G500-G550 MTM V1

FlightSafety international

GULFSTREAM G500/G550 MAINTENANCE TRAINING MANUAL VOLUME 1

FlightSafety International, Inc. Marine Air Terminal, LaGuardia Airport Flushing, New York 11371 (718) 565-4100 www.flightsafety.com

Courses for the G500/G550 and other Gulfstream aircraft are taught at the following FlightSafety learning centers: FlightSafety International Gulfstream Learning Center 301 Robert B. Miller Road Savannah, Georgia 31408 (912) 644-1000 (800) 625-9369

FlightSafety International Long Beach Learning Center Long Beach Municipal Airport 4330 Donald Douglas Drive Long Beach, CA 90808 (562) 938-0100 (800) 487-7670

FlightSafety International Greater Philadelphia/Wilmington Learning Center New Castle County Airport 155 North duPont Highway New Castle, Delaware 19720 (302) 221-5100 FlightSafety International DFW Learning Center 3201 East Airfield Drive DFW Airport, TX 75261-9428 (972) 534-3200

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

F O R T R A I N I N G P U R P O S E S O N LY

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s Pilot Manuals and Maintenance 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 conflict between data provided herein and that in publications issued by the manufacturer or 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 any suggestions you might have for improving this manual or any other aspect of our training program.

F O R T R A I N I N G P U R P O S E S O N LY

CONTENTS VOLUME 1 Chapter Title

ATA Number

INTRODUCTION ATA 100 AIRCRAFT GENERAL

5–12

AIR CONDITIONING

21

AUTOFLIGHT

22

COMMUNICATIONS

23

ELECTRICAL POWER

24

EQUIPMENT AND FURNISHINGS

25

FIRE PROTECTION

26

FLIGHT CONTROLS

27

FUEL

28

HYDRAULIC POWER

29

ICE AND RAIN PROTECTION

30

INDICATING AND RECORDING SYSTEMS

31


GULFSTREAM G500/G550 MAINTENANCE TRAINING MANUAL

LIST OF EFFECTIVE PAGES Dates of issue for original and changed pages are: Original ......0 ....... January 2005 NOTE: For printing purposes, revision numbers in footers occur at the bottom of every page that 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:

Page No.

*Revision No.

Cover—6 ................................................ LEP.......................................................... 1-2 .......................................................... 2-i—2-24 .................................................. 5-i—12-46 ................................................ 21-i—21-90 .............................................. 23-i—23-14 .............................................. 24-i—24-150 ............................................ 25-i—25-20 .............................................. 26-i—26-50 .............................................. 27-i—27-112 ............................................ 28-i—28-62 .............................................. 29-i—29-50 .............................................. 30-i—30-46 .............................................. 31-i—31-64 ..............................................

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

*Zero in this column indicates an original page.

Revision 2

FOR TRAINING PURPOSES ONLY

LEP-1



INTRODUCTION

INTRODUCTION This training manual provides a description of the major airframe and engine systems installed in the Gulfstream G500/G550. This information is intended as an instructional aid only; it does not supersede, nor is it meant to substitute for, any of the manufacturer’s maintenance or operating manuals. This material has been prepared from the basic design data, and all subsequent changes in airplane appearance or system operation will be covered during academic training and by subsequent revisions to this manual.

GENERAL In addition to the basic Maintenance Training Manual, FlightSafety provides a supplemental Maintenance Schematic Manual (MSM) to be used concurrently. The MSM, commonly called the “flat manual,” is printed in an 11 x 17-inch format and contains schematics to be used only as a tool in understanding a system. They are not kept current. The corresponding schematic(s) in the manufacturer’s Maintenance Manual must be used when performing maintenance.

The second chapter of this manual, “ATA 100/Publications,” is an introduction to the Air Transport Association format for aircraft maintenance manuals. It is intended to describe simply the basic format for all ATA 100 maintenance manual chapters and also to explain where variations may exist from one manufacturer to another. In addition, it includes information on various Gulfstream Aerospace publications useful to the maintenance technician.


1-1



Following “ATA 100/Publications,” each chapter of this book has listed on the divider tab t h e ATA c h a p t e r i n c l u d e d , s u c h a s “ 2 4 Electrical Power.” In some cases it was appropriate, for training purposes, to include more than one ATA chapter in one chapter of this book, such as Chapter 51–57, “Structures,” w h i c h i n c l u d e s i n f o r m a t i o n f r o m ATA Chapters 51 through 57. The goal of this course is to provide the very best training possible for the clients in our maintenance initial program. So that there is no uncertainty about what is expected of the client, the following basic objectives are presented for this course.

• Perform selected normal and emergency cockpit procedures as required for engine start/run-up, APU start, battery check, airplane taxiing, etc. (requires use of a simulator). The FlightSafety instructor will modify the stated overall objective conditions and criteria to satisfy selected performance requirements, when appropriate. The performance levels specified will not vary from those directed by the FlightSafety Director of Maintenance Training.

Given the Gulfstream G500/G550 Aircraft Maintenance Manual (AMM), class notes, and this training manual (as specified by the FlightSafety instructor), the client will be able to pass a written examination upon completion of this course to the grading level prescribed by the FlightSafety Director of Maintenance Training. The maintenance technician will be able to: • Outline the ATA 100 system of maintenance documentation, including the major chapter headings and symbology. • Describe the meaning and application of each piece of manufacturer’s maintenance documentation and use the documentation in practical applications. • Perform routine servicing. • Outline the recommended maintenance schedule applicable options and locate routine procedures in the manufacturer’s Maintenance Manual. • Locate major components without reference to documentation and other components with the aid of documentation. • Describe the operation of all major systems in the normal and various abnormal operating modes.

