DISTRIBUTION SYSTEMThe heart of the distribution system is theevaporator and blower assembly, installed inthe cabin ceiling above the baggage com-partment (Figure 11-9). The assembly housesthe ducting, the cabin blower assembly, thecockpit fan assembly, the Freon system evap-orator, and the optional electrical heating el-ements (when installed).

Cabin Blower DistributionThe cabin blower assembly consists of twosquirrel-cage blowers driven by a single DCmotor. The blowers draw air from the bag-gage compartment area though the evaporatorand force it through separate ducts to a lou-vered grille at the front the ducts. The air isdiffused as it blows out directly into the cabin.When installed, the optional heating elementsare located within these ducts.

Diverter doors are installed in the ductingforward of the cabin blower. On SNs 35-002through 642 and 36-002 through 063, the doorsare located in the bottom of each duct and aremanually controlled and actuated by theOPEN–CLOSE knob adjacent to the louveredgrill (Figure 11-10). On SNs 35-643 through35-646, electrically controlled and actuated di-verter doors are installed in the top of each ductalong with the mechanically controlled doorson the bottom. On SNs 35-647 and subse-quent, and 36-064 and subsequent, only theelectrically controlled doors are installed.

On airplanes with the manual diverter doors,when the knob is rotated to the OPEN position,the diverter doors are raised up into the airflowfrom the cabin blower and divert the air downinto the baggage compartment. When the knobis positioned to CLOSE, the doors are flushwith the bottom of the ducting and the airflowfrom the cabin blower is directed into the cabin.




Figure 11-10. Cabin Blower Grille Outlet




Figure 11-9. Evaporator and Blower Assembly

On airplanes with the electric diverter doors,the doors are controlled by a two-position,ON–OFF, rocker switch located below thecabin blower air outlet. When the switch is po-sitioned to OFF, the diverter doors are loweredinto the airflow from the cabin blower and di-rect the air into the space between the cabinheadliner and the fuselage skin. When theswitch is positioned to ON, the diverter doorsare flush with the top of the ducting and theairflow from the cabin blower is directed intothe cabin. On SNs 35-643 through 35-670,the doors may also be controlled by the aux-iliary internal windshield defog system. SeeChapter 10, “Ice and Rain Protection,” foradditional information.

When used simply for additional air circula-tion, the cabin blower is turned on by selectingFAN on the three-position FAN–OFF–COOLswitch, located on the CABIN CLIMATECONTROL panel (Figure 11-7). DC electri-cal power is provided by the CAB BLO circuitbreaker on the left main bus. On SNs 35-113and subsequent and 36-032 and subsequent,variable blower speed control is affordedthrough the CABIN BLOWER rheostat, lo-cated on the copilot’s sidewall panel (Figure11-11). Earlier airplanes do not have this feature unless AMK 77-4 is incorporated.




Figure 11-11. COCKPIT AIR and CABIN BLOWER Rheostats

Cockpit Fan DistributionBetween the two ducts fed by the cabin blowersis another duct which encloses the axial cockpitfan. This fan draws air from the baggage com-partment area through the evaporator, but itsoutput is furnished directly to four smallerducts concealed in the cabin overhead panel-ing. Two of these ducts run directly to the twolouvered overhead outlets in the cockpit(Figure 11-12). On SNs 35-092 and 36-025 andsubsequent, two additional ducts (one on eachside) are connected to the individual overheadWEMAC outlets above each of the passengersea t s (F igu re 11 -13 ) . A i r vo lume anddirectional control is provided at each outlet.The fan motor is cooled by the air it movesthrough the ducting.

The cockpit fan is controlled by the COCKPITAIR rheostat on the copilot’s sidewall panel(Figure 11-11) using DC power from the CABBLO circuit breaker on the left main bus.

On SNs 35-002 through 35-112, and 36-002through 36-031, the OFF detent is at the fullclockwise posit ion, and the fan speed isincreased by rotating the rheostat in a counter-clockwise direction. On SNs 35-113 andsubsequent and 36-032 and subsequent, theOFF detent is at the full counterclockwiseposition, and speed is increased by rotating therheostat in the clockwise direction.

If all the cockpit and overheat outlets areclosed, the cockpit fan must be operatedbecause no cooling airflow for the motor isavailable and the motor will overheat.




Figure 11-12. Cockpit Upper Air OutletsFigure 11-13. Passenger Overhead

Outlets (WEMACS)

AUXILIARY COOLING SYSTEMA freon refrigeration system (auxiliary cooler)is installed to provide supplemental coolingfor ground and in-flight operations and can alsobe used for dehumidification.

System components, identified schematically(Figure 11-14), are conventional. The com-pressor (belt-driven by a 3 3/4 horsepowermotor), the condenser, and the dehydrator arelocated inside the tailcone. The compressormotor is cooled by air from the tai lconeventilation airscoop on the left side of thefuselage. The evaporator and expansion valveare located inside the evaporator and blowerassembly (above the baggage compartment).

OperationElectrical power for system operation mustbe supplied by either a GPU or an engine-driven generator. The system is turned on byselecting the COOL position on the FAN–OFF–COOL switch. DC power is appliedsimultaneously to the compressor motor andthe cabin blower motor.

For best results on the ground, the CABINAIR switch should be off to keep warm bleedair from entering the cabin while the enginesare running.

Cool air is drawn through the evaporator andcirculated as already described in CabinBlower Distribution, except that the blowermotor will run continuously at its maximumspeed (the CABIN BLOWER rheostat, ifins ta l led , i s rendered inopera t ive) . Thecompressor motor is powered from the batterycharging bus through a 150-amp current limiterand a control relay powered from the FREONCONT circuit breaker on the left main bus.

The diverter doors may be positioned as desiredto control airflow into the cabin through thelouvered grille above the divan seat. If desired,the cockpit fan may also be used to providewider circulation of the cooled air to the cockpitand passenger WEMAC outlets.

When the cooling system is being poweredby a GPU, it is possible in some conditions forthe airplane batteries to be depleted if GPUfailure occurs.

The compressor motor i s au tomat ica l lydeenergized when the START–GEN switch isselected to START. However, normal operatingprocedures require that the FAN–OFF–COOLswitch be in the OFF or FAN position prior toengine start to preclude possible electricalsystem damage.







Figure 11-14. Freon Refrigeration System Schematic.

AUXILIARY HEAT SYSTEMS(OPTIONAL)Two optional electric auxiliary heat systemsare available: one for the cabin and one for thecockpit. Both systems may be used to provideadditional heating on the ground or in flight.

Auxiliary Cabin Heat System

General

The auxiliary cabin heat system uses the cabinblower to circulate heated air. It also incor-porates two, dual-coil heating elements, onelocated in each of the cabin blower ducts(Figure 11-15). Each heating element containsa thermoswitch set for high and low limits(150˚ F and 125˚ F), and a thermal fuse foroverheat protection.

On SNs 35-002 through 35-646, and 36-002through 36-063, if the manual diverter doorsare open (air being diverted into the baggagecompar tment) , the cabin heat sys tem isinoperative. On SNs 35-643 and subsequent,and 36-064 and subsequent, if the electricaldiverter doors are open (air being divertedabove the headliner), the diverter doors willclose when the auxiliary cabin heat system isturned on.

Because of the high amperage required by theheating coils, they cannot be powered withonly airplane battery power. Either a groundpower unit or an engine-driven generator mustbe supplying power to operate the auxiliarycabin heat system.

The auxiliary cabin heat system will not au-tomatically shut down when a START–GENswitch is positioned to START. Therefore, itis recommended that the system be turned offduring engine start to avoid possible 275-ampcurrent limiter failure.

The Freon air-conditioning system has prior-ity over the auxiliary cabin heat system. If theFreon air-conditioning system is operating,

the auxiliary cabin heat is inoperative. If theauxiliary cabin heat system is operating, turn-ing on the Freon air-conditioning system willturn off the auxiliary cabin heat system.

Operation

On SNs 35-002 through 35-670, and 36-002through 36-063, the auxiliary cabin heatsystem is controlled by a three-position (LO–OFF–HI) AUX HT switch on the copilot'slower, right switch panel. Selecting the LOposition powers the cabin blower and one heat-ing coil on each element. The HI position pow-ers the cabin blower and all four coils.

On SNs 35-671 and subsequent and 36-064 andsubsequent, the cabin auxiliary heat system iscontrolled by a three-position (OFF–CREW–CAB & CREW) AUX HT swi tch on thecopilot's lower, right switch panel. The CREWposition of the switch energizes the crewauxiliary heater explained later in this section.Selecting the CAB & CREW position ener-gizes the cabin blower and all four auxiliarycabin heating coils.

Initially, the cabin blower will run at one-tenthof its normal speed until one of the thermo-switches senses a high limit. At that time, thecabin blower will come up to full speed andelectrical power to the heating coils will beremoved. The coils will cool until the ther-moswitch senses a low limit. Electrical powerwill then be reapplied to the heating coils andthey will continue to cycle, on and off, betweenthe high and low limits controlled by the ther-moswitch. The cabin blower will continue tooperate at full speed as long as the auxiliarycabin heat system is in operation.

DC electrical power to the heating coils isprovided by the same 150-amp current limiteron the battery charging bus used to power theFreon air-conditioning compressor motor.Control power for the auxiliary cabin heatingsystem is provided by the AUX CAB HT circuitbreaker on the left main DC bus.







Figure 11-15. Auxiliary Heating System Components

Auxiliary Cockpit Heat System(SNs 35-643 and Subsequent,and 36-064 and Subsequent)

General

The auxiliary cockpit heat system providesadditional heat for crew comfort and interiorwindshield defogging. It includes an electricheater installed in the forward end of the right-hand cabin bleed-air duct, where it connectsto the cockpit air distribution ducting and usescondition bleed airflow to circulate heated air(Figure 11-15).

Operation

The heating element for the auxiliary cockpitheat system requires bleed airflow through itfor cooling. Because of this, on SNs 35-671and subsequent, and 36-064 and subsequent,the CABIN AIR switch must be ON, at leastone engine must be running and its bleed airshut off and regulator valve must be openbefore electrical power can be applied to theheating element. If only the left engine isrunning, the left emergency pressurizationvalve must be in normal.

Despite these safeguards, on all airplanes,the crew should ensure the CABIN AIR switchis ON, at least one engine is running, andthere is adequate airflow in the right-handcabin bleed-air duct to cool the heating ele-ment before activating the auxiliary cockpitheating system.

On SNs 35-643 through 35-670, the auxiliarycockpit heating system is controlled by a three-position (OFF–CKPT–W/S AUX DEFOGHEAT) switch on the ANTI–ICE control panel.Selecting either CKPT or W/S AUX DEFOGHEAT will power the heater element. (SeeChapter 10, “Ice and Rain Protection,” foraddit ional information on the W/S AUXDEFOG HEAT function.)

On SNs 35-671 and subsequent, and 36-064and subsequent, the auxiliary cockpit heatingsystem is control led by a three-posi t ion(OFF–CREW–CAB & CREW) AUX HTswitch, located on the copilot’s lower, rightswitch panel. Selecting either CREW or CAB& CREW will power the heater element, aslong as the CABIN AIR switch is ON and theother conditions described above are met.

With the heater element powered, all the aircoming through the bleed-air outlets in thecockpit will be heated. A thermoswitch, locatedbetween the windshield defog diffusers andthe center footwarmer, monitors the tempera-ture of the airflow. The thermoswitch will cycleelectrical power to the heater element off andon between approximately 155° and 160° F. Inthe event of an overheat, a 295° thermoswitchin the heater should remove power to the ele-ment. Finally, a thermal fuse on the heater willmelt at approximately 415° F and remove powerto the element.

Power for the auxiliary cockpit heat elementis provided by two 20-amp current limitersfrom the battery charging bus. Control powerfor the auxiliary cockpit heat system is pro-vided by a circuit breaker on the left essentialA bus. On SNs 35-643 through 670, the cir-cuit breaker is labeled AUX DEFOG. On SNs35-671 and subsequent, and 36-064 and sub-sequent, it is labeled AUX CREW HT.




QUESTIONS




1. The manual diverter doors must be fullyclosed:A. To operate the cockpit fanB. To operate the Freon systemC. To opera te the auxi l iary heat ing

systemD. Airplane does not have manual di-

verter doors

2. Equipment which can be operated withairplane battery power only is:A. The auxiliary defog systemB. The Freon air-conditioning system C. The cabin blower and cockpit fanD. The auxiliary heating system

3. When the airplane is unpressurized onthe ground, air circulation is provided by:A. Ram airB. The cockpit fan and the cabin blowerC. Bleed-air systemD. Auxiliary defog system

4. The primary air conditioning in flight isprovided by:A. Engine bleed air B. The heat pumpC. The auxiliary heaterD. The Freon refrigeration system

5. When using the auxiliary cabin heater,the heated air blows out through:A. The conditioned air outletsB. The louvered grille above the divan

seatC. The overheat cockpit air outletsD. The overheat passenger WEMAC

outlets

6. The Freon system should not be usedabove:A. 5,000 feetB. 8,000 feetC. 18,000 feetD. 35,000 feet

7. The F reon sy s t em au toma t i ca l l y disengages:A. During engine startB. Upon touchdownC. When unpressurizedD. If the main door is opened

8. When the Freon system is operating, itcools:A. Ram airB. Cabin airC. Outside airD. Bleed air

9. When operating the Freon system on theground with engines running, the switchthat should be in OFF for maximum cool-ing effectiveness is the:A. GEN–STARTB. CABIN BLOWERC. CABIN AIRD. COCKPIT AIR

10. In order to operate the auxiliary cabinheater:A. Engines cannot be running.B. CABIN AIR switch must be off.C. Either a GPU or an engine-driven gen-

erator is required.D. Airplane must be on the ground.




11. If DC power fails, the flow control valve:A. Closes completelyB. Modulates from open to closedC. Remains openD. Is bypassed

12. The temperature control indicator shows:A. Cabin air temperatureB. Cockpit air temperatureC. The temperature of the bleed air in

the plenum chamberD. The position of the H-valve

12-i

CHAPTER 12PRESSURIZATION

CONTENTS

Page

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

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

MAJOR COMPONENTS ..................................................................................................... 12-3

Cabin Outflow Valve...................................................................................................... 12-3

Vacuum Jet Pump and Regulator Assembly .................................................................. 12-3

Pressurization Control Module ...................................................................................... 12-3

Cabin Safety Valve......................................................................................................... 12-6

CABIN AIR Switch ....................................................................................................... 12-6

Indicators ....................................................................................................................... 12-7

NORMAL SYSTEM OPERATION...................................................................................... 12-7

Before Takeoff ............................................................................................................... 12-7

Flight Operation—Automatic ........................................................................................ 12-7

Flight Operation—Manual............................................................................................. 12-8

Descent........................................................................................................................... 12-8

Landing .......................................................................................................................... 12-8

EMERGENCY PRESSURIZATION .................................................................................... 12-8

System Operation—SNs 35-002 through 35-112 and 36-002 through 36-031 ............. 12-8

System Operation—SNs 35-113 and Subsequent and 36-032 and Subsequent ............ 12-9

Emergency Pressurization Override Switches ............................................................. 12-11

QUESTIONS....................................................................................................................... 12-12




12-iii

ILLUSTRATIONS

Figure Title Page

12-1 Pressurization System Control ............................................................................... 12-2

12-2 Pressurization Control Module .............................................................................. 12-3

12-3 HORN SILENCE Switch and Test Control ........................................................... 12-5

12-4 CABIN ALT and DIFF PRESS Indicator .............................................................. 12-7

12-5 Emergency Pressurization Override Switches ..................................................... 12-11

TABLES

Table Title Page

12-1 Automatic Protection and Warning Features—SNs 35-002 through35-112 and 36-002 through 36-031 ..................................................................... 12-10

12-2 Automatic Protection and Warning Features—SNs 35-113 and Subsequentand 36-032 and Subsequent .................................................................................. 12-10




INTRODUCTIONThe Lear 35/36 airplane incorporates a pressurization system which maintains a speci-fied level of pressure consistent with built-in limits. Cabin pressure is regulated bycontrolling the exhaust of conditioned bleed air supplied by the engines. During normaloperation, the system functions automatically to provide crew and passenger comfortwithin the operational envelope of the airplane. Cabin pressure is controlled by an out-flow valve which is pneumatically operated to maintain a specified differential betweencabin and ambient pressures. Inward and outward relief for both negative and excess pos-itive differential conditions are incorporated to protect the airplane’s structure. A con-trol module provides a full range of manual control in the event of a malfunction of theautomatic controls. The purpose of the pressurization system is to ensure crew and pas-sengers comfort at all altitudes.

GENERALThe p r e s su r i za t i on con t ro l sy s t em i scompletely pneumatic during normal in-flightautomatic operation. Pneumatic pressure isprovided by a vacuum jet pump. Controlpressure (vacuum) is applied to the outflowvalve through the pressurization controlmodule. The pressurization controller provides

for both automatic and manual capabilities.Electrically actuated solenoid valves andswitches are incorporated for ground andmanual operation.

During climbs and descents the controllerregulates the outflow discharge rate. This ratecontrol is necessary to maintain a cabin changerate that is comfortable regardless of the air-planes rate of climb or descent.

CHAPTER 12PRESSURIZATION




12-2FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

FlightSafety

international

LE

AR

JE

T 3

0 S

erie

s

PIL

OT

TR

AIN

ING

MA

NU

AL

Figure 12-1. Pressurization System Control

ENGINE BLEED AIR

LEGEND

VACUUM CONTROL PRESSURE

SOL VALVE (NC)ENERGIZED OPEN ON GND

SOL VALVE (NO)ENERGIZED CLOSEDIN MANUAL, ON GROUNDOR ABOVE 8,750 ± 250 FTCABIN ALT LIGHT4

AUTO

MAN

AUTO

MAN

FILTER

NCR

UP

DN

CABIN

AIR

ON

OFF

UP

DN

RATE

MAXNORMOFF

CABINAIR

CURRENT AIRPLANES

EARLY AIRPLANES

PRESSURIZATION

RATE

DECR INCR

SOL VALVE (NO)ENERGIZED CLOSED

ON GND

CONTROL PRESSURE(VACUUM) SOURCE

SOLENOID VALVES DEPICTEDIN FLIGHT AUTO POSTIION

NOTE:


18.7 PSID28.9 PSID39.2 PSID410,000 FT (NO CABIN ALT LIGHT)5 AIRPLANES INCORPORATIONG AMK 78-5 ONLY611,000 FT ± 1,000 FT735-099/36-029 AND SUBSEQUENT

ENG BLEED AIR

SAFETYVALVE

CAB ALTLIMITER11,500 FT.± 1, 500 FT6

ORIFICE

STATICPORT

TAILCONE

STATICFILTER

REGJET PUMP

FILTER

CABIN

PRESS DIFFRELIEF 9.7 PSID3

PRESS DIFFRELIEF 9.4 PSI2

OUTFLOWVALVE

9.2 PSID1

STATICPRESS

CABINPRESS

STATIC PORT

UP

DN

ALTERNATESTATICPORT

NOSE CABIN

SOL VALVE (NC)ENERGIZED OPEN

ON GND WITHCAB AIR OFF7

CABINPRESS

STATIC PRESSURE

CABIN PRESSURE

MODIFIED CONTROL PRESSURE

CONTROLLERCABIN

CAB

INC

ABIN

32 1 0

ALT-FT

30 25

X1000

AIRCRA

FTA

LTX

10

00FT

CABIN ALT X 1000 FT.

SL

1

2

3 4 56

7

8

9

10

24

2624

CAB ALT LIM5

11,500 FT ± 1,500 FT6

The “Pneumatics” and “Air Conditioning”chapters describe how the cabin and cockpit arepressurized, heated, and cooled. This chapterdeals primarily with how the pressure is regulated.

MAJOR COMPONENTSThe pressurization control system (Figure 12-1)incorporates the following major components:

• Cabin outflow valve

• Vacuum jet pump and regulator assembly

• Pressurization control module

• Cabin safety valve

CABIN OUTFLOW VALVEThe pneumatically operated outflow valve islocated on the forward pressure bulkhead infront of the copilot’s position. Excess cabinair pressure is relieved into the unpressurizednose section through the outflow valve asnecessary to maintain the desired cabin pressure.

VACUUM JET PUMP ANDREGULATOR ASSEMBLYThe pneumatic pressure source for control of theoutflow valve is established by a vacuum jetpump and regulator assembly in the tailconesection of the airplane. Engine bleed air is routedthrough a venturi (jet pump) which generates anegative pressure, while a regulator ensuresthat the negative pressure maintains a constantdifferential pressure with respect to cabin pres-sure. This negative pressure (vacuum) is fur-nished to the pressurization control modulewhich uses it to control the outflow valve.

PRESSURIZATIONCONTROL MODULE

GeneralThe pressurization control module is located onthe copilot’s lower instrument panel. The con-trols on the front of the module are located on

what is referred to as the pressurization controlpanel. Figure 12-2 illustrates a typical airplanepressurization control module configuration.

AUTO-MAN SwitchPressurization control is normally accomplishedin the automatic mode. With the AUTO-MANswitch in the AUTO posi t ion , the cabincontroller automatically adjusts the pneumaticpressure sent to the outflow valve to regulatecabin pressure. If there is a malfunction in thecabin controller, the automatic pneumatic cir-cuit can be isolated from the outflow valve byselecting MAN. The outflow valve is thenmanually controlled with the UP–DN controlknob to regulate cabin pressure.

Cabin ControllerIn the AUTO mode of operation, the cabin con-troller regulates cabin pressure in relation to thealtitude that is set on the altitude selector knob.Rotating the knob on the face of the cabin con-troller either turns a dial or aligns a window toindicate two scales with a fixed index betweenthem. The outer scale represents cabin altitude,and the inner scale represents airplane altitude.




Figure 12-2. Pressurization Control Module

CURRENT

EARLY

For the current ECS, the cabin controller iscapable of maintaining the cabin pressure at sealevel with airplane altitudes up to approximately24,000 feet. If the airplane is flown to an alti-tude higher than 24,000 feet, the cabin altitudemust increase in order to maintain the samepressure differential. At an altitude of 45,000feet, the cabin altitude will normally be ap-proximately 6,700 feet.

Rate ControlA RATE knob is installed to the lower left ofthe CABIN CONTROLLER to control the rateat which the cabin climbs and descends. TheRATE control knob allows variable controlw i th in t he app rox ima te l im i t s o f 175feet/minute and 2,500 feet/minute. In AUTOmode, the CABIN CONTROLLER maintainsthe desired rate of climb or descent until theselected altitude is attained.

Manual Cabin AltitudeControl ValveThe UP–DN lever can be used to pneumaticallycontrol the outflow valve. Because of the redknob on the end of the lever, it is frequentlyreferred to as the “cherry picker.”

The UP–DN lever is spring-loaded to the centerposition and is wire guarded on later airplanesto prevent inadvertent activation.

The UP–DN lever can be used to increase ordecrease cabin altitude in either the AUTO orMAN mode. However, if it is used in the AUTOmode, the CABIN CONTROLLER will also at-tempt to control the outflow valve, and, assoon as the UP–DN lever is released to neu-tral, the cabin controller will return the cabinpressure to the original value.

Differential Pressure Relief Valve (Primary)The primary differential pressure relief valvefunctions in association with the CABINCONTROLLER. Its purpose is to relieveexcessive control pressure to the outflow valvein the event that cabin pressure should exceedthe normal limit when operating in AUTO mode.

NOTEThe primary differential pressurerelief valve does not function inMAN control.

On airplane SNs 35-002 through 35-112 and36-002 through 36-031, the relief valve is setfor 8.9 psid.

On airplane SNs 35-113 and subsequent and 36-032 and subsequent, the valve is set for 9.4 psid.

During a rapid airplane climb, with a lowsetting on the RATE knob, it is possible toreach the differential pressure relief settingprior to reaching the selected airplane altitude,a t which t ime the cabin c l imb ra te wi l lapproximate the airplane climb rate.

Cabin Altitude Limiter (For Outflow Valve)A cabin altitude limiter is installed on airplaneSNs 35-113 and subsequent, and 36-032 andsubsequent, and earlier SNs incorporatingAMK 78-5. It functions to limit the loss ofcab in p r e s su re due t o ma l func t i on ingcontroller or inadvertent operation of theprimary differential pressure relief valve.

If cabin altitude reaches 11,000 ± 1,000 feeton early airplanes, or 11,500 ± 1,500 feet onlater airplanes, the altitude limiter forcesmodulation of the outflow valve by introduc-ing cabin pressure into the control line, therebyrestricting cabin altitude to the listed level.

Controller Solenoid ValvesThree solenoid-operated valves are installedin the controller which are used to control therouting of pneumatic control pressure to theoutflow valve. All three valves are energizedon the ground by the squat switch relay box,causing the outflow valve to open, therebydepressurizing the cabin.

One of the valves is used in flight to effectmanual control of the outflow valve, and isreferred to as the “manual-mode solenoid valve”(see Flight Operation-Manual, this section).




For normal automatic in-flight operation, allthree solenoid valves are deenergized.

On the early SNs, valve actuation requires DCpower from the AIR BLEED circuit breaker onthe left essential bus. Later SNs require DCpower from the CAB PRESS circuit breakeron the right essential bus.

Aneroid SwitchesEither one or two aneroid switches are installedin the pressurization system depending onairplane serial number. Early airplanes use asingle aneroid switch for both warning horn andmanual solenoid operation. Later airplanes usetwo aneroid switches, one for the warning hornand another for manual solenoid operation.

Manual PressurizationAneroid Switch

The pressurization aneroid switch is locatedinside the pressurization module.

On airplane SNs 35-002 through 35-112 and36-002 through 36-031, if cabin alti tudeincreases to 10,000 feet or above, the aneroidswitch completes a power c i rcui t to thenormally open manual control solenoid valve.The solenoid valve is energized c losed,isolating all automatic pneumatic circuits fromthe outflow valve. The outflow valve, nowisolated holds its last attained position. Whencabin altitude decreases to 8,000 feet or below,the aneroid resets and deenergizes the solenoidvalve open, provided the AUTO MAN switchis in AUTO.

On airplane SNs 35-113 and subsequent and36-032 and subsequent, the description ofoperation is the same as early SNs, exceptthat the aneroid switch actuates at 8,750±250feet and resets at 7,200 feet and, when actuated,causes the amber CAB ALT annunciator toilluminate (Annunciator Panel section). Whenthe aneroid resets, the annunciator extinguishes.

Should the above cabin altitudes be reachedor exceeded, the “cherry picker” is the onlyway to control the outflow valve.

Cabin Altitude Warning HornAneroid Switch

Early SNs use the manual pressurizationaneroid just described, while later SNs use aseparate 10,100-foot cabin aneroid to sounda cabin altitude warning horn. A spring-loadedHORN SILENCE switch on the center switchpanel (Figure 12-3) may be used to silencethe horn. However, the horn will reactivateapproximately 30 seconds after being silencedwith the HORN SILENCE switch.

The horn will continue to reactivate after eachuse of the HORN SILENCE switch until theaneroid resets at a cabin altitude of 8,000 feet(early SNs) or 8,590 feet (later SNs).

The rotary system TEST switch on the centerswitch panel (Figure 12-3) is used to test thecabin altitude warning horn. Rotating theswitch to CAB ALT and depressing the TESTbutton provides a ground, simulating the al-titude warning horn aneroid switch actuation.This test does not illuminate the CAB ALTlight (if installed). During the test sequence,HORN SILENCE switch operations shouldalso be checked.




Figure 12-3. HORN SILENCEand Test Control

CABIN SAFETY VALVE

GeneralA pneumatically operated cabin safety valveis installed in the aft pressure bulkhead. Itspurpose is to relieve a cabin overpressure ora negative pressure differential caused by amalfunction in the normal control system. Inflight it normally remains fully closed unlessacted upon by the secondary differentialpressure relief valve, causing it to open (dueto an overpressure). In the case of a negativedifferent ial pressure condit ion, ambientpressure unseats the safety valve, allowing aninward flow to equalize the differential.

OperationOperation of the safety valve is automatic inflight; there is no crew control. On SNs 35-099and subsequent and 36-029 and subsequent, afourth solenoid valve is instal led in thepneumatic control circuit to allow control ofthe safety valve on the ground only (enginerunning and BLEED AIR switches at ON).The solenoid valve is energized open when theCABIN AIR switch is turned OFF (to open thesafety valve), and is deenergized closed 10seconds after the CABIN AIR switch is turnedto ON (to close the safety valve). The solenoidis deenergized in flight regardless of CABINAIR switch position.

On the earlier SNs, the safety valve does notopen on the ground.

Differential Pressure Relief Valve (Secondary)The secondary pressure relief valve functionsin association with the safety valve. Should theprimary pressure relief valve not functionproperly, the secondary pressure relief valveforces the safety valve open to limit cabinpressure. The safety valve will relieve pressureat 9.2 psid on SNs 35-002 through 35-112 and36-002 through 36-031. On SNs 35-113 andsubsequent and 36-032 and subsequent, thepressure is relieved at 9.7 psid.

Cabin Altitude Limiter(Secondary)The cabin altitude limiter for the cabin safetyvalve serves the same purpose as the cabin al-titude limiter for the outflow valve. If the sec-ondary differential pressure rel ief valvemalfunctions (causing the safety valve to open)and cabin altitude reaches 11,000 + 1,000 feeton early airplanes (11,500 + 1,500 feet on cur-rent airplanes), the cabin altitude limiter in-troduces cabin air pressure into the safetyvalve. This causes the valve to modulate andmaintain cabin altitude at the listed value.

CABIN AIR SWITCHThe CABIN AIR switch primarily controlsthe flow control valve as previously describedin Chapter 11 (“Air Conditioning”). Addi-tionally, the ON position (for current airplanes)provides electrical power for the cabin tem-perature sensor blower. Selecting the OFF po-sition on airplanes subsequent to SNs 35-098and 36-028, opens the safety valve if the air-plane is on the ground. Early airplanes havea MAX position which opens the valve to fullflow for smoke and fume elimination. Currentairplanes have no position equivalent to MAX;increased airflow is achieved by positioningthe BLEED AIR switches to EMER. TheCABIN AIR switch uses DC power from theAIR BLEED circuit breaker on the left es-sential bus.




INDICATORS

CABIN ALT and DIFFPRESS IndicationCabin altitude and differential pressure areindicated on a single indicator incorporatingtwo scales and two pointers (See Figure 12-4).

Cabin altitude is indicated by a single pointerand a circular scale on the outer edge withCABIN ALT markings from 0 to 50,000 feet.

The cabin differential pressure is indicatedby a circular scale on the inner portion of theindicator and a single pointer. The scalerepresents differential pressure from 0–10 psi,with a green band from 0–8.9 psi on earlyairplanes and 0–9.4 psi on current airplanes,a yellow band from 8.9–9.2 psi on earlyairplanes and 9.4–9.7 psi on current airplanes,and a red band above 9.2 psi on early airplanesand 9.7 psi on current airplanes. Cabin altitudeshould always be equal to or less than the air-plane altitude; therefore, cabin pressure should

always be equal to or greater than atmosphericpressure at the airplane altitude. The indica-tor should normally read approximately .2 psibelow the yellow arc.

Cabin Vertical Speed IndicatorThe cabin vertical speed indicator is positionedto the right of the cabin altimeter. It providesan indication of cabin climb or descent ratesof between 0 and 6,000 feet per minute.

NORMAL SYSTEM OPERATION

BEFORE TAKEOFFDuring ground operation, the CABIN AIRswitch is normally not turned to ON until justprior to takeoff unless engine bleed air isdesired for cabin heating.

When accomplishing the Before Starting En-gines checklist in the approved AFM, the crewwill normally (1) set the AUTO-MAN switchto AUTO, (2) set the expected cruise altitudeon the ACFT (inner) scale of the CABIN CON-TROLLER dial, and (3) set the RATE knob toapproximately the 9 o’clock position.

When the CABIN AIR switch is turned onprior to takeoff, the flow control valve isopened, allowing engine bleed air to enter thecabin. On SNs 35-099 and subsequent and36-029 and subsequent, there is a delay ofapproximately 10 seconds before the safetyvalve closes.

FLIGHT OPERATION—AUTOMATICAt l i f t o ff , t h e squa t sw i t ch r e l ay boxdeenergizes all pneumatic solenoids andpressurization begins. The cabin altitudebegins to climb at a rate based on the RATEknob se t t i ng . I t shou ld be ad jus t ed , a snecessary, to maintain a comfortable cabinaltitude climb rate of approximately 600 feet




Figure 12-4. CABIN ALT and DIFFPRESS Indicator

per minute. As the airplane climbs to cruisealtitude, the cabin controller automaticallyadjusts the outflow valve to give the desiredcabin climb rate until the cabin altitude reachesthe altitude set on the cabin controller dial. Asthe airplane continues its climb, the differentialpressure increases while the cabin altituderemains constant until the airplane arrives atthe selected ACFT altitude. If it is observedthat the DIFF PRESS indicator is riding on theyellow/red line, a slightly higher cabin altitudeshould be selected. Adjust the cabin controlleras necessary when changing cruise altitude.

Moni tor cabin pressure and di fferent ia lpressure throughout the flight.

FLIGHT OPERATION—MANUALIf the cabin controller is not functioningproperly, follow the Manual Mode Operationprocedures in Section 2 of the approved AFM.

Manual mode operation is established when theAUTO–MAN switch is placed to MAN. Thiscloses the manual mode solenoid valve,blocking the automatic pneumatic circuit. TheUP–DN lever (cherry picker) is then used tocontrol the outflow valve directly by using thestatic air source or existing cabin pressure tochange position of the outflow valve, causingthe cabin to climb or descend, respectively.

The manual control valve is very sensitive,and even small, momentary displacements ofthe lever will generate significant cabin climbor descent rates.

In manual mode, the cabin altitude must bemon i to r ed much more c lo se ly t han i nautomatic mode, and the outflow valve positionmust be adjusted frequently during climbs anddescents and when making power adjustments.

DESCENTDuring descent for landing, destination fieldelevation should be set on the CABIN scale ofthe CABIN CONTROLLER dial. The airplanerate of descent should be controlled so that thedescent rate is comfortable (approximately600 feet/minute).