1-2




CHAPTER 2 ATA 100/PUBLICATIONS CONTENTS Page INTRODUCTION ................................................................................................................... 2-1 GENERAL............................................................................................................................... 2-1 AIR TRANSPORT ASSOCIATION (ATA) NUMBERING SYSTEM .................................. 2-3 General ............................................................................................................................. 2-3 Aircraft Maintenance Manual Format.............................................................................. 2-3 Revisions and Service Bulletins....................................................................................... 2-5 AIRCRAFT MAINTENANCE MANUAL ............................................................................. 2-9 General ............................................................................................................................. 2-9 Types of Information........................................................................................................ 2-9 Fault Isolation Manual ................................................................................................... 2-13 WIRING DIAGRAM MANUAL.......................................................................................... 2-15 General........................................................................................................................... 2-15 Equipment Locator ........................................................................................................ 2-15 Schematic Diagrams ...................................................................................................... 2-19 Wire Lists....................................................................................................................... 2-21 Termination Lists ........................................................................................................... 2-21 ADDITIONAL PUBLICATIONS ...........................................................................................2-21 Illustrated Parts Catalog................................................................................................. 2-21 Operating Manual .......................................................................................................... 2-21 Airplane Flight Manual.................................................................................................. 2-21


2-i



Structural Repair Manual............................................................................................... 2-21 Weight and Balance Manual .......................................................................................... 2-22 Master Minimum Equipment List.................................................................................. 2-22 Configuration Deviation List ......................................................................................... 2-22 OPERATOR COMMUNICATIONS..................................................................................... 2-23

2-ii




ILLUSTRATIONS Figure

Title

Page

2-1

ATA System Code Example ..................................................................................... 2-2

2-2

System Description................................................................................................... 2-8

2-3

Adjustment/Test Example ...................................................................................... 2-10

2-4

Fault Isolation Manual Example ............................................................................ 2-12

2-5

Equipment Locator Example (System-Referenced)............................................... 2-14

2-6

Equipment Locator Example (Aircraft Assembly-Referenced)............................. 2-16

2-7

Schematic Diagram Example................................................................................. 2-18

2-8

Schematic Symbols ................................................................................................ 2-19

2-9

Wire List Example ................................................................................................. 2-20

2-10

Termination List Example...................................................................................... 2-20


2-iii



CHAPTER 2 ATA 100/PUBLICATIONS 24

25

26

31

30

29

28

INTRODUCTION The purpose of this chapter is to describe the arrangement, numbering system, and special features of the Air Transport Association format for aircraft maintenance manuals. To take advantage of all the material presented in an ATA 100-format manual, the maintenance technician must become thoroughly familiar with the outline and contents presented for any given airplane. In addition, the various types of publications and operator communications for the Gulfstream G500/G550 aircraft are discussed.

GENERAL ATA Specification No. 100 is issued by the Air Transport Association of America as the Specification for Manufacturers’ Technical Data. It establishes a standard for the presentation of certain data produced by aircraft, engine, and component manufacturers required for the support of their respective products.

Under this format, the Aircraft Maintenance Manual is broken down into standard chapters as defined by ATA 100. Each chapter covers a specific area of maintenance information, such as Chapter 10, “Parking and Mooring”, or a specific system, such as Chapter 32, “Landing Gear”. All data pertaining to a given system is located within its chapter, regardless of whether it is mechanical, hydraulic, or electrical in nature. The chapters are arranged in alphabetical order.


2-1



27 . 51 . 05

CHAPTER/SYSTEM FLIGHT CONTROLS

SUBSYSTEM/SECTION FLAPS

UNIT/SUBJECT REMOVAL/INSTALLATION

Figure 2-1. ATA System Code Example

2-2




Besides the Aircraft Maintenance Manual, the maintenance technician uses many other publications and communiqués, such as the Wiring Manual, the Illustrated Parts Catalog,Fault Isolation Manual, and Service Bulletins, to perform aircraft maintenance and keep his knowledge up to date.

AIR TRANSPORT ASSOCIATION (ATA) NUMBERING SYSTEM

following items (as applicable) filed at the front: • Effectivity code cross-reference list • Highlights page(s) for each normal revision • List of effective pages • List of effective temporary revisions • List of service bulletins • Table of contents

Standard Numbering System GENERAL All maintenance publications are formatted according to the Air Transport Association (ATA) numbering system, which identifies chapter/system, subsystem/section, and unit/subject for each assigned item of equipment (Figure 2-1). The following information is a general discussion of the ATA 100 system and is not specific to Gulfstream Aerospace publications. The Aircraft Maintenance Manual (AMM) is prepared from the manufacturer’s technical data in accordance with the Air Transport Association and conforms to ATA 100 Revision 32. The AMM provides sufficient information to enable a mechanic who is unfamiliar with the airplane to service, test, adjust, and repair systems and to remove and install any unit normally requiring such action on the line or in the maintenance hangar.

AIRCRAFT MAINTENANCE MANUAL FORMAT Division of Subject Matter The introduction to the AMM lists the chapters from the ATA 100 format which are included in the manual. Each chapter has the

The numbering system identifies and segregates subject matter by chapter (system), section (subsystem), and subject (unit) (Figure 2-1). The system is a conventional dash-number breakdown, and each number is composed of three elements consisting of two digits each. When referred to as a unit, the three-element number (chapter/section/subject) is called the “chapter/section” number. The chapter/section number is located in the lower-right corner of each page with the page number and date. Each system, subsystem, and unit is allocated a block number. A page numbering system allows rapid location of information for retrieval. All maintenance information is separated into specific types of information (topics), and blocks of page numbers are assigned to each.

Chapter Numbering System The chapter numbering system provides a functional breakdown of the entire aircraft. It uses a three-element number, with the elements separated by dashes. Each element contains two digits, corresponding to chapter/ system, section/subsystem, and subject/unit.


2-3



Chapter/System The first pair of digits is assigned by ATA Specification 100 and designates the chapter/system. For example, 21-xx-xx identifies the air-conditioning system.