LANDING

As the airplane descends and reaches thepreselected cabin altitude, the outflow valvemodulates toward the open position. The cabinshould be unpressur ized a t l and ing . Attouchdown, the squat switch relay box actuatesthe three pneumatic solenoid valves in thecontroller, causing the outflow valve to opencompletely to ensure cabin depressurization. Inaddition, when the CABIN AIR switch is placedto OFF, the flow control valve closes, and onSNs 35-099 and subsequent and 36-029 andsubsequent, an additional solenoid valve is en-ergized open, causing the safety valve to open.

EMERGENCYPRESSURIZATIONAn emergency source of pressurization bleedair is provided to increase the flow of air intothe cabin in the event of a leak.

SYSTEM OPERATION—SNS 35-002 THROUGH 35-112AND 36-002 THROUGH 36-031Emergency pressurization is provided by useof the windshield ant i - ice/defog system(Chapter 10). This is accomplished by push-ing the IN–NORMAL/OUT–DEFOG knob in,t h e n p o s i t i o n i n g t h e W S H L D H TAUTO–MAN switch to AUTO. This causes thedefog shutoff valve to fully open and also il-luminates the WSHLD HT light. These actionsintroduce air directly into the cabin areathrough the pilot’s foot warmer and bypasspossible leaks in the conditioned bleed-airdistribution system. To isolate such a leak, theCABIN AIR switch must then be selectedOFF to close the flow control valve (Chapter11, AIR CONDITIONING, Figures 11-1through 11-3).

On SNs 35A-082, 35A087 through 35A-112,and 36A-023 through 36A-031, and earlierairplanes incorporating AMK 76-7, the flowcontrol valve is located downstream of the




heat exchanger. Engine bleed air is availableto the heat exchanger whenever an engine isoperating and the BLEED AIR switches are on.Because of this, a pressure switch is installedin the tailcone ducting prior to the heat ex-changer. Should this pressure switch actuate(at approximately 47 psi), both red BLEEDAIL L and R annunciator lights illuminate toindicate the overpressure condition.

To deactivate emergency pressurization, selectMAN and toggle the spring-loaded WSHLDHT switch to OFF until the valve is closed.

SYSTEM OPERATION—SNS 35-113 AND SUBSEQUENTAND 36-032 AND SUBSEQUENTEmergency pressurization is accomplished byrouting bleed air directly into the cabin fromeither (or both) engine(s) through the emer-gency pressurization valves. This air com-pletely bypasses the entire manifold andconditioned bleed-air distribution system. SeeChapter 9, “Pneumatics.”

The valves are spring-loaded to the emergencyposition and require both servo bleed-airpressure and DC power to cause them to positionto normal. Cockpit control of the valves is pro-vided by the three-position (OFF–ON–EMER)BLEED AIR switches, while automatic posi-tioning occurs as a result of excessive cabin al-titudes or DC power failure.

With the BLEED AIR switches on, a solenoidon each emergency valve is energized, causingservo bleed-air pressure to move the valve tothe NORMAL position.

Positioning either BLEED AIR switch toEMER deenergizes the respective solenoid,causing the servo bleed-air pressure to beblocked; the valve repositions to emergencyby spring pressure. At the same time, HP airinput to the shutoff and regulator valve isblocked so that only LP air is allowed to enterthe cabin.

The emergency pressurization valves are alsocontrolled by two cabin aneroid switches (one

for each valve). The aneroids are set to operateat 9,500 feet ±250 feet cabin altitude. Shouldthe cabin altitude reach 9,500 feet ±250 feet,the aneroid switches deenergize the solenoidson the emergency pressurization valves and thevalves move to the emergency position. Theane ro id s r e se t when t he cab in a l t i t udedecreases to approximately 8,300 feet; how-ever, the approved AFM requires that the cabinaltitude be at, or below 7,200 feet beforeattempting to reset the emergency pressur-ization valves.

To reset the emergency pressurization valvesafter they have been positioned to emergency,the BLEED AIR switches, one at a time, mustbe positioned to OFF momentarily, then backto ON.

On airplane SNs 35-113 through 35-658, and36-032 through 06-063, not incorporatingAMK 90-3, the emergency pressurizationvalves are powered by the L and R MOD VALcircuit breakers on the left and right main DCbuses. These circuit breakers also provideelectrical power to the L and R bleed-airshutoff and regulator valves. With a MODVAL circuit breaker open, the emergencypressurization valve will position to emer-gency, the bleed-air shutoff and regulator valvewill fail open, and HP air to the shutoff andregulator valve will be blocked so only LP airwill be allowed to enter the cabin. In this case,positioning the BLEED AIR switch to OFF willnot stop airflow into the cabin since DC elec-trical power is required to close the bleed-airshutoff and regulator valve.

On airplane SNs 35-659 and subsequent, 36-064 and subsequent, and earlier airplanesmodified by AMK 90-3, the emergency pres-surization valves are powered by the L and REMER PRESS circuit breakers on the left andright main DC buses. On these airplanes, thebleed-air shutoff and regulator valves are pow-ered by separate circuit breakers labeled Land R BLEED AIR, also located on the left andright main DC buses. With an EMER PRESScircuit breaker open, the emergency pressur-ization valve will position to emergency andthe bleed-air shutoff and regulator valve willremain open. In this case, positioning the




BLEED AIR switch to OFF will stop airflowinto the cabin since DC electrical power, fromthe BLEED AIR circui t breaker, wil l beavailable to close the bleed air shutoff andregulator valve.

See Chapter 9, “Pneumatics,” for additional information on the bleed-air shutoff and regu-lator valves.

During the first engine start, the valves willautomatically shift position from emergencyto normal as HP servo air pressure from theengine becomes available. A slight rush of airinto the cabin is normal during start.

Tables 12-1 and 12-2 provide a description ofthe automatic protection and warning featuresfor cabin depressurization.

Table 12-1. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 35-002THROUGH 35-112 AND 36-002 THROUGH 36-031




Table 12-2. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 35-113AND SUBSEQUENT AND 36-032 AND SUBSEQUENT

CABIN ALTITUDE PROTECTION AND WARNING

10,000 ± 250 feet • Pressurization aneroid automatically switches the system to manual control. • Cabin altitude warning horn sounds—initiate emergency descent.

11,000 ± 1,000 feet • Cabin altitude limiters actuate.

14,000 ± 750 feet • Passenger oxygen masks are deployed and cabin overhead lights are illuminated.

* The differential pressure relief for the outflow valve is 8.9 psid, and the differential pressure relief for the safety valve is 9.2 psid.

CABIN ALTITUDE PROTECTION AND WARNING

8,700 ± 250 feet • Pressurization aneroid automatically switches the system to manual control. • CABIN ALT caution light illuminates.

9,500 ± 250 feet • Emergency pressurization valves are activated by aneroid switches, directing engine bleed air directly into the cabin.

10,100 ± 250 feet • Cabin altitude warning horn sounds—initiate emergency descent.

11,500 ± 1,500 feet • Cabin altitude limiters actuate.

14,000 ± 750 feet • Passenger oxygen masks are deployed and cabin overhead lights are illuminated.

* The differential pressure relief for the outflow valve is 9.4 psid, and the differential pressure relief for the safety valve is 9.7 psid.

EMERGENCYPRESSURIZATIONOVERRIDE SWITCHESOn SNs 35-605 and subsequent and 36-056 andsubsequent and earlier SNs incorporating AAK84-4, two emergency pressurization overrideswitches (Figure 12-5) allow the crew tooverride the 9,500-foot cabin aneroids tofacilitate landing at high-elevation airports.

The guarded switches are labeled “L” and “REMER PRESS” and have positions labeled“OVERRIDE” and “NORMAL.” With theguards down, the switches are in the NORMALposition. Lifting the guards and moving theswitches to the OVERRIDE position

disconnects the 9,500-foot aneroids from thesystem. The switches can also be used:

• To reset an emergency valve which hasinadvertently positioned to emergencydue to a malfunctioning aneroid

• To reset the emergency valves in orderto restore windshield and stab/wing anti-icing (at any altitude)

In either case, selecting OVERRIDE must befo l l owed by cyc l i ng t he BLEED AIRswitch(es) to OFF and then to ON, providedDC power is available and the MOD VAL (orEMER PRESS , a s app l i c ab l e ) c i r cu i tbreaker(s) are in.




Figure 12-5. Emergency Pressurization Override Switches

1. To regulate cabin pressure, the cabin con-troller modulates the:A. Cabin safety valveB. Flow control valveC. Outflow valveD. Primary differential pressure relief

valve

2. Illumination of the amber CABIN ALTlight (if installed) indicates:A. Cabin altitude is at or above 8,750

±250 feet, and the pressurizationcontrol system is in manual mode.

B. Cabin altitude is at or above 8,750±250 feet, and the pressurization control system may be in either AUTOor MAN mode.

C. Cabin altitude is at or above 9,500±250 feet, and the emergency pres-surization mode has activated.

D. The CABIN AIR switch is in the OFF position.

3. On airplanes with emergency pressuriza-tion valves, if DC power fails:A. Cabin pressurization must be con-

trolled manually with the UP–DNknob.

B. Cabin pressure will dump.C. The emergency pressurization valves

automatically actuate to provide emer-gency cabin pressure.

D. The flow control valve fails closed.

4. The cabin altitude warning horn soundswhen cabin altitude reaches approxi-mately:A. 8,750 feetB. 9,500 feetC. 10,100 feetD. 11,500 feet

5. To dump residual cabin pressure on touch-down:A. The outflow valve opens automati-

cally.B. The cabin safety valve opens auto-

matically.C. The flow control valve closes auto-

matically.D. The bleed-air shutoff and regulator

valves close automatically.

6. On airplanes without the emergency pres-surization valves, if DC power fails:A. The windshield anti-ice/defog system

can be used in the event of a pressur-ization failure.

B. The cabin will remain pressurized,but emergency pressurization capa-bility will be lost.

C. The flow control valve fails closed.D. The bleed-air shutoff and regulator

valves fail closed.

7. On all airplanes, if DC power fails:A. Pressurization control reverts to man-

ual control.B. The manual mode of pressurization

control cannot be selected or main-tained.

C. Cabin pressure is not controlled.D. The cabin slowly depressurizes.

QUESTIONS




13-i

CHAPTER 13HYDRAULIC POWER SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 13-1

GENERAL ............................................................................................................................ 13-1

HYDRAULIC SYSTEM OPERATION................................................................................ 13-3

HYDRAULIC SUBSYSTEMS............................................................................................. 13-4

QUESTIONS......................................................................................................................... 13-5




13-iii

ILLUSTRATIONS

Figure Title Page

13-1 Controls and Indicators .......................................................................................... 13-2

13-2 Hydraulic System Schematic ................................................................................. 13-3




INTRODUCTIONTwo engine-driven pumps normally provide hydraulic pressure for operation of the landinggear, wheel brake, flap, spoiler/spoileron, and Dee Howard TR-4000 thrust reverser (if in-stalled) subsystems. An electrically driven auxiliary pump incorporated for use in the eventof system failure is normally used only on the ground for operation of the brakes and flapswhen both engines are shut down. It cannot be used to operate the spoiler/spoileron system.

GENERALA 1.9 gallon reservoir pressurized by regu-lated engine bleed air ensures a positive sup-ply of MIL-H-5606 fluid to both engine-drivenpumps and to the auxiliary pump. The 4-gpm,variable-volume, engine-driven pumps are sup-plied from supply lines connected to the sideof the reservoir at approximately the .4-gallon

level. This limits the amount of fluid the engine-driven pumps can deliver to a system leak,reserving fluid for the auxiliary pump that isconnected to the bottom of the reservoir.

Hydraulic shutoff valves installed at the reser-voir in each engine-driven pump supply linecan be closed from the cockpit in the event offire or when maintenance is to be performed.

CHAPTER 13HYDRAULIC POWER SYSTEMS




An accumulator precharged with dry air ornitrogen dampens pressure surges and helpsmaintain system pressure. A direct-readingindicator on the center instrument paneldisplays system pressure. An amber annun-ciator light warns of low pressure.

There are three filters in the system—one ineach pressure line, and one in the return line.

A system relief valve set to relieve at 1,700 psiprevents system damage by relieving excessivepressure into the return line.

Pressure from the engine-driven pumps is avail-able to actuate the spoilers/spoilerons, flaps,

landing gear, brakes, and TR-4000 thrust reversers(if installed). A check valve prevents auxiliarypump actuation of the spoilers/spoilerons.

The reservoir and the accumulator are locatedin the tailcone. Reservoir fluid level should bejust above the sight glass with zero systempressure. Fluid is low if the level can be seenin the glass or if fluid is not visible.

Accumulator precharge, indicated by the gageon the accumulator, should be 850 psi whenhydraulic pressure is zero.

Controls and indicators for the system areshown-in Figure 13-1.




LOWHYD

(STANDARD)

Figure 13-1. Controls and Indicators

HYDRAULIC SYSTEMOPERATIONUnless there is residual hydraulic systempressure, the auxiliary hydraulic pump mustbe operated to provide pressure for setting theparking brakes prior to engine start. Placingthe HYD PUMP switch in the on (HYD PUMP)position starts the auxiliary pump, assumingboth engines a re shut down and sys tempressure is below 1,125 psi (Figure 13-2).

As pressure increases, a pressure switchactuates at 1,250 psi to extinguish the amberLOW HYD light on the annunciator panel.(All annunciators are shown in AnnunciatorPanel section.) At approximately 1,250 psi, thepressure switch stops the auxiliary pump. TheHYD PUMP switch should then be positionedto OFF, where it normally remains unless flapoperation is required prior to engine start. TheLOW HYD light will illuminate if pressuredrops below 1,125 psi.

If the HYD PUMP switch is lef t on, thepressure switch will cycle the pump between1,125 psi and 1,250 psi.




FILTER

ENGINEBLEED

AIR

REGULATOR

PRESSURE RELIEF VALVE (20 PSI)

OVERBOARD

VACUMRELIEFVALVE

GROUNDSERVICE

AUXILIARYPUMP

ENGINE-DRIVENPUMPS

CASE DRAIN LINE

LEGEND

PRESSURE

SUPPLY

RETURN

AIR/NITROGEN

REGULATED BLEED AIR

SPOILERS/ SPOILERONS

FLAPS GEARDOORS GEAR BREAKS

LO HYDPRESS

GROUNDSERVICE

AUXILIARYHYDRAULIC

PUMPSWITCH OFF

ON VDC50A

28

28 VDC

1,250 PSI

1,125 PSI

ACCUMULATOR(850-PSI AIR)

1,700-PSIRELIEF

THRUSTREV DEE HOWARD

(IF INSTALLED)

Figure 13-2. Hydraulic System Schematic

In the event of engine fire or when maintenanceis to be performed, the DC motor-driven shut-off valves can be closed by pulling the ap-propriate FIRE handle on the glareshield.Pulling either handle also arms the fire-ex-tinguisher system; therefore, if valve closingis to facilitate maintenance, the applicableFIRE EXT circuit breakers(s) should be pulledto prevent accidental discharge of the bottles.The shutoff valves are opened by pushing inthe appropriate handle(s). The shutoff valvesoperate on DC power supplied through the Land R FW SOV circuit breakers on the leftand right essential buses, respectively.

Af ter s ta r t ing the fi r s t engine , the HY-DRAULIC PRESSURE indicator should bechecked to verify engine-driven pump opera-tion. Pressure should stabilize at 1,550 ±25 psi,indicating that the engine-driven pump is op-erating properly.

When the second engine is started, there is nochange in pressure indication, but capacity isdoubled. There is no positive indication that thesecond pump is operating properly; therefore,after landing, operation of the second pump canbe verified by shutting down the engine startedfirst and actuating a hydraulic subsystem.

If an engine-driven pump fails in flight, theother engine-driven pump is capable of meet-ing system demands.

Loss of fluid due to a system leak is the mostprobable cause of complete loss of hydraulicpressure. If all hydraulic system pressure islost, the LOW HYD light will illuminate aspressure decreases below 1,125 psi. Do not op-erate the auxiliary pump until alternate land-ing gear extension procedures have beenaccomplished, as directed by the approvedAFM. Otherwise, the auxiliary pump may dis-charge the .4 gallon of reserve fluid throughthe same leak.

There is no circuit-breaker protection in thecockpit for the auxiliary pump; it is powereddirectly from the battery charging bus througha 50-ampere current limiter.

HYDRAULICSUBSYSTEMSOperation of the hydraulic subsystems is pre-sented in Chapter 14, “Landing Gear andBrakes,” Chapter 15, “Flight Controls” (flapsand spoi le r / spoi le rons) , and Chapter 7 ,“Powerplant” (Dee Howard TR-4000 thrustreversers).




QUESTIONS




1. Normal hydraulic system pressure withthe engine-driven pumps operating is:A. 1,400 ±50 psiB. 1,550 ±25 psiC. 1,650 psiD. 1,700 psi

2. The hydraulic shutoff valves are closed:A. By pulling the engine FIRE handlesB. Automatically when the FIRE warn-

ing light comes onC. By the GEN switch in the OFF posi-

tionD. By the BLEED AIR switches.

3. In the event of hydraulic system pressurefailure in flight:A. Immediately turn the HYD PUMP

switch on.B. Turn the HYD PUMP switch on when

the LOW HYD light illuminates.C. Refer to the Abnormal Landings

checklist.D. Re fe r t o t he Hydrau l i c Sys t em

Failure/Alternate Gear Extensionchecklist.

4. In the event of hydraulic system failure,the LOW HYD light will illuminate at:A. 1,125 psiB. 1,500 psiC. 1,250 psiD. 850 psi

5. During a hydraulic system failure, do notoperate the following subsystem usingthe auxiliary hydraulic pump:A. Landing gearB. SpoilersC. BrakesD. Flaps

6. The approved fluid for the hydraulicsystem is:A. MIL-H-5606B. MIL-O-M-332C. SkydrolD. MIL-H-2380

7. The operational time limit of the auxiliarypump is:A. 5 minutes on, 15 minutes offB. 5 minutes on, 25 minutes offC. 3 minutes on, 20 minutes offD. 2 minutes on, 30 minutes off

8. The auxiliary hydraulic pump will provideapproximately:A. 1,200 psiB. 1,550 psiC. 1,700 psiD. 1,250 psi

9. If DC electrical power is applied to theairplane and residual hydraulic pressureis 1,450 psi:A. The auxiliary hydraulic pump will not

operate when the HYD PUMP switchis on.

B. The LOW HYD light will be out.C. 1 ,450 p s i w i l l be shown on t he

HYDRAULIC PRESSURE indicator.D. All the above

14-i

CHAPTER 14LANDING GEAR AND BRAKES

CONTENTS

Page

INTRODUCTION ................................................................................................................. 14-1

GENERAL ............................................................................................................................ 14-1

LANDING GEAR................................................................................................................. 14-2

Indicating System .......................................................................................................... 14-2

General........................................................................................................................... 14-3

Operation ....................................................................................................................... 14-5

BRAKES ............................................................................................................................... 14-9

General........................................................................................................................... 14-9

Normal Operation .......................................................................................................... 14-9

Antiskid Operation....................................................................................................... 14-11

Emergency Brakes ....................................................................................................... 14-12

Parking Brakes............................................................................................................. 14-12

NOSEWHEEL STEERING ................................................................................................ 14-12

General......................................................................................................................... 14-12

Operation ..................................................................................................................... 14-14

QUESTIONS....................................................................................................................... 14-15




14-iii

ILLUSTRATIONS

Figure Title Page

14-1 Gear Position Indicator Lights ............................................................................... 14-2

14-2 Gear Position Indications ....................................................................................... 14-3

14-3 Main Gear............................................................................................................... 14-3

14-4 Nose Gear............................................................................................................... 14-4

14-5 Nose Gear Centering Cams.................................................................................... 14-5

14-6 Landing Gear Retracted ......................................................................................... 14-6

14-7 Landing Gear Extended.......................................................................................... 14-7

14-8 Emergency Air Pressure Indicator ......................................................................... 14-7

14-9 Alternate Extension Controls ................................................................................. 14-8

14-10 Alternate Landing Gear Extension......................................................................... 14-8

14-11 Brake System Schematic...................................................................................... 14-10

14-12 Nosewheel Steering System................................................................................. 14-13

14-13 Nosewheel Steering System Controls .................................................................. 14-14






INTRODUCTIONThe retractable landing gear is electrically controlled and hydraulically operated. The maingear incorporates dual wheels equipped with individual hydraulic brakes and retractsinboard. The single-wheel, self-centering nose gear incorporates an electrical steering systemand retracts forward. Alternate gear extension and emergency braking are pneumatic. Anantiskid system is incorporated into the normal hydraulic braking system.

CHAPTER 14LANDING GEAR AND BRAKES


GENERALThe landing gear has three air-hydraulic shockstruts. The main gear outboard doors aremechanically linked to the gear and move withit. The inboard doors are hydraulically oper-ated and close when the gear is fully extendedor retracted. An air bottle is provided for al-ternate gear extension and emergency braking.The gear actuators incorporate integral down-locking devices; downlock pins are not re-quired. Gear position indications are displayedon the copilot’s instrument panel.

The hydraulic brake system is controlled byfour valves (two for each pilot) linked to the

rudder pedals. Hydraulic system pressure ismetered to the self-adjusting multiple disc brakeassemblies in proportion to pedal deflection.

The antiskid system provides maximum de-celeration without skidding the tires. Whent h e s y s t e m i s o p e r a t i n g , w h e e l s p e e dtransducers (generators) furnish wheel speedinformation to a control box which signalsthe antiskid servo valves to modulate brakingpressure. The parking brake is set by pullinga h a n d l e o n t h e t h r o t t l e q u a d r a n t , a n ddepressing the brake pedals, trapping hy-draulic pressure in the brake assemblies.

The variable-authority, electric nosewheel steer-ing system operates only on the ground. Whenthe system is engaged, a computer determinesthe amount of nosewheel deflection allowable,based on rudder pedal movement and taxi speed,and uses a DC electric motor to deflect the nose-wheel accordingly. Maximum authority is 45ºeither side of center at slow speeds, decreasingas speed increases.

LANDING GEAR

INDICATING SYSTEM

GeneralThe landing gear position indicating systemconsists of three red lights and three greenlights, a test switch, and an aural warning horn.

Gear Position LightsThe three green LOCKED DN lights (Figure14-1) are il luminated by their respectivedownlock switches on the gear actuators.

As each gear locks down, the correspondinggreen LOCKED DN light illuminates. Duringgear retraction the lights extinguish when thedownlocks are hydraulically released.

The nose gear red UNSAFE light is illuminatedwhen the nose gear is in transit (neither downand locked nor up and locked). When the nosegear is locked in either the up or the downposition, the light extinguishes.

The two main gea r r ed UNSAFE l igh t silluminate whenever the respective main geardoor is unlocked. As each inboard door latchesup (dur ing extension or re t rac t ion) , thecorresponding red light extinguishes.

Indications for gear down-and-locked, up-and-locked, and in-transit conditions areshown in Figure 14-2.

If the gear is extended with the alternate(pneumatic) system, all three green lightsand the two main gear red lights will be il-luminated (both main gear doors will remainfully extended).

The position lights are tested by holding theTEST/MUTE switch on the landing gear controlpanel in the TEST position. All six lights will il-luminate and the warning horn will sound. Thelights can be dimmed with the dimming rheostat(Figure 14-2), provided the navigation lights areon; otherwise they will be at maximum intensity.

Circuitry related to the left and right maingear green position lights is common with thelanding/taxi light for that side. Confirmationof main gear downlocking (after bulb testing)can be made by switching on the respectiveLDG LTS switch.

Nose gear green light circuitry is common withthe engine synchronizing system (if installed).Confirmation of nose gear downlocking (after




Figure 14-1. Gear Position IndicatorLights

bulb testing) is made by positioning the ENGSYNC switch on the pedestal to ENG SYNC(on) and observing that the amber ENG SYNClight on the annunciator panel illuminates.

Landing Gear Warning SystemThe aural warning horn will sound and threered UNSAFE lights will illuminate when allof the following conditions are present:

• Landing gear is not down and locked.

• Altitude is less than 14,500 ±500 feet.

• Either thrust lever is retarded belowapproximately 55-60% N1.

• Airspeed is below 170 KIAS (FC 530 air-planes only)

At altitudes above 14,500 ±500 feet, the hornwill not sound when the thrust levers arer e t a rded , and t he UNSAFE l i gh t s mayilluminate. The horn also sounds when theflaps are extended beyond 25º if the landinggear is not down and locked, regardless ofthrust lever position or altitude.

Holding the TEST/MUTE switch in the TESTposition illuminates all six position indicatorlights and sounds the horn. Momentarilypositioning the switch to MUTE silences thehorn when the thrust levers are retarded, andthe gear is not down and locked.

The horn cannot be muted when the gear is notdown and locked and the flaps are extendedbeyond 25º.

GENERAL

Main Gear ComponentsEach main gear consists of a conventional air-hydraulic shock strut, dual wheels, scissors,squat switch, main gear actuator, inboard andoutboard doors, and an inboard door actuator(Figure 14-3).




UP AND LOCKED

IN TRANSIT

DOWN AND LOCKED

DIMMINGRHEOSTAT

UNSAFE LOCKED DN

TEST

MUTE

UP

DNLANDING GEAR

BRT

UNSAFE LOCKED DN

TEST

MUTE

UP

DNLANDING GEA

BRT

UNSAFE LOCKED DN

TEST

MUTE

UP

DNLANDING GEAR

BRT

Figure 14-2. Gear Position Indications

Figure 14-3. Main Gear

The main gear hydraulic actuator also serves asa side brace when the gear is extended. It fea-tures an integral downlock mechanism that canbe unlocked only by hydraulic pressure on theretract side; therefore, no downlock pins areprovided. Each main gear scissors link actuatesa squat switch.

The main gear is hydraulically held in theretracted position and is enclosed by an outboardand an inboard door. The outboard door ismechanically linked to, and travels with, thegear. The inboard door is hydraulically actuated,electrically sequenced by microswitches, andheld retracted by hydraulic pressure and aspring-loaded, overcenter uplatch that is releasedby a hydraulic actuator.

Proper shock strut inflation is an important con-sideration. When the airplane weight is on thegear, the amount of strut extension will vary withthe airplane load. With a full fuel load and nopassengers or baggage aboard, 3 to 31⁄2 inchesof bright surface should be visible on the lowerportion of the main gear strut.

Main Gear Wheel and Tires

Each main gear wheel incorporates a fusibleplug that prevents tire blowout caused by ex-cessive heat resulting from hard braking. Tiresmust be changed when the tread has worn to thebase of any groove at any location or if the cordis exposed. Main gear tire pressure is determinedby airplane gross weight certification.

Nose Gear ComponentsThe nose gear consists of an air-hydraulic shockstrut incorporating a self-centering device, anosewheel steering actuator, and mechanicallyoperated doors (Figure 14-4).

The nose gear strut is conventional, with twoexceptions: it does not have a scissors and thenosewheel steering actuator motor is mountedon top of the strut housing.

The nose gear actuator incorporates an integraldownlock mechanism to maintain a positivedownlocked condition; therefore, a downlockpin is not required. As with the main gear

actuator, the locking mechanism can bereleased only by hydraulic pressure on theretract side. The gear is held retracted byhydraulic pressure and a spring-loaded up-latch hook that engages the uplatch roller onthe forward side of the strut. The uplatch hookis released by a hydraulic actuator.

When retracted, the nose gear is enclosed bytwo doors that are linked to, and travel with,the gear.

An improperly centered nosewheel could jam inthe wheel well; therefore, the nose strut incor-porates a self-centering mechanism. At liftoff,two cams within the strut are engaged by strutair pressure to center the wheel (Figure 14-5).

Since nosewheel centering depends on air pres-sure in the strut, proper inflation of the strut isespecially important. When the airplane weightis on the gear, the amount of strut extension willvary with airplane load. With a full fuel loadand no passengers or baggage aboard, 51⁄4 to 53⁄4inches of bright surface should be visible on thelower portion of the nose gear strut.

Because the cams cannot center the wheel ifit is swiveled 180º from the normal position,the nose gear should be checked on the exteriorinspection to ascertain that the gear uplatchroller (Figure 14-4) is facing forward.




Figure 14-4. Nose Gear

Nose Gear Wheel and Tire

The nosewheel tire is chined to deflect wateror slush spray (up to 3⁄4-inch deep) away fromthe engine intakes during takeoff or landing.

Nosewheel tire pressure should be maintainedat from 104 to 114 psi when the airplane isloaded and the crew is in the cockpit.

OPERATIONThe landing gear system incorporates twosolenoid-operated hydraulic control valves—one for operation of the main gear inboarddoors and one for gear operation. Both in-board doors must be fully open before the gearcan be extended or retracted.

The door control valve is energized to thedoor–open position when the landing gear se-lector switch is placed in either the UP or theDN position. It is energized to the door–closeposition by main gear-operated switches whenboth gear are fully retracted or down and locked.

The gear control valve is energized to the extendor the retract position by switches sensing thefull open position of both main gear inboarddoors. During retraction, the circuit is routedthrough both squat switches to ensure that theairplane is off the ground before the valve canbe energized to the retract position.

Normal landing gear operation requires DCpower supplied through the gear circuit breakeron the right essential bus.

Normal RetractionPositioning the landing gear selector switch toUP energizes the door control valve to the openposition, directing pressure to release the maingear inboard door uplatches and to open thedoors. The two red main gear UNSAFE lightsilluminate simultaneously with uplatch release.

When both inboard doors are fully open, thedoor–open switches are actuated. When bothdoor–open switches are actuated and bothsquat switches are in the airborne position, thegear control valve energizes to the retract po-sition, and hydraulic pressure is directed to re-tract the landing gear (Figure 14-6). The threegreen LOCKED DN lights extinguish, and thered nose gear UNSAFE light illuminates.

When the nose gear has fully retracted, the rednose gear UNSAFE light extinguishes. Whenboth main gear are fully retracted, two gear-up trunnion switches are actuated to energizethe door control valve to the closed position.Pressure closes the gear inboard doors, whichlock in position by spring tension on the dooruplatches, and the two red main gear UNSAFElights extinguish.




Figure 14-5. Nose Gear Centering Cams

Normal ExtensionPositioning the landing gear selector switch toDN energizes the door control valve to the openposition, directing pressure to release the maingear inboard door uplatches and to open thedoors. The two red main gear UNSAFE lightsilluminate simultaneously with uplatch release.

When both inboard doors are fully open, thedoor–open switches are actuated to energizethe gear control valve to the down position.This directs pressure to release the nose gearuplatch and extend the nose and main gear(Figure 14-7). The red nose gear UNSAFElight illuminates.

When the gear is fully down and locked, thethree green LOCKED DN lights illuminate andthe red nose gear UNSAFE light extinguishes.

Circuitry is completed by both main geardownlock switches to energize the door con-trol valve to the closed position. Pressurecloses the gear inboard doors (Figure 14-7),which lock in position by spring tension on thedoor uplatches, and the two red main gearUNSAFE lights extinguish.

Alternate Extension

General

The a l t e rna te gea r ex tens ion sys t em i spneumatically operated by a bottle charged to1,800–3,000 psi with dry air or nitrogen. Bottlepressure is shown on the di rect - readingEMERGENCY AIR indicator on the center in-strument panel (Figure 14-8). The bottle alsoprovides pressure for emergency braking.




SOL

RETRACTEXTENDSOL

SOL SOL

EMERAIR

BOTTLE

DOORACTUATOR

DOORCONTROLVALVE

LEGENDSYSTEM HYDRAULIC PRESSURE

RETURN

AIR PRESSURE

TO BRAKE

SYSTEM

TO

EMERGENCY

BRAKES

OVERBOARD

GEAR ALTERNATE EXTENSION CONTROL VALVE

PRIORITYVALVE

UPLATCH

UPLATCHACTUATOR

NOSEGEARACTUATOR

GEARCONTROLVALVE

MAIN GEARACTUATOR

UPLATCHACTUATOR

UPLATCH

GEAR INBOARD DOOR

UPLATCH

DOORACTUATOR

UPLATCHACTUATOR

MAIN GEARACTUATOR

Figure 14-6. Landing Gear Retracted




PRIORITY

MAINGEAR

GEARCONTROL

UPLATCHACTUATOR

OVERBOARD


UPLATCH

VALVE

ACTUATOR

VALVE

SOL

RETRACTEXTENDSOL

SOL SOL

TOEMER

BRAKES

EMERAIR

BOTTLE

UPLATCHACTUATOR

NOSEGEARACTUATOR

MAINGEARACTUATOR

UPLATCHACTUATOR

DOORACTUATOR

DOORACTUATOR

UPLATCH

DOORCONTROLVALVE

UPLATCH

GEAR INBOARD DOORLEGEND

TOBRAKE

SYSTEM

SYSTEM HYDRAULIC PRESSURE

RETURN

AIR PRESSURE

Figure 14-7. Landing Gear Extended

Figure 14-8. Emergency Air Pressure Indicator

Prior to using the system, the landing gear selec-tor switch (Figure 14-2) should be placed inthe DN position and the GEAR circuit breakeron the right essential bus should be pulled. Thiswill prevent inadvertent gear retraction sub-sequent to a successful extension. The systemis activated by pushing down the emergencygear lever on the right side of the pedestal(Figure 14-9). The lever has a ratchet to keepit in the down position, once activated, and canbe raised only by manually actuating the re-lease tab while simultaneously lifting theemergency gear lever.

Operation

Pushing the emergency gear lever down opens avalve to release air bottle pressure to position thegear control and door control valves to the ex-tend position (Figure 14-10). This provides areturn flow path for fluid in the retract side ofthe gear and door actuators. The air pressure




OVERBOARD


UPLATCH

GEAR DOOR

SOL

RETRACTEXTEND

SOL

SOL SOL

AIR PRESSURE

RETURN

LEGEND

TOEMERGENCYBRAKE VALVE

TOBRAKE

SYSTEM

EMERAIR

BOTTLE

PRIORITYVALVE

UPLATCHACTUATOR

NOSEGEARACTUATOR

UPLATCHACTUATOR

UPLATCH

MAINGEARACTUATOR

UPLATCH

UPLATCHACTUATORDOOR

ACTUATOR

DOORACTUATOR

MAINGEAR

ACTUATOR

DOORCONTROLVALVE

GEARCONTROLVALVE

Figure 14-10. Alternate Landing Gear Extension

EMERGENCYGEARLEVER

RELEASETAB

Figure 14-9. Alternate Extension Controls

also repositions the shuttle valves to accom-plish the following:

• Release the nose gear uplatch and themain gear door uplatches.