Section/Subsystem The second pair of digits designates the section/subsystem. Only the first digit will be assigned by ATA Specification 100. When the second pair is -00-, it shows the matter will be treated in general and applies to the chapter/system as a whole. The second digit of the pair is used when it is convenient to break down the section/subsystem.

When applicable, the effectivities of service bulletins are differentiated through the following indications: • Pre-Mod SB—Aircraft covered by the service bulletin effectivity that do not have the relevant modification(s) incorporated • Post-Mod SB—Aircraft whose operator has accomplished the service bulletin or that have the relevant modification(s) factory-incorporated The following page number blocks are used in the AMM: • 1 to 99 .... Description and Operation • 201 to 299 .... Maintenance Practices

For example, 21-20-xx identifies the air distribution subsystem, and 21-22-xx identifies the passenger cabin distribution subsystem.

• 301 to 399 ............................ Servicing • 401 to 499 ........ Removal/Installation • 501 to 599 ................ Adjustment/Test

Subject/Unit The third pair of digits designates a component or functions of chapters and sections covered by the previous elements.

• 601 to 699 .............. Inspection/Check • 701 to 799 ............ Cleaning/Painting • 801 to 899 .............. Approved Repairs

When the third pair is -00-, it shows that the matter will be dealt with in general and applies to the section as a whole, without treating specifics concerning components or functions.

Each new topic of information starts with page 001, 101, 201, 301, etc., and continues within the page numbering block as necessary; unused page number blocks are omitted.

As an example, 21-24-01 identifies the recirculation fans of the air-conditioning distribution subsystem.

Illustrations and tables use the same numbering system as the page block in which they appear—for example, Figure 403 is the third figure in the Removal/Installation topic. If an illustration requires more than one page unit, whether it is a foldout or multiple-sheet presentation, each page unit is assigned a sheet number.

Effectivity In the lower left corner of each page is information on effectivity. When a page applies to all aircraft, the word “ALL” is printed in the effectivity box. If the information does not apply to all aircraft, the particular aircraft to which the information does apply are specified. Effectivity differences are reflected within the text or figures through references, callouts, or even by adding specific page blocks.

2-4




REVISIONS AND SERVICE BULLETINS General ATA 100 allows the manufacturer a great deal of leeway or freedom in the area of AMM revisions and their dissemination. Virtually every aircraft manufacturer has a system different from any other manufacturer; some differences are great while others are barely noticeable, but all are intended to get maintenance information, routine or vital, to the field in a timely manner. Because changes, particularly new temporary changes, may be vital to ground and/or airborne safety, the maintenance technician should be thoroughly familiar with the methodology used by a particular manufacturer to incorporate changes into an AMM. The manufacturer’s methods are listed in detail in the AMM “Introduction” for a given airplane.

temporary revision. The changes in the temporary revisions will be incorporated in the first permanent revision following their release.

List of Effective Temporary Revisions Temporary revisions are recorded on the list of effective temporary revisions. The page has columns for writing in the temporary revision number and issue date.

Permanent Revisions General Permanent revisions are printed on white paper and are issued to qualified holders as required to update the AMM. Additions, deletions, or revisions to the text are identified on the text page by a black bar in the left margin of the page adjacent to the revision.

Letter of Transmittal Two types of revisions are issued for the AMM: permanent and temporary. Service bulletins are also issued and disseminate information which may be of a routine nature or require special handling and prompt compliance. When text or art in illustrations is revised, a black bar appears on the page outside the margin beside the revised, added, or deleted material. A bar beside the page number or the section title and the printing date indicates that neither the text nor the illustration has been changed but that the material has been relocated to a different page or a totally new page has been added.

A letter of transmittal accompanies each permanent revision published by the manufacturer. The letter gives filing instructions and the reason for issue. Listed in the filing instructions are the temporary revisions which are incorporated in the permanent revision. Those temporary revisions are removed from the manual.

List of Effective Pages A new list of effective pages is issued with each permanent revision. The AMM uses three codes to identify pages for update: • R—Revised (to be replaced) • D—Deleted (to be removed) • N—New (to be added)

Temporary Revisions General Temporary revisions are urgent in nature.They are printed on yellow paper and notify operators of changes or provide advance information on some equipment or modifications. A temporary revision is filed in the manual as instructed in the


2-5



Effectivity Code Cross-Reference List

Record of Service Bulletins

The aircraft serial number effectivity for an AMM page is listed in the lower-left corner of the page. It says “ALL” or gives a specific effectivity code. The code is listed on the effectivity code cross-reference list, which is printed on green paper and lists the specific aircraft serial numbers affected by that code.

The record of service bulletins is filed in the front of the AMM and provides columns which list the service bulletin identification number and the subject.

Highlights A highlights page is printed on white paper and is issued with each permanent revision. It lists the pages in each chapter which are changed and the reasons for change. It also states “No revised page for this revision” if a permanent change does not affect that particular chapter.

Record of Revisions The record of revisions is filed in the front of the AMM and provides a place for the responsible individual to record the successive revision numbers, dates inserted, and his initials against the appropriate revision number. If the revision is inserted by the factory for a reprint of the manual, the revision record shows the revisions already incorporated.

Service Bulletins General Service bulletins are printed on either white or blue paper. White paper indicates routine handling with a specified time limit for compliance. Blue paper indicates special handling with a specified time limit for compliance, which may be immediate. This information is incorporated in the normal revisions.

Service Bulletin List The service bulletin list has columns which give the service bulletin number, the revision in which it is incorporated, and the service bulletin subject.