• Open the main gear inboard doors.

• Extend all three gear.

Since no provision is made to close the maininboard doors, the two main gear red UNSAFElights will remain illuminated. The three greenLOCKED DN lights will illuminate.

In a hydraulic failure situation, after the gearis down and locked, air pressure must be bledfrom the gear system by lifting the release tab(Figure 14-9) and raising the emergency gearlever to the normal position. This closes thevalve on the emergency air bottle and isolatesair pressure from the gear system, preventinga poss ible leak in the gear sys tem fromdepleting air pressure that might be requiredfor emergency braking.

If alternate extension is required due to anelectrical fault, the emergency gear lever mustremain in the down pos i t ion to preventsubsequent inadvertent retraction of the gear.

BRAKES

GENERALThe brake system is powered by hydraulicsystem pressure from the nose gear down(extend) line. The brakes can be applied byeither pilot. The system has four multidisc,self-adjusting brake assemblies, one for eachmain gear wheel, operated by power brake

valves linked to the top section of the rudderpedals. The left pedals control both brakeassemblies on the left gear; the right pedalscontrol the brake assemblies on the right gear.Braking force is in direct proportion to pedalapplication unless modulated by the antiskidsystem. The antiskid system, monitored by thered ANTISKID GEN warning lights, permitsstopping in the shortest possible distance for anygiven runway condi t ion . (Warning andannunciator lights are shown in AnnunciatorPanel section.) Parking brakes can be set bypulling a handle on the center pedestal. Apneumatic emergency brake system is used tostop the airplane if hydraulic pressure is lost.Neither antiskid protection nor differentialbraking is available during emergency braking.

NORMAL OPERATIONWhen either pilot depresses a brake pedal, theassociated brake valve meters system hydraulicpressure through shuttle valves (one in eachmain pressure line), parking brake valves,antiskid valves, brake fuses, and a second setof shuttle valves, one for each of the four brakeassemblies (Figure 14-11). The first set of shut-tle valves determines whether the pilot or thecopilot has control of the brakes (highestpressure predominating).

P is tons in each brake assembly move apressure plate, forcing the stationary androtating discs together against a backing plateto produce the braking action. Depressing onepedal applies both brakes on the correspondingmain gear; therefore, differential braking isavailable, if required.

Releasing pedal pressure repositions the brakevalve, and springs in the brake assembly forcefluid back through the brake valves to thereservoir, thereby releasing the brakes.







OVERBOARD

SERVO

MECHANICAL

COPILOTBRAKE VALVE

PILOTBRAKE VALVE

BRAKE FUSE

PARKINGVALVES

GEARALTERNATEEXTENSIONCONTROL VALVE

BRAKEAIR BOTTLE

EMERGBRAKEVALVE

TORESERVOIR

FROM NOSEGEAR DOWNLINE

PILOTBRAKE VALVE

COPILOTBRAKE VALVE

ANTISKIDDISCONNECTSWITCH

WARNLIGHTCB

SOLENOIDSHUTOFF

ON

OFF

ANTISKID

ANTI-SKID GEN*PARKING BRAKE LIGHT SNs 35-626 627 , 630, AND SUBSEQUENT, 36-056 AND 059 AND SUBSEQUENT

LEGENDEMERGENCY BRAKEAIR PRESSURE

ELECTRICAL

SQUAT SWITCH

ANTISKIDCONTROL BOX

SERVO

PARKBRAKE

ANTISKIDVALVE

TORESERVOIR

SQUAT SWITCH

SOLENOIDSHUTOFF

SERVO

ANTISKIDVALVE

SERVO

SYSTEMPRESSUREMETERED BRAKEPRESSURE

RETURN

Figure 14-11. Brake System Schematic




During gear retraction, a restrictor in the nosegear return line creates back pressure on thebrakes which is sufficient to stop the wheels fromrotating prior to their entering the wheel well.

A priority valve, also in the nose gear downline,ensures proper gear sequencing during retractionby restricting hydraulic pressure applied to thenose gear actuator while full system pressure isbeing applied to the main gear actuators.

When taxiing through slush or snow, frequentbrake application creates friction heat whichmay prevent the brakes from freezing.

If a takeoff is made in slush or snow, the wheelsshould be allowed to spin down for approxi-mately 1 minute prior to gear retraction. Thisminimizes the possibility of brake freezing byslinging off accumulated slush. If frozen brakesare suspected after the gear is extended forlanding, the ANTISKID switch should bepositioned to OFF, and the brakes applied 6 to10 times to break up any possible ice forma-tions. The ANTISKID switch should be turnedback to ON prior to landing.

ANTISKID OPERATION

GeneralOne of two antiskid systems may be installed.The early system was standard on airplane SNs35-002 through 35-066 and 36-002 through36-017. The later system is standard on airplaneSNs 35-067 and subsequent and 36-018 andsubsequent, and may be retrofitted to earlyairplanes by AAK 76-4. The two systems aresimilar and are discussed together with thedifferences being noted.

The antiskid system limits braking on eachmain gear wheel independently to allow max-imum braking under all runway conditionswithout tire skidding.

The system consists of four wheel speed trans-ducers (one on each main wheel), two antiskidcontrol valves, a control box, monitor lights,and a lever-locking ANTISKID switch on thecenter instrument panel.

Airplanes with the early antiskid system havetest provisions on the system’s rotary test switch.On these airplanes, the system is tested duringthe Before Taxi check in accordance with theapproved AFM. The ANTISKID switch shouldbe positioned to OFF after testing unless the airplane incorporates AAK 75-1 or AMK 76-3, in which case it can be left on. On airplaneswith the later system, no testing is required andthe switch is normally left in the ON position.

The antiskid system is not required to beoperational for flight. However, if a mal-function is indicated by illumination of a redANTI-SKID GEN light(s), it must be assumedthat antiskid protection is lost on the associ-ated wheel. Takeoff and landing data must becomputed accordingly.

The system uses DC power from the ANTISKIDcircuit breaker on the right main DC bus.

OperationThe fol lowing condit ions must exist foroperation of the antiskid system:

• The ANTISKID switch must be on.

• Both squat switches must be in theground mode (left for outboard, rightfor inboard).

• The parking brake must be released.

• Taxi speed must be above 8 to 10 knots(wheel speed, 150 rpm).

At high speed, with the ANTISKID switch onand brakes applied, the control box receives andanalyzes wheel speed inputs from the transduceron each main wheel (Figure 14-11). If anywheel deceleration rate reaches a predeter-mined limit, the applicable servo valve willmodulate braking force on the correspondingbrake by diverting pressure into the return line.

With the gear extended in flight, the braking system is disabled. When the main gear squatswitches go airborne, all braking pressure isdiverted into the return line (as though all wheels




were in a full-skid condition). This precludesthe possibility of touching down on the nextlanding with brakes applied inadvertently.Further, at the moment of touchdown, the squatswitches initiate a requirement for a 150-rpmwheel spinup or a 1- to 2-second delay, thusenabling the control box to sense realisticwheel speeds before normal braking can begin.

If the brakes are to be applied in flight to breakup suspected accumulations of ice on thebrakes, the ANTISKID switch must first bepositioned to OFF. Position the switch to ONprior to landing.

At low taxi speeds (wheel speed below 150 rpm,8–10 knots), the antiskid system is inoperative.The system is automatically disconnected whenthe parking brakes are set; however, the redANTI-SKID GEN lights will not illuminate.

Four red ANTI-SKID GEN lights monitor cir-cuitry from each wheel speed transducer andwil l individual ly i l luminate i f a faul t isdetected. Cycling the ANTISKID switch toOFF then back to ON may clear the fault. Allfour lights illuminate if power to the controlbox is lost or if the ANTISKID switch is off.

EMERGENCY BRAKESPneumatic emergency brakes are provided foruse in the event of normal brake system fail-ure. Antiskid protection, differential braking,and parking brakes are not available whileusing the emergency brakes.

To apply brakes with the emergency system,the EMER BRAKE handle must be pulled outof its recess (Figure 14-11) and pressed down-ward. This meters pressure from the emer-gency air bottle through four shuttle valves tothe brake assemblies in proportion to handlemovement. Releasing the handle stops flowfrom the bottle and allows applied air pressureto be vented overboard, releasing the brakes.

PARKING BRAKESNormal hydraulic system pressure from eitherengine-driven pump or the auxiliary pump canbe used to set the parking brakes. Pulling thePARKING BRAKE handle on the centerpedestal mechanically closes both parkingbrake valves (Figure 14-11). The closed valvesfunction as one-way check valves, allowingpressure from the pilot or copilot brake valvesto be trapped in the brake assemblies.

To set the parking brakes, pedal pressure mustbe applied and the PARKING BRAKE handlepulled out (but not necessarily in that order).Setting the parking brake opens the antiskid dis-connect switch (Figure 14-11) to disconnect theantiskid system and prevent inadvertent loss ofbrake pressure.

To release the parking brakes, the PARKINGBRAKE handle must be pushed in all the wayto the stop. If the PARKING BRAKE handleis not pushed in to the stop, the parking brakesmay be released but the antiskid disconnectswitch may not actuate to enable the antiskidsystem. The ANTI-SKID GEN lights will notilluminate, and subsequent heavy braking willresult in wheel skids.

Airplanes SNs 35-626, 627, 630 and subse-quent, 36-056, and 059 and subsequent havean additional PARK BRAKE light just abovethe ANTI-SKID GEN l ights . The PARKBRAKE light will illuminate if the parkingbrake handle is not in the completely forward(released) position.

NOSEWHEEL STEERING

GENERALThe electrical actuated nosewheel steering sys-tem has variable authority, as determined by sig-nals from the left inboard and both right wheel




28 VDC

AC

AC

115 VAC

WHEEL SPEEDTRANSDUCERS

STEER LOCKBUTTON

CONTROL WHEELMASTER SWITCHES

STEER ON

NOSE GEARUPLOCK SWITCH

(RELEASED)

RUDDER PEDALFOLLOW-UP

SQUAT SWITCHRELAY BOX

REVERSIBLEMOTOR

CLUTCH

NOSEWHEEL STRUTFOLLOW-UP

NOSEWHEEL STRUT

*LEFT MAIN GEARDOWNLOCK SWITCH

*SNs 35-134 AND SUBSEQUENT,36-036 AND SUBSEQUENT. NOSEWHEEL DOWNLOCK SWITCHEARLIER AIRPLANES.

LEFT INBOARDRIGHT INBOARD

RIGHT OUTBOARD

NOSEWHEEL STEERINGCOMPUTER

Figure 14-12. Nosewheel Steering System




Figure 14-13. Nosewheel Steering System Controls

speed transducers. System components also in-clude a rudder pedal follow-up, a computer-am-plifier, and a DC steering actuator motor (Figure14-12). AC and DC power is supplied throughthe NOSE STEER circuit breakers on the left ACand left main buses, respectively.

The steering actuator, mounted on top of thenose strut, steers the nosewheel through agearbox and an electrical clutch. When theairplane is on the ground, the clutch engageswhenever DC power is applied to the elec-trical system allowing the steering actuator tofunction as a shimmy damper, even with steer-ing disengaged. If DC power is lost or the DCNOSE STEER circuit breaker is out, the nose-wheel is free to swivel, and the shimmy damperis inoperative.

Prior to towing, electrical power should beremoved from the airplane. It is possible tomisalign the nosewheel more that 90º fromnormal during towing; therefore, the nose gearuplock roller on the nose gear strut must bepointing forward prior to flight.

Steering authority varies from a maximum of45º either side of center at speeds below 10knots and decreases as ground speed increases.At the maximum steering speed of 45 knots, au-thority has been reduced to approximately 8º.

OPERATIONWith the squat switches in the ground mode,nosewhee l s t e e r ing can be engaged bymomentarily depressing the STEER LOCKswitch or by depressing and holding the con-trol wheel master switch (MSW) on either con-trol wheel (Figure 14-13). STEER LOCK isdisengaged by momentarily depressing ei-ther control wheel master switch.

When steering engages, the green STEER ONannunciator illuminates. A rudder pedal fol-l owup p rov ide s t he d i sp l acemen t anddirectional signals modified by the computer-amp l i f i e r i npu t f r om the whee l speedtransducers. The computer-amplifier drivesthe s teer ing ac tuator in the appropr ia tedirection until it is stopped by a signal froma follow-up located in the drive gearbox.

If the nosewheel steering system is inopera-tive, differential power and braking can beused to taxi the airplane.

Since variable-authority steering is depen-dent upon wheel speed transducer signals,steering should not be used above 10 knots ifany two of the following three ANTI-SKIDGEN lights are illuminated: two inboard andright outboard.

CENTERPEDESTAL

CONTROL WHEELS

l Emergency air pressure can be used for:A. Gear extension and parking brakeB. Gear, flaps, spoilers, and brakesC. Gear extension and brakesD. Gear extension, flaps, and brakes

2. Prior to takeoff, the EMERGENCY AIRpressure indicators should indicate:A. 1,800 to 3,000 psiB. Minimum 1,700 psiC. 3,000 to 3,350 psiD. Maximum 1,750 psi

3. During normal gear operation, main gearinboard doors and the main gear are se-quenced by:A. MicroswitchesB. Emergency air pressureC. Mechanical linkageD. Both A and B

4. Automatic brake snubbing is providedduring gear retraction by restricting re-turn fluid from the:A. Antiskid systemB. Engine-driven pumpsC. Squat switchesD. Landing gear system

5. After an emergency gear extension, thegear position light indication should be:A. Three greenB. Three green, two redC. Three red, two greenD. Three red, three green

6. Three gear UNSAFE lights will be on andthe gear warning horn sounding when the:A. Gea r i s r e t r a c t ed and no g r een

LOCKED DN lights are on.B. Gear is down, thrust levers are above

approximately 70% N1, and altitudeis below 14,500 ±500 feet.

C. Gear is up, thrust levers are below ap-proximately 55%–60% N1, altitudeis below 14,500 ±500 feet and, on FC530 airplanes, airspeed is below 170KIAS.

D. Flaps are extended below 25º, re-gardless of altitude.

7. With the flaps extended beyond 25º andthe gear not down and locked, the warn-ing horn:A. Will sound but can be mutedB. Will not soundC. Will sound but cannot be mutedD. None of the above

8. Illumination of a red main gear UNSAFElight indicates:A. The corresponding main gear is not

down and locked.B. The corresponding main gear is not up

and locked.C. The corresponding main gear inboard

door is not fully closed.D. The corresponding main gear inboard

door is locked in the closed position.

9. The red nose gear UNSAFE light will beon when:A. The nose gear is unsafe or in transit.B. Nosewheel steering is inoperative.C. The nose gear doors are open.D. The nose gear doors are closed.




QUESTIONS

10. Parking brakes can be set with the:A. Pilot’s brake pedals onlyB. Copilot’s brake pedals only when the

ANTISKID switch is onC. Pilot’s or copilot’s brake pedalsD. Pilot’s or copilot’s brake pedals only

with the ANTISKID switch off

11. If the first three ANTI-SKID GEN lightsare illuminated:A. Takeoff weight is limited to 17,000

pounds.B. Nosewheel steering should not be en-

gaged above 10 kts.C. Takeoff (VR) will be affected.D. Both A and B are correct.

12. Normal brake pressure is provided by:A. Main hydraulic system pressure from

the nose gear down lineB. Brake accumulatorC. Emergency air bottle through the an-

tiskid control valvesD. Emergency air bottle

13. Related to nosewheel steering, the pre-cautions that should be taken prior totowing the airplane are:A. Keep rudder pedals centered.B. Do not exceed the 55º turning limits.C. Pull the NOSE STEER DC circuit

breaker if the battery switches are on.D. Turn off the ANTISKID switch.

14. If the green main gear LOCKED DN lightis burned out, positive down-and-lockedcondition can be confirmed by:A. GND IDLE light illuminatedB. ENG SYNC light illuminatedC. Illumination of the corresponding land-

ing light when the switch is turned onD. Red UNSAFE lights illuminate.

15. The electrical requirements for nosewheelsteering are:A. 24 VAC and 28 VDCB. Only 28 VDCC. Only 115 VACD. 28 VDC and 115 VAC

16. When STEER LOCK is engaged:A. Nosewheel steering is engaged and

full steering is available up to 45 kts.B. The nosewheel is locked in whatever

position it is in at the time.C. Up to 45º left or right steering is avail-

able, with decreasing authority athigher speeds.

D. Nosewheel becomes free swiveling.

17. STEER LOCK is disengaged by:A. Depressing the OFF buttonB. Depressing the STEER LOCK button

a second timeC. Momentarily depressing either wheel

master switchD. Depressing the ANTISKID release

button




15-i

CHAPTER 15FLIGHT CONTROLS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 15-1

GENERAL ............................................................................................................................ 15-1

PRIMARY FLIGHT CONTROLS........................................................................................ 15-2

Elevators......................................................................................................................... 15-2

Ailerons.......................................................................................................................... 15-3

Rudder............................................................................................................................ 15-4

TRIM SYSTEMS .................................................................................................................. 15-4

General .......................................................................................................................... 15-4

Rudder (Yaw) Trim ........................................................................................................ 15-4

Aileron Trim .................................................................................................................. 15-4

Pitch Trim ...................................................................................................................... 15-6

Mach Trim ................................................................................................................... 15-10

SECONDARY FLIGHT CONTROLS................................................................................ 15-12

Flaps ............................................................................................................................ 15-12

Spoilers ........................................................................................................................ 15-14

YAW DAMPERS ................................................................................................................ 15-18

General ........................................................................................................................ 15-18

Yaw Damper Control Panel ......................................................................................... 15-18

Operation (Airplanes with FC 200 AFCS) ................................................................... 15-19

Operation (Airplanes with FC 530 AFCS) ................................................................... 15-19




STALL WARNING SYSTEMS.......................................................................................... 15-20

General......................................................................................................................... 15-20

Operation ..................................................................................................................... 15-22

MACH OVERSPEED WARNING/STICK PULLER......................................................... 15-23

General ........................................................................................................................ 15-23

Operation ..................................................................................................................... 15-23

QUESTIONS....................................................................................................................... 15-24




15-iii

ILLUSTRATIONS

Figure Title Page

15-1 Flight Control Surfaces .......................................................................................... 15-2

15-2 Flight Controls Gust Lock...................................................................................... 15-2

15-3 Aileron Tabs ........................................................................................................... 15-4

15-4 Trim Systems Controls and Indicators ................................................................... 15-5

15-5 Pitch Trim System Schematic (Airplanes with FC 200 AFCS) ............................. 15-7

15-6 Pitch Trim System Schematic (Airplanes with FC 530 AFCS) ............................. 15-8

15-7 Mach Trim System Schematic ............................................................................. 15-10

15-8 Flap System.......................................................................................................... 15-13

15-9 Spoiler System ..................................................................................................... 15-15

15-9A Spoiler Operation ................................................................................................. 15-16

15-10 Spoileron Operation (Left Aileron Up)................................................................ 15-17

15-11 Yaw Damper Systems .......................................................................................... 15-19

15-12 Stall Warning System........................................................................................... 15-21




INTRODUCTIONThe manually operated primary flight controls incorporate electrical trim in all three axes.Secondary flight controls consist of hydraulically actuated spoilers/spoilerons and flaps.Other systems related to flight controls are the yaw damper, stall warning, Mach over-speed warning, and Mach trim systems.

GENERALThe primary flight controls (ailerons, elevator,and rudder) are mechanically operated throughthe dual control columns, control wheels, andrudder pedals. They are incorporated into boththe FC 200 and the FC 530 automatic flightcontrol system (AFCS). Both systems alsoincorporate a rudder/aileron interconnect.

The ailerons incorporate mechanical balancetabs to provide aerodynamic assistance. Trimsystems (roll, yaw, and pitch) are electricallyoperated and controlled. Trim tabs are installedon the left aileron and the rudder. The mov-able horizontal stabilizer provides pitch trim.

The flaps and spoilers are hydraulically actu-ated and electrically controlled.

Aileron augmentation is provided by a spoil-eron system which increases roll authoritywhen the airplane is configured for landing.

A dual yaw damper system provides yawstability.

A dual stall warning system provides an in-dication of impending stall by vibrating thecontrol column and, if no corrective action istaken, induces a forward control column move-ment to reduce the airplane angle of attack.

20

20 20

105

510 10

5

5

LO

C

GS

CHAPTER 15FLIGHT CONTROLS




A Mach overspeed warning system warns ofoverspeed and induces an aft control columnmovement to raise the nose of the airplane.

A Mach trim system provides automatic pitchtrim to compensate for Mach tuck.

All flight control surfaces are shown in Fig-ure 15-1.

A flight controls gust lock is provided to pre-vent wind gust damage to the primary flightcontrol surfaces. When installed, the lockholds full left rudder, full left aileron, and fulldown elevator displacement (Figure 15-2).

PRIMARY FLIGHTCONTROLS

ELEVATORSThe elevators are hinged to the aft edge of thehorizontal stabilizer and are positioned by

fore-and-aft movement of the control column.Three scuppers are located near the aft edgeof each elevator for moisture drainage, andthree static dischargers are attached to thetrailing edge of each elevator.




Figure 15-1. Flight Control Surfaces

Figure 15-2. Flight Controls Gust Lock

The elevators can also be positioned by anelectrically actuated pitch servo.

A bob weight attached to the control columnand a downspring assembly in the elevatorcontrol linkage are incorporated to enhancepitch stability.

Pitch ServoThe pitch servo (torquer) is DC operated. It ismechanically connected to the elevator con-trol linkage through a capstan mechanism in-corporating an electric clutch and a mechanicalslip clutch. Three flight control systems usethe pitch servo to operate the elevators:

• Autopilot—When engaged, the autopi-lot can alter noseup or nosedown attitudeby commanding the servo to torque theelevator up or down, as required.

• Both stall warning systems—Either sys-tem will cause the servo to torque the el-evator nose down in the event of animpending stall (stick pusher). On FC530 models, pulsating nose down torquesignals are used for the “nudger.”

• Mach overspeed warning system—Op-erating through the L STALL WARN-ING switch, the system will commandthe servo to torque the elevator nose up(stick puller) due to an overspeed.

On FC 200 AFCS airplanes, the electricclutch must be engaged to couple the servoto the elevator linkage. The clutch engageswhen any one of the following switches is inthe ON position:

• L STALL WARNING

• R STALL WARNING

• AUTOPILOT master

With all three of the above switches in theOFF position, the electric clutch is disengaged,disconnecting the servo from the elevators.This enables the pilot to gain manual controlof the elevator by eliminating the servo in theevent of a malfunction.

By exerting sufficient force on the controlcolumn to slip the mechanical clutch, the pilotcan also override any undesirable servo in-puts to the elevators, if necessary.

On FC 530 AFCS airplanes, the electric clutchremains deenergized until the servo is signalledby either the autopilot, L or R stall warningsystem, or the overspeed puller system.

On these airplanes, the servo can be eliminatedas a cause of malfunction by simply depressingand holding the wheel master switch. The pilotcan also, by exerting the required force on thecontrol column to slip the mechanical clutch,override any undesirable servo operation.

Autopilot operation is described in Chapter 16,“Avionics.”

AILERONSThe ailerons, mechanically positioned with eithercontrol wheel, provide primary roll control.Aileron effectiveness is augmented by spoileronswhen the airplane is configured for landing.

Spoileron (aileron augmentation) operationis automatically activated when the flaps arelowered beyond 25˚. In the spoileron mode,when an aileron is moved up to initiate airplaneroll, the spoiler on the same wing automaticallyrises the same number of degrees to provideadditional roll.

Roll Servo (Autopilot Function Only)The ailerons can also be positioned by the au-topilot roll servo. The roll servo is similar tothe pitch servo but does not incorporate anelectric clutch. A mechanical slip clutch allowsthe pilot to override undesired roll servo inputs;the servo can also be disconnected by disen-gaging the autopilot.

Balance TabThe balance tab on each aileron (Figure 15-3)provides aerodynamic assistance in moving theaileron, thus reducing control wheel forces.




Trim TabThe electrically operated aileron trim tab is at-tached to the inboard trailing edge of the leftaileron (Figure 15-3). The tab is positioned byeither the pilot’s or copilot’s control wheeltrim switch. Aileron trim tab posit ion isindicated on the cockpit center pedestal.

Aileron Follow-upsAileron follow-up mechanisms, driven by theaileron control linkage, provide aileron dis-placement information to the spoileron com-puter, yaw damper, and the autopilot.

RUDDERThe rudder can be manually positioned witheither set of rudder pedals, or by either of twoyaw damper servos (primary and secondary).The crew can manually override the yaw damperthrough a mechanical slip clutch in the event ofa malfunction. The yaw damper can be disen-gaged by depressing either wheel master switchor the corresponding yaw damper OFF button.

Rudder Trim TabA trim tab, mounted on the bottom trailingedge of the rudder, is controlled by a trimswitch on the center pedestal. Trim positionis also indicated on the center pedestal.

TRIM SYSTEMS

GENERALThe ailerons and rudder are trimmed with con-ventional tabs on the control surfaces aspreviously described.

The airplane is trimmed in the pitch axis bychanging the angle of incidence of the movablehorizontal stabilizer. A dual-motor (primaryand secondary) actuator moves the leadingedge of the horizontal stabilizer up or downin response to pitch trim inputs. Controls andindicators for the trim systems are shown inFigure 15-4.

The trim position indicators for pitch, roll,and yaw are all DC powered through the TAB& FLAP POSN circuit breaker on the rightessential bus.

RUDDER (YAW) TRIM

ControlRudder (yaw) trim is controlled by the ruddertrim switch on the center pedestal, (Figure15-4) spring-loaded to the OFF position.

The switch knob is split into an upper and alower half . Both halves must be rota tedsimultaneously to initiate rudder trim tabmotion. This is a safety feature to reduce thepossibility of inadvertent trim actuation. Therudder trim system is DC powered through theYAW circuit breaker on the left essential bus.

Rudder Trim IndicatorRudder trim tab position indication is providedby the RUDDER TRIM indicator (Figure 15-4).

AILERON TRIM

ControlAileron (roll) trim is controlled with eithercontrol wheel trim switch located on theoutboard horn of each control wheel (Figure15-4). Each control wheel trim switch is a




AILERONWING

BALANCETAB

TRIM TAB

Figure 15-3. Aileron Tabs

dual-function (trim and trim arming) switchwhich controls roll and primary pitch trim.Each switch has four positions—LWD, RWD,NOSE UP, and NOSE DN, and is spring-loaded

to the neutral position. The arming button ontop of the switch must be depressed and heldwhile simultaneously moving the trim switchin the direction of desired trim action. Actu-ation of either control wheel trim switch to




PILOT'S CONTROL WHEEL(COPILOT'S SIMILAR)

OR

PITCH TRIMSELECTOR

SWITCH

WHEELMASTERSWITCH

ARMINGSWITCH

CONTROL WHEELTRIM SWITCH

SECONDARY PITCHTRIM SWITCH

Figure 15-4. Trim Systems Controls and Indicators

LWD or RWD (with arming button depressed)will signal the trim tab actuator motor in the leftaileron to move the trim tab in the appropriatedirection. Actuation of the pilot’s trim switchwill override actuation of the copilot’s switch.

The aileron trim motor is DC powered throughthe ROLL circuit breaker on the left essential bus.

Aileron Trim IndicatorAileron trim tab position indication is providedby the AIL TRIM indicator (Figure 15-4).

PITCH TRIM

GeneralPitch trim is accomplished by repositioning thehorizontal stabilizer to the desired trim settingwith a dual-motor (primary and secondary)actuator that operates in four modes:

1. Primary pitchtrim mode } Primary trim motor

2. Mach trim mode

3. Secondary pitchtrim mode } Secondary trim

4. Autopilot pitchmotor

trim mode

The pilot-operated primary pitch trim and sec-ondary pitch trim systems are electrically in-dependent systems. Mode selection (primaryor secondary) is made with the PITCH TRIMselector switch (Figure 15-4).

Primary pitch trim is pilot-controlled througheither of the control wheel trim switches;secondary pitch trim is controlled through thesecondary pitch trim toggle switch on the centerpedestal (Figure 15-4).

Airplanes with the FC 530 automatic flightcontrol system (AFCS) incorporate a two-speedprimary trim motor, a trim monitor system, andan audible clicker that signals trim in motion.

Mach trim automatically engages at approxi-mately 0.69 M1 if the autopilot is not engaged,and uses the primary trim motor to adjust pitch

trim. Autopilot operation uses the secondarymotor to adjust pitch trim.

NOTEThe PITCH TRIM selector switchmust be in the PRI position to enablethe Mach trim system. It may be ineither the PRI or SEC position duringautopilot operation.

Horizontal stabilizer position is displayed onthe PITCH TRIM indicator (Figure 15-4).

Pitch Trim ActuatorThe pitch trim actuator is operated by eitherof two DC-powered motors, either of whichcan move the horizontal stabilizer. On FC-200 AFCS airplanes, the primary trim motorand control circuits are powered through thePITCH circuit breaker on the left essentialbus. On FC-530 AFCS airplanes, the motor ispowered by the battery charging bus, and thePITCH circuit breaker on the left essentialbus controls a relay in the power circuit. Thesecondary trim motor and control circuits arepowered through the SEC PITCH TRIM (orSEC P TRIM) circuit breakers on the rightessential bus.

On FC 200 AFCS airplanes, the secondarytrim motor operates at approximately one-halfthe speed of the primary trim motor.

On airplanes with the FC 530 AFCS, the two-speed pr imary t r im motor opera tes a t aconsiderably slower rate (approximately one-fourth speed) with the flaps up. A 3˚ flap switchis used for speed switching. On these airplanes,operating speed of the secondary tr im isapproximately the same as the speed of theprimary trim with flaps up.

PITCH TRIM Selector SwitchThe PITCH TRIM selector switch providesthe primary and secondary mode selections(Figure 15-4). In the PRI (forward) position,primary pitch trim is available from both of thecontrol wheel trim switches and from the Mach







CONTROL WHEELTRIM SWITCH

PITCHTRIMPRI

NOSEDN

NOSEUPSEC

PRI

SEC

WHEEL MASTER SWITCH(MSW)

AUTOPILOTPITCH

COMPUTER AUTOPILOTPUSHERPULLER

PITCH SERVO

T.O.TRIM

ELECTRICAL

LEGEND

OFF

ANNUNCIATOR

Figure 15-5. Pitch Trim System Schematic (Airplanes with FC 200 AFCS)




Figure 15-6. Pitch Trim System Schematic (Airplanes with FC 530 AFCS)

CONTROL WHEEL TRIM SWITCH

FAST

SLOW

ANNUNCIATOR

PITCHTRIM

PRIMARY TRIM

PITCHTRIMPRI

NOSEDN

NOSEUPSEC

TRIMMONITOR

SECONDARY TRIM

PRI

SEC

WHEEL MASTER SWITCH(MSW)

3 FLAP SWITCH

AUTOPILOT TRIM

AURALTRIM INMOTION

AUTOPILOTPITCH

COMPUTER AUTOPILOTPUSHERPULLERNUDGER

PITCH SERVO

T.O.Trim

Electrical

LEGEND

o

OFF

ANNUNCIATOR

trim system. In the OFF position, both trimmotors and control circuits are deenergized. Inthe SEC (aft) position, secondary pitch trim isavailable from the secondary trim switch (Fig-ure 15-4); this renders the pilot’s primary trimand Mach trim inoperative. The secondary pitchtrim switch is spring-loaded to the OFF position.

The autopilot always uses the secondary trimmotor whether the PITCH TRIM selectorswitch is in the PRI or SEC position; however,if either control wheel trim switch is actuatedwith the arming button depressed (Figures 15-5 and 15-6) or if the secondary trim switch isactuated, the autopilot will disengage.

In the event of inadvertent primary pitch trimoperation (runaway trim), depressing and hold-ing the wheel master switch will:

• Stop only the primary pitch trim motor(airplanes with FC 200 AFCS)

• Stop both the primary and the sec-ondary trim motors (airplanes with FC530 AFCS)

The control wheel trim switches (Figure 15-4) were descr ibed in th is chapter underAileron Trim.

Pitch Trim IndicatorHorizontal stabilizer trim position indicationis provided by one of two types of PITCHTRIM indicators (Figure 15-4). On eachindicator, a T.O. (takeoff) trim segment ismarked to indicate the takeoff trim limits forcenter-of-gravity extremes. On early airplanes,the segment is marked by a green band on theedge of the indicator; on later airplanes, bywhite lines. Late model indicators may beretrofitted on early airplanes. In either case,whenever the pitch trim is not set within theT.O. trim segment, the amber T O TRIM annun-ciator light will illuminate (on the ground only).(All annunciator lights are shown in Annuncia-tor Panel section.)

Pitch Trim Monitor System(Airplanes with FC 530 AFCS)

General

A monitor system incorporated in theseairplanes provides a visual indication of cer-tain faults in the primary trim system.

Though not physically a part of the monitorsystem, a clicker provides audible evidence oftrim in motion (primary or secondary trimsystem) when the flaps are up.

Operation

The monitor system monitors the primary trimsystem, the 3˚ flap switch, and the horizontalstabilizer actuator mechanism. Faults are in-dicated by illumination of the amber PITCHTRIM light.

With flaps up (slow trim required), the mon-itor system will illuminate the PITCH TRIMlight if it senses that primary trim is runningat the fast rate (trim overspeed).

Regardless of flap position, the monitor systemwill also illuminate the PITCH TRIM light if itsenses certain electrical faults in the primary sys-tem that create the potential for uncommandedmotion of the stabilizer actuator.

When the PITCH TRIM light illuminates, thesecondary trim system must be selected byplacing the PITCH TRIM selector switch in theSEC position (unless it illuminates whileholding the wheel master switch depressed,which is normal).