2-6


NOTES



NOTES

NOTES


2-7



Figure 2-2. System Description

2-8




AIRCRAFT MAINTENANCE MANUAL GENERAL The primary publication used for maintenance is the Aircraft Maintenance Manual (AMM). The purpose of the AMM is to acquaint maintenance technicians with the systems and components of the Gulfstream G500/G550 and to direct them in the proper procedures for maintaining the aircraft in an airworthy condition. All Gulfstream G500/G550 maintenance publications are also available as electronic manuals (CD ROM).

The System Description section is provided for each airframe and powerplant system and describes the system on multiple levels necessitated by the system. It provides a central location for the description of all the aircraft systems, including location, configuration, function, operation, and control of the complete system and its subsystems. As an example of the function of this section, the nosewheel steering system description includes its purpose and a general description, the major subsystems, outstanding system features, and a description of system operation, as well as an operational summary.

NOTES

NOTE Only the installations made in the aircraft during manufacture have been reflected in this manual.

TYPES OF INFORMATION The AMM provides two types of information: system description, fault isolation, and maintenance practices.

NOTE The Engine Maintenance Manual will be incorporated in the G u l f s t re a m G 5 0 0 / G 5 5 0 A i rc ra f t Maintenance Manual. The Structural Repair Manual also contains the procedures for corrosion prevention and corrosion treatment.

System Description The first type of information contained in the AMM is System Description. This information is located in page block 1 to 99 and is used by maintenance technicians to quickly gain an overview knowledge of any particular system.


2-9



Figure 2-3. Adjustment/Test Example

2-10




Maintenance Procedures

NOTES

The second type of information is Maintenance Procedures, which are located in page block 201 to 299. The Maintenance Procedures are used by maintenance technicians to perform daily aircraft maintenance and servicing. Pa g e b l o c k 2 0 1 t o 2 9 9 i s u s e d w h e n a l l subtopics of Maintenance Procedures are relatively brief. When individual subtopics become so lengthy as to require a number of pages, the following page number blocks are used (Figure 2-3): • 301 to 399 ............................ Servicing • 401 to 499 ........ Removal/Installation • 501 to 599 ................ Adjustment/Test • 601 to 699 .............. Inspection/Check • 701 to 799 ............ Cleaning/Painting • 801 to 899 ................................ Repairs


2-11



Figure 2-4. Fault Isolation Manual Example

2-12




FAULT ISOLATION MANUAL

NOTES

Even though the Gulfstream G500/G550 Fault Isolation Manual is a stand-alone publication, it is a vital part of the AMM and has been prepared in accordance with Air Tr a n s p o r t a t i o n Association ( ATA ) Specification No. 100, Revision 32. The Fault Isolation Manual provides a central location for the detailed breakdown of the various aircraft systems fault isolation procedures; it also includes AMM references to assist the technician in the resolution of system malfunctions. Figure 2-4 is an example of a fault isolation index for AC electrical load distribution, for an AC crosstie bus failure. The Fault Isolation Manual is written for use by experienced technicians and contains troubleshooting trees needed to isolate system problems to the LRU level. Fault isolation is used by maintenance personnel to locate and determine the cause of any particular maintenance malfunctions and the possible solution.


2-13

2-14 WIRING DIAGRAM

Detail F

Detail H

RIGHT MAIN WHEEL STRUT

RIGHT ENGINE

Detail D RIGHT ELECTRONIC EQUIPMENT RACK (REER)

REF DES 053S1 053S2 292DS5 292DS6


d) e) f) g)

2D 2C

3E 3D

a) b) c)

PEDESTAL AREA a) b) c) d)

REF DES REER-C27 140A1 155A2 A7+8 031A1 031A2 031A3 035A2

NOMENCLATURE Left Fire Handle Switch Right Fire Handle Switch Left Fuel Shutoff Switch Right Fuel Shutoff Switch

1D 1E 1C

1F

2A

COCKPIT OVERHEAD PANEL (COP) REF DES a) 031DS1 b) 031DS2 c) 031DS3

NOMENCLATURE APU Generator Switch Right Generator Switch Left Generator Switch

Detail A

NOMENCLATURE Right Main Gear Downlock Sw

4E

5E

6E

7E

8E

9E

10E

11E

12E

4D

5D

6D

7D

8D

9D

10D

11D

12D

3C

4C

5C

6C

7C

8C

9C

10C

11C

12C

3B

4B

5B

6B

7B

8B

9B

10B

11B

12B

2B

1G

2A 1A

REF DES a) 112S2

NOMENCLATURE R GCU PWR c/b Annun Lights Dim & Test Box Modular Avionics Unit #2 Dual Generic I/O 2 Module APU Generator Control Unit Right Generator Control Unit Right Bus Power Control Unit Right Power Distribution Box

REF DES a) 031G2

14D 14F

14C

14E

NOMENCLATURE Right Integrated Drive Generator

14K

14J

14A 3A

4A

5A

6A

7A

8A

9A

10A

11A

12A

14B

14H

LEFT ELECTRONIC EQUIPMENT RACK (LEER) REF DES LEER-G11 LEER-G16 031A5 031A4 155A1 A9+10 f) 035A1 g) 181K6EL a) b) c) d) e)

NOMENCLATURE APU GCU PWR c/b L GCU PWR c/b Left Bus Power Control Unit Left Generator Control Unit Modular Avionics Unit #1 Dual Generic I/O 1 Module Left Power Distribution Box APU Ready to Load Relay (347 Pnl)