The audio clicker will sound anytime the sta-bilizer actuator is in motion with flaps up,whether trimming is being accomplished withthe primary or secondary motor. However, topreclude the clicker from sounding every timetrim is commanded, a delay of approximately1/4 second must follow each in-motion signal,thereby eliminating nuisance signals whenthe pilot uses short trim inputs.




The monitor system and trim-in-motion clickerare tested in accordance with procedures out-lined in Section 2 of the approved AFM. Ei-ther a three-position switch decaled “TRIMOVSP–OFF–TRIM MON” (spring-loaded toOFF) or the TRIM OVSP and TRIM MON po-sitions of the rotary systems test switch areused to perform the test.

MACH TRIM

GeneralThe Mach trim system is an automatic pitchtrim system that uses the primary trim motorto enhance longitudinal stability during ac-ce le ra t ions /dece le ra t ions a t h igh Machnumbers to compensate for Mach tuck. Thereis no switch to engage the system; it auto-matically becomes active at approximately0.69 MI if the autopilot is not engaged.

Since the Mach trim system requires the useof the primary pitch trim motor, the PITCHTRIM selector switch must be in the PRIposition for system operation.

If the autopilot is engaged, the Mach trim sys-tem assumes a passive (standby) mode. In thiscase, the PITCH TRIM selector switch can bein either the PRI or SEC position, since theautopilot can utilize the secondary trim motorin both switch positions.

The Mach trim system consists of a computer,an air data sensor, a follow-up on the horizontalstabilizer, and a red MACH TRIM annunciatorlight, Mach overspeed warning horn and amonitor circuit. The system is powered by115 VAC supplied by the MACH TRIM circuitbreaker on the left AC bus, and DC power sup-plied by the PITCH circuit breaker on the leftessential bus.




PRI

SEC

MACH TRIMMON

ELECTRICAL

LEGEND

PITCHTRIMPSI

SEC

OVERSPEED WARNINGHORN

MACH TRIMCOMP

AIR DATA

SENSOR

STATIC

PITOT

MACHTRIM

MECHANICAL

MACH TRIMFOLLOW-UP

Figure 15-7. Mach Trim System Schematic

OperationDuring flight, the air data sensor receivesstatic pressure inputs from the left and rightshoulder static pressure ports (FC 200 AFCS)and a pitot pressure input from the right pitottube (Figure 15-7). On FC 530 AFCS air-planes, static pressure is provided by the RHstatic 1 and LH static 2 lines. This will beshown in Chapter 16, “Avionics.”

The air data sensor electrically transmits thisinformation to the Mach trim computer. Withthe autopilot disengaged, the Mach trim sys-tem becomes active at approximately 0.69 MI.The Mach trim computer will command the ap-propriate pitch trim changes (noseup trim forincreasing Mach, nosedown for decreasingMach) through the primary motor of the pitchtrim actuator. The follow-up on the horizon-tal stabilizer provides the nulling signal tothe computer.

Mach trim is interrupted whenever the air-plane is manually trimmed. The system resyn-chronizes to function about the new horizontalstabilizer position when manual trim is re-leased. In flight, synchronization may also beaccomplished by selecting the MACH TRIMposition on the SYS TEST switch and de-pressing the TEST button (applies to airplaneSNs 35-247 and subsequent, 36-045 and sub-sequent, and earlier airplanes incorporating SB35/36 22-4).

Mach Trim MonitorThe Mach trim monitor circuit continuouslymonitors input signals and power to the Machtrim computer, and compares signal inputsfrom the air data sensor (Mach) and the Machtrim follow-up on the horizontal stabilizer. Amalfunction exists if the Mach trim monitordoes not receive a corresponding signal changefrom the Mach trim follow-up when the airdata sensor signals change (Mach change). Amalfunction is also indicated in the event ofpower loss to the Mach trim computer, loss ofinput signals, or a Mach number/horizontalstabilizer trim position error. In either case, theMach trim monitor will disengage Mach trimand illuminate the MACH TRIM light. If speedis above 0.74 MI, the Mach overspeed warninghorn will also sound. If the fault clears orpower is restored, the system can be resyn-chronized by selecting the MACH TRIM po-sition on the SYS TEST switch and depressingthe TEST button (applies to airplane SNs 35-247 and subsequent, 36-045 and subsequent,and earlier airplanes incorporating SB 35/36-22-4). If the warning horn continues to sound,airspeed must be reduced below 0.74 MI orthe autopilot (if operational) may be engaged.Engaging the autopilot will cancel all warnings,and the airplane can be accelerated to MMO.




SECONDARYFLIGHT CONTROLS

FLAPS

GeneralThe single-slotted Fowler flaps are electri-cally controlled and hydraulically actuated.The left and right flaps are interconnected bycable to minimize asymmetrical effects in theevent of a malfunction.

Position switches mechanically connected toeach flap provide flap position information tothe landing gear warning, stall warning, spoilerwarning, spoileron, and autopilot systems. Onairplanes SNs 35-067 and subsequent, SNs36-018 and subsequent, and earlier airplanesincorporating AAK 76-4, the flap positionswitches actuate at 3°, 13°, and 25° of flapextension. On earlier airplanes the switches ac-tuate only at 13° and 25°. On airplanes with thepreselect flap system, flap limit switches au-tomatically maintain flap position at the se-lected setting.

If hydraulic system pressure is lost, the flapswill probably remain in their last position.However, if the flaps are extended and hydrau-lic pressure is lost due to a leak in the flapdownline, airloads on the flaps may causesome flap retraction.

The flaps can also be operated from EMERBAT 1 (ON position) in the event of electri-cal failure; however, the flap indicator is notpowered by the emergency battery.

Flap Selector SwitchThe flap selector switch may be one of threetypes. On SNs 35-002 through 35-010, theswitch has three positions (up, neutral, anddown), and is spring-loaded to the neutralposition. The selector switch on later airplanesis not spring-loaded to neutral and will remainin the selected position. Airplane SNs 35-417,419, 477, 479, and 483 and subsequent, andSNs 36-051 and subsequent incorporate the

preselect flap system. On these airplanes theflap selector switch has four positions: UP,8°, 20°, and DN (40°), with detents at the 8°,and 20° positions (Figure 15-8). The flapsystem is powered by the FLAPS circuitbreaker on the right essential bus. EarlierSNs may be retrofitted with the preselect sys-tem by AAK 83-7.

Flap Position IndicatorA vertical-scale FLAP position indicator ismounted on the center switch panel (Figure15-8).

Left flap position is electrically transmitted tothe indicator. The indicator is DC powered bythe TAB FLAP POSN circuit breaker on theright essential bus. The indicator will indicateDN with loss of electrical power, regardless ofactual flap position.

Operation (Preselect Flaps)When the flap selector switch is placed in theDN position, the down solenoid positions theflap control valve to direct pressure to the ex-tend side of both flap actuators. The downsolenoid remains energized, and the controlvalve maintains down pressure on the flapactuators to hold the flaps full down (40°). Acheck valve at the control valve inlet preventsf lap re t ract ion in the event of upst reamhydraulic system failure.

Moving the selector switch to an intermediate(8° or 20°) position energizes the down or upsolenoid, as appropriate, which repositions thecontrol valve to extend or retract the flaps. Theappropriate flap limit switch deenergizes theaffected solenoid and the control valve closes,thereby stopping flap motion (9° and 21° duringextension, 19° and 7° during retraction).

When extended, the flaps are protected fromexcessive airloads (due to excessive airspeed)by a relief valve in the downline. Pressure isrelieved into the return line, causing the flapsto creep upward. The limit switches will en-ergize the down solenoid to return the flaps tothe selected position when the airspeed is re-duced appropriately.







RETRACTEXTEND

RELIEF VALVE(FLAP BLOWUP)

LEGEND

NORMAL HYDRAULICSYSTEM PRESSURE

RETURN

STATIC

MECHANICAL

ELECTRICAL

FLAP CONTROLVALVE

POSITIONTRANSMITTER

SWITCHES(PRESELECT)

FLAPLIMIT19 /21

7 /9

FLAP POSITIONSWITCHES

FLAP POSITIONSWITCH

INTERCONNECTCABLE

FLAPACTUATOR

*3 , 13 , 25 ON AIRPLANES SNs 35-067 AND SUBSEQUENT, 36-018AND SUBSEQUENT, AND EARLIERAIRPLANES INCORPORATING AAK76-4. 13 AND 25 ON PREVIOUSAIRPLANES

PRESELECT NONPRESELECT

*

*

Figure 15-8. Flap System

When the selector switch is moved from theDN position toward the UP position, an inter-mediate stop is encountered at the 20˚ positionto facilitate retraction in a go-around situation.Further movement of the selector switch to-ward UP or 8˚ requires that the switch lever bepulled out to clear the stop.

When the flap selector switch is placed in theUP position, the up solenoid positions the flapcontrol valve to direct pressure to the retracts ide of both f lap ac tuators . In the ful lyretracted position, the up solenoid remainsenergized and the control valve maintainspositive pressure on the retract side of bothflap actuators.

Operation (Nonpreselect Flaps)When the flap selector switch is placed in theDN position, the down solenoid positions theflap control valve to direct pressure to theextend side of both flap actuators. The flapsmay be stopped in any intermediate positionby placing the selector switch in the centerneutral position. This deenergizes the downsolenoid which repositions the control valveto the neutral position, trapping fluid betweenthe control valve and the actuators to hold theflaps in the selected position.

When extended, the flaps are protected fromexcessive airloads (due to excessive airspeed)by a relief valve in the downline, and theflaps will creep up until the airspeed is re-duced appropriately.

If the flap selector switch is left in the DN po-sition, the down solenoid remains energized, andthe control valve maintains extend pressure onthe flap actuators. A check valve at the controlvalve inlet prevents flap retraction in the eventof an upstream hydraulic system failure.

Placing the selector switch in the UP positionenergizes the up solenoid, and the controlvalve repositions to direct pressure to theretract side of both actuators. In the fullyretracted position, the up solenoid remainsenergized, and the control valve maintainsretract pressure on the flap actuators. Re-

turning the selector switch to the neutral po-sition deenergizes the up solenoid and thecontrol valve repositions to neutral.

SPOILERSThe spoilers, located on the upper surface of thewings forward of the flaps, may be extendedsymmetrically for use as spoilers (spoiler mode)or asymmetrically for aileron augmentationwhen the f laps are extended beyond 25˚(spoileron mode).

The spoilers are hydraulically actuated by asolenoid-operated spoiler selector valve andtwo servo valves (one for each spoiler). Elec-trical control of the system is accomplished bythe SPOILER switch (for spoiler mode) or bythe spoiler computer (spoileron mode).

Both modes require DC and 115-VAC electri-ca l power through the SPOILER andSPOILERON circuit breakers, respectively, onthe right essential and AC buses. If either cir-cuit breaker is pulled or either power source islost in flight, the spoilers will “slam down” (ifextended) and will be inoperative in both modes.Spoiler mode operation does not require 115-VAC power on the ground.

A spoiler annunciator light illuminates duringnormal spoiler deployment or when an uncom-manded unlocked condition exists on eitherspoiler. On FC 200 AFCS models, the light isred. On FC 530 AFCS models, the light is amber.

In the event of main system hydraulic failure,the spoilers, if extended, will blow down andbe inoperative. Spoilers cannot be operatedwith hydraulic pressure from the auxiliaryhydraulic pump.

The spoiler mode, when selected, will over-ride the spoileron mode if it is operating.

While airborne, flaps and spoilers should notbe extended simultaneously. To do so maycause damage to the flaps and create excessivedrag and loss of lift, resulting in increasedstall speed for which the stall warning systemis not compensated. If the spoilers are extended




whi le the f laps a re be ing ex tended , theSPOILER annunciator light will flash as theflaps extend beyond the 13˚ position.

Operation (Spoiler Mode)The spoilers can be symmetrically extendedor retracted with the SPOILER switch (Figure15-9).

When the SPOILER switch is positioned toEXT, the spoiler selector valve is energized,the servo valves meter pressure to the extendside of the spoiler actuators, and the SPOILERlight illuminates steady. Full extension islimited to approximately 40˚. Returning theswitch to RET deenergizes the spoiler selectorvalve which repositions to route pressure to theretract side of the actuators, and the servo valvesneutralize. The SPOILER light extinguisheswhen both spoilers are locked down by lockswithin the actuators (Figure 15-9A).

Spoiler extension and retraction times vary,depending on whether the airplane is airborneor on the ground, and whether the FC 200AFCS or the FC 530 AFCS is installed. Grounddeploy and retract times (all airplanes) is 1–2seconds and 3–4 seconds, respectively. In-flight deployment times are 3–4 seconds (FC

200) and 5–7 seconds (FC 530). Retract timesare 3–4 seconds for all airplanes.

Spoiler deployment and retraction causes sig-nificant nosedown and noseup pitching (re-spectively). This should be anticipated andoffset by application of elevator control pres-sure and pitch trim, as necessary.

Operation (Spoileron Mode)During the spoileron (aileron augmentation)mode of operation, the spoilers are indepen-dently extended and retracted on a one-to-oneration with the upgoing aileron to increaselateral control in the landing configurations.Aileron augmentation (spoilerons) increasesroll control authority up to 50%.

The spoileron mode is automatically engagedwhen the flaps are lowered beyond 25˚ and theSPOILER switch is in RET position. The spoil-eron computer continuously monitors aileronposition. When the ailerons are displaced fromneutral, the computer signals the servo valveto extend the spoiler on the wing with theraised aileron. The spoiler on the oppositewing is held retracted by its servo valve.Spoiler extension is limited to approximately




Figure 15-9. Spoiler System




SPOILER SPOILERON

(FC 530)

R AC BUS

SPOILER SWITCH

13 FLAP SWITCHSPOILERONCOMPUTERAMPLIFIER

SPOILERON RESET SWITCH

LSPOIL

SPOILSELECTVALVE

AC

LEFTSERVOVALVE

ACTUATOREXTENDED

AUGAIL

EXTEND

FOLLOW-UPRIGHTSERVOVALVE


ENGINE-DRIVENHYDRAULIC PUMP

PRESSURE

DC

TO COMPUTER

RSPOIL

FOLLOW-UP

. .

NORMAL HYDRAULICSYSTEM PRESSURE

EXTEND

RETURN

ELECTRICAL

LEGEND

(FC 200)

R ESS BUS

DC

SPOILER

ACTUATOREXTENDED

SPOILER

EXTEND

Figure 15-9A. Spoiler Operation

15˚ dur ing spoi le ron opera t ion ( fu l l upaileron). The SPOILER light will not illumi-nate during spoileron operation.

Spoileron operation is shown in Figure 15-10.

Spoileron Monitor System

The computer monitors spoiler and spoileronmodes of operation by a followup in eachspoiler and each aileron. In flight, if a split ofmore than 6˚ occurs between the two spoilers(spoiler mode) or between the aileron andspoiler (spoileron mode), the amber AUG AILlight will illuminate and the spoilers will slamdown. Both modes will remain inoperative (inflight) as long as the AUG AIL light is illu-minated. However, the spoiler mode may beoperative on the ground.

Spoileron Reset Switch

The SPOILERON RESET switch (Figure 15-9)is spring-loaded to the OFF position. If a mal-function occurs in either mode (AUG AIL lighton), moving the SPOILERON RESET switchmomentarily to RESET may restore spoil-er/spoileron operation, provided the malfunc-tion has cleared. If the AUG AIL light does notgo out, both modes are inoperative in flight.

The SPOILERON RESET switch is also usedduring the spoileron/spoiler preflight check ofmonitor circuit operation. On the ground withflaps down, holding the switch in RESET in-duces a fault that inhibits spoileron movement.Therefore, if the control wheel is turned whileholding the switch in RESET, the AUG AILlight should come on after the aileron hasdeflected approximately 6˚. The system can bereset by releasing the SPOILERON RESET




SPOILER SPOILERON

R ESS BUS R AC BUS

SPOILER SWITCH RETRACT

25 FLAP SWITCHSPOILERONCOMPUTERAMPLIFIER

SPOILERON RESET SWITCH

LAIL

RETRACT

ACSPOIL

SELECTVALVE

AC DC

LSPOIL

ENGINE-DRIVENPUMP HYDRAULIC

PRESSURE

ACTUATOREXTENDED

AugAIL

EXTEND

FOLLOW-UPS

DC

RIGHTSERVOVALVE

ACTUATOREXTENDED


DC TO COMPUTER

RSPOIL

RAIL

FOLLOW-UPS

. .

NORMAL HYDRAULIC PRESSURE SYSTEM

EXTEND

RETRACT

RETURN

ELECTRICAL

LEGEND

LEFTSERVOVALVE

Figure 15-10. Spoileron Operation (Left Aileron Up)

switch to OFF and then momentarily movingit back to RESET. Refer to the approved AFMfor the complete spoileron/spoiler check.

YAW DAMPERS

GENERALEither of two yaw damper systems may be in-stalled, depending on whether the airplane isequipped with the FC 200 AFCS, or the FC 530AFCS. Both systems are described herein.

Two separate, independent (dual) yaw dampersystems are installed in all airplanes to pro-vide yaw stability. Either system providesfull-time yaw damping in flight (whether or notthe autopilot is engaged) by applying rudderagainst transient motion in the yaw axis, whilecoordinating the rudder during turns. Switch-ing logic is such that only one yaw dampermay be engaged at a time. Each system con-sists of a yaw rate gyro, a lateral accelerom-e t e r, a compu te r- amp l i f i e r, an a i l e ronfollow-up, and a DC rudder servo-actuator.Additionally, FC 530 AFCS models use a yawdamper force sensor, a calibration assembly,and a three-axis disconnect box.

The rudder servo actuator incorporates a cap-stan mechanism (slip clutch) which allowsthe pilot to override the yaw damper at anytime, if required, by applying sufficient rud-der pedal force.

When the stall warning indicators are in theshaker range, the yaw damper effectiveness isreduced. The reduction signal for the primaryyaw damper comes from the left stall warningsystem and for the secondary yaw damperfrom the right stall warning system.

The primary yaw damper uses DC and ACpower supplied by the AFCS YAW and PRIYAW DAMP circuit breakers, respectively,on the left AC and essential buses. The sec-ondary yaw damper uses DC and AC powersupplied by the SEC AFCS and SEC YAWDAMP circuit breakers, respectively, on theright AC and essential buses.

Both yaw dampers must be operational forflight, with one engaged at all times whileairborne. The yaw damper should be disen-gaged while trimming the rudder, then reen-gaged. Ground testing of the yaw dampersmust be accomplished in accordance with theapproved AFM, Section 2.

YAW DAMPERCONTROL PANELThe yaw damper control panel located on thecenter pedestal (Figure 15-11) provides theyaw damper selection, test, and indicatingfunctions. The dual systems are independent,but share a common control panel.

On airplanes with the FC 200 AFCS, twoPWR/TEST buttons, one for each yaw damper,are used to apply power to the respective con-troller-amplifier, and for system testing. Thetwo green PWR/TEST lights illuminate to in-dicate that the associated system is powered.The two ENG buttons provide the means of en-gagement. The two green ENG lights illumi-nate to indicate an engaged yaw damper. Yawdamper disengagement may be accomplishedby depressing the associated inboard OFF but-ton, while power may be removed from the sys-tems by depressing the associated outboardOFF button. A single servo force indicatorprovides indication of the amount of rudderforce being applied by whichever yaw damperhappens to be engaged, with clockwise de-flection indicating a right rudder force.

On airplanes with the FC 530 AFCS, a singleTST button provides simultaneous testing ofboth yaw damper systems. Two PWR buttons,one for each yaw damper, are used to applyand remove power to their respective controller-amplifiers. Two ENG buttons, one for each yawdamper, are used to engage and disengage theselected yaw damper. The two green ON an-nunciators illuminate to indicate that theassociated system is powered. The two greenENG annunciators illuminate to indicate anengaged yaw damper. A servo force indicatoris provided for each yaw damper, providingindication of rudder force being applied by itsrespec t ive yaw damper, wi th c lockwisedeflection indicating right rudder force.




OPERATION (AIRPLANESWITH FC 200 AFCS)When the AUTOPILOT master switch is on,electrical power is applied to both yaw damperamplifiers, causing both green PWR/TEST lightsto illuminate. However, if the AUTO-PILOTmaster switch is off, the PWR/TEST buttons,when individually depressed, will apply powerto their respective systems, causing the asso-ciated PWR/TEST light to illuminate.

With power on (PWR/TEST lights illuminated),depressing either ENG button will engage thecorresponding yaw damper and illuminate theassociated green ENG light. If one yaw damperis engaged, depressing the opposite ENG buttonwill automatically disengage the first yawdamper and engage the second.

Disengagement of either yaw damper may beaccomplished by depressing the correspond-ing OFF button or by momentarily depressingeither pilot’s wheel master switch (MSW). Onthese airplanes, there is no audible annunci-ation of disengagement.

When a PWR/TEST button is held depressed(dur ing g round tes t ing) , the respec t ivePWR/TEST and ENG lights should illumi-nate. Simultaneously, the force indicatorshould suddenly move toward the side beingtested, then slowly drift past neutral. Releas-ing the PWR/TEST button should extinguishthe ENG light, and the force indicator shouldsuddenly move in the opposite direction, thenslowly drift back to neutral. The sudden move-ment of the force indicator tests the rate gyrocircuitry, while the slow drift of the indicatortests the lateral accelerometer. A 5-secondwaiting period should be observed if retestingis desired.

OPERATION (AIRPLANES WITH FC 530 AFCS)On these airplanes, the PWR buttons must bedepressed in order to apply power to the in-dividual amplifiers. Depressing a PWR buttona second time will remove power from theamplifiers.




PRIMARYPOWER ON

ANNUNCIATOR

PRIMARYENGAGED

ANNUNCIATOR

SERVO FORCEINDICATOR

PRIMARYPOWER

BUTTON

PRIMARYENGAGEBUTTON

CONTROLLER PANEL(FC 530 AFCS)

SERVOFORCE

INDICATOR

CONTROLLER PANEL(FC 200 AFCS)

Figure 15-11. Yaw Damper Systems

With power on (PWR annunciators illuminated),depressing either ENG button the first time willengage the corresponding yaw damper and il-luminate the associated ENG annunciator. De-pressing the ENG button a second time willdisengage the yaw damper. If one yaw damperis engaged, depressing the opposite ENG but-ton will automatically disengage the first yawdamper and engage the second.

Disengagement of either yaw damper mayalso be accomplished by momentarily de-pressing either pilot’s wheel master switch(MSW). On these airplanes, the audible au-topilot disconnect tone will always sound tosignal yaw damper disengagement.

The TST button provides simultaneous test-ing of both yaw dampers. With power on (PWRannunciators illuminated), depressing andholding the TST button should illuminate bothENG annunciators. Simultaneously, both forceindicators should suddenly move to the right,then slowly drift toward the left. Releasing theTST button should extinguish both ENG an-nunciators. The sudden movement of the forceindicators tests the rate gyro circuitry, whilethe slow drift of the indicators tests the lateralaccelerometers. A 5-second waiting periodshould be observed if retesting is desired.

On these airplanes, when flaps are extendedbeyond 25˚, the amount of rudder pedal forcerequired to override the yaw damper is sig-nificantly reduced. This enables the pilot toapply cross-control pressures without en-countering noticeable yaw damper opposi-tion. Because of this, the yaw damper must beengaged all the way to touchdown except whenlanding must be made with 0˚, 8˚ or 20˚ flaps,in which case it should be disengaged in theflareout prior to touchdown.

STALL WARNINGSYSTEMS

GENERALOne of two stall warning systems may be in-stalled on the airplane. Airplane SNs 35-067and subsequent, and 36-018 and subsequent,and earlier airplanes incorporating AAK 76-4, have the “Alpha Dot” system. Earlier un-modified airplanes have the non-Alpha Dotsystem. Both are dual systems that providevisual and tactile warning of an impendingstall and are equipped with the following dual(left and right) components: stall vane/trans-ducer assemblies, computer-amplifiers, redSTALL warning lights, stick shaker motors,ANGLE OF ATTACK indicators, and STALLWARNING switches. Both systems use theelevator pitch servo for stick pusher/nudger op-eration (Figure 15-12).

The Alpha Dot system uses flap positionswitches, aneroid switches, and rate sensorsto provide “bias” information to the computer,which accounts for changes in stall speed inrelation to flight conditions and flap config-urations. Flap bias is provided by flap switchesat the 3˚, 13˚, and 25˚ positions. Altitude biasis provided by the aneroid switches at 22,500feet. The rate sensors establish the rate ofchange of increasing angle of attack, as in anaccelerated approach to a stall. The non-AlphaDot system is biased only for flap position at13˚ and 25˚ and is not equipped with theaneroid switches or rate sensors.

The left and right systems are completely in-dependent. The systems operate on DC powersupplied from the L and R STALL WARN cir-cuit breakers on the left and right batterybuses; therefore, each system can be poweredeven when the battery switches are off. The Land R STALL warning lights are the only com-ponents that do not take power directly fromthe battery buses.







ELECTRICAL

MECHANICAL

LEGEND

LSTALL

R STALL

L STALLWARNING

COMP/AMP

STALL WARNINGVANES

OFF

L R

BIAS INPUTS: FLAP POSITION ALTITUDE RATE SENSOR ACCELEROMETER

SHAKERMOTOR

PRIMARYYAW

DAMPER

STALL WARNING

ANGLE OF ATTACK

SHAKERMOTOR

TORQUEREDUCTION

SIGNAL

SECONDARYYAW

DAMPER

ELEVATOR

* ALPHA DOT AIRPLANES ONLY ** FC 530 AFCS ONLY

PITCH SERVONUDGERPUSHER

ANGLE OF ATTACK

R STALLWARNING

COMP/AMP

**

*

**

TORQUEREDUCTIONSIGNAL

*

Figure 15-12. Stall Warning System

ANGLE OF ATTACK IndicatorsThe computers translate signals from the stallvane transducers into visual indications of stallmargin on the ANGLE OF ATTACK indicators.The face of the indicators is divided into threecolor segments—green, yellow, and red. Thegreen segment represents the normal operat-ing range. The yellow segment warns of an ap-proaching stall condition. Tactile warningoccurs in this area, alerting the pilot to takepositive action. The red segment signifies thataerodynamic stall is imminent or has occurred.The stick pusher is engaged in this area,thereby forcing a reduction in angle of attack.

Warning LightsThe L and R STALL warning lights begin toflash when the respective ANGLE OF AT-TACK indicator pointers enter the shakerrange, as described above. The STALL WARNlights illuminate steady in the red segment(pusher range). Steady illumination of thelights at any other time indicates a computerpower loss or a circuitry malfunction. Cyclingthe STALL WARNING switch(es) off, thenon, may restore normal operation. The lightsi l luminate whenever the STALL WARNswitches are OFF.

Stick ShakerStick shaker motors are attached to the frontside of each control column. Actuation of theshakers causes a high-frequency vibration inthe control columns.

PusherThe stick pusher function utilizes the elevatorpitch servo to reduce angle of at tack bydecreasing airplane pitch attitude. Pusheractivation provides elevator down motion,causing a sudden abrupt forward movement ofthe control column. The mechanical slip clutchon the pitch servo allows the pilot to overridean inadver t en t pushe r ac tua t ion due tomalfunct ion. Addi t ional ly, on a i rplanesequipped with the FC 530 AFCS, depressing

and holding the wheel master switch willcancel an inadvertent pusher. See the approvedAFM for appropriate corrective action.

Nudger (Airplaneswith FC 530 AFCS)On these airplanes, a nudger is incorporatedinto the stall warning system. As angle ofattack increases slightly beyond the point ofshaker motor operation (but prior to pusheroperation), a gentle pulsating forward pushcommand is applied to the pitch servo (thesame servo that operates the pushers).

If the nudger fails to operate, a pulsatingnudger monitor horn will sound to alert thepilot. In this case, angle of attack must bedecreased immediately because the pusherhas also failed.

OPERATIONDuring flight, the stall warning vanes alignwith the local airstream. The vane-operatedtransducers produce a voltage proportional toairplane angle of attack. These signals, biasedby information from the flap position switches,altitude switches, and rate sensors (as appli-cable) are sent to the respective computer. Asangle of attack increases, the indicator pointmoves to the right. As it crosses the green/yel-low line, activation of the flashing STALLlights, st ick shaker, and st ick nudger (ifinstalled) begins. If angle of attack is allowedto increase further, the pusher is activated asthe pointer crosses the yellow/red line. As-suming an unaccelerated entry to a stall con-dition of altitudes below 22,500 feet, thegreen/yellow line approximates 7 knots or 7%above pusher speed, whichever is higher. Theyellow/red line approximates 5% above stallspeed (non-Alpha Dot); 1 knot above stallspeed (Alpha Dot, except FC 530 AFCS air-planes) or; stall speed ±3 knots (Alpha Dot air-planes with FC 530 AFCS). The 22,500-footaneroids on all Alpha Dot airplanes causewarn ing and pushe r func t ions to occurapproximately 15 knots earlier at high altitudesin the flaps-up configuration.




MACH OVERSPEEDWARNING/STICKPULLER

GENERALThe Mach overspeed warning system providesaudible overspeed warning in the event air-plane speed reaches VMO or MMO. The stickpuller function signals the pitch servo to torquethe elevator noseup if MMO is exceeded. OnFC 530 AFCS models, the puller also operatesif high-altitude VMO is exceeded.

The stick puller utilizes the autopilot pitchaxis circuitry to control the elevator servoforce applied. The resultant noseup force onthe control column during puller actuation isapproximately 18 pounds. If the autopilot isengaged, puller actuation cancels any selectedf l ight di rector pi tch modes and inhibi tsautopilot use of the pitch servo until the pulleris released. System control circuits require28 VDC and 115 VAC supplied through the LSTALL WARN and AFCS PITCH circuit

breakers, respectively, on the left essentialand AC buses. Power for the stick puller systemis controlled through the L STALL WARNswitch. The system will be inoperative if theswitch is in the OFF position.

OPERATIONThe overspeed warning horn is functionalwhenever the airplane electrical system ispowered and either WARN LTS circuit breakeris engaged (essential buses). The stick pullersystem becomes functional when the L STALLWARN switch is positioned to the on (STALLWARN) position. The STALL WARN switchesshould remain on at all times in flight exceptas directed by the approved AFM EmergencyProcedures and Abnormal Procedures sec-tions. With the stick puller inoperative, speedis limited to 0.74 MI. The mechanical slipclutch on the pitch servo allows the pilot tooverride an inadvertent puller actuation due tomalfunct ion. Addi t ional ly, on a i rplanesequipped with the FC 530 AFCS, depressingand holding the wheel master switch will can-cel an inadvertent puller. See the approvedAFM for appropriate corrective action.







1. The airplane systems that use the pitchservo to position the elevator are:A. Autopilot, Mach trim, stick pullerB. Autopilot, stick pusher, stick pullerC. Pusher, stick puller, Mach trimD. Yaw damper, stick pusher, stick puller

2. The airplane is trimmed in the pitch axisby:A. The elevator trim tabB. CanardsC. The movable horizontal stabilizerD. The elevator downspring

3. To enable pitch trim through the controlwheel trim switches, the PITCH TRIMselector switch must be in the:A. PRI or SEC positionB. PRI, OFF, or SEC positionC. PRI positionD. SEC position

4. Illumination of the red MACH TRIM lightindicates:A. Mach trim is not operating.B. The secondary trim motor is inoper-

ative.C. The autopilot is engaged above 0.74

MI.

D. The trim speed controller/monitor hasdetected a trim speed error.

5. The systems that can function with thePITCH TRIM selector switch in the SECposition are:A. Primary pitch trim and Mach trimB. Secondary pitch trim and Mach trimC. Secondary pitch trim and primary

pitch trimD. Secondary pitch trim and autopilot

pitch trim

6. In the event of runaway trim, both trimmotors can be disabled by:A. Depressing and holding either control

wheel master switchB. Moving the PITCH TRIM selector

switch of OFFC. Moving the PITCH TRIM selector

switch to EMER positionD. A or B

7. The MACH position on the rotary systemtest switch is used to test:A. Mach trim and Mach trim monitorB. Mach overspeed warning horn and

stick pullerC. Mach monitorD. The HORN SILENCE switch

8. In the event of airplane electrical fail-ure, the flap position indicator will:A. Be powered by the EMER BAT and

indicate actual position of the flapsB. Not be powered and will freeze at last

flap positionC. Fail, indicating DN regardless of flap

positionD. None of the above

9. A flashing SPOILER light indicates:A. Spoilers are split more than 6˚.B. Spoi le r-a i le ron re la t ionsh ip has

exceeded 6˚.C. Spoiler system is inoperative.D. Spoilers are extended, and flaps are

down more than 13˚.

QUESTIONS




10. The SPOILERON RESET swi tch i s used to:A. Retract the spoilers in the event of a

malfunction.B. Extend the spoilers in the event of a

malfunction.C. Reset the spoiler/spoileron system

when the AUG AIL light illuminates.D. Test the monitor system in flight.

11. If one yaw damper is found inoperativeprior to takeoff:A. The airplane may be flown, but alti-

tude is restricted to 20,000 feet.B. The airplane may be flown, but alti-

tude is restricted to 41,000 feet.C. The airplane may be flown, but the

YAW DAMP circuit breaker for the in-operative system must be pulled.

D. The airplane must not be dispatched.

12. When the ANGLE OF ATTACK indica-tor pointers are in the yellow segment:A. The pusher engages, and the horn

sounds.B. STALL WARN l igh ts i l lumina te

steady.C. The shakers (and nudgers on FC 530)

activate and the STALL WARN lightsflash.

D. The shakers activate and the stallwarning horn sounds.

13. The electrical power source for the stallwarning system is provided by:A. Battery busesB. Battery-charging busC. Main DC busesD. Emergency battery

14. If either L or R stall warning system isfound to be inoperative before takeoff:A. The airplane can be flown provided

the STALL WARN circuit breaker ispulled for the inoperative system.