Detail C

APU ENCLOSURE (TAIL) REF DES a) 031G3

Detail I

LEFT MAIN WHEEL STRUT REF DES a) 112S1

NOMENCLATURE Left Main Gear WOW Switch

Detail E

NOMENCLATURE APU Generator

LEFT ENGINE REF DES a) 031G1

NOMENCLATURE Left Integrated Drive Generator

ELECTRICAL POWER SOURCE Aircraft: 5001-9999 Equipment Locator

Figure 1, Sheet 1 of 1

Figure 2-5. Equipment Locator Example (System-Referenced)

international

24-00-00

FlightSafety

Detail G


Detail B



WIRING DIAGRAM MANUAL GENERAL Many factors have influenced the design of the Gulfstream G500/G550 Aircraft Wiring Manual, among them page size, system complexity, and system design.The requirement to provide the maintenance technician with an efficient and precise set of informational tools resulted in an electrical system presentation unlike any that Gulfstream has presented in the past. This manual is formatted to ATA 100 specifications and provides detailed illustrations of the aircraft electrical systems and how they interface with other aircraft systems. The purpose of the Gulfstream G500/G550 Aircraft Wiring Manual is to provide all the aircraft systems wiring information needed to perform troubleshooting, fault isolation of the electrical circuits and the repair of specific electrical systems and components. The complexity of Gulfstream G500/G550 electrical/avionics systems has prompted non-conventional methods of presentation. Systems are represented in four complementary drawing types:

EQUIPMENT LOCATOR The equipment locators (EL) are designed to provide navigation to component locations throughout the aircraft. The locator indicates an approximate area, such as cockpit overhead or tail compartment. Equipment locators fall into two general categories: system-referenced and aircraft assembly-referenced.

System-Referenced The system-referenced equipment locators (Figure 2-5) precede system schematics. They list those components, which by virtue of their reference designation and function are essential components of the system. Also cont a i n e d o n t h e s y s t e m - l eve l l o c a t o r s a r e references and general locations of separate system major components, which interface with the system being presented. An often-included example is the Annunciator Lights Dim and Test Box, which appears on the locator of each system providing input to, or receiving output from, this component (most system drawings are of this category).

1. Equipment Locator 2 Schematic Diagram 3. Wire List 4. Te r m i n a t i o n L i s t ( r e p r e s e n t e d i n Chapter 91) Each of these drawing types is designed to provide the maintenance technician with information essential to maintain the wiring and component integrity of the aircraft systems. All drawing types contained within the wiring manual are designed for ease of use and understanding. Symbology is included to clearly indicate the logic states (ON/OFF, energized/de-energized, etc) and the interconnections of all included systems. Refer to the illustrations through the remainder of this section for specific examples. FOR TRAINING PURPOSES ONLY

2-15

2-16 WIRING DIAGRAM

Left Junction and Relay Panel – 347A1

Equipment Locator

91-02-10 Figure 1, Sheet 1 of 1

international

FlightSafety

Aircraft: 5001-9999


FOR TRAINING PURPOSES ONLY 1)

LEFT JUNCTION AND RELAY PANELS

Figure 2-6. Equipment Locator Example (Aircraft Assembly-Referenced)



Aircraft Assembly-Referenced

NOTES

Aircraft assembly-referenced equipment locators (Figure 2-6) include all junctions and connections, which are not specific to any one system. Aircraft interconnections are referenced by the junction/relay panel or connector cluster into which they are installed. All assembly-referenced equipment locators are located in Chapter 91 of the Gulfstream G500/G550 Aircraft Wiring Manual.

Panel Assembly Equipment Locator The panel assembly locator serves to provide a visual presentation of panel components and connectors. Components such as relays which have a system designation are identified by nomenclature and the drawing(s) upon which they appear. Full reference designation is provided for those components whose decal does not provide such. Schematic diagrams and the electrical equipment list do not include the relay socket designation “X” which is inserted in the relay reference designation. Example: 032K3 = Relay 032XK3 = Relay Socket


2-17

2-18 035A2P1A5 P113-20

L BATT CONT LEER - G15 L BATT BUS A To 24-60-00 DC PWR DIST

5

P82-16

1 A

ON

3

P101-22

CH 122

39

1P156-22

7 8 10

9

P101-22

11 J C F


29

140A1

EXT BATT SW ON

14

1

NC2 1

2

2

ENERGIZED

P143-20

035A2P1A5

P143-20

P143-20

A1

P144-20

L BATT CONTACTOR STAT

28

1P109-22/24

L BATT CHARGER FAIL

27

1P135-22/24

P143-20

C2

P114-20/22

G H J

P143-22

032K7P1

DUAL GENERIC I/O 1 155A1

155A1A9+10

Loc: LEER

+

POS

P143-22

A3

P144-22

B6

Loc: REER

LEDXC 1

A1

BATT CHARGER OUTPUT +

1P124-8

NEG

BATT CHARGER OUTPUT -

1P125-8N 032BC1P1

FAIL ANNUN OUT

9

1P135-22/24

CHGR MODE CONT

8

1P108-20/22

GROUND

2

1P134-20N

OVERTEMP SWITCH

1

1P130-20

OVERTEMP SWITCH RTN

3

1P129-20

THERMISTOR RETURN

12

1P131-20 (BLU)

THERMISTOR INPUT

11

1P132-20 (WHT)

E4120A

1

2

1P119-16/20

E4121D

E

1

2

NC3

12

OVERTEMP SWITCH

5

1P121-22N

A

CB369 15 P97-12 LEFT BAT BUS B 385 PNL (Tail)

F

X2

A1 A3

CB368 P96-14 LEFT BAT BUS A 385 PNL (Tail)