B. The airplane can be flown providedthe pilot has an ATP rating.

C. The airplane may be flown providedthe autopilot and yaw damper sys-tems are operating.

D. The airplane must not be flown.

15. The switch used to turn the stick pullersystem on and off is the:A. STICK PULLER switchB. AUTOPILOT master switchC. L STALL WARN switchD. R STALL WARN switch

Revision .03 16-i

CHAPTER 16AVIONICS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 16-1

GENERAL ............................................................................................................................ 16-1

NAVIGATION SYSTEM ...................................................................................................... 16-3

Pitot-static System (FC 200 AFCS)............................................................................... 16-3

Pitot-static System (FC 530 AFCS)............................................................................... 16-4

Air Data.......................................................................................................................... 16-6

RAM AIR TEMP Indicator ........................................................................................... 16-7

AUTOFLIGHT SYSTEM ..................................................................................................... 16-7

General........................................................................................................................... 16-7

Flight Director Systems ................................................................................................. 16-8

Autopilot/Flight Director............................................................................................. 16-10

COMMUNICATION SYSTEM.......................................................................................... 16-22

Static Discharge Wicks................................................................................................ 16-22

RSVM SYSTEM................................................................................................................. 16-22

General......................................................................................................................... 16-22

Learjet RSVM Installation........................................................................................... 16-24

West Star RSVM Installation....................................................................................... 16-24

QUESTIONS....................................................................................................................... 16-40

Navigation System....................................................................................................... 16-40

Autoflight System........................................................................................................ 16-41

Communication System............................................................................................... 16-41




Revision .03 16-iii

ILLUSTRATIONS

Figure Title Page

16-1 Pitot-static System (FC 200 AFCS) ..................................................................... 16-2

16-2 Pitot Head (Typical) ............................................................................................. 16-3

16-3 Static Ports (Typical) ............................................................................................ 16-3

16-4 ALTERNATE STATIC SOURCE Valve .............................................................. 16-4

16-5 Pitot-static Head (Typical) ................................................................................... 16-4

16-6 Pitot-static System (FC 530 AFCS) ..................................................................... 16-5

16-7 STATIC PORT Switch ...........................................................................................16-6

16-8 RAM AIR TEMP Indicator .................................................................................. 16-7

16-9 ADI and HSI (Typical) ......................................................................................... 16-8

16-10 Remote Heading and Course Selector (Typical) .................................................. 16-8

16-11 Autopilot and Flight Director Control Panels ...................................................... 16-9

16-12 ADI and HSI Indications ...................................................................................... 16-9

16-13 Control Wheel Switches (Typical) ..................................................................... 16-12

16-14 Altitude Display ................................................................................................. 16-21

16-15 Static Wicks (Typical) ........................................................................................ 16-22

16-16 Rosemount Pitot and Static Probe........................................................................ 16-24

16-17 Static Source/Static Port Switch .......................................................................... 16-24

16-18 Right Side Pitot Static Probe................................................................................ 16-25

16-19 Pilot and Copilot Altimeters ................................................................................ 16-25

16-20 Air Data Switch Panel.......................................................................................... 16-26

16-21 Emergency Battery Power System....................................................................... 16-27

16-22 Learjet Electrical Diagram for Altimeter/ADDU and AIU ................................. 16-28

16-23 Standby Altimeter ................................................................................................ 16-29




16-24 Rosemount Pitot and Static Probe........................................................................ 16-30

16-25 Shoulder Static Port ............................................................................................. 16-30

16-26 Pitot Static System Schematic for AFCS FC 200 Aircraft .................................. 16-31

16-27 Pitot Static System Schematic for AFCS FC 530 Aircraft .................................. 16-32

16-28 Static Port/Source Switch .................................................................................... 16-33

16-29 West Star Air Data Computer (ADC) .................................................................. 16-33

16-30 West Star Learjet 35/36 RSVM Avionics Block Diagram................................... 16-36

16-31 West Star Pilot Altimeter ..................................................................................... 16-36

16-32 West Star Copilot Altimeter................................................................................. 16-37

16-33 Altitude Alerter .................................................................................................... 16-37

16-34 Standby Altimeter ................................................................................................ 16-38

16-35 RH Airspeed Static Valve .................................................................................... 16-38

TABLES

TABLE Title Page

16-1 FC 200 Autopilot System Modes and Annunciators ......................................... 16-14

16-2 FC 530 Autopilot System Modes and Annunciators ......................................... 16-17

16-3 West Star ADC Failure Indications Chart for FC 200 Aircraft ........................... 16-34

16-3 West Star ADC Failure Indications Chart for FC 530 Aircraft ........................... 16-35

Revision .0316-iv FOR TRAINING PURPOSES ONLY



INTRODUCTIONThe Learjet 35/36 avionics consists of, but is not limited to, the navigation system, the auto-matic flight control system (AFCS), and the comm/nav system. This chapter includes the stan-dard avionics used in the Learjet 35/36. The user should consult applicable supplements in theapproved AFM and vendor manuals for additional information and information on specificsystems not included in this chapter.

GENERALThe basic navigation system consists of the pitot-static system and air data sensor, and the ram-airtemperature gage.

The AFCS includes the flight director, autopilot,dual yaw damper, and Mach trim system. Thestandard automatic flight control systemsinstalled on the Learjet 35/36 are the JetElectronics and Technology, Inc. (J.E.T.) FC 200on the early models, and the FC 530 on the latemodels. The flight directors can be usedindependently with the pilot steering the airplane

to satisfy the flight director commands asprogrammed, or the autopilot may be engaged toautomatically steer the airplane to satisfy flightdirector commands as programmed. The dualyaw damper system operates independently ofthe autopilot and may be engaged with or withoutthe autopilot engaged. The Mach trim systemoperates at high Mach numbers when theautopilot is disengaged. The yaw damper andMach trim systems are described in Chapter 15,“Flight Controls.”

The Communication System section of thischapter discusses the static discharge wicks.

CHAPTER 16AVIONICS







Figure 16-1. Pitot-static System (FC 200 AFCS)

LH SHOULDER STATIC PORT

DRAIN VALVE

LH PITOT HEAD ALTITUDEPRESSURESWITCH*

FLAP BLOWUP AIRSPEED SWITCH **

RH SHOULDER STATIC PORT

RH PITOT HEAD

DRAIN VALVE

RH FWDSTATIC PORT

DRAIN VALVE

RH CENTERSTATIC PORT

RH AFT STATICPORT

MACH TRIM AND HIGHALTITUDE OVERSPEEDSWITCH

ALTIMETER(COPILOT)

RATE OF CLIMB(COPILOT)

AIRSPEED AND MACHNUMBER INDICATOR (COPILOT)

REAR PRESSUREBULKHEAD STATIC PORT

ALTERNATE STATIC PORT(IN NOSE COMPARTMENT)

ALTITUDEPRESSURE SWITCH*

AIRSPEED AND MACHINDICATOR (PILOT)

MACH WARNING AND LOW ALTITUDE OVERSPEED SWITCH

FORWARDPRESSUREBULKHEADSTATIC PORTINSTRUMENT ALTERNATE

STATIC SOURCE VALVE

DIFFERENTIAL PRESSURE-

RELIEF VALVE

RATE OF CLIMB(PILOT)

ALTITUDEALERTER

ALTIMETER(PILOT)

DRAIN VALVE

LH AFTSTATIC PORT

STATIC DEFECTCORRECTION

MODULE*

DRAIN VALVELH FWD

STATIC PORT

LEGEND

PILOT's PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

ALTERNATE STATIC

OTHER STATIC

AIR DATASENSOR

*SNs 35-067 AND SUBSEQUENT, 36-018 AND SUBSE- QUENT, AND EARLIER AIRPLANES WITH AAK 76-4

**SNs 35-002 THROUGH 35-059 AND 36-002 THROUGH 36-017

PRESSURIZATIONMODULE

NAVIGATION SYSTEM

PITOT-STATIC SYSTEM(FC 200 AFCS)The pitot-static system supplies pitot and staticair pressure for operation of the airspeed andMach indicators, the high- and low-altitudeoverspeed switches, the air data sensor, and thestatic defect correction module. Static pressureis also supplied to the copilot’s vertical ve-locity indicator, both altimeters, the pressur-ization control module, and the aft differentialpressure relief valve (Figure 16-1).

A heated pitot head is located on each side ofthe fuselage just forward of the cockpit (Fig-ure 16-2). Pitot heat switches are located onthe pilot’s anti-icing control panel. They alsosupply heat to both stall warning vanes. Referto Chapter 10, “Ice and Rain Protection,” foradditional information.

The normal static system provides independentsources of static pressure to the pilot’s andcopilot’s instruments. Each static source (pilotor copilot) has one static port on each side ofthe airplane nose (Figure 16-3). The dual staticports are provided for redundancy and to re-duce sideslip effects on the instruments whichuse static air.

The left front and right center static ports(both heated) are connected to the pilot’s in-struments. The left rear and right front staticports (both heated) are connected to the copi-lot’s instruments. The right rear static port(not heated) is connected with an alternatestatic port inside the nose compartment toprovide the pressurization module with a staticsource. Refer to Chapter 12, “Pressurization,”for additional information.

Two heated shoulder static ports are locatedon top of the fuselage nose in front of thewindshield. These ports provide static pressureto the air data sensor and the copilot’s FD108/FD 109 altitude controller (if installed).




Figure 16-2. Pitot Head (Typical)

Figure 16-3. Static Ports (Typical)

An ALTERNATE STATIC SOURCE valve islocated below the pilot’s instrument panel(Figure 16-4). For normal operation, the leverremains down (CLOSED); for alternate air, thelever is moved up (OPEN).

When the ALTERNATE STATIC SOURCEvalve is positioned to OPEN, the pilot’sinstruments are connected to an alternate portinside the unpressurized nose section. WithOPEN selected, the altimeter and Machindicators will read slightly lower than normal.

Condensation drain valves for the pitot andstatic air lines are located adjacent to the nosewheel well doors.

PITOT-STATIC SYSTEM(FC 530 AFCS)Pitot and static pressure for instruments andsystems is obtained from two pitot-staticprobes, one on each side of the nose section

(Figure 16-5). Each probe contains a pitotport in the tip and two static ports on the side.The probes also contain electrical heating el-ements controlled by the L and R PITOT HEATswitches. (Refer to Chapter 10, “Ice and RainProtection.”) Four drain valves located near theaft end of the nose gear doors (two on eachside) are installed at the system’s low pointsto drain moisture from the system.

The pitot systems (Figure 16-6) are indepen-dent. The left probe provides pitot pressurefor the pilot’s Mach/airspeed indicator, andthe right probe head provides pitot pressurefor the copilot’s Mach/airspeed indicator, theMach switch (0.74 MI), gear warning air-speed switch (170 KIAS), air data unit, andother optional equipment.

There are four static ports in the main pitot-static system—two on each pitot-static probe.The ports on one probe are interconnectedwith those on the other probe to provide re-dundance. Four solenoid-operated shutoffvalves enable the pilot to select the source ofstatic pressure.




Figure 16-4. ALTERNATE STATICSOURCE Valve Figure 16-5. Pitot-static Head (Typical)




Figure 16-6. Pitot-static System (FC 530 AFCS)

GEAR WARNINGAIRSPEED SWITCH

GEAR WARNINGALTITUDE SWITCH

OPTIONALEQUIPMENT

AIR DATA UNIT

MACHSWITCH

RATE-OF-CLIMBINDICATOR (COPILOT)

ALTIMETER(COPILOT)

MACH/AIRSPEEDINDICATOR(ALTITUDE/OVERSPEEDSWITCHES)

PRESSURIZATIONMODULE

RATE-OF-CLIMBINDICATOR (PILOT)

ALTIMETER(PILOT)

MACH/AIRSPEED INDICATOR (ALTITUDE/ OVERSPEED SWITCHES)

CLOSE

CLOSE CLOSE

CLOSE

SOLENOIDVALVES

SOLENOIDVALVES

STATIC PORT

L

R

DRAINS DRAINS

DIFFERENTIALPRESSURE-

RELIEF VALVE

REAR PRESSUREBULKHEAD STATIC PORT

BOTH

CopilotsPitot-StaticHead

PILOT'SPITOT-STATIC

HEAD

PITOT

STATIC 1

STATIC 2

PITOT

STATIC 1

STATIC 2

LEGEND

PILOT'S PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

OTHER

The source of static pressure is controlled withthe STATIC PORT switch located on the pilot’sswitch panel. The STATIC PORT toggle switchhas three positions: L (left), BOTH, and R(right). This switch is normally set to the BOTHposition except in the event one of the pitot-static heads becomes inoperable or unreliable(Figure 16-7).

In the BOTH position, the pilot’s instrumentsreceive static pressure from the forward port onthe left head and the aft port on the right head.The copilot’s instruments, the Mach switch, thegear warning altitude switch (14,500 feet), thegear warning airspeed switch, the air data unit,and other optional equipment receive staticpressure from the front port on the right headand the aft port on the left head. This crossconnection eliminates yaw error.

When the STATIC PORT switch is placed inthe L or R position, solenoid-operated shut-off valves are energized to shut off the staticsource from the opposite side static ports.(See Figure 16-6.)

When the STATIC PORT switch is in the Lposition, static pressure is provided to alluser systems only from the two static ports

on the left pitot-static head. In the R position,static pressure is provided to all user systemsonly from the two static ports on the rightpitot-static head.

The shutoff valves operate on DC power sup-plied through the STATIC SOURCE circuitbreaker on the left main bus. In the event of elec-trical failure, all shutoff valves will be open re-gardless of the STATIC PORT switch position.

A separate unheated s ta t ic por t i s f lushmounted on the right side of the nose sectionto provide static pressure to the pressurizationcontrol module. Refer to Chapter 12, “Pres-surization,” for additional information.

AIR DATAThe air data sensor provides air data to the auto-pilot computer and to the Mach trim computer.On airplanes equipped with the FC 200automatic flight control system, static input to theair data sensor is from the shoulder static airports. The FC 530-equipped airplanes use thecopilot’s static air system for air-data-unit input.On all airplanes, the pitot input is from thecopilot’s pitot system. The unit is located insidethe nose compartment.




Figure 16-7. STATIC PORT Switch

RAM AIR TEMP INDICATORRam-air temperature is displayed on the RAMAIR TEMP indicator located on the centerinstrument panel (Figure 16-8). The indicator iscalibrated in degrees celsius and requires DCpower from the RAM AIR TEMP circuit breakeron the left essential bus. For conversion to outsideair temperature (OAT), refer to the Ram Air ToOutside Air Temperature Conversion (RAT toOAT) figure in Section V of the approved AFM.

AUTOFLIGHT SYSTEM

GENERALEither the J.E.T. FC 200 or the J.E.T. FC 530AFCS may be installed, depending on pro-duction serial number. The FC 530 AFCS is

installed on SNs 35-408, 35-447, 35-468, 35-506 and subsequent, and 36-054 and subse-quent, and earlier SNs incorporating AAK 83-2.

NOTEThe yaw axis is controlled by the dualyaw damper system which operatesindependently of the autopilot andflight director.

Both systems incorporate a dual-channel AFCScomputer which integrates the autopilot pitch androll axes with the customer-specified flight directorsystem. The AFCS control panel, located in thecenter of the glareshield, provides pilot access tothe autopilot and to the AFCS computer for theflight director programming (mode selection).

The AFCS computer processes informationreceived from the primary vertical and directionalgyros, horizontal situation indicator (HSI), theNAV 1 receiver, and the air data sensor. Theresulting computed roll and/or pitch command(s)is applied by the computer to the flight directorindicator (FDI) command bars, which are builtinto the pilot’s attitude director indicator (ADI).

When engaged, the autopilot is always coupled tothe flight director command bars. The pilot hasthe option of using the flight director with theautopilot disengaged.

Additional controls available to the pilot forcontrol of the autopilot and flight directorfunctions are:

• Both four-way trim switches

• Both maneuver control switches

• The pilot’s pitch SYNC switch

• The go-around switch (left thrust leverknob)

• The pilot’s HSI heading (HDG) andCOURSE selector knob

• The altitude alerter and pilot’s altimeter(FC 530 AFCS only)

All of these controls are described in detail inthis section.




Figure 16-8. RAM AIR TEMP Indicator

FLIGHT DIRECTOR SYSTEMS

GeneralSeveral different flight directors are availablefor installation on the Learjet 35/36. The mostcommon installations are the Collins FD 108,FD 109, FIS 84, and FDS 85. Either system in-cludes an attitude director indicator (ADI) anda horizontal situation indicator (HSI) whichprovide conventional, “raw-data” attitude andheading reference, and glide slope and coursedeviation displays. The basic airplane attitudeand heading references are energized wheneverDC and AC power is applied to the airplane.

The flight director system is connected to theAFCS when the AUTO PILOT master switchis turned on.

When the auto pilot master switch is posi-tioned to auto pilot (on), the PWR annuncia-tor illuminates on the AFCS control panel,indicating that power is available to the auto-pilot and flight director. The AFCS controlpanel provides for flight director mode se-lection and annunciation whether the autopi-lot is engaged or disengaged. Autopi lotengagement is accomplished by depressingthe ENG button.

Refer to Figures 16-9 through 16-11 for typ-ical installations.

Attitude Director Indicator (ADI)The pilot’s ADI provides a visual presentationof the airplane attitude, as furnished by the re-mote primary vertical gyro. The flight direc-tor indicator (FDI) is built into the ADI andconsists of a set of computer-positioned com-mand bars which provide a single-cue com-mand reference for both pitch and roll. Thebars move up or down to command pitch, androtate counterclockwise and clockwise to com-mand roll. When a flight director mode(s) hasbeen selected, the command bars appear inview to provide the computed pitch and rollcommands. When the autopilot is engaged, itautomatically responds to the command bars.If the autopilot is disengaged, the pilot must




Figure 16-9. ADI and HSI (Typical)

Figure 16-10. Remote Heading and CourseSelector (Typical)




TRK

APPR

HDG

ARM CAPT

NAV REV LVL

PWR ROLL

TEST ENG

PITCH

SOFT

IAS MACH ARM CAPT

SPD V/S G/S ALT

G/A

FNL

PITCH SOFT

IASMACH ON

ARMCAPT

ARMCAPT ON

FNL

G/AALTHLD

ALTSEL

G/SV/SSPDSFTENGTSTLVLBCNAV 1/2BNK

HDGMON

TRKON ON ON ON

ROLL PWR

AFCS

FC 200 AFCS

FC 530 AFCS

ARMCAPT

Figure 16-11. Autopilot and Flight Director Control Panels

Figure 16-12. ADI and HSI Indications

20 20

1010

10 10

20 20

SLOW

FAST

DECISION HEIGHTANNUNCIATOR

SPEED DEVIATIONDISPLAY

AIRPLANE SYMBOL

RUNWAY SYMBOL

RATE-OF-TURNINCLINOMETER

PUSH-TO-TESTSWITCH

COMMAND BARS

GLIDE-SLOPEPOINTER

HORIZON

ATTITUDE TAPE

TEST

DH

HEADING MARKER(HEADING BUG)

LUBBERLINE

INS TRACKPOINTER

COURSEDISPLAY

COURSEARROW

INSANNUNCIATOR

LATERALDEVIATIONBAR

AZIMUTH CARDBEARING POINTER

AIRPLANEREFERENCESYMBOL

GLIDE-SLOPEPOINTER

TO-FROMPOINTER

DISTANCEDISPLAY

01 1COURSE

INS

MILES

33

30W

24

21

3

15S

N

612

E

perform the roll and pitch maneuvers neces-sary to align the airplane symbol with thecommand bars. Figure 16-12 illustrates thevisual indications provided by the ADI andHSI. The ADI also provides for indication oflocalizer and glide-slope deviation and turnand slip.

Horizontal Situation Indicator(HSi)The HSI provides a pictorial presentation ofairplane position relative to VOR radials andlocalizer and glide-slope beams. Heading ref-erence with respect to magnetic north is pro-vided by a remote directional gyro which isslaved to a remote fluxgate compass. TheSLAVE-FREE switch on the lower instrumentpanel allows unslaved operation by selectingFREE, in which case the magnetic reference(flux-gate compass) is removed.

The HSI provides the AFCS computer infor-mation regarding existing heading, headingmarker reference, selected course, and coursedeviation. The heading marker (bug) is used todirect the airplane to turn to and maintain theheading selected with the heading (HDG) con-trol knob. The course deviation indicator is usedto intercept and track a VOR or LOC coursewhich is set with the course control knob.

AUTOPILOT/FLIGHT DIRECTOR

GeneralThe autopilot will automatically fly the air-plane to, and hold, desired heading, attitudes,and altitudes. The autopilot system can alsocapture and track VOR/LOC/ILS radio beams.The system provides modes for speed controland vertical rate control as well.

On Learjet 35/36 airplanes with the standardavionics installation, the flight director is in-tegrated with the autopilot by a computerthrough the AFCS cont ro l pane l on theglareshield. Autopilot and flight directormodes are engaged by depressing the appli-cable mode selector buttons on the control

panel. Flight-director-only mode selection isaccomplished by depressing the desired modeselectors on the control panel (Figure 16-11),but with the autopilot disengaged.

When the autopilot is not engaged, the ADIcommand bars indicate the deviation from thedesired flight path, enabling the pilot to manu-ally fly the airplane in response to the flight di-rector system. When the autopilot is engaged,it will align the airplane with the command barsautomatically to maintain the desired flight path.

DescriptionAirplane SNs 35-462, 35-447, 35-506 and sub-sequent and 36-054 and subsequent a reequipped with the FC 530 AFCS. Earlier SNsare equipped with the FC 200. Both are manu-factured by J.E.T. AAK 83-2 is available toretrofit the earlier models with the FC 530.Both systems include: an autopilot/flight di-rector computer, an electric box, and interface,all located under the pilot’s seat; the AFCScontrol panel mounted in the center glareshield;the roll and pitch servoactuators and follow-ups;the customer-specified flight director system;a roll-rate gyro; the NAV 1 receiver; the primary(pilot’s) vertical gyro, directional gyro andHSI; and the air data sensor. The FC 530 alsouses the altitude alerter and the pilot’s altime-ter for its altitude preselect feature.

AFCS Control Panel

The control panel (Figure 16-11) is mounted inthe center of the glareshield. It is accessible toboth pilots and provides the switches requiredfor autopilot engagement and flight directormode selection. Annunciator lights are green,amber, blue, or white and appear above the modeselect switches. The legend (white lettering) onthe panel is back-lighted. On FC 200 models, in-tensity of the annunciator lights and the legendlighting is controlled by the PEDESTAL lightsrheostat on the copilot’s right sidewall. On FC530 models, intensity of the annunciator lightsis fixed so that they are legible in daylight, whilethe NAV LTS switch must be turned on for fixedillumination of the legend lighting.




The autopilot engage (ENG) pushbutton is usedonly to engage the autopilot, while all otherpushbutton switches operate with alternate ac-tion. The first depression engages a mode; thesecond depression cancels it. Automatic can-cellations also occur. Annunciation of the modeselected appears above the pushbutton. Anyoperating mode not compatible with a newlyselected mode is automatically canceled infavor of the latest selection. This allows thepilot to advance along the flight sequencewithout the inconvenience of having to deselectmodes manually.

Computer

The two-channel (roll and pitch) computer con-tinuously monitors input signals from all AFCScomponent sensors. The computer is pro-grammed by depressing the desired mode se-lector button(s) on the AFCS control panel.The computer computes the roll and pitch at-titudes necessary to comply and signals theflight director V-bars to position accordingly,while also applying simultaneous signals tothe roll and pitch servoactuators (if the au-topilot is engaged).

Operation

The autopilot and flight director system con-trols airplane movement about two axes (pitchand roll). The yaw damper provides indepen-dent, automatic control of the yaw axis in thesame way as when the airplane is being flownmanually.

Pitch Axis Control

The compute r p i t ch channe l p rocessesinformation furnished by the primary (pilot’s)vertical gyro, which establishes the basic pitchreference; the air data sensor, which suppliesaltitude, vertical velocity, and airspeed/Machinformation; glide-slope signals from the NAV1 receiver; and a follow-up device in the pitchservoactuator which signals elevator move-ment. The FC 530 also uses the altitude alerterand pilot’s altimeter for its altitude preselect fea-ture and a vertical accelerometer which mon-itors G forces.

When a pitch mode is selected on the AFCS con-trol panel, the computer responds by position-

ing the flight director V-bars accordingly. Ifthe autopilot is engaged, a signal is also ap-plied to the elevator pitch servo which adjustselevator position. Feedback of elevator move-ment is provided by the servo follow-up. Whenthe new pitch attitude has been established, thecomputer zeroes the servo effort by applyinghorizontal stabilizer trim via the secondarypitch trim motor, thereby preventing any airplanepitching motion when disengaging the autopi-lot. Pitch changes can also be induced by eitherpilot’s wheel trim switch (without depressingthe center button).

The computer uses the servo follow-up to con-trol pitch changes to a rate of 1º per second,and limits pitch attitudes to ±25º (FC 200),and +20º and –10º (FC 530).

Roll Axis Control

The computer roll channel processes infor-mation furnished by the primary (pilot’s) ver-tical gyro, which establishes the basic rollreference; the primary (pilot’s) directionalgyro and HSI, which supply heading andcourse references; VOR bearing and ILS/LOCcourse references from the NAV 1 receiver; aroll rate gyro, which provides roll rate data;and a follow-up on the left-hand aileron sec-tor, which signals aileron position.

When a roll mode is selected on the AFCScontrol panel, the computer responds by posi-tioning the flight director V-bars accordingly.If the autopilot is engaged, a signal is also ap-plied to the aileron roll servo which adjustsaileron position. Feedback of aileron positionis provided by the aileron followup. Rollchanges can also be induced by either pilot’swheel trim switch when moved to LWD orRWD (without depressing the center button).

The autopilot does not apply trim in the rollaxis as it does in the pitch axis. Therefore, ifthe airplane is out of trim in the roll axis, theautopilot must apply continuous roll servo ef-fort to hold the desired roll attitude. This con-dit ion will be noticed by a continuouslydeflected roll force meter and control wheel.

The computer uses the roll rate gyro to con-trol roll rates to 6º per second (FC 200), and4º to 5º per second (FC 530). Bank angles arelimited to a maximum of 30º.




The FC 200 uses a 13º flap position switch toincrease autopilot roll authority when the air-plane is configured for approach. This providesmore lateral authority at the slower speedsand is annunciated by the green APPR light onthe AFCS control panel. The FC 530 uses a 3ºflap position switch to desensitize VOR andLOC signals which enhances close-in stabilityduring approaches. It does not affect auto-pilot roll authority nor is it annunciated.

Electrical Requirements

The autopilot requires DC and AC electricalpower. The DC power is supplied through theAFCS, AFCS PITCH, and AFCS ROLL circuitbreakers on the left essential bus; 115 VAC issupplied through AFCS PITCH and AFCSROLL circuit breakers on the left AC bus. Allautopilot circuit breakers are located on thepilot’s circuit-breaker panel; however, on FC 200AFCS airplanes, there are three circuit breakerson the front side of the autopilot electric boxunder the pilot’s seat for autopilot and yawdamper annunciator lights and edge lights.

Controls and Indicators

The autopilot and flight director control panelcontains most of the controls and indicators usedfor the autopilot system. Additional controls andindicators are found on the control wheels, thepilot’s switch panel, the HSI, the remote headingand course selector, the ADI, the altitude alerter,and the thrust levers.

Autopilot Master Switch

Power is provided to the autopilot and flightdirector systems when the AUTO PILOT masterswitch (located on the pilot’s lower switch panel)is placed in the AUTO PILOT position, the greenPWR (power) annunciator on the autopilotcontroller illuminates, and the red CMPTR flagon the pilot’s ADI goes out of view.

Control Wheel Trim Switch

Either control wheel trim switch (NOSEUP/NOSE DN/LWD/RWD) functions as amanual autopilot controller when moved in anyone of four directions without depressing the

trim arming button (Figure 16-13). When anattitude change is made in this manner, theappropriate servo changes the attitude of theairplane and disengages any modes previouslyselected in the affected axis (except NAV ARM,G/S ARM, and ALT SEL ARM). The autopilotreverts to basic attitude hold in the affected axiswhen the switch is released.

Depressing the trim arming button and moving thetrim switch in any of the four directionsdisengages the autopilot, and the autopilotdisengagement tone will sound. This is consideredthe normal means of disengaging the autopilotsince it does not disengage the yaw damper.Previously selected flight director modes are notdisengaged when the autopilot is disengaged.Autopilot disengagement is further described inthis chapter under “Autopilot Disengagement.”

Control Wheel Master Switch

Depressing either pilot’s control wheel masterswitch (MSW) disengages the autopilot and theyaw damper. The switch is referred to as theautopilot release/nose steer switch on the FC200 AFCS.

Control Wheel Maneuver Switch

The control wheel maneuver control switch isreferred to as the MANEUVER switch on the FC200 AFCS and as the MANUV/RP switch on theFC 530 AFCS.




Figure 16-13. Control Wheel Switches(Typical)

On FC 200 airplanes, depressing and holdingeither the pilot’s or copilot’s MANEUVERswitch (see Figure 16-13) temporarily releasesautopilot access to the pitch and roll servos,biases the command bars out of view, and cancelsthe ROLL and PITCH modes if engagedpreviously. This enables either pilot to change theairplane attitude in both pitch and roll axesmanually. When the switch is released, theautopilot assumes basic attitude hold functions.

During flight-director-only operation, themaneuver switch will simply cancel all selectedflight director modes and bias the command barsout of view.

On FC 530 airplanes, depressing and holdingeither the pilot’s or copilot’s MANUV/RP switchtemporarily releases autopilot access to the pitchand roll servos and extinguishes the green ROLLand PITCH annunciators, but does not cancel anypreviously selected flight director roll or pitchmodes. This enables either pilot to change theairplane attitude in both pitch and roll axesmanually. When the switch is released, theautopilot will resynchronize to and hold theoriginal roll mode and the existing (new) valuesin the SPD, V/S, or ALT HLD modes, and thegreen ROLL and PITCH annunciators willilluminate again.

Control Wheel SYNC Switch

On FC 200 airplanes, the pilot's pitch SYNCswitch:

• Releases autopilot access to the pitch servo

• Allows the pilot to use manual elevatorcontrol to establish a new pitch attitude

• Cancels any selected pitch modes (except G/SARM), but does not affect any roll modes

• Causes the command bars to synchronizeto the new pitch attitude

• Causes the autopilot to hold the pitch attitudeexisting at the moment of switch release

On FC 530 airplanes, the pilot’s PITCH SYNCswitch:

• Is a flight director function only, and hasno effect if the autopilot is engaged

• Will cancel any selected pitch modes(except G/S ARM and ALT SEL ARM)

• Synchronizes the command bars to theexisting pitch attitude

In the case of a dual flight director installation,the copilot’s pitch SYNC switch synchronizesonly the copilot’s command bars to the existingattitude and cancels the copilot’s G/A mode, ifselected. It does not affect the autopilot in anyway (as the maneuver switch does).

Autopilot Engagement

The AUTO PILOT master switch must be placedon to accomplish system ground checks prior toflight and normally remains on throughout theflight. When the PWR annunciator is illuminated,the autopilot can then be engaged at any time(except during takeoff and landing) by depressingthe ENG button. Illumination of the PITCH andROLL annunciators indicate engagement of therespective axes.

On FC 200 airplanes, initial autopilot engage-ment will cancel all previously selected flightdirector modes (if bank angle happens to be morethan 5°), the command bars will disappear, andthe autopilot will hold the existing roll and pitchattitudes (if within normal limits). If bank angleis less than 5° at the moment of initialengagement, the LVL light will illuminate andthe command bars will be presented,commanding the autopilot to maintain wingslevel at the existing pitch attitude. If the roll orpitch attitude(s) happens to be beyond the normallimit(s), the autopilot will (at normal rates) rolland/or pitch the airplane to the normal limit(s).

If the PITCH TRIM selector switch is in the OFFposition, the autopilot may engage, but willdisengage when it attempts to adjust secondarypitch trim and cannot.

On FC 530 airplanes, autopilot engagement willautomatically couple to any previously selectedflight director mode(s) (except G/A, in whichcase the G/A light will extinguish and theautopilot will maintain the existing attitude at themoment of engagement). If the autopilot isengaged without any previously selected flightdirector mode(s), the autopilot will maintain theexisting roll and pitch attitudes (if within normallimits), and the command bars will remain out ofview. If bank angle is less than 5° at the momentof engagement, the LVL light will annunciate andthe command bars will be presented,commanding the autopilot to maintain wings




level at the existing pitch attitude. The autopilotwill not engage at bank angles in excess of 38°±2° (regardless of pitch attitude). However, ifbank angle happens to be between 30° and 38°±2° and/or pitch angle is greater than –10° or+20°, the autopilot will (at normal rates) rolland/or pitch the airplane to the normal limit(s).

If the PITCH TRIM selector switch is in the OFFposition, the autopilot will not engage.

Attitude Hold Mode

The autopilot is in pitch attitude hold when thePITCH annunciator is illuminated and all otherpitch axis annunciators are extinguished (exceptG/S ARM and, for FC 530, ALT SEL ARM).The autopilot is in roll attitude hold when theROLL annunciator is illuminated and all otherroll axis annunciators are extinguished (exceptNAV ARM). When the autopilot is in both pitchand roll attitude hold, the flight directorcommand bars will be out of view. Autopilot roll(bank) limit is a nominal 30°, while pitch limitsare +25° and –25° (FC 200 airplanes), and +20°and –10° (FC 530 airplanes).

Extended autopilot operation in roll attitudehold or LVL will cancel the automatic erec-tion feature of the vertical gyro. As the verti-cal gyro precesses, the autopilot will bank theairplane to maintain a zero-bank indication onthe attitude indicator.

When the autopilot is in the basic attitude holdmode, attitude commands are accepted by theautopilot through either pilot’s control wheeltrim switch (arming button not depressed), and

the autopilot will hold the attitude that existswhen the command is released.

Autopilot/Flight Director Mode Selection

Autopilot and flight director modes are engagedby depressing the applicable mode selectorbutton on the autopilot control panel. Theengaged modes may be disengaged by depressingthe selector button (except for the SPD mode onthe FC 530 AFCS) a second time or by selectinganother pitch mode.

Flight-director-only mode selection is made bydepressing the applicable mode selector with theautopilot disengaged.

The roll axis modes are LVL (level), HDG(heading), NAV (navigation), VOR or LOC (usedin conjunction with the NAV mode), BC (backcourse—FC 530), REV (back course—FC 200),and 1⁄2 BNK (half bank—FC 530 only).

The pitch modes are SPD (speed), V/S (verticalspeed), G/S (glide slope), ALT SEL (altitudeselect—FC 530 only), ALT HOLD (altitudehold), and SFT (soft). The SPD submodes of IASand MACH, and the V/S, G/S CAPT, ALT SELCAPT, and ALT HLD modes cancel each otherwhen one is selected. G/S ARM is compatiblewith a previously selected SPD. V/S or ALTmode, while ALT SEL is compatible with apreviously selected SPD, or V/S mode.

Refer to Tables 16-1 and 16-2 for a furtherdescription of each mode, the applicableannunciator, and the function of each modeselector switch and annunciator.




MODE ANNUNCIATOR FUNCTION

PWR Indicates electrical power is available for autopilot/flight director operation (circuit breakers are in and the AUTO PILOT master switch is in the ON position.

TEST When pressed during ground check, all autopilot controller annun-ciators illuminate. Failure to light indicates a malfunction in the AFCS or a burned out lamp. Force meters oscillate. When pressedin flight, only the annunciators illuminate.

ENG ROLL When depressed, the autopilot engages and the ROLL and PITCHPITCH annunciators illuminate.

Table 16-1. FC 200 AUTOPILOT SYSTEM MODES AND ANNUNCIATORS





SOFT SOFT When depressed, the autopilot provides softer response in the pitchand roll axes for flying through turbulence. No function during flight-director-only operation.

NOTESOFT mode is locked out when an ILS frequency istuned on NAV 1.

HDG ON When selected, flight director commands are generated tomaneuver the airplane to fly a heading selected with the pilot’s HSIheading “bug” using up to 25° of bank.

NOTEThe turn will be commanded in the shortest direction.It is recommended that the heading “bug” initially beset to not more than 135° in the direction of thedesired turn when the turn is more than 135°.

NAV When selected, it activates the flight director function that capturesand tracks VOR and LOC. Functional only when the NAV 1 receiveris tuned to the appropriate frequency, NAV flag is out of view, anddesired course is set on the pilot’s HSI. The HDG mode may beused to intercept the course provided the intercept angle is lessthan 90°.

ARM Illuminates when NAV mode is selected. Goes out when the CAPTlight illuminates. The ARM light will flash if NAV CAPT disengagesdue to a noisy or failed receiver signal, and in the cone of silenceover VOR stations.

NOTEWhen the ARM light is flashing, the flight director willassume a heading hold.

CAPT (Capture) Illuminates when the airplane approaches the desired course.Extinguishes if the receiver signal becomes noisy or fails, or while inthe cone of silence over VOR stations.

TRK In the NAV CAPT mode, illuminates to indicate the airplane hasacquired the center of a VOR or LOC beam. Crosswindcompensation begins and maximum bank angle will be limited to15° when it illuminates.

APPR The APPR light illuminates when the flaps are lowered beyond 13°and increases the autopilot roll torque limit to compensate for slowerairspeed.

REV Functional only with NAV mode selected for localizer backcourse(BACK approach with ILS frequency tuned in. When selected, course in-COURSE) formation to the flight director is reversed and the glide-slope signal

is locked out. The published inbound (front) course must be set inthe pilot’s HSI course window.

ON Indicates that the backcourse mode is selected.

NOTEREV may also be used to fly outbound on an ILSfront course.

Table 16-1. FC 200 AUTOPILOT SYSTEM MODES AND ANNUNCIATORS (Cont)





LVL When engaged, wings level is commanded by the flight director only (LEVEL) if the autopilot is engaged.

ON Indicates the level mode is engaged. It is also a function of G/Amode, but has no other flight-director-only functions.

SPD When selected, the flight director will command a pitch attitude that (SPEED) will maintain the airspeed existing at the time of mode selection.

Power must be set by the pilot.

IAS Illuminates at altitudes up to approximately 29,000 feet.

MACH Illuminates at altitudes above approximately 29,000 feet.

V/S (VER- When selected, the flight director commands a pitch attitude thatTICAL will maintain the existing vertical speed. Power must be set by the SPEED) pilot.

ON Illuminates when V/A mode is selected.NOTE

Before engaging this mode, maintain the desiredrate long enough (approximately 15 seconds) forvertical speed indicator lag to diminish.

G/S When selected, activates the flight director function that captures(GLIDE the glide slope.SLOPE)

Functional only when the NAV 1 receiver is tuned to an ILS fre-quency, an active glide-slope signal is present, the G/S flag is out ofview, and the REV mode is not selected.

ARM Illuminates when the G/S mode is selected and the airplane is noton the glide-slope beam. Goes out when the airplane captures thebeam.

CAPT Illuminates when the airplane intercepts and captures the glide-slope beam.

FNL (FINAL) Illuminates during an ILS or a localizer approach when the beamsignal is being desensitized for close-in stability.

NOTEThe FNL mode will be activated when passing overthe outer marker. If the outer marker signal is notavailable, depressing the NAV 1 TEST buttonmomentarily will activate the FNL mode. This shouldbe accomplished at the final approach fix. The flapsmust be down 13° or more to initiate FNL.

ALT When selected, the flight director will command an airplane pitch(ALTITUDE attitude that will maintain the existing altitude.HOLD)

ON Illuminates when ALT hold is engaged.

G/A (GO- Flight-director-only mode, selected by depressing the GO-AROUNDAROUND) button on the left thrust lever knob. Illuminates the G/A and LVL

annunciators, and positions command bars to 9° pitch up, wingslevel.

On SNs 35-002 through 35-009 and 36-002 through 006, the G/Amode is coupled to the autopilot when N1 is above 80%.






PWR Indicates electrical power is available for autopilot/flight directoroperation (circuit breakers are in and AUTO PILOT master switch isin ON position).

TST (TEST) When depressed, all autopilot controller annunciators illuminate(light test only). When depressed simultaneously with ENG button,a system self-test is performed.

MON (MONITOR) Illuminates during self-test. Flashes if fault is detected.

ENG ROLL When depressed, the autopilot engages and the ROLL and PITCHPITCH annunciators illuminate.

SFT SOFT When depressed, the autopilot provides softer response in the pitchand roll axes for flying through turbulence. No function during flight-director-only operation.

NOTESFT mode is locked out when in NAV localizerCAPT, NAV VOR APPR, and ALT SEL CAPT.

HDG ON When selected, flight director commands are generated tomaneuver the airplane to fly a heading selected with the pilot’s HSIheading “bug” using up to 25° of bank.

NOTEThe turn will be commanded in the shortest direction.It is recommended that the heading “bug” initially beset to not more than 135° in the direction of thedesired turn when the turn is more than 135°.

1⁄2 BANK ON Functional only with HDG or NAV VOR mode selected. Limits bankto a maximum of 13°.

NAV When selected, it activates the flight director function that capturesand tracks VOR and LOC courses. Functional only when the NAV 1receiver is tuned to the appropriate frequency, NAV flag is out ofview, and desired course is set on the pilot’s HSI. The HDG modemay be used to intercept the course provided the intercept angle isless than 90°.

Illuminates when NAV mode is selected. Goes out when the CAPT light illuminates. The ARM light will flash if NAV CAPT disengages due to a noisy or failed receiver signal, or while in the cone of silence

ARM over VOR stations.NOTE

When the ARM light is flashing, the flight director willcommand a heading equal to the selected course plusthe computed wind drift correction angle.

Illuminates when the airplane approaches the desired course.Extinguishes if the receiver signal becomes noisy or fails, or while inthe cone of silence over VOR stations.

NOTEWhen flying in VOR approach, the flaps must be set at8° or more in order to achieve signal desensitizationfor close-in stability. This function is provided by the 3°flap switch.

TRK In the NAV CAPT mode, illuminates to indicate the airplane is nearingthe VOR or LOC beam. Crosswind compensation begins andmaximum bank angle will be limited to 15° when it illuminates.

Table 16-2. FC 530 AUTOPILOT SYSTEM MODES AND ANNUNCIATORS

CAPT(Capture)





BC (BACK- Functional only with NAV mode selected for localizer backcourseCOURSE) approach. When selected, course information to the flight director is

reversed and the glide-slope signal is locked out. The publishedinbound (front) course must be set in the pilot’s HSI course window.

ON Indicates that the backcourse mode is selected. Is also a function ofG/A mode.

NOTEBC may also be used to fly outbound on an ILS frontcourse.

LVL When the LVL button is depressed (autopilot engaged or not), the(LEVEL) flight director will command wings level, and any previously selected

roll mode will be canceled. If a pitch mode happens to be engaged,pitch commands for that mode will not be affected; otherwise, thecommand bars will assume the existing pitch attitude.

ON Indicates the level mode is engaged.

NOTEDuring flight-director-only operation, selecting SPD,V/S, or ALT HLD without a prior roll mode selectionwill automatically engage the LVL mode.

SPD When selected, the flight director will command a pitch attitude(SPEED) that will maintain the airspeed existing at the time of mode selection.

Power must be set by the pilot.

IAS Illuminates when the SPD mode selector is first depressed. Theexisting IAS is maintained.

MACH Illuminates when the SPD mode selector is depressed a secondtime. The existing Mach number is maintained.

NOTEThe switch will cycle between IAS and MACH,always starting with IAS upon initial engagement.Therefore, to disengage the mode, another pitchmode must be engaged, or momentarily move eithercontrol wheel trim switch (without depressing armingbutton) in the noseup or nosedown direction. In theflight-director-only mode, SPD is disengaged withactivation of the pitch sync switch.

V/S When selected, the flight director commands a pitch attitude that(VERTICAL will maintain the existing vertical speed.SPEED)

ON Illuminated when V/S mode is selected.

NOTEBefore engaging this mode, maintain the desiredrate long enough(approximately 15 seconds) forvertical speed indicator lag to diminish.






G/S When selected, activates the flight director function that captures(GLIDE and tracks glide slope.SLOPE)

Functional only when the NAV 1 receiver is tuned to an ILS fre-quency, an active glide-slope signal is present, the G/S flag is out ofview, and the BC mode is not selected.

ARM Illuminates when the G/S mode is selected and the airplane is noton the glide-slope beam. Goes out when the airplane captures thebeam.

CAPT Illuminates when the airplane captures the glide-slope beam.

FNL (FINAL) illuminates during an ILS or a localizer approach when the LOC andG/S beam signals are being desensitized for close-in stability.

NOTEIf the radio altimeter signal is valid, the FNL light willilluminate at approximately 1,200 feet AGL. If theradio altimeter is not valid, the FNL mode will beactivated when passing over the outer marker. If theradio altimeter and outer marker are not valid,depressing the NAV 1 TEST button will activate theFNL mode. This should be accomplished at the finalapproach fix. The flaps must be down 3° or more toinitiate desensing (FNL) manually.

ALT HLD When selected, the flight director will command an airplane pitch(ALTITUDE attitude that will maintain the existing altitude. Vertical velocityHOLD) should be less than 1,000 ft/min.

ON Illuminates when ALT HLD is engaged.

ALT SEL When selected, the flight director will capture preselected(ALTITUDE altitudes.SELECT)

ARM Illuminates when ALT SEL is activated. The desired altitude is seton the altitude alerter and any pitch mode (except ALT HLD) may beused to attain that altitude. Upon nearing the selected altitude, theARM light goes out and any other pitch mode in use disengages.

CAPT Illuminates when an altitude interception begins. When the airplaneis within 20 feet of the selected altitude and vertical speed withinlimits, the ALT HLD mode engages, the ALT HLD ON lightilluminates, and the ALT SEL CAPT light extinguishes.

G/A (GO-AROUND) Flight-director-only mode, selected by depressing the GO-AROUNDbutton on the left thrust lever knob. Disengages autopilot (ifengaged), illuminates the G/A and LVL annunciators, and positionscommand bars to 9° pitch up, wings level.


Autopilot Disengagement

Whenever the autopi lot and/or rol l axesdisengage, the applicable PITCH and/or ROLLannunciators will extinguish and the autopilotdisengage tone will sound, as defined below:

• Either control wheel trim switch, witharming button depressed and moved inany of the four directions (NOSE UP,NOSE DN, LWD, or RWD), will disen-gage both autopilot axes.

• Either control wheel master switch(MSW), when depressed, will disengageboth autopilot axes and the yaw damper.

• The AUTO PILOT master switch, whenset to OFF, will disengage both autopi-lot axes.

• The PITCH TRIM selector switch, whenmoved to the OFF position, will disen-gage both autopilot axes, but only whenit attempts to trim the horizontal stabi-lizer and cannot (FC 200). On FC 530 air-planes, autopi lot disengagement isimmediate.

• With the PITCH TRIM selector switchin either the PRI or SEC position, movingthe pedestal NOSE DN–OFF–NOSE UPswitch to NOSE UP or NOSE DN willdisengage both autopilot axes.

• Individual axes may be disengaged bypulling the applicable axis AC or DCcircuit breakers (pilot’s AC and essen-tial buses).

NOTEOn the FC 530 AFCS, if the AC AFCSPITCH circuit breaker is out, the pullersystem is also rendered inoperative andairspeed must be limited to 0.74 M1.

• Depressing the pilot’s VG ERECT buttonor actuating the pilot’s L-R SLAVE switchwill disengage both autopilot axes.

• On the FC 530 AFCS, depressing the GO-AROUND button (left thrust lever knob,will disengage the autopilot and selectflight director G/A (go-around) and LVL

modes. This positions the command barsat a wings level 9° noseup pitch position.

NOTEOn SNs 35-002 through 35-009, and36-002 through 36-006, the G/A modeis coupled to the autopilot if engagedwhen power is advanced toapproximately 80% N1.

Servo Force Meters

Two servo force meters are located in the cen-ter of the control panel. The indicators providean indication of what autopilot servo forces arepresent when the autopilot is engaged. Theleft one indicates roll force and the right, pitchforce. If the force meter(s) is deflected, the ap-propriate axis should be trimmed to center themeter(s) prior to engaging the autopilot. Ifthe autopilot is engaged, and the meter(s) in-dicates a steady deflection, the autopilot shouldbe disengaged and the appropriate axis re-trimmed. Small deflections before and after en-gagement are normal.

Roll Monitors

The computer uses the roll rate gyro and thepilot’s vertical gyro to control the rate of roll andbank angle, respectively.

On FC 200 airplanes, excessive roll rate willdisengage the roll axis, sound the disengage tone,and extinguish the ROLL light.

On FC 530 airplanes, excessive roll rate or bankangle in excess of approximately 40° willdisengage both axes, sound the disengage tone,and extinguish the ROLL and PITCH lights.

Pitch Trim Monitor

The autopilot maintains pitch trim using theairplane’s secondary pitch trim system. Wheneverthe autopilot is engaged and the secondary trimruns in a direction opposite the elevator servoforce, a monitor will disengage both axes, soundthe disengage tone, and extinguish the ROLL andPITCH lights.




Out-of-trim Monitors (FC 530 AFCS Only)

With the autopilot engaged, the out-of-trimmonitors cause the applicable PITCH or ROLLannunciator to flash if an out-of-trim conditionexists to a degree that servo force is continuouslyapplied for more than approximately 20 seconds.The light continues to flash until either the trim isrestored or the axis is disengaged.

G-force Monitor (FC 530 AFCS Only)

G forces are sensed by the vertical accelerometerwith the autopilot engaged. The G-force monitorcauses the elevator to streamline whenever the Glevel reaches 1.6 G or 0.6 G. The pitch axis remainsengaged, but keeps the elevator streamlined.Previously engaged pitch modes also remain on.When the airplane is within the G limits, the pitchaxis resumes normal elevator inputs.

Autopilot/Stick Nudger/Pusher/Stick Puller Interface

If the autopilot is engaged and the stick nudger(FC 530 AFCS), pusher, or puller actuates, anyselected pitch mode disengages. The autopilotthen maintains a synchronous standby modeuntil the nudger, pusher, or puller releases. Uponthis release, the autopilot maintains the existingpitch attitude.

Altitude Alerter

The altitude alerter provides automatic visual andaural signals announcing approach to anddeparture from a selected altitude. The alerter is adirect-reading instrument with a five-digit display(Figure 16-14).

The altitude alerter located in the centerinstrument panel functions in conjunction with thepilot’s altimeter. An OFF flag adjacent to thealtitude display will be in view whenever power isnot available to the alerter. During flight, as theairplane passes within approximately 1,000 feet ofthe selected altitude, the amber ALT annunciatorson the pilot’s and copilot’s altimeters willilluminate and an alert bell will sound. The point

at which the approach to the preselected altitude isannunciated depends upon airplane vertical speed.The annunciators will extinguish when theairplane is within 300 feet of the preselectedaltitude. Should the altitude subsequently deviatemore than ±300 feet from the selected altitude, theALT annunciators will illuminate and the alert bellwill sound.

The altitude alerter is also used to program theflight director altitude select (ALT SEL) mode onthe FC 530 AFCS.




Figure 16-14. Altitude Display

COMMUNICATIONSYSTEM

STATIC DISCHARGE WICKSA static electrical charge, commonly referredto as “P static” (precipitation static), builds upon the surface of an airplane while in flight andcauses interference in radio and avionicsequipment operation. The charge may be dan-gerous to persons disembarking after landingas well as to persons performing maintenanceon the airplane. The static wicks are installedon all trailing edges (Figure 16-15) to dissi-pate static electricity.

RVSM SYSTEM

GENERALIn the late 1950’s, vertical separation for air-craft in upper airspace was 1,000 feet. How-ever, in the early 1960’s, as more and moreaircraft were entering the airspace above29,000 feet, a determination was made to in-crease the vertical separation above 29,000feet to 2,000 feet. Starting in the late 1970’s,a series of studies was conducted to deter-mine the feasibility of reducing the current2,000 foot vertical separation between FL290and FL410 to 1,000 feet. These studies con-tinued through the late 1980’s. The studiesconcluded that the reduction to 1,000 footseparation was feasible, providing the aircraftwere equipped with an altimetry system withincreased accuracy, which would also pro-duce increased accuracy in the altitude re-porting system.

The first implementation of Reduced VerticalSeparation Minimum (RVSM) began in theNorth Atlantic Region in March, 1997. Sincethen, it has successfully expanded to includethe South Atlantic, the Pacific, the South ChinaSea, the West Atlantic Route and the conti-nental airspace of Australia.

In September, 2004 senior FAA Managers metwith their counterparts from Canada and Mex-ico. After reviewing significant implementa-tion factors, made the decision to proceed withRVSM implementation in North America. TheDomestic Reduced Vertical Separation Mini-mum (D-RVSM) implementation date was Jan-uary 20, 2005 for altitudes between FL290 toFL410 (inclusive) in the airspace of the lower48 States, Alaska, Atlantic High and Gulf ofMexico High Offshore Airspace, and the SanJuan Flight Information Region. Also includedwere Southern Canadian Domestic Airspaceand the Airspace in Mexico. RVSM had al-ready been implemented in Northern Cana-dian Domestic Airspace in 2002.




Figure 16-15. Static Wicks (Typical)

Af te r J anua ry 20 , 2005 , a i r c r a f t NOTEQUIPPED with special RVSM equipmentmust be granted special permission to transi-tion through block altitudes FL290 to FL410,or maintain an altitude of FL290 or lower.

All Learjet models 35-35A/36-36A are eligi-ble for RVSM modification. However, in somecases specific aircraft modifications must havebeen already successfully completed and doc-umented in the aircraft log book, or compliedwith concurrent with the RVSM modification.A maintenance log check must be accom-plished to insure all necessary modificationshave been completed or scheduled.

There are currently two Supplemental TypeCertificate (STC) holders that can accomplishthe necessary aircraft modifications for RVSMfor the Learjet 35-35A/36-36A group. One isAero Mech, Inc. (AMI) under their STC Num-ber s ST 00952SE, ST 00952SE-D, ST01199NY and ST 01199NY-D. To simplify infuture discussion this will be referred to as theLearjet RVSM Installation. The other is WestStar/Honeywell under their STC Numbers ST01524LA, ST 01525LA and ST 01526LA.Again, for simplicity this will be referred to asthe West Star RVSM Installation.

Each one accomplishes the same end task, butgets there differently. The Rosemount pitot-static probe system is installed in the affectedmodel in accordance with STC ST 00321WIor ST 00321WI-D. Limitations and other pro-cedures have also changed in some areas.Therefore, ensure you have the proper Air-plane Flight Manual Supplements in your Air-plane Flight Manual and you MUST refer tothem for the proper LIMITATIONS, NOR-MAL, EMERGENCY and ABNORMAL pro-cedures for operating your equipment.

With the implementation of D-RVSM, the fol-lowing are areas of significant importance andthese checks should be closely monitored:

1. Altimeter Checks—Prior to takeoff forflights planned into RVSM airspace, pri-mary altimeters must be within 75 feetof a known elevation. While withinRVSM airspace, primary altimeters mustbe within 200 feet of each other.

2. Altitude Awareness—To preclude er-rors in hearing clearances and/or incor-rectly setting the altitude pre-select, thefollowing technique/SOP is suggested:

a. Pilot flying is manually flying the air-craft—Pilot monitoring sets altitudepre-selector, both pilot point to the al-titude set in the altitude pre-selector,and both verbally state that altitude.

b. Pilot flying is flying the aircraft on au-topilot—Pilot flying sets the altitudepre-selector, both pilots point to the al-titude set in the altitude pre-selector,and both verbally state that altitude.

3. Climbs and Descents—To preclude un-warranted TCAS TAs or RAs, limit climband descent rates to 1,500 feet per minuteor less during the last 1,000 feet of an al-titude change (AIM 4-4-9[d]).

4. Respond immediately and appropriatelyto any TCAS RAs.




LEARJET RVSM INSTALLATION

Rosemount Pitot-Static ProbesFor FC 200 autopilot equipped aircraft, the tra-ditional pitot tubes and static ports are re-moved and replaced by Rosemount pitot staticprobe (Figure 16-16). Earlier FC 200 autopi-lot aircraft that have already been modifiedwith the Rosemount pitot static probe systemand FC 530 autopilot are described under theFC 530 modification (see AFMS W1266). TheALTERNATE STATIC SOURCE valve locatedat the bottom of the left side of the instrumentpanel has been removed. The pressurizationstatic port installation has not changed.

The Rosemount pitot static probes are mountedon each side of the nose section and provideboth pitot and static pressure to designatedsystems. The probes also contain heating el-ements for anti-icing and are controlled bythe L or R PITOT HEAT switches (refer toChapter 10, Ice and Rain Protection). Fourdrain valves are located near the end of the nosegear doors and are installed at the system lowpoints to drain moisture.

The pitot pressure is sensed separately fromthe front of each probe. The left pitot pressureis plumbed to the captain’s airspeed indicator.Pitot pressure from the right pitot probe isplumbed to all other systems that need pitot

pressure. Static pressure is sensed by twosources on each probe; static 1 (S1) and static2 (S2). Static 1 on the left probe is cross con-nected to static 2 on the right probe and static1 on the right probe is cross connected to static2 on the left probe.

Four solenoid-operated isolation shutoff valvesenable the pilot to select the source of staticpressure. The source of static pressure is con-trolled by a static source/static port switch(Figure 16-17) located on the top of the throt-tle quadrant or on the anti-ice control panel.

Probe sensing is extremely accurate and onlya minor correction must be made in the airdata display unit (ADDU)–air data computer(ADC). There are altitude correction chartsthat must be used if a malfunction occurs ei-ther in the ADDU system or the pitot static sys-tem. These correction factors are included inthe applicable Airplane Flight Manual Sup-plement.

There is a critical nose section area adjacentto each pitot-static probe (left and right) thatmust be checked on every preflight beforeflight into RVSM airspace (Figure 16-18).The inspection area is also identified by four




Figure 16-17. Static Source/StaticPort Switch

Figure 16-16. Rosemount Pitotand Static Probe

90° angle marks painted on the fuselage ineach corner. The preflight walkaround checkis to ensure that no obvious skin damage or de-formation has occurred in that area. Also,check the pitot static probe heads for any de-formation or obstruction around the inlet orstatic ports.

Static Source/Static Port SwitchA static source/static port switch is installedeither on the pilot anti-ice switch panel, or onthe forward part of the throttle quadrant (Fig-ure 16-17). The switch is a three position tog-gle switch and the positions are identified asBOTH (center), L (left), or R (right). Theswitch is used to select the static source to beused. For example, when BOTH is selected,both S1 and S2 on both sides are being usedand are cross connected. When the L positionis selected, the RIGHT probe STATIC is iso-lated and STATIC pressure from the LEFTprobe only is being used. When the R positionis selected, STATIC pressure from the RIGHTside is only being used and the LEFT probeSTATIC is isolated. Normally, both systems areoperative so the switch remains in the BOTHposition. Whatever switch position is selecteddetermines the static source that is used. The

BOTH position is required for normal flightinto RVSM airspace. However, if the systemmalfunctions and either the L or R position isused in RVSM airspace, refer to the AFM Sup-plement for procedures/guidance.

IS&S Altimeter/ADC System

General

On the FC 200 and FC 530 autopilot aircraft, thepilot (servo pneumatic or pneumatic) and copi-lot (pneumatic) altimeters are replaced with theIS&S combination self sensing alt imeter(ADDU–Air Data Display Unit/ADC–Air DataComputer) (Figure 16-19).

On FC 200 autopilot aircraft, an analog in-terface unit (AIU) is installed and convertsdigital data from the altimeters to analog sig-nals that interface with the existing FC 200 au-topilot. The AIU also provides outputs forVMO/MMO overspeed warning, gear horn warn-ings, aircraft speed data for the mach trimcomputer and air data information to otheroptional aircraft systems (e.g. long range nav,SAT/TAS indicator).

On FC 530 autopilot aircraft the AIU convertsdigital data from the altimeters to analog sig-nals to interface with the existing FC 530 au-topilot and to provide air data information toother optional aircraft systems (e.g. long rangenav, SAT/TAS indicator).




Figure 16-18. Right SidePitot-Static Probe

Figure 16-19. Pilot andCopilot Altimeters

On both the FC 200 and FC 530 autopilot air-craft, the altitude alerter panel was removed andreplaced with an air data switch panel (Figure16-20). This panel consists of green ADC1and ADC2 pushbutton switchlights and a redAIU FAIL annunciator light. The altime-ter/ADC combines the function of the basic al-timeter with those of the traditional altitudealerter and is also a self-sensing unit with pitotand static connections. A new standby altimeteris installed and plumbed to the copilot staticsource. The installation utilizes three inde-pendent sources of electrical power to the al-timeters, one of which is the emergency battery.

Altimeter OperationThe Learjet RVSM installation chose to installIS&S altimeters/ADDU as shown in Figure 16-19. They provide altitude indications and alsogenerate and indicate additional data. WhenADC1 (pilot) or ADC2 (copilot) switch is de-pressed on the air data switch panel, the switchpanel light will illuminate and the selected al-timeter/ADDU will have an active master A il-luminated. This now becomes the masteraltimeter/ADDU and is used for transponder,altitude pre-select, altitude alerting, air datainput to the AIU, and other auxiliary outputs.

The altimeter that does not have the A illu-minated is referred to as the slave unit.

NOTEThe autopilot must be disengagedwhen swi tch ing f rom one ADCsource to another.

To toggle between IN.HG or hPa, press theBARO select knob located to the lower right onthe altimeter. If the BARO knob is held de-pressed for longer than four seconds, unit se-lection mode is entered and each additionalpress of the knob for four seconds will togglethe altimeter display between IN.HG and hPa.

If the BARO select knob is depressed andheld for 8 seconds or longer, the altitude unitdisplay will toggle between feet and meters.

Barometric pressure is set by rotating theBARO select knob. Momentarily depressing theBARO knob for less than 2 seconds will set29.92 IN.HG or 1013 hPa. Note that the mas-ter A and the slave baro set knobs are totallyindependent and different units (IN.HG or hPa)and different baro settings are possible.

Additional information may be displayed onthe alt imeter bezel (e.g. ALT and/or DHlights). Failure of either altimeter is indicatedby a blank display or the word OFF displayed.There is a COM and STBY light indication onthe face of each altimeter (ADDU). An illu-minated COM indication indicates that thedata bus communication between the pilot andcopilot ADDU is lost. An illuminated STBYindication indicates SSEC corrections are notbeing applied. Should the COM or STBY lightsdisplay or AIU FAILURE indicator illumi-nate, consult your appropriate AFM Supple-ment as the ABNORMAL procedures aredifferent between FC 200 and FC 530 aircraft.If the AIU FAIL light, located on the air dataswitch panel illuminates, select the other ADCon the switch panel and refer to the AFM Sup-plement ABNORMAL procedures.




Figure 16-20. Air Data Switch Panel

Power Source/FailureElectrical power for the pilot altimeter is sup-plied by the ALTM or PRI ALTM circuitbreaker located on the left essential bus (LESS BUS). It may also be powered by theemergency battery through the EMER ALTMcircuit breaker located on the left circuitbreaker panel as illustrated in Figure 16-21.

If normal electrical power is lost to the pilotaltimeter (ADDU) and it is being powered bythe emergency battery, the pilot ADDU willfunction using the emergency battery power,but the PWR and COM indication will illu-minate on the pilot ADDU (altimeter) display(Figure 16-22).

On FC 200 aircraft, a pilot altimeter (ADDU)emergency lighting (PLT ALTM EMER LTG)switch may be installed and is located on thepilot side panel. If the switch is installed andnormal electrical power is lost to the pilot al-timeter, the ADDU back lighting will remainON and the pilot may select desired intensityof the digital display by using this switch. Ifthis switch is not installed, the ADDU backlighting will remain ON if the pilot INSTRPNL dimmer knob (pilot side panel) is turnedON (out of detent) and the altitude displaywill be dimmed. For daylight conditions, theINSTR PNL dimmer knob should be turnedOFF (in the OFF detent position) which willcause the back lighting to be off and the alti-tude display to be bright.




RESS

BBUS

BATTERY OUTPUTVDC

28 VDCINPUT

INVERTER

EMERGENCYBATTERY

EMERGENCYBAT CB

GEAR

FLAPS

STBY ALT VIBRATOR

EMER ALTM (PILOT ADDU ALTIMETER)

PULL

C

A G E

TO

CLI MB

DI VE

STBY

EMERBAT

OFF

LIGHTINGSTBY ALT

115 VAC OUTPUT

VAC OUTPUT

4.6 VAC OUTPUT

EMRPWR

ATTITUDE GYRO

Figure 16-21. Emergency Battery Power System

For FC 530 aircraft, this switch is not in-stalled and the ADDU back lighting will re-main ON if the pilot INSTR PNL (pilot sidepanel) dimmer knob is turned ON (out of de-tent) and the altitude display will be dimmed.For daylight conditions, the INSTR PNL dim-mer knob should be turned OFF (in the OFFdetent position) which will cause the backlighting to be off and the altitude display to bebright. If normal electrical power is lost to thecopilot altimeter, the copilot ADDU will beinoperative.

Altitude Alerter OperationSelect the desired alerter altitude by rotatingthe ALT SEL knob on the face of the altime-ter (ADDU) (see Figure 16-19). Clockwiserotation causes the selected altitude to in-crease and counter-clockwise to decrease.Knob sensitivity is 100 feet per detent (30meters in metric mode). As long as the sameunits (feet or meters) are selected, rotatingthe ALT SEL knob on the master ADDU (A il-luminated) will change the selected altitude onboth the master and the slave ADDU. If dif-ferent units are selected, the display on theslave unit will blank and its ALT SEL knobwill be disabled. Momentarily depressing theALT SEL knob will extinguish the altitudealarms until the appropriate approach condi-tions are met again.




PRI ALTML ESS BUS A

SEC ALTMR ESS B BUS

AIU PWR 1L ESS B BUS

AIU PWR 2L ESS B BUS

AIU REF26 VAC L AC BUS

1

1

ANALOGINTERFACE UNIT

(AIU)

OUTPUT ANALOGSIGNALS FROM AIUUSED BY:• GEAR WARNING ALTITUDE (FC 200)• AUTOPILOT (FC 200 AND FC 530)• VMO/MMO OVERSPEED WARNING (FC 200)• MACH TRIM (FC 200)• LONG RANGE NAV (FC 200 AND FC 530)• SAT/TAS (FC 200 AND FC 530)

1

1

1

EMER ALTML ESS BUS

1

STBY ALTM R ESS BUS B

STANDBYALTIMETER(VIBRATOR)

1

(EMERGENCY BATTERY ON)

INSTRUMENTS LIGHT FROM EMERGENCY BATTERY

ADC1

ADC2AIU FAIL

ADC SWITCHAND ANNUNCIATOR

PANEL

PILOT ALTIMETER

COPILOT ALTIMETER

Figure 16-22. Learjet Electrical Diagram for Altimeter/ADDU and AIU

Altitude ReportingAltitude reporting data may be supplied fromeither air data display unit (ADDU). Select-ing ADC-1 on the air data switch panel (seeFigure 16-20) provides altitude informationfrom the pilot ADDU for either transponder.Selecting ADC-2 on the switch panel providesaltitude information from the copilot ADDUfor either transponder.

The TFR 1-2 switch (if installed) is located onthe transponder control panel. Selecting TFR-1 transmits altitude information from the LEFTtransponder supplied by the selected ADDU.Selecting TFR-2 transmits altitude informa-tion from the RIGHT transponder supplied bythe selected ADDU.

System Checks/Tests

System Operational Check

An operational check of the altimeter/ADCsystem is outlined in the appropriate AirplaneFlight Manual Supplement. Refer to your sup-plement for information on how and when toperform this system operation check and forproper display information during the check.