L ESS DC BUS

032K3 L BATT BUS B RELAY Loc: 347A1

X1

P143-20

035A1P1B6

P107-12

P167-20N

A2

R BAT BUS B To 24-60-00 DC PWR DIST

3 3 3

L BAT BUS B To 24-60-00 DC PWR DIST

8

1P130-20

6

H797-20

H797-20

E

SHEET 6

P133-20

P133-20

D

SHEET 6

H797-20

3

L ESS DC CONT #2 COPILOT - G7 R ESS DC CONT #1 COPILOT - G8 R ESS DC CONT #2 PILOT - G8

P80-20

F

P87-20

G

F

A

A9-

A

1P117-22N

385A1E3C

B

1P116-22N 1P109-22/24

1P110-22

1

P112-22 P113-20

P113-22 P111-22

P111-22

P112-20

032K2P1 P112-22 P111-22

032K2P1

RELAXED

15 11

17 13

P114-22 P113-22

H

1P119-22

3

1P118-22N

5

+ -

385A1E1A

G

B

P111-20

BATTERY CONTACTOR #2 032K2

P128-00

Loc: 385A1

D

F BATT TIE BUS

1P127-00 P133-20

K

E

385TJ1C Loc: 385A1

Q182-20

E

BATTERY CONTACTOR #1

2

032K1

032HE7

P133-20

E4159A

SHEET 6

To 49-60-00 APU CONTROL

1) 2) 3) 4)

Left Battery Bus Control Battery Contactor #1 Left Battery Charger Left Battery DC POWER SOURCE

D

A

C

B

Schematic 032HE7P1

24-30-00 Figure 1, Sheet 5 of 7

F

SHEET 7

Figure 2-7. Schematic Diagram Example

international

Loc: Tail

Loc: 385A1

G

Aircraft: 5001-9999

LEFT BATTERY 032BT1

Loc: LEER

19 21 18 20 16 14 15 17 11 13

HALL EFFECT SENSOR 1P140-00N

031A5

ENERGIZED

P81-20

Loc: 385A1 FROM 49-60-00 APU CONTROL 28 VDC FROM APU CONT #1 AND/OR APU CONT #2 WHEN STARTER ENGAGED SIGNAL IS OUTPUT BY APU ECU

LEFT BUS POWER CONTROL UNIT

P123-00 T4

Loc: LEER

385A1E1A

P112-22

385TJ1C

P128-00

L ESS DC Battery Contactor

1P330-22 P114-22

P112-20

P111-20

15E LEDBC AUX RTN

032K1P1

Loc: 385A1 P88-20

P235-22 (BLU)

L POWER DISTRIBUTION BOX

385TJ1B B

15F LEDBC AUX

8 T3

LEDCB

1P108-20/22

A

P234-22 (WHT)

A1

A1+

035A1

To 24-60-00 DC POWER DISTRIBUTION

1P330-22

L ESS DC CONT #1 PILOT - G7

A1

A2

1

FlightSafety

1P129-20

1 NC4

E2165A

To 24-60-00 DC POWER DISTRIBUTION

10

031A5P1A

035A1P1A5 2

1P108-20/22 L BAT BUS B To 24-60-00 DC PWR DIST

1P131-20 (BLU)

NEG

032K5 L BATT CONTROL RELAY Loc: 347A1

Y1

G

1P132-20 (WHT)

032BT1P1 POS 1P127-00

3

1P153-20/22

L BATT CHGR CONTACTOR

9

4

X1 X2

1

NC1 ENERGIZED

032K7 Loc: 385A1

032BT1P2

11

1P330-22

385A1E1A

Loc: Tail

THERMISTOR

1P115-22N

19

347A1E3A G P120-20N

F

D

P144-22

2

NC2

385A1E3C

P102-10

L BATTERY CHARGER 032BC1

P82-16 P322-20N

LEDC

RELAXED

2

NC1

A3

21

Loc: REER

A2

A2 P82-16

H

RIGHT BUS POWER CONTROL UNIT 031A3

035A1P1A6

APU STARTER ENGAGED

347A1E3A

REDBC AUX

T2

REDCB

035A2

Loc: TAIL 385TJ1L

C3 C1 1P119-20

181K2 APU PILOT RELAY Loc: 381A1 (Ref: 49-60-00)

1

15E REDBC AUX RTN

R POWER DISTRIBUTION BOX

P82-16

1

15F

P280-22 (BLU) P123-00

035A1P1B6 APU STARTER NOT ENGAGED

4

P279-22 (WHT)

R ESS DC Battery Bus Contactor

A9-

E2166A

085K3 AUX HYD PUMP CONT RELAY #1 Loc: 381A1 (Ref: 29-21-00)

L M K

5

16

031A3P1A

035A2P1A5

25 20

A2

A1+ P199-20N

PUMP ON

Loc: TAIL 385TJ1L

C

A2

1

NC3

1 NC4

A1

A3 PUMP OFF

13

R ESS DC BUS

1

NC1 2

NC1

P143-20

REDXC

RELAXED

2

2

1P153-20/22 P146-20

6 P143-20 OFF 032S1 EXTERNAL BATTERY SWITCH Loc: Radome

A U

REDC

1P110-22

P143-20

15

P145-22/24

23

Loc: REER

14 16 1P154-20N C ON L MAIN BATT 318A1E5C SWITCH 032DS1 Loc: COP 10 ON 11 12 P82-20 P143-20 P143-20

PS 1B CH B M

24

P143-20

ANNUN LTS DIM & TEST

SHEET 6

1P151-22/24

155A1A10P4 155A1A10P4E1 LEFT BATT SW OFF

12

Q182-20

64

O/GND

WIRING DIAGRAM

P144-20


6

P82-20 Q182-22

140A1P1C

TEST DIM

O/GND

OFF 5

Loc: Tail

P R O C 2

P1C

2

A

381TJ1K

P S 1 B

W154-22

4

Q182-20

PS 1A CH A

From 33-11-00 ANNUNCIATOR LIGHTS

A

FROM 29-21-00 AUXILIARY HYDRAULIC PUMP CONTROL 28 VDC WHEN AUXILIARY HYDRAULIC PUMP IS ON

2

3

A

L BATT CHGR CONT LEER - J14 L ESS 28 VDC To 24-60-00 DC PWR DIST

23

035A2P1A6



SCHEMATIC DIAGRAMS Schematic diagrams (Figure 2-7) represent all Gulfstream G500/G550 systems. The schematics are designed to functionally describe the syst e m s w i t h a m i n i m u m o f o ff - s h e e t o r off-drawing reference. Entire circuits are repeated on different system drawings when circuit function and operation is significant to more than one system. The schematic diagrams are designed to be used in conjunction with their respective wire lists. Schematics do not show the wires’ segment letters. These are found on the wire lists only. Schematics function primarily to perform initial system troubleshooting, though in many cases they will be all that is necessary to fault isolate the system. Due to the compact nature of the system schematics, they are also highly effective teaching aids.

The schematics depict all active components (switches, relays, circuit breakers, etc.) within the subject system (Figure 2-8). All terminal junctions are also illustrated. With the exception of vendor-supplied disconnects and cables, no aircraft assembly-referenced connectors are shown. Aircraft assembly connectors are purposely omitted so that entire system functions can be shown on a single sheet. All G500/G550 Wiring Manual schematic circuits and relays are shown in the de-energized (relaxed) state. Annunciator switches are depicted in the extended position. Toggle type switches are shown in the OFF position. Threeposition switches are ordinarily shown in the center-resting position.

ON-SHEET, OFF-SHEET & OFF-DRAWING REFERENCES

CIRCUIT BREAKERS

1 L BATT BUS A TO: 24-60-00 DC PWR DIST

L BATT CONT LEER-G15

5

All circuit breakers presented indicate the electrical bus which provides primary power and the source's system drawing. All circuit breakers are presented with the names as they appear on the circuit breaker panel, the circuit breaker panel name, and the grid location of the circuit breaker. RELAYS

C2

APU STARTER NOT ENGAGED C3 C1

APU STARTER 181K2A ENGAGED APU PILOT RELAY LOC: 381A1 (REF: 49-60-00)

P120-20N P143-20 P37-12

X2

032K3R L BATT BUS B RELAY LOC: 347A1

X1 A1 A3

P107-12 A2

Relays may be depicted in several alignments, thought their logic (de-energized, relaxed) remains constant throughout the schematics. Energized and de-energized coil and contact logic states (on/off, air/gnd, etc.) are included to minimize the need to consult other drawings. When partial relays are shown, as is the case with relay 181K2A above, the system schematic number for the relay is indicated.

TO; 032K1P1-20 (THIS SHEET)

TO: 105A3P1A-15 (THIS SHEET)

1

On-sheet references are sometimes used to reduce the clutter of wires on complex drawings. The on-sheet reference is a triangular figure with a number in it. The corresponding, matching reference is the same shape and number. The point of the triangle always points in the direction of the matching reference. FROM 181K2-D1 (APU PILOT RELAY): 28VDC FROM APU CONT NO. 1 AND/OR APU CONT NO. 2 WHEN STARTER ENGAGED SIGNAL IS OUTPUT BY THE APU ECU (SEE 49-80-00): APU CONTROL

H797-20

Off-drawing references point to another system drawing. These references include, at a minimum, the drawing ATA designation and name of the system. If the referenced wire provides logic to the circuit being depicted, a text block will explain the source of the signal and what controls the logic states. Ann Lts Pwr and Warn Lts Pwr references include circuit breaker and location. The majority of off-drawing references include the system ATA and sheet number of the corresponding drawing. G

TO SHEET 6 032BT2P2-4

TO SHEET 5 032BT1P2-4

G

Off-sheet references are indicated by this symbol. The continuation of the signal path on the indicated sheet is the same symbol and letter.

Figure 2-8. Schematic Symbols


2-19



From Wire Number

Item

From Ref Des.

Pin No.

Item

Ref Des.

Pin No.

Cable Number

Effectivity

X9A20

SOCKET

040XKT

A2

STUD, GND

E4018E

NOPIN

5001-9999

1P108A22

CONN, PLUG

032K1P1

16

CONN, RCPT

385A1J1

D*

5001-9999

1P108B20

CONN, PLUG

032BC1P1

8

CONN, PLUG

385A1P1

*

5001-9999

1P109A22

CONN, PLUG

032K1P1

20

CONN, RCPT

385A1J1

C*

5001-9999

1P109B22

CONN, PLUG

385A1P1

C*

CONN, PLUG

4000P9

35

5001-9999

1P109C22

CONN, PLUG

2000P9

35

CONN, RCPT

4000J9

35

5001-9999

1P109D22

CONN, RCPT

2000J9

35

CONN, PLUG

331A1P35

10

5001-9999

1P109E24

CONN, PLUG

155A1A10P4

28

CONN, RCPT

331A1J35

10

5001-9999

Figure 2-9. Wire List Example

From Item

Wire No.

328CB354-2 328CB11-2 328CB50-2 328CB280-2 328CB262-2 328CB296-2 328CB34-2 328CB335-2 328CB333-2 328CB332-2 328CB328-2 328CB326-2 328CB24-2 328CB306-2 328CB198-2 328CB197-2 328CB73-2 328CB277-2 328CB278-2 328CB279-2 328A1E3B 328CB340-2 328CB20-2 328CB321-2 328CB320-2 328CB323-2 328CB365-2 328C B315-2

2F15A22 L230A20 G50A20 P33A22 2F27A20 F10A22 2H177A20 2C44A20 1C44A20 C48A20 W172A22 W170A22 2SA2A20 E104A22 1E20A22 1E21A22 W140A22 2X238B22(WHT) 2X239B22(BLU) 2X240B22(ORN) 2X241B22(GRN) FD59A22 2L90A22 2E20A22 2E21A22 2Q280A22 FR51A22(WHT) 2Q281A20

Pin No. 2 3 4 5 7 8 9 10 11 12 13 14 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30