Initiated Built-In Test (BIT)

Both the pilot and copilot ADDU/ADC con-tain a built-in test feature that may be per-formed as desired by the crew. The aircraftmust be below 40 knots to activate this test.The test is initiated by depressing the recessedTEST button located on the lower left front sideof the ADDU bezel (see Figure 16-19). Whendepressed, the ADDU and the AIU will begintheir BIT tests. Your AFM Supplement de-scribes exactly how to perform this test andwhat indications to look for. If an error is de-tected during the test, the AIU fail light on theair data switch panel remains illuminated.

Standby AltimeterThe standby altimeter is a pure static altime-ter and is plumbed to the copilot static system(Figure 16-23). Electrical power for the al-timeter lighting and vibrator is supplied fromthe aircraft emergency battery when the switchis placed in the ON position. The standby al-timeter is not powered when the EMER BATswitch is in STBY. There is an OFF flag on theleft lower corner to indicate that the vibratoris not operating.

Altitude Position CorrectionChartsThere are altitude position correction chartsin the Airplane Flight Manual and also in theAFM Supplements. For the FC 200 autopilotaircraft, the altitude position correction chartssupplied with the FC 200 autopilot retrofitwith Rosemount pitot static probes supple-ment (AFMS W1266) must be applied to thestandby altimeter. They also must be appliedto the indicated altitude when the STBY in-dicator light is illuminated on a IS&S ADDU(primary altimeter). The correction charts sup-plied with AFMS W1483 are used for altitudecorrection on the primary altimeters when theSTBY indicator light is not illuminated.

For FC530 autopilot aircraft, the altitude po-




Figure 16-23. Standby Altimeter

sition correction charts supplied with the Air-plane Flight Manual must be applied to thestandby altimeter. The AFM corrections mustalso be applied to the indicated altitude whenthe STBY indicator light is illuminated on theIS&S ADDU (primary altimeters). The cor-rection charts supplied with AFMS W1484are used for altitude correction on the primaryaltimeter (ADDU) when the STBY indicatorlight is not illuminated.

WEST STAR RVSMINSTALLATION

Rosemount Pitot-Static ProbesFor FC 200 autopilot equipped aircraft, the tra-ditional pitot tubes and static ports are re-moved and replaced with Rosemount pitot andstatic probes (Figure 16-24).

Earlier FC 200 aircraft that have already beenmodified with the Rosemount pitot static probesystem and the FC 530 autopilot aircraft aredescribed under FC 530 modification (SeeAFMS W1266). The alternate static sourcevalve, located at the bottom of the left side ofinstrument panel, has been removed. The pres-surization static port has not changed (SeeChapter 12, Pressurization). The shoulder staticports remain installed and will be used for thestandby altimeter and an alternate static sourcefor the airspeed indicators (Figure 16-25).

The Rosemount pitot static probes are mountedon each side of the nose section and provide bothpitot and static pressures to designated sys-tems. The probes also contain heating elementsfor anti-icing and are controlled by the L andR PITOT HEAT switches (refer to Chapter 10,Ice and Rain Protection). Drain valves are lo-cated near the end of the nose gear doors andare installed at the system low points to drainmoisture during preflight. The left pitot pres-sure is plumbed to the captain airspeed indicatorand the Honeywell AZ-252 air data computer(ADC). The right pitot pressure is plumbed toall other systems that use pitot pressure. Staticpressure is sensed by two sources on each probe:static 1 (S1) and static 2 (S2). Static 1 on theleft probe is cross-connected to static 2 on theright probe and static 1 on the right probe iscross-connected to static 2 from the left probe.Either probe can furnish static pressure to allsystems except the standby altimeter and the al-ternate static pressure to the airspeed indica-tors (Figures 16-26 and 16-27).

Probe sensing is extremely accurate and onlyminor corrections must be made. These cor-rection factors are included in your applicableAFM Supplement. There is a critical nose sec-tion area adjacent to each pitot static probe(left and right) that must be checked on everypreflight before flight into RVSM airspace.The preflight check is to ensure that no dents,paint chips or distortions are present in theprobe area that would disrupt or distort airflowin that area. Also, check the pitot static headsfor any deformation or obstruction around theinlet or static ports. FC 530 autopilot West Starmodified aircraft will utilize the already existingstatic port/source switch.




Figure 16-24. Rosemount Pitot andStatic Probe

Figure 16-25. Shoulder Static Port

Revision .03

16-31FO

R T

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ININ

G P

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ON

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FlightSafety

international

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T 3

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PIL

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MA

NU

AL

EXISTING NOSELOWER SKIN

LH ROSEMONTP/S PROBE

EXISTING COPILOTSTALL WARNING

22,500' SWITCH

PILOT AIRSPEEDSTATIC VALVE

PITOT

F.S.160.77FR 5

PILOT—S1COPILOT—S2 STATIC

COPILOT—S1PILOT—S2 STATIC

REFERENCE DESTINATIONS

SHOULDER STATIC

SHOULDERSTATIC PORTS

NEW AZ-252AIR DATA COMPUTER

EXISTING DRAINS5 PLACES

RH ROSEMONTP/S PROBE

COPILOT AIRSPEEDSTATIC VALVE

EXISTING PILOT STALLWARNING 22,500' SWITCH

PILOTAIRSPEEDINDICATOR

PILOTVERTICALSPEED IND

COPILOTAIRSPEEDINDICATOR

COPILOTAM-250

ALTIMETER

COPILOTVERTICALSPEED IND

STBYALTIMETER

Figure 16-26. Pitot Static System Schematic for AFCS FC 200 Aircraft

Revision .03

16-32FO

R T

RA

ININ

G P

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ON

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FlightSafety

international

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AR

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LH ROSEMONTP/S PROBE

ISOLATIONVALVES (REF) ISOLATION

VALVES (REF)

PITOT

F.S.160.77FR 5

PILOT—S1COPILOT—S2 STATIC

COPILOT—S1PILOT—S2 STATIC

REFERENCE DESTINATIONS

NEW AZ-252AIR DATA COMPUTER

DRAINS4 PLACES

RH ROSEMONTP/S PROBE

PILOTAIRSPEEDINDICATOR

PILOTVERTICALSPEED IND

COPILOTAIRSPEEDINDICATOR

COPILOTAM-250

ALTIMETER

COPILOTVERTICALSPEED IND

STBYALTIMETER

MACH SWITCH

Figure 16-27. Pitot Static System Schematic for AFCS FC 530 Aircraft

Static Port/Source SwitchOn FC 530 autopilot aircraft, including earlierFC 200 aircraft that have been modified withthe Rosemount pitot static probe system, astatic port/source switch is installed (Figure16-28). This switch is installed either duringproduc t ion o r i s ins ta l l ed by STCs: ST00321WI, or ST 00321WI-D-FC 200 autopilotretrofit with Rosemount pitot-static probes.The function of this switch does not change withthe West Star installation.

The L, R or BOTH position of this switch isthe static source being used. It must be in theBOTH position for flight into RVSM airspace.However, if the system malfunctions and ei-ther the L or R position is selected while inRVSM airspace, refer to your AFM Supple-ment for the procedures that must be followed.The FC 200 autopilot aircraft modified bythe West Star installation does not have thisswitch installed.

Air Data ComputerThe West Star RVSM installation chose to useHoneywell equipment. The Honeywell AZ-252 advanced air data computer (ADC) sys-tem consists of a RVSM capable advanceddigital air data computer with analog outputsfor both the FC 200 and FC 530 autopilot sys-tems (Figure 16-29).

The ADC receives total pressure input fromthe left pitot probe and static pressure from S1(left probe) and S2 (right probe) static port.The ADC receives total air temperature (TAT)from the TAT probe.

The ADC provides outputs to the pilot altime-ter, altitude alerter, autopilot, Mach trim, air-speed warning, landing gear warnings, ATC#1 and the following optional equipment:SAT/TAS/TAT indicator, IDC/Kolsman verti-cal speed indicator and long range navigationsystem (FMS). An ADC BITE (built-in-test) isinitiated upon system power up. The ADC hasno failure annunciations. Failure is indicatedby fault indications in the associated indicatorsand controls.




Figure 16-28. Static Port/Source Switch

Figure 16-29. West Star Air DataComputer (ADC)

Refer to Tables 16-3 and 16-4 for a partial listof cockpit indications should the ADC fail.The AZ-252 air data computer requires 115VAC electrical power and it is supplied from theLeft AC Bus.

See Figure 16-30 for the West Star AvionicsBlock Diagram.




ADC FAILURE

EQUIPMENT INDICATION REMARKS

PILOT MACH/AIRSPEED

PILOT ALTIMETER

PILOT VERTICAL SPEED OR**PILOT IDC VSI

ALTITUDE ALERTER

AUTOPILOT

MACH TRIM

ATC TRANSPONDER

**SAT/TAS/TAT

**FMS/GPS ETC.

– –

DASHES IN ALL LCD DISPLAYFIELDS, POINTER PARKS AT 8

– – ORVSI POINTER PARKED AT 0

0 DISPLAYED

DASHES DISPLAYED

AUTOPILOT VERTICAL MODES WILLENGAGE BUT ARE UNRELIABLE

MACH TRIM ILLUMINATED

LOSS OF ALTITUDE REPORTING

ADC FAIL MESSAGE

M/ASI IS OPERATIVE, AURAL OVER-SPEED WARNING INOPERATIVE

ALTIMETER IS INOPERATIVE, USESTANDBY ALTIMETER OR CROSS SIDE ALTIMETER

PNEUMATIC VSI IS OPERATIVE ORIDC VSI IS INOPERATIVE

INOPERATIVE

INOPERATIVE

DO NOT USE AUTOPILOT VERTICALMODES

LIMIT MACH NO. TO ≤MO.74

SELECT ATC 2

LOSS OF ADC INPUTS, USEMANUAL INPUTS IF APPLICABLE

**OPTIONAL EQUIPMENT WHICH MAY OR MAY NOT BE INSTALLED, SEE THE AIRCRAFT EQUIPMENT LIST

Table 16-3. WEST STAR ADC FAILURE INDICATIONS CHART FOR FC 200 AIRCRAFT




ADC FAILURE

EQUIPMENT INDICATION REMARKS

PILOT MACH/AIRSPEED

PILOT ALTIMETER

PILOT VERTICAL SPEED OR**PILOT IDC VSI

ALTITUDE ALERTER

AUTOPILOT

MACH TRIM

ATC TRANSPONDER

**SAT/TAS/TAT

**FMS/GPS ETC.

– –

DASHES IN ALL LCD DISPLAYFIELDS, POINTER PARKS AT 8

– – ORVSI POINTER PARKED AT 0

0 DISPLAYED

DASHES DISPLAYED

MACH TRIM ILLUMINATED

LOSS OF ALTITUDE REPORTING

ADC FAIL MESSAGE

M/ASI IS OPERATIVE, THE 300 KIAS AURAL OVERSPEED WARNING INOPERATIVE

ALTIMETER IS INOPERATIVE, USESTANDBY ALTIMETER OR CROSS SIDE ALTIMETER

PNEUMATIC VSI IS OPERATIVE ORIDC VSI IS INOPERATIVE

INOPERATIVE

INOPERATIVE

VERTICAL MODES ARE INOPERATIVEVERTICAL MODES ARE CANCELED

LIMIT MACH NO. TO ≤MO.74

SELECT ATC 2 ORSELECT ENCODE ALT-XFER

LOSS OF ADC INPUTS, USEMANUAL INPUTS IF APPLICABLE

**OPTIONAL EQUIPMENT WHICH MAY OR MAY NOT BE INSTALLED, SEE THE AIRCRAFT EQUIPMENT LIST

Table 16-4. WEST STAR ADC FAILURE INDICATIONS CHART FOR FC 530 AIRCRAFT

Pilot Altimeter—BA-250Barometric AltimeterThe BA-250 barometric altimeter, installed atthe pilot position, incorporates an analog/LCDdigital display of baro-corrected pressure alti-tude, baro-correction displays and an amberaltitude alert light. It is both English and Met-ric capable (Figure 16-31).

Barometric pressure is set manually with theBARO knob and is displayed in inches of mer-cury and hectoPascals (hPa) on the baro-cor-rected displays. A STD pushbutton selectsstandard barometric pressure (29.92 or 1013hPa). When ADC inputs are lost, dashes are dis-played in all LCD display fields and the pointermoves to the 8 on the numerical dial. The al-timeter requires 28-VDC electrical power andit is supplied by the ALTM 1 circuit breaker onthe left main bus.




LEAR 35/36 FC-200/FC-530 AFCS

MACH TRIMLANDING GEAR WARNING

OTHER EQUIPMENT

AIR DATA SIGNALS AIR DATA LOGIC

LEARJET 35/36AZ-252 AIR DATA COMPUTER

ALTITUDE ALERT LIGHT

PS

PT ATC #2

IDC VSI*

*OPTIONAL

DISPLAY ANDSELECT DATA

BAROSET A429

PS

PT

BA-250BAROMETRIC

INDICATOR

AL-800ALTITUDEALERTER

AM-250BAROMETRIC

INDICATOR

ALTITUDEALERT HORN

ATC #1

Figure 16-30. West Star Learjet 35/36 RVSM Avionics Block Diagram

Figure 16-31. West Star Pilot

Copilot Altimeter—AM-250Barometric AltimeterThe AM-250 barometric altimeter, installed inthe copilot position, is a fully RVSM capablealtimeter with an integrated air data computer(Figure 16-32). It is a self-contained unit andis not connected to the AZ-252 air data com-puter. It incorporates an analog/LDC display ofbaro-corrected pressure altitude, baro-correcteddisplays and an amber altitude alert light.

Barometric pressure is manually set with theBARO knob and displayed in inches of mercuryand hectoPascals on the baro-correction dis-plays. A STD pushbutton selects standard baro-metric pressure. It also provides an output toATC #2. When silicon pressure sensor inputsare lost, dashes are displayed in all LCD fieldsand the pointer moves to the 8 on the numeri-cal dial. This altimeter requires 28-VDC elec-trical power and it is supplied by the ALTM 2circuit breaker on the right main bus.

AL-800 Altitude AlerterThe AL-800 altitude alerter system providesboth visual and aural signals for altitude aware-ness (Figure 16-33). The desired altitude is se-lected by slewing the displayed altitude to thedesired value. During flight, when approach-ing the preselected altitude, at 1,000 feet priorto reaching that altitude, the amber altitudealert light in each altimeter is illuminated andan aural alert is sounded. The altitude alertlight remains illuminated until the aircraft iswithin 200 feet of the selected altitude whereit extinguishes.

If the aircraft should subsequently deviate fromthe selected altitude by 200 feet, the altitudealert light illuminates and the aural warning willsound again. The altitude alert light remains il-luminated until the aircraft returns to within 200feet of the selected altitude, or a new altitudeis selected. An 0 will be displayed on the alti-tude alerter display when ADC inputs are in-val id . The al t i tude aler ter uses 28-VDCelectrical power and it is supplied by the ALTALERT circuit breaker on the left main bus.




Figure 16-32. West Star CopilotAltimeter

Figure 16-33. Altitude Alerter




Standby AltimeterThe standby altimeter is a pure static altimeter(Figure 16-34). On the FC 200 autopilot aircraftit is plumbed to the shoulder static ports. Thesestatic ports are heated anytime there is electri-cal power on the aircraft. On the FC 530 au-topilot aircraft, the standby altimeter is plumbedto the copilot Rosemount probe static system.There is a vibrator installed in the standby al-timeter and an OFF flag in the upper left cor-ner to indicate that the vibrator is not operating.Electrical power for the vibrator is suppliedby the STANDBY ALT circuit breaker locatedon the Right Main Bus. See Figure 16-30 forthe West Star RVSM installation avionics blockdiagram.

Airspeed Static ValvesThere are two airspeed static valves installedbelow the instrument panel on each side (Fig-ure 16-35). These manual valves are providedto supply an alternate static source to the re-spective airspeed indicator. The valves havetwo positions—NORMAL and ALTERNATE.When NORMAL is selected the respective air-speed indicator receives static pressure fromthe normal Rosemount static source. When AL-TERNATE is selected, the valve blocks the nor-mal static pressure and connects that airspeedindicator to the shoulder port static source.Both valves must be in the NORMAL positionfor flight into RVSM airspace. If ALTERNATEis ever selected while in RVSM airspace, con-sult your AFM Supplement for corrective action.

Altitude Position and AirspeedCorrection ChartsThe new Rosemount pitot static probe instal-lation changes the static source position errorfor the basic aircraft. New charts are includedthe the AFM Supplement, Document Number30A04002, and have been developed from flighttest calibrations. The chart numbering systemin the supplement matches the basic aircraftAFM to the maximum extent possible.

The new charts include aircraft weights up to19,600 lbs to accommodate the increased grossweights that may be applicable to some Lear-jet 35/35A and 36/35A aircraft that have beenaltered by Avcon Division gross weight in-crease modifications.

The pilot and copilot altimeters are electrical,with the pilot BA-250 altitude display beingdriven by the AZ-252 air data computer and thecopilot having an AM-250 barometric altime-ter. The AZ-252 Air Data Computer and thecopilot AM-250 barometric altimeter have staticsource correction curves incorporated into thedisplay, so the pilot and copilot altimeters havenegligible errors in cruise flight.

The standby altimeter is connected to the shoul-der ports and has a static source error. Whenusing the standby altimeter, the static sourcecorrection factor must be applied to obtain theproper indication. When an airspeed staticsource valve, located under the instrument

Figure 16-34. Standby Altimeter

Figure 16-35. RH Airspeed Static Valve

panel, is selected to ALTERNATE, it appliesshoulder port static pressure to the applicableairspeed indicator. Airspeed indicator and Machposition correction chart valves must be ap-plied. These correction charts are located in theAFM Supplement. A cross-reference betweenFigures in the AFM Supplement and the Lear-jet AFMs (AFM-019—Model 35/36) and AFM-102—Model 35A/36A with FC200 Autopilot) ispresented in the AFM Supplement. In somecases, charts in the West Star Supplement arenew and did not exist in the Learjet AFM.







NAVIGATION SYSTEM

1a. FC 200 AFCS—The static ports for flightinstrument operation are located:A. In the unpressurized nose sectionB. On the top and bottom of the pitot-

static headsC. Flush mounted on the left and right

sides of the fuselage nose sectionD. On both sides of the aft fuselage

1b. FC 530 AFCS—The static ports for flightinstrument operation are located:A. In the unpressurized nose sectionB. In the pitot-static headsC. Flush mounted on the left and right

sides of the nose sectionD. On both sides of the aft fuselage

2a. On Learjet modified FC 200 AFCS—Thepilot controls the static pressure sourcefor the pilot’s flight instrument opera-tion:A. Electrically with the STATIC PORT

switchB. Mechanically with the STATIC PORT

switchC. Electrically with the ALTERNATE

STATIC SOURCE switchD. Mechanically with the ALTERNATE

STATIC SOURCE valve lever

2b. FC 530 AFCS—The pilot controls thestatic pressure source for the pilot’s flightinstrument operation:A. Electrically with the STATIC PORT

switch B. Mechanically with the STATIC PORT

switchC. Electrically with the ALTERNATE

STATIC SOURCE switchD. Mechanically with the ALTERNATE

STATIC SOURCE switch

3a. FC 200 AFCS—The air data sensor re-ceives pitot information from:A. The left pitot headB. The right pitot headC. Both pitot-static headsD. The right pitot-static head

3b. FC 530 AFCS—The air data unit receivespitot information from:A. The left pitot headB. The right pitot headC. Both pitot-static headsD. The right pitot-static head

4a. FC 200 AFCS—The air data sensor re-ceives static information from:A. The shoulder static air portsB. The pressurization module static air

portC. The right pitot-static headD. Both pitot-static heads

4b. FC 530 AFCS—The air data unit receivesstatic information from:A. The shoulder static air portsB. The pressurization module static air

portC. The right pitot-static headD. Both pitot-static heads with static

source switch in BOTH

QUESTIONS

AUTOFLIGHT SYSTEM

5. During flight-director-only operation,depressing the pilot’s SYNC switch:A. Disengages G/S ARM and ALT SEL

ARM (FC 530)B. Inhibits the roll and pitch axesC. Disengages any pitch mode except

G /S ARM and ALT SEL ARM (FC 530)

D. Cages the ADI to airplane centerlinereference

6. The ADIs and HSIs are energized when:A. An inverter is turned on.B. The AUTO PILOT master switch is

positioned to ON.C. The TEST switch is depressed.D. The VG ERECT switch is depressed.

7. To control the airplane in the pitch axis,the autopilot uses the:A. Pitch servo onlyB. Pi tch servo and t r im tabs on the

elevatorsC. Horizontal stabilizer trim actuator

onlyD. Pitch servo and secondary pitch trim

motor

8. If the stick nudger or puller engages dur-ing autopilot operation:A. Selected pitch modes will be canceled.B. The autopilot maintains a synchronous

standby mode in the pitch axis until thenudger or puller releases.

C. Selected roll modes remain engaged.D. All the above

9. When using the flight director REV (orBC) mode during a localizer back courseapproach, the:A. Reciprocal of the front course must be

set in the HSI course window.B. Glide-slope receiver signal is cap-

tured.C. Published inbound (front) course must

be set in the HSI course window.D. Both B and C are correct.

10. When using the autopilot, the followinglimitation applies:A. The pilot and copilot must be in their

r e spec t ive s ea t s w i th s ea t be l t sfastened.

B. The pilot or copilot must be in his re-spective seat with seat belt fastened.

C. The autopilot must be operative forairplane flight if the Mach trim systemin inoperative.

D. Do not extend or retract gear or flapswith autopilot engaged.

COMMUNICATION SYSTEM

11. The static wicks are important becausethey:A. Collect static electricityB. Function as an aerodynamic aidC. Dissipate lightning strikesD. Dissipate static electricity




17-i

CHAPTER 17MISCELLANEOUS SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 17-1

GENERAL ............................................................................................................................ 17-1

OXYGEN SYSTEM.............................................................................................................. 17-2

Oxygen Cylinder............................................................................................................ 17-3

Overboard Discharge Indicator...................................................................................... 17-3

OXYGEN PRESSURE Gage ........................................................................................ 17-4

Crew Distribution System.............................................................................................. 17-4

Passenger Distribution System ...................................................................................... 17-6

DRAG CHUTE...................................................................................................................... 17-8

General........................................................................................................................... 17-8

Operation ....................................................................................................................... 17-9

SQUAT SWITCH SYSTEM ................................................................................................. 17-9

General........................................................................................................................... 17-9

Squat Switches............................................................................................................... 17-9

Squat Switch Relay Box .............................................................................................. 17-10

QUESTIONS....................................................................................................................... 17-11




17-iii

ILLUSTRATIONS

Figure Title Page

17-1 Oxygen System ...................................................................................................... 17-2

17-2 Oxygen Cylinder and Overboard Discharge Indicator........................................... 17-3

17-3 OXYGEN PRESSURE Gage................................................................................ 17-4

17-4 Crew Oxygen Mask................................................................................................ 17-4

17-5 OXY-MIC Panel (Typical) ..................................................................................... 17-5

17-6 Passenger Distribution System............................................................................... 17-6

17-7 Passenger Mask...................................................................................................... 17-7

17-8 Drag Chute Components Location......................................................................... 17-8




INTRODUCTIONMiscellaneous systems covered in this section include the oxygen system, the dragchute, and the squat switch system. The airplane uses high-pressure oxygen stored in acylinder located in either the right nose section or the dorsal fin. Optional long-rangeoxygen installations are available. The drag chute is offered as optional equipment. Thesquat switch system provides the airborne and ground signals which activate or deacti-vate certain systems during takeoff and landing.

GENERALThe 35/36 series oxygen system consists of thecrew distribution system and the passengerdistribution system. Oxygen is available tothe crew at all times and can be made avail-able to the passengers either automaticallyabove 14,000 feet cabin altitude or manuallyat any altitude by the cockpit controls. Thesystem is primarily designed for use in theevent of rapid decompression or pressurizationsystem failure. It is not designed for planned

extended unpressurized flight at high cabinaltitudes requiring the use of oxygen.

The optional drag chute is used to improvedeceleration on the ground. It is most effec-tive when deployed at higher speeds, but canstill be effective when deployed at speedsbelow 60 knots.

The squa t sw i t ch sy s t em inc ludes twoswitches, one located on each of the main gearscissors, and a relay box.

1612

8

40

NO 1 FUELTRANS

NO 1 FUELLOWBATTHOT

ANTI-ICEON

OIL

BLOWEROFF

ENG 1CHIP

NO 1 FUELFILTER

NO 1 BATTSYS

XMSNOIL

90° BOXOIL

GEN 1HOT

TEST

RESET

CHAPTER 17MISCELLANEOUS SYSTEMS




OXYGEN SYSTEMThe oxygen system components include anoxygen s t o r age cy l i nde r and a shu to ff

valve-regula tor assembly, an overboard discharge indicator, an oxygen pressuregage, and the distribution systems for thecrew and passengers. Figure 17-1 depictsthe oxygen system.




DISCHARGEINDICATOR

TO COPILOT‘SMASK

PILOT’SMASK

SUPPLY PRESSURE

LEGEND

REGULATED PRESSURE

MASK VALVE/FLOW REGULATOR

QUICKDISCONNECT

PASSENGER MASKSTORAGE COMPARTMENT

PASS MASKVALVE

LANYARD PIN

MASK

DOORLATCH

SOLENOIDVALVE

ANEROID SWITCH(14,000 FEET)

PRESSURE REGULATORAND SHUTOFF VALVE

PASS OXYVALVE

FILLERVALVE

OXYGENCYLINDER

030

OXYGENPRESSURE

PSI X 10

155195

200

Figure 17-1. Oxygen System

OXYGEN CYLINDERThe system is supplied with oxygen from astorage cylinder located in the right nosesection on airplane SNs 35-002 through 35-491and 36-002 through 36-050 (Figure 17-2). Ona i rp l ane SNs 35 -492 and 36 -051 andsubsequent, the cylinder is located in the dorsalfin . An opt ional long-range insta l la t ionincorporat ing two cyl inders is avai lable(location of the cylinders varies).

Each oxygen cylinder has a storage capacityof 38 cubic feet at 1,800 psi. The shutoff valveand pressure-regulator assembly is attached tothe storage cylinder and provides for pressureregulation, pressure indication, and servic-ing. Oxygen pressure for the passenger andcrew distribution system is regulated at 60 to80 psi. The cylinder, along with its shutoffvalve and regulator assembly, can be reachedthrough an access door. Under normal condi-tions, this valve should always be left in the

on (open) position (a specified item on theexterior preflight inspection). The pilot shouldbe aware that if the oxygen cylinder shutoffvalve is closed, oxygen pressure will still beread on the OXY PRESS gage in the cockpit.During the interior preflight inspection, ensurethat the shutoff valve is open by checking foroxygen flow through both crew oxygen masks,using the 100%, or EMER, position.

OVERBOARD DISCHARGEINDICATORThe overboard discharge indicator (green blowoutdisc) (Figure 17-2) provides the pilot with a vi-sual indication that there has not been an over-pressure condition in the oxygen storage cylinder.The disc blows out if the cylinder pressure reaches2,700 to 3,000 psi, releasing all oxygen pres-sure. System pressure should normally be be-tween 1,550 and 1,850 psi. The green blowoutdisc is located on the right side of the dorsal finor the lower right side of the nose section.




Figure 17-2. Oxygen Cylinder and Overboard Discharge Indicator



OXYGEN PRESSURE GageThe OXYGEN PRESSURE gage (Figure 17-3) provides a direct reading of oxygen cylin-der pressure, which is necessary to ensure thatan adequate supply of oxygen is aboard. Thegage is marked as follows:

• Yellow arc ..............................0–300 psi

• Green arc ......................1,550–1,850 psi

• Red line ....................................2,000 psi

The gage is located on the pilot’s side panelon late model airplanes; on early models it ismounted on the instrument panel.

CREW DISTRIBUTION SYSTEMThe crew distribution system (see Figure 17-1)consists of the pilot’s and copilot’s oxygenmasks with mask-mounted regulators for diluter-demand or 100% operation. Oxygen is availableto the crew anytime the storage bottle shutoffvalve is open and the mask are plugged in.

The crew masks (Figure 17-4) are stowed onthe pilot’s and copilot’s sidewalls. The maskoxygen lines are connected to quick-discon-nect receptacles located on the cockpit side-walls. Optional oxygen-flow detectors maybe installed in the mask oxygen lines.

NOTEHeadsets, eyeglasses, or hats worn bycrewmembers may interfere with thequick-donning capabilities of theoxygen mask.

Three d i ffe ren t oxygen mask / regu la to rconfigurations are available on the 35/36model airplanes.

The ZMR 100 series diluter-demand maskregulator has a NORMAL-100% oxygenselector lever. With NORMAL selected, theregulator delivers diluted oxygen, on demand,up to 20,000 feet cabin altitude. Above 20,000feet cabin altitude, the 100% oxygen positionmust be selected. With the selector in the 100%position, 100% oxygen is delivered at anycabin altitude. The 100% position should beused when smoke or fumes are present in thepressurized compartment.


***

* LATE MODELS**EARLY MODELS

Figure 17-3. OXYGEN PRESSURE Gage

Figure 17-4. Crew Oxygen Mask

The Robertshaw diluter-demand mask/regu-lator has two controls—the NORMAL-EMER-GENCY selector and the 100% lever. WithNORMAL selected, the regulator delivers dilutedoxygen on demand, up to 30,000 feet cabinaltitude. Above 30,000 feet, the regulator delivers100% oxygen under a slight positive pressure.Depressing the 100% lever will deliver 100%oxygen at any t ime. With EMERGENCYselected (at any altitude) and the 100% leverdepressed, the regulator delivers 100% oxygenand maintains a slight positive pressure forrespiratory protection from smoke and fumes.

The Puritan-Bennett pressure demand mask/reg-ulator incorporates a three-position selectorknob labeled “NORM,” “100%,” and “EMER.”With NORM selected, the regulator deliversdiluted oxygen on demand, up to 33,000 feetcabin altitude. Above 33,000 feet, the regula-tor automatically delivers 100% oxygen. At39,000 feet, it provides positive-pressure breath-ing. To obtain 100% oxygen at any time, 100%must be selected on the pressure-regulatorcontrol. With EMER selected, the regulator de-livers 100% oxygen and maintains a slight pos-itive pressure in the mask cup at all times forrespiratory protection from smoke and fumes.

The Scott ATO MC 10-15-02 mask, in the nor-mal pressure regulator position with the 100%lever extended, will deliver diluted oxygenup to 30,000 feet cabin altitude, 100% oxygenabove 30,000 feet cabin altitude, and auto-matic pressure breathing above approximately37,000 feet cabin altitude. To obtain 100%oxygen at any time, depress the 100% lever onthe mask pressure regulator. With EMER-GENCY selected, the mask will deliver 100%oxygen and maintain a positive pressure in themask cup at all times for respiratory protectionfrom smoke and fumes.

Each mask assembly includes a microphoneand has an electrical cord which is plugged intothe OXY-MIC jack on the respective OXY-MIC panel (Figure 17-5) located on each sidepanel. To operate the mask microphone, theOXY-MIC switch must be in the ON positionand the microphone keyed, using the micro-phone switch on the outboard horn of thecontrol wheel . Communicat ion betweencrewmembers can be accomplished by usingthe INPH function of the audio control paneland increasing the MASTER VOL level.




Figure 17-5. OXY-MIC Panel (Typical)

PASSENGER DISTRIBUTIONSYSTEMThe passenger distribution system (Figure 17-6) is used to provide oxygen to the passengersin case of a pressurization system failure or any

other time that oxygen is required. Oxygen isavailable in the crew oxygen distribution lineswhenever the oxygen cylinder shutoff valve isopen; however, oxygen is not available to thepassenger distribution system until required.




Figure 17-6. Passenger Distribution System

PASS MASK VALVEMAN—AUTONORMALLY CLOSED (AUTO)

FROMCREW

OXYGENSYSTEM

PASS OXY VALVENORM—OFF

NORMALLY OPEN (NORM)

SOLENOID VALVENORMALLY CLOSED

DUAL-MASKSTORAGE

COMPARTMENT

OXYGENTRANSFER

TUBE

TO OTHER

PASSENGERMASK

COMPARTMENTS

SINGLE-MASK STORAGECOMPARTMENT

REGULATED PRESSURE

LEGEND

Oxygen supply to the passengers’ system iscontrolled with three valves. Two valves aremanually operated with control knobs on thepilot’s sidewall, and the third is solenoid-op-erated by an aneroid switch. The manuallycontrolled PASS OXY valve is normally inthe NORM (open) position, which allowsoxygen up and to the manually controlled PASSMASK valve and to the aneroid-controlledsolenoid valve. Oxygen can be admitted to thepassenger distribution system through either ofthese passenger mask valves, both of whichare normally in the closed position.

With the PASS OXY valve is the OFF (closed)position, oxygen will not be available to the pas-senger distribution system in any event. Thisposition may be used only when no passengersare being carried.

With the PASS OXY valve in the NORM(open) position, oxygen will be automaticallyadmitted to the passenger distribution systemthrough the aneroid-controlled solenoid valveif the cabin reaches 14,000 ±750 feet. Theaneroid switch opens the solenoid valve anddeploys the passenger masks. It also illumi-nates the cabin overhead lights.

In the event of airplane electrical failure, au-tomatic deployment of the passenger masks isnot possible. The oxygen solenoid valve re-quires DC power through the OXY VAL cir-cuit breaker on the left essential bus forautomatic mask deployment.

With the PASS OXY valve in the NORM (open)position, rotating the PASS MASK valve fromAUTO to the MAN position admits oxygeninto the passenger distribution system andcauses the passenger oxygen masks to drop.This position can be used to deploy the pas-senger masks at any altitude, but will not causethe cabin overhead lights to illuminate.

The passenger oxygen masks (Figure 17-7)are stowed in compartments in the conveniencepanels above the passenger seats. The com-partments may contain as many as three masks,depending on the airplane seating configura-tion. There will be at least one spare mask.

The passenger mask storage compartmentdoors are held closed by latches. When oxy-gen is admitted into the passenger distributionsystem, the oxygen pressure causes the doorlatches (plungers) to open each compartmentdoor. When the doors open, the passenger




Figure 17-7. Passenger Mask

masks fall free and are available for passen-ger use. As the passenger pulls down on hismask to don it, an attached lanyard withdrawsa pin from the supply valve which releasesoxygen into the mask breather bag at a re-stricted, constant-flow rate. The rebreatherbag may seem to inflate slowly, but this isnormal. When inhaling, 100% oxygen is de-livered to the mask cup. The breath is then ex-hausted into the rebreather bag.