To Item

Ref ATA

2F15B22 L230B20 5G0B20 P33B22 2F27B20 F10B22 2H177AB20 2C44B20 1C44B20 C48B20 W172B22 W170B22 2SA2820 E104B22 1E20B22 1E21B22 W140B22 2X238C22(WHT) 2X239C22(BLU) 2X240C22(ORN) 2X241C22(GRN) FD59B22 2L90B22 2E20B22 2E21B22 2Q280AB20 2Q281AB20 FR52B22(WHT)

30-30-00 33-30-00 32-50-00 24-40-00 30-30-00 30-30-00 21-60-00 27-50-00 27-50-00 27-50-00 31-50-00 31-51-00 34-43-00 77-31-00 28-40-00 28-40-00 52-70-00 24-20-00 24-20-00 24-20-00 24-20-00 31-60-00 31-41-00 28-40-00 28-40-00 28-26-00 28-26-00 31-31-00

Figure 2-10. Termination List Example

2-20


Remarks



WIRE LISTS

OPERATING MANUAL

Each individual wire “belonging” to a system is found on the wire list. The graphical pointto-point wire list displays the source, all interim disconnects (connectors) and the destination. All junctions (splices and terminal junction modules) are also shown. Twisted or twisted-shielded wires are indicated by specific symbols, and the cable name (i.e., 2/203, 3S1, etc.) is provided for each segment. All components which terminate a wire are shown, with the component’s reference designation and nomenclature clearly indicated. Figure 2-9 provides a sample wire list.

The purpose of the Operating Manual is to provide the flight crew with the aircraft’s necessary operating limitations, procedures, performance, and systems information required for safe and efficient operation..

TERMINATION LISTS The termination lists (Figure 2-10) graphically depict wire connections for connector and panel assembly components. Wires that terminate at interim or panel assembly-designated connections are found in Chapter 91 of the Gulfstream G500/G550 Aircraft Wiring Manual.

ADDITIONAL PUBLICATIONS As an aid to the Aircraft Maintenance Manual, Gulfstream Aerospace also publishes other documentation.

The Operating Manual serves as a comprehensive reference for use during transition and recurrency training and proficiency checks on the aircraft. The manual provides necessary operational data from the FAA-approved G u l f s t re a m G 5 0 0 / G 5 5 0 A i r p l a n e F l i g h t Manual and standardized procedures and practices to enhance airplane operation. Maintenance technicians use the Operating Manual for engine run-up and taxi operations.

AIRPLANE FLIGHT MANUAL The Gulfstream G500/G550 Airplane Flight Manual (AFM) serves as a comprehensive reference for all flight operations. The AFM provides operators with numbered sections that contain limitations, procedures, and performance data for the aircraft and aircraft systems. All performance limitations and information listed are in compliance with FAA regulations, Part 25, and must be on board the aircraft for all flight operations.

ILLUSTRATED PARTS CATALOG

STRUCTURAL REPAIR MANUAL

The Illustrated Parts Catalog provides a pictorial/part number breakdown of the aircraft and ground support equipment. The Illustrated Parts Catalog is the only approved part number listing for the aircraft. Part effectivity is provided via aircraft serial number ranges or by notes at the bottom of the particular parts list page.

The Structural Repair Manual provides information for general repairs of simple and common structural components, repair materials, and their specifications and processes.

Maintenance technicians use the Illustrated Parts Catalog to locate and determine the proper part number when replacing and/or inspecting hardware and components. The Illustrated Parts Catalog can also be used in conjunction with the AMM system descriptions to better understand a particular system.

The Structural Repair Manual includes general repair procedures that must be performed concurrently with structural repairs (such as sheet metal forming, fastener installation, corrosion treatment, and sealing), along with required skin thickness diagrams, detailed illustrations, and part number identification.


2-21



WEIGHT AND BALANCE MANUAL The purpose of the aircraft Weight and Balance Manual is to provide the operator with guidelines to ensure that the maximum weight and center-of-gravity limits are not exceeded during operation. The Weight and Balance Manual provides weight and balance data, loading graphs and CG envelopes, manufacturer’s bare empty weight, operating gross weight (in pounds), typical loading summary, outfitting weight allowance, and aircraft weighing procedures.

MASTER MINIMUM EQUIPMENT LIST The Master Minimum Equipment List (MMEL) is initiated by Gulfstream on their equipment and the number of components allowed to be inoperative to release an aircraft for flight operations. All manufacturer’s MMELs are approved by the FAA, per the operating FAR. The maintenance activity then reviews the MMEL and submits a Minimum Equipment List (MEL) to the local FSDO for approval. The MEL then becomes a legal means to release the aircraft with an inoperative system or component. The MEL must remain with the aircraft at all times and will include instructions on its use by the maintenance technicians. The Maintenance Operational Placarding (MOP) procedures manual is a supplement to the MEL that contains maintenance, operational, and placarding procedures. The MMEL from the aircraft manufacturer provides guidelines for operators to develop an individual aircraft MEL.

2-22

NOTE The MEL is intended to permit safe aircraft operation with inoperative items for a period of time until repairs can be accomplished. Absolute compliance is required.

CONFIGURATION DEVIATION LIST The Configuration Deviation List (CDL) allows the aircraft to maintain safe flight operations without certain parts, as listed. The CDL provides an additional certification limitations listing of the type and number of components allowed to be inoperative for safe flight operations and lists required placarding and logbook entry information for maintenance technicians and aircrews when operating under CDL limitations. The CDL is used to determine operating limitations due to missing or removed aircraft equipment and imposes performance penalties on the aircraft. The CDL does not include parts which do not affect the airworthiness of the aircraft and is based on the aircraft’s configuration as originally manufactured.

NOTE All Configuration Deviation List penalties and limitations are approved by the FAA and must be followed when operating with a configuration deviation. The CDL is located in Appendix B of the Gulfstream G500/G550 Airplane Flight Manual.




OPERATOR COMMUNICATIONS The purpose of Gulfstream G500/G550 operator communications is for Gulfstream