Should the doors be inadvertently opened fromthe cockpit, oxygen pressure must be bledfrom the passenger distribution system beforethe masks can be restowed. This is accom-plished by pulling one of the passenger masklanyards after ensuring that the PASS MASKvalve is closed (AUTO). If the doors open dueto malfunction of the solenoid-operated valve,the PASS OXY valve must be turned off topermit stowage of the passenger masks.

The compar tmen t door s can be openedmanually for mask cleaning and servicing.

DRAG CHUTE

GENERALThe optional drag chute may be used to shortenstopping distances. The greatest decelerationrate is produced at the highest speed; however,the chute is still effective at speeds below 60knots. The chute is stored in a removable can-ister which is mounted inside the tailcone ac-cess door. The canister lid is released from thecanister when the drag chute handle is pulled,allowing the pilot chute to deploy. The pilotchute then pulls the main chute canopy out ofthe canister.

The main chute riser attaches to the airplaneat the chute control mechanism just forwardof the canister (Figure 17-8). The loop at theend of the main riser slips over a recessedmetal pin which is held in position by springpressure when the drag chute handle is stowed.Therefore, if the chute should inadvertently de-ploy (handle in stowed position), the mainchute riser will slip free of the pin and sepa-rate from the airplane. When the drag chutehandle is pulled, the pin is mechanically locked




DRAG CHUTECANISTER-LIDASSEMBLY

DRAG CHUTE CONTROLMECHANISM

DRAG CHUTECONTROL CABLE

DRAG CHUTE CONTROL HANDLE(RH SIDE OF PEDESTAL)

Figure 17-8. Drag Chute Components Location

in position to retain the chute riser while themechanical canister control mechanism op-erates to release the canister lid, thereby de-ploying the chute.

The drag chute can be used:

• When landing on wet or icy runaway

• During any landing emergency in-volving no-flap hydraulic or brakefailure, or loss of directional control

• During takeoff if the decision ismade to abort

Do not deploy the drag chute under the fol-lowing conditions:

• In flight

• If the nose gear is not on the ground

• When the indicated airspeed is above150 knots

• With thrust reversers deployed

OPERATIONAs the nosewheel touches down, the copilot,on the pilot’s command, deploys the drag chuteby squeezing the drag chute control handle(Figure 17-8) and pulling it up to its full ex-tension (a pull force of approximately 50pounds will be required). With the chute de-ployed, the pilot should keep the airplane wellclear of the runway and taxiway lights, mark-ers, and obstructions on the upwind side. Taxi-ing downwind should always be avoided.

The d r ag chu t e can be j e t t i soned a f t e rdeployment at anytime. Normally, the pilotheads the airplane into the wind as much aspossible to jettison the chute after the airplaneclears the runway. The copilot jettisons thedrag chute by squeezing the control handle gripsafeties and pushing the handle down to thestowed position to release the chute. If thechute has collapsed prior to jettisoning, thechute riser must be pulled free after stowing

the handle. Because the possibility always ex-ists that jettisoning the chute might be re-quired during the landing roll, any planneddeployment should be coordinated with thecontrol tower.

SQUAT SWITCH SYSTEM

GENERALSome airplane systems operate only on theground while others operate only in the air. Thesquat switch system is designed to providethe necessary ground or airborne signals tothese systems. The squat switch system con-sists of two squat switches, one on each mainlanding gear strut scissors, and a relay box lo-cated under the cabin floor. When the airplaneis on the ground, and the main landing gearstruts are compressed, the squat switches closeto provide a ground mode signal. When the air-plane lifts off the ground, and the main land-ing gear struts extend, the squat switches open,interrupting the ground mode signals, therebyshifting to air mode.

SQUAT SWITCHESEach squat switch provides ground or air sig-nals to the following components:

• Stall Warning System

• The switches disable the stall warn-ing test feature in the air.

• The switches disable the stall warningrate sensor on the ground. The ratesensor remains disabled for approxi-mately five seconds after lift-off.

• The left squat switch controls the leftstall warning system while the rightsquat switch controls the right stallwarning system.

• Antiskid System

• The switches disable the wheel brakesin the air with the antiskid system on.




The wheel brakes remain inoperativeuntil the wheels’ spinup requirementshave been met on landing.

• The left squat switch controls the out-board wheel brakes while the rightsquat switch controls the inboardwheel brakes.

• Gear Control Valve

• The switches disable the gear-upsolenoid on the ground to prevent in-advertent landing gear retraction.

• Either squat switch in ground modewill disable the gear-up solenoid. Bothsquat switches must be in the air modeto allow landing gear retraction.

• Squat Switch Relay Box

• Either squat switch in the ground modeputs the relay box in ground mode.

• Both squat switches must go to airmode to put the relay box in air mode.

The position of the SQUAT SW circuit breakerhas no effect on landing gear, antiskid, or stallwarning system operation. These systemsreceive signals directly from the squat switchesas explained previously.

SQUAT SWITCH RELAY BOXThe squat switch relay box is necessary becauseof the limited number of electrical contactsavailable on the main landing gear squatswitches. Sensing signals from both squatswitches, the relay box provides ground or airmode signals to the components listed below.The squat switch relay box uses DC powerfrom the SQUAT SW circuit breaker on theleft main DC bus to provide ground mode sig-nals. With the SQUAT SW circuit breaker open,all the relay box functions go to air mode.

The squat switch relay box provides groundor air mode signals to the following:

• Nosewheel steering—Disabled in the air

• Spoiler/spoileron system—Disables themonitor system on the ground. Slowsthe rate of spoiler deployment in the air.

• Cabin pressurization

• Safety valve vacuum solenoid closes inthe air (SNs 35-099 and subsequent and36-029 and subsequent only).

• Amber CAB ALT light (if installed) isdisabled on the ground.

• Control module solenoids shift fromground to air mode.

• Amber TO TRIM light—Disabled in theair

• Windshield heat system—Shifts fromground to air mode. (See Chapter 10,“Ice and Rain Protection,” for additionalinformation.)

• Hourmeter and Davtron clock flight timefunction (if installed)—Disabled on theground

• Mach trim test—Operates only on theground

• Thrust reversers—Operates only on theground

• Generator load limiting—Limits the out-put of a single, engine-driven generatoron the ground only (SNs 35-148 andsubsequent and 36-036 and subsequentonly).

• Air data unit—TAS disabled on theground (AFCS 530 autopilot airplanesonly).

• Mach overspeed warning/stick puller—Test function disabled in the air (AFCS530 airplanes only).

• Yaw damper—Disconnects at touch-down (AFCS 530 airplanes only).




QUESTIONS1. During preflight the pilot can determine

if the oxygen bottle is turned on by:A. Reading the pressure indicated on the

oxygen pressure gage in the cockpitB. selecting 100% on the mask regulator

and t ak ing severa l deep b rea thsthrough the mask

C. Placing the OXY-MIC switch to theOXY position

D. Visually checking for the green flowindicator on the mask supply hose

2. With the PASS OXY valve in the NORMposition, selecting MAN on the PASSMASK valve:A. Causes passenger masks to drop and

turns on the cabin overhead lightsB. Prevents oxygen from entering the

passenger oxygen distribution linesC. Disarms the 14,000-foot cabin aneroidD. Admits oxygen to the passenger

dis t r ibut ion l ines and causes thepassenger oxygen masks to drop

3. With the PASS OXY valve in the NORMposition and the PASS MASK valve inthe AUTO position:A. Oxygen is supplied to the passenger

masks if the cabin altitude reaches10,000 feet.

B. Passenger masks will automaticallydeploy in the event of electrical failure.

C. Passenger masks will automaticallydeploy and the cabin overhead lightswi l l i l lumina te i f cab in a l t i tudereaches 14,000 feet.

D. The aneroid-controlled passengermask drop valve is disabled.

4. The OXY PRESS gage reads:A. Direct pressure of the cylinderB. Electrically derived system high pres-

sureC. Direct pressure of the pilot’s supply

lineD. Electrically derived system low pres-

sure

5. The maximum demonstrated crosswindcomponent for drag chute deployment is:A. 10 knotsB. 15 knotsC. 20 knotsD. 25 knots

6. The drag chute is deployed by:A. Squeezing the control handleB. Rotating the control handle fully

clockwise and pulling it up to its fullextension

C. Squeezing the control handle andpulling it up to its full extension

D. Squeezing the control handle andpushing it completely forward

7. The maximum indicated airspeed for dragchute deployment is:A. 120 knotsB. 130 knotsC. 140 knotsD. 150 knots

8. If either main landing gear squat switchremains in ground mode after takeoff:A. The landing gear will not retract.B. The airplane will not pressurize.C. The ambe r TO TRIM l i gh t may

illuminate.D. All of the above






WA-1FOR TRAINING PURPOSES ONLY

WALKAROUNDThe following section is a pictorial walkaround.It shows each item called out in the exteriorpower-off preflight inspection. The fold-outpages, WA-2 and WA-15, should be unfolded be-fore starting to read.

The general location photographs do not specifyevery checklist item. However, each item is por-trayed on the large-scale photographs that follow.

WA-2 FOR TRAINING PURPOSES ONLY





Revision .01FOR TRAINING PURPOSES ONLY

3 5 1 2

9 10

7

8 6

4

83 84 79

81 82

77

76 78 74

86

87

88

75

85

1. PILOT’S WINDSHIELD ALCOHOL DISCHARGEOUTLETS AND PILOT’S DEFOG OUTLET—CLEAR OFOBSTRUCTIONS

2. LEFT SHOULDER STATIC PORT (FC 200) —CLEAR OFOBSTRUCTIONS

3a. LEFT PITOT-STATIC PROBE (FC 530)—COVERREMOVED, CLEAR OF OBSTRUCTIONS

4. LEFT STALL WARNING VANE—FREEDOM OFMOVEMENT, LEAVE IN DOWN POSITION

5. LEFT STATIC PORTS (2) (FC 200)—CLEAR OFOBSTRUCTIONS

6. SHOULDER STATIC (1) (FC 200) AND LEFT PITOT-STATIC (2) DRAIN VALVES—DRAIN

7. NOSE GEAR AND WHEEL WELL—HYDRAULICLEAKAGE AND CONDITION

WALKAROUND INSPECTION

3. LEFT PITOT HEAD (FC 200)—COVER REMOVED, CLEAROF OBSTRUCTIONS

29

20

26

34

1210a

2223

24

25

27

18

13

1615

17 11

21 19

1428303233

31

35



Revision .01 WA-3FOR TRAINING PURPOSES ONLY

FC 200 AND FC 530 FC 200 ONLY

8. NOSEWHEEL AND TIRE—CONDITION AND NOSE GEARUPLOCK FORWARD

10a. OXYGEN BOTTLE SUPPLY VALVE—ONOXYGEN PRESSURE RELIEF DISC—INTACT

12. RIGHT STALL WARNING VANE—FREEDOM OFMOVEMENT, LEAVE IN DOWN POSITION

13. RIGHT STATIC PORTS FC 200 (3) OR FC 530 (1)—CLEAR OF OBSTRUCTIONS

14. RIGHT PITOT-STATIC DRAIN VALVES (2)—DRAIN

9. RADOME ALCOHOL DISCHARGE PORT—CLEAR OFOBSTRUCTION

10. RADOME AND RADOME EROSION SHOE—CONDITION

11. RIGHT PITOT HEAD (FC 200) AND TEMPERATURE PROBE—COVERS REMOVED, CLEAR OF OBSTRUCTIONS

13a. RIGHT PITOT-STATIC PROBE AND TEMPERATUREPROBE (FC 530)—COVER REMOVED, CLEAR OFOBSTRUCTIONS




WA-4

15. RIGHT SHOULDER STATIC PORT—CLEAR OFOBSTRUCTIONS (FC0-200)

16. COPILOT’S WINDSHIELD DEFOG OUTLET—CLEAR OFOBSTRUCTIONS

17. LOWER FUSELAGE ANTENNAE, ROTATING BEACONLIGHT AND LENS—CONDITION

18. EMERGENCY EXIT—SECURE19. UPPER FUSELAGE ANTENNAE—CONDITION

20. ROTATING BEACON LIGHT AND LENS (ON VERTICALFIN)—CONDITION

21. RIGHT ENGINE INLET AND FAN—CLEAR OFOBSTRUCTIONS AND CONDITION

22. FUEL CROSSOVER, LEFT WING SUMP, LEFT ENGINEFUEL, RIGHT WING SUMP, AND RIGHT ENGINE FUELDRAIN VALVES—DRAIN

23. RIGHT MAIN GEAR AND WHEEL WELL—HYDRAULIC/FUEL LEAKAGE AND CONDITION

27. RIGHT WING ACCESS PANELS (UNDERSIDE OFWING)—CHECK FOR FUEL LEAKAGE

26. STALL STRIP, WING LEADING EDGE, AND STALLFENCE—CONDITION

28. RIGHT FUEL VENT (UNDERSIDE OF WING)—PLUGREMOVED, CLEAR OF OBSTRUCTIONS

24. RIGHT MAIN GEAR LANDING LIGHT—CONDITION

25. RIGHT MAIN GEAR WHEELS, BRAKES, AND TIRES—CONDITION

29. VORTEX GENERATORS OR BOUNDARY LAYERENERGIZERS—CONDITION







WA-6

30. RIGHT WING HEAT SCUPPER (UNDERSIDE OF WING

FORWARD)—CLEAR OF OBSTRUCTIONS

33. RIGHT TIP TANK SUMP DRAIN VALVE—DRAIN

31. RIGHT TIP TANK—CONDITION 34. RIGHT TIP TANK FUEL CAP—CONDITION AND

SECURE

35. RIGHT TIP TANK NAVIGATION LIGHT, STROBE LIGHT,

AND LENS—CONDITION

32. RIGHT TIP TANK RECOGNITION LIGHT AND

LENS—CONDITION36. RIGHT TIP TANK FIN AND STATIC DISCHARGE

WICKS (2)—CONDITION

38. SCUPPER (UNDERSIDE OF RIGHT WING AFT)—CLEAROF OBSTRUCTIONS, NO FUEL LEAKAGE

39. RIGHT AILERON—CHECK FREE MOTION,BALANCE TAB LINKAGE, BRUSH SEAL CONDITION

40. RIGHT SPOILER AND FLAP—CONDITION

41. RIGHT ENGINE OIL QUANTITY—CHECK FILLER CAP AND ACCESS DOOR—SECURE

42. RIGHT ENGINE OIL BYPASS VALVE INDICATOR—CHECK, NOT EXTENDED

37. RIGHT TIP TANK FUEL JETTISON TUBE—CLEAR OFOBSTRUCTIONS







WA-8

43. RIGHT ENGINE THRUST REVERSER—CONDITIONAND STOWED (AERONCA)

45. RIGHT ENGINE FUEL BYPASS VALVE INDICATOR—CHECK, NOT EXTENDED

43A. RIGHT ENGINE THRUST REVERSER—CONDITIONAND STOWED (DEE HOWARD)

46. FUEL VENT DRAIN VALVE, TRANSFER LINE DRAINVALVE, FUSELAGE TANK SUMP DRAIN VALVE—DRAIN

44. RIGHT ENGINE TURBINE EXHAUST AREA—CONDITION, CLEAR OF OBSTRUCTION, BLOCKERDOORS STOWED (AERONCA)

47. LEFT AND RIGHT FUEL FILTER DRAIN VALVES—DRAIN



48. TAILCONE ACCESS DOOR—OPEN 49B. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDITION OF INSTALLEDEQUIPMENT

49. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDITION OF INSTALLEDEQUIPMENT

50. DRAG CHUTE—CHECK FOR PROPER INSTALLATION

49A. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDITION OF INSTALLEDEQUIPMENTHYDRAULIC ACCUMULATOR AIR CHARGE—750PSI MINIMUM

50A. DRAG CHUTE—CHECK FOR PROPER INSTALLATION





WA-10

51. TAILCONE ACCESS DOOR—CLOSED AND SECURE 55. RIGHT FUEL COMPUTER DRAIN VALVE—DRAIN(DRAIN VALVES ARE RECESSED ON AIRPLANESEQUIPPED WITH DRAG CHUTE.)

52. OXYGEN BOTTLE SUPPLY VALVE—OPEN 56. RIGHT VOR/LOC ANTENNA—CONDITION57. VERTICAL STABILIZER, RUDDER, HORIZONTAL

STABILIZER, AND ELEVATOR—CONDITION,DRAIN HOLES CLEAR

58. STATIC DISCHARGE WICKS (6 ON HORIZONTALSTABILIZER, 1 ABOVE NAV LIGHT, 1 ONVERTICAL FIN)—CONDITION

59. VERTICAL FIN NAVIGATION LIGHTS, STROBE LIGHTAND LENS—CONDITION

60. VLF H-FIELD ANTENNA—CONDITION61. LEFT VOR/LOC ANTENNA—CONDITION

53. OXYGEN SERVICING DOOR—SECURE54. OXYGEN DISCHARGE DISC—CONDITION

62. LEFT FUEL COMPUTER DRAIN VALVE—DRAIN (DRAINVALVES ARE RECESSED ON AIRPLANES EQUIPPEDWITH DRAG CHUTE.)

65A. LEFT ENGINE THRUST REVERSER—CONDITIONAND STOWED (DEE HOWARD)

63. FIRE EXTINGUISHER DISCS—CONDITION 65. LEFT ENGINE TURBINE EXHAUST AREA—CONDITION, CLEAR OF OBSTRUCTIONS ANDBLOCKER DOORS STOWED (AERONCA)

64. LEFT ENGINE OIL BYPASS VALVE INDICATOR—CHECK, NOT EXTENDED

66. LEFT ENGINE TRUST REVERSER —CONDITION ANDSTOWED (AERONCA)




Revision .01WA-12 FOR TRAINING PURPOSES ONLY



69. LEFT SPOILER AND FLAP—CONDITION 72. LEFT TIP TANK FUEL JETTISON TUBE—CLEAR OFOBSTRUCTIONS

67. LEFT ENGINE FUEL BYPASS VALVE INDICATOR—CHECK, NOT EXTENDED

70. LEFT AILERON—CHECK FREE MOTION, BALANCE,AND TRIM LINKAGE, AND BRUSH SEAL CONDITION

71. SCUPPER (UNDERSIDE OF LEFT WING AFT)—CLEAR OF OBSTRUCTIONS, NO FUEL LEAK

68. LEFT ENGINE OIL QUANTITY—CHECKFILLER CAP AND ACCESS DOOR—SECURE




76. LEFT TIP TANK SUMP DRAIN VALVE—DRAIN 79. LEFT WING HEAT SCUPPER (UNDERSIDE OF WINGFORWARD)—CLEAR OF OBSTRUCTIONS

73. LEFT TIP TANK FIN AND STATIC DISCHARGEWICKS (2)—CONDITION

77. LEFT TIP TANK RECOGNITION LIGHT AND LENS(IF INSTALLED)—CONDITION

78. LEFT TIP TANK—CONDITION74. LEFT TIP TANK NAVIGATION LIGHT, STROBE LIGHTAND LENS—CONDITION

75. LEFT TIP TANK CAP—CONDITION ANDSECURE

Revision .01WA-14 FOR TRAINING PURPOSES ONLY



82. LEFT FUEL VENT (UNDERSIDE OF WING)—PLUGREMOVED, CLEAR OF OBSTRUCTIONS

85. LEFT MAIN GEAR AND WHEEL WELL—HYDRAULIC/FUEL LEAKAGE AND CONDITION

80. VORTEX GENERATORS OR BOUNDARY LAYERENERGIZERS—CONDITION

83. STALL STRIP (IF INSTALLED) AND WING LEADINGEDGE—CONDITION

84. STALL FENCE (IF INSTALLED)—CONDITION81. LEFT WING ACCESS PANELS (UNDERSIDE OFWING)—CHECK FOR FUEL LEAKAGE

WA-15FOR TRAINING PURPOSES ONLY



Revision .01 FOR TRAINING PURPOSES ONLY



86. LEFT MAIN GEAR LANDING LIGHT—CONDITION

87. LEFT MAIN GEAR WHEELS, BRAKES, AND TIRES—CONDITION

88. LEFT ENGINE INLET AND FAN—CLEAR OFOBSTRUCTIONS AND CONDITION 57

58

59

60

56 48

49

50

51

55 47 43

45

42

46

39

38

37 36

4144

5452

53

40

75 73 80

68 67 66

74 71 72 70 69 64 63 65 62

61

WA-16 FOR TRAINING PURPOSES ONLY



NOTE:THE NUMBERS ON THISDIAGRAM CORRESPOND TOTHE PREFLIGHT POSITIONSDEPICTED IN THE AIRPLANEFLIGHT MANUAL.

APP-i

This appendix contains the following conversion tables:

Table Title Page

APP-1 Conversion Factors ............................................................................................. APP-1

APP-2 Farenheit and Celsius Temperature Conversion................................................. APP-2

APP-3 Inches to Millimeters.......................................................................................... APP-3

APP-4 Weight (Mass): Ounces or Pounds to Kilograms .............................................. APP-4

APP-5 Weight (Mass): Thousand Pounds to Kilograms............................................... APP-5




APP-1

Table APP-1. CONVERSION FACTORS




MULTIPLY BY TO OBTAIN

centimeters 0.3937 incheskilograms 2.2046 poundskilometers 0.621 statute mileskilometers 0.539 nautical milesliters 0.264 gallonsliters 1.05 quarts (liquid)meters 39.37 inchesmeters 3.281 feetmillibars 0.02953 in. Hg (32°F)feet 0.3048 metersgallons 3.7853 litersinches 2.54 centimetersin. Hg (32°F) 33.8639 millibarsnautical miles 1.151 statute milesnautical miles 1.852 kilometerspounds 0.4536 kilogramsquarts (liquid) 0.946 litersstatute miles 1.609 kilometersstatute miles 0.868 nautical miles

AP

P-2

FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

FlightSafety

international

LE

AR

JE

T 3

0 S

erie

s

PIL

OT

TR

AIN

ING

MA

NU

AL

Table APP-2. FARENHEIT AND CELSIUS CONVERSION

APP-3FOR TRAINING PURPOSES ONLY



Table APP-3. INCHES TO MILLIMETERS (0.0001 INCH TO 10 INCHES)

INCHES 0.0000 0.0001 0.0002 0.0003 0.0004 0.0005 0.0006 0.0007 0.0008 0.0009

MILLIMETERS

0.000 0.0025 0.0050 0.0076 0.0101 0.0127 0.0152 0.0177 0.0203 0.02280.001 0.0254 0.0279 0.0304 0.0330 0.0355 0.0381 0.0406 0.0431 0.0457 0.04820.002 0.0508 0.0533 0.0558 0.0584 0.0609 0.0635 0.0660 0.0685 0.0711 0.07360.003 0.0762 0.0787 0.0812 0.0838 0.0863 0.0889 0.0914 0.0939 0.0965 0.09900.004 0.1016 0.1041 0.1066 0.1092 0.1117 0.1143 0.1168 0.1193 0.1219 0.1244

0.005 0.1270 0.1295 0.1320 0.1346 0.1371 0.1397 0.1422 0.1447 0.1473 0.14980.006 0.1524 0.1549 0.1574 0.1600 0.1625 0.1651 0.1676 0.1701 0.1727 0.17520.007 0.1778 0.1803 0.1828 0.1854 0.1879 0.1905 0.1930 0.1955 0.1981 0.20060.008 0.2032 0.2057 0.2082 0.2108 0.2133 0.2159 0.2184 0.2209 0.2235 0.22600.009 0.2286 0.2311 0.2336 0.2362 0.2387 0.2413 0.2438 0.2463 0.2489 0.2514

INCHES 0.000 0.001 0.002 0.003 0.004 0.005 0.006 0.007 0.008 0.009

MILLIMETERS

0.00 0.025 0.050 0.076 0.101 0.127 0.152 0.177 0.203 0.2280.01 0.254 0.279 0.304 0.330 0.355 0.381 0.406 0.431 0.457 0.4820.02 0.508 0.533 0.558 0.584 0.609 0.635 0.660 0.685 0.711 0.7360.03 0.762 0.787 0.812 0.838 0.863 0.889 0.914 0.939 0.965 0.9900.04 1.016 1.041 1.066 1.092 1.117 1.143 1.168 1.193 1.219 1.244

0.05 1.270 1.295 1.320 1.346 1.371 1.397 1.422 1.447 1.473 1.4980.06 1.524 1.549 1.574 1.600 1.625 1.651 1.676 1.701 1.727 1.7520.07 1.778 1.803 1.828 1.854 1.879 1.905 1.930 1.955 1.981 2.0060.08 2.032 2.057 2.082 2.108 2.133 2.159 2.184 2.209 2.235 2.2600.09 2.286 2.311 2.336 2.362 2.387 2.413 2.438 2.463 2.489 2.514

INCHES 0.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09

MILLIMETERS

0.0 0.254 0.508 0.762 1.016 1.270 1.524 1.778 2.032 2.2860.1 2.540 2.794 3.048 3.302 3.556 3.810 4.064 4.318 4.572 4.8260.2 5.080 5.334 5.588 5.842 6.096 6.350 6.604 6.858 7.112 7.3660.3 7.620 7.874 8.128 8.382 8.636 8.890 9.144 9.398 9.652 9.9060.4 10.160 10.414 10.668 10.922 11.176 11.430 11.684 11.938 12.192 12.446

0.5 12.700 12.954 13.208 13.462 13.716 13.970 14.224 14.478 14.732 14.9860.6 15.240 15.494 15.748 16.002 16.256 16.510 16.764 17.018 17.272 17.5260.7 17.780 18.034 18.288 18.542 18.796 19.050 19.304 19.558 19.812 20.0660.8 20.320 20.574 20.828 21.082 21.336 21.590 21.844 22.098 22.352 22.6060.9 22.860 23.114 23.368 23.622 23.876 24.130 24.384 24.638 24.892 25.146

INCHES 0.00 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9

MILLIMETERS

0. 2.54 5.08 7.62 10.16 12.70 15.24 17.78 20.32 22.861. 25.40 27.94 30.48 33.02 35.56 38.10 40.64 43.18 45.72 48.262. 50.80 53.34 55.88 58.42 60.96 63.50 66.04 68.58 71.12 73.663. 76.20 78.74 81.28 83.82 86.36 88.90 91.44 93.98 96.52 99.064. 101.60 104.14 106.68 109.22 111.76 114.30 116.84 119.38 121.92 124.46

5. 127.00 129.54 132.08 134.62 137.16 139.70 142.24 144.78 147.32 149.866. 152.40 154.94 157.48 160.02 162.56 165.10 167.64 170.18 172.72 175.267. 177.80 180.34 182.88 185.42 187.96 190.50 193.04 195.58 198.12 200.668. 203.20 205.74 208.28 210.82 213.36 215.90 218.44 220.98 223.52 226.069. 228.60 231.14 233.68 236.22 238.76 241.30 243.84 246.38 248.92 251.46

APP-4 FOR TRAINING PURPOSES ONLY



Table APP-4. WEIGHT (MASS): OUNCES OR POUNDS TO KILOGRAMS

0 1 2 3 4 5 6 7 8 9

kg kg kg kg kg kg kg kg kg kg

oz0 – 0.028 0.057 0.085 0.113 0.142 0.170 0.198 0.227 0.255

10 0.283 0.312 0.340 0.369 0.397 0.425 0.454 0.482 0.510 0.539

lb0 – 0.45 0.91 1.36 1.81 2.27 2.72 3.18 3.63 4.08

10 4.5 5.0 5.4 5.9 6.4 6.8 7.3 7.7 8.2 8.620 9.1 9.5 10.0 10.4 10.9 11.3 11.8 12.2 12.7 13.230 13.6 14.1 14.5 15.0 15.4 15.9 16.3 16.8 17.2 17.740 18.1 18.6 19.1 19.5 20.0 20.4 20.9 21.3 21.8 22.250 22.7 23.1 23.6 24.0 24.5 24.9 25.4 25.9 26.3 26.860 27.2 27.7 28.1 28.6 29.0 29.5 29.9 30.4 30.8 31.370 31.8 32.2 32.7 33.1 33.6 34.0 34.5 34.9 35.4 35.880 36.3 36.7 37.2 37.6 38.1 38.6 39.0 39.5 39.9 40.490 40.8 41.3 41.7 42.2 42.6 43.1 43.5 44.0 44.5 44.9

100 45 46 46 47 47 48 48 49 49 49

0 10 20 30 40 50 60 70 80 90

200 91 95 100 104 109 113 118 122 127 132300 136 141 145 150 154 159 163 168 172 177400 181 186 191 195 200 204 209 213 218 222500 227 231 236 240 245 249 254 259 263 268600 272 277 281 286 290 295 299 304 308 313700 318 322 327 331 336 340 345 349 354 358800 363 367 372 376 381 386 390 395 399 404900 408 413 417 422 426 431 435 440 445 449

1000 454 458 463 467 472 476 481 485 490 494

(1 oz = 0.028 349 52 kg) (1 lb = 0.453 592 4 kg)

APP-5

lb 0 100 200 300 400 500 600 700 800 900

(000)* kg kg kg kg kg kg kg kg kg kg

1 454 499 544 590 635 680 726 771 816 8622 907 953 998 1043 1089 1134 1179 1225 1270 13153 1361 1406 1451 1497 1542 1588 1633 1678 1724 17694 1814 1860 1905 1950 1996 2041 2087 2132 2177 22235 2268 2313 2359 2404 2449 2495 2540 2585 2631 26766 2722 2767 2812 2858 2903 2948 2994 3039 3084 31307 3175 3221 3266 3311 3357 3402 3447 3493 3538 35838 3629 3674 3719 3765 3810 3856 3901 3946 3992 40379 4082 4128 4173 4218 4264 4309 4354 4400 4445 4491

10 4536 4581 4627 4672 4717 4763 4803 4853 4899 494411 4990 5035 5080 5126 5171 5216 5262 5307 5352 539812 5443 5488 5534 5579 5625 5670 5715 5761 5806 585113 5897 5942 5987 6033 6078 6123 6169 6214 6260 630514 6350 6396 6441 6486 6532 6577 6622 6668 6713 675915 6804 6849 6895 6940 6985 7031 7076 7121 7167 721216 7257 7303 7348 7394 7439 7484 7530 7575 7620 766617 7711 7756 7802 7847 7893 7938 7983 8029 8074 811918 8165 8210 8255 8301 8346 8391 8437 8482 8528 857319 8618 8664 8709 8754 8800 8845 8890 8936 8981 902620 9072 9117 9163 9208 9253 9299 9344 9389 9435 9480




Table APP-5. WEIGHT (MASS): THOUSAND POUNDS TO KILOGRAMS

*Multiply lb value by 1000

(1 lb = 0.453 592 4 kg)




ANNUNCIATORS

The Annuncia tor Sec t ion presents a co lor representation of all the annunciator lights in the airplane.

Please unfold page ANN-1 to the right and leaveit open for ready reference as the annunciatorsare cited in the text.

ANN-3FOR TRAINING PURPOSES ONLY






CLOSECROSS FLOW

PITCHTRIM

CURLIM

L CURLIMITERR CUR

LIMITER

LOWHYD

FUELXFLO

R LOOIL

L LOOIL

FMIZ

MMOM

MSTRWARN

EMERPWR 1

EMERPWR 2

FIRE PULL FIRE PULL

ARMED ARMED ARMED ARMED

LH ENGCHIP

RH ENGCHIP

L PITOTHEAT

R PITOTHEAT

RIGHTTHRUST REVERSER

ARM ARM

OFF

TEST TEST

DEPLOY DEPLOYARM ARM

UNLOCK DEPLOY BLEEDVALVE UNLOCK DEPLOY BLEED

VALVE

TEST THRUST NORM EMER STOW REVERSER EMER STOW

AERONCA

HDG REV GA FNL

GSCAPT

GSARM

NAVCAPT

NAVARM

FM/Z

OM MM

MSTRWARN

ANTI-SKIDGEN

L R

TRK

APPR

HDG NAV REV LVL TEST ENG SOFT SPD V/S G/S ALT

AMR CAPT. FNL

G/A

PITCH IAS MACHPWR ROLLARM CAPT

BRT UPTEST

MUTE DN

LOCKED DNUNSAFE

EMPTYXFER

OFFFILL

FULL

OPEN

CLOSEFUS VALVE

FUSTANK

AIR IGN L ON

OFF OFF OFF

BAT 2

START 1 START R

GEN

OFF

OFFOFF

OFF

GEN

R BUS

AUX INVERTER INVERTERL BUS PRI SEC AIR IGN R

R GENRESET

L GENRESET

OFF

OFF

LEFT

DEE HOWARD TR 4000

NOTE: FOR FC-530 AUTOPILOT/FLIGHT DIRECTOR PANEL SEE CHAPTER 16

OR

OR

OR

AMK 80-17

FUELQUANTITY

LBSI x 1000

L ON R

L ON R

L ON R

JET PUMP

OPEN

FUEL TSN

EMPTYXFER

OFFFILL

FULL

FUS

TANK

0 0 0 03

2

1

0

45

6

78

L WING1254

R WING1254

L TIP1215

R TIP1175

TOTAL6238LBS

FUS1340

OR

DH

PRIINV

LOWFUEL

SECINV

L FUELPRESS

AUXINV

R FUELPRESS

LO OILPRESS

SPOILER

STABOV HT

DOOR

WSHLDOV HT

AUGAIL

STEERON

PITOTHT

BLEEDAIR L

FUELFILTER

BLEEDAIR R

L ENGICEL

GEN

R ENGICE

RGEN

L FUELCMPTR

CABALT

R FUELCMPTR

WINGOV HT

LSTALL

WSHLDHT

RSTALL

ALCAI

L VGMON

BAT140

R VGMON

BAT160

MACHTRIM

ENGSYNC

DH

TOTRIM

BAT 1

Figure ANN-1. Annunciators

learjet 35 manual

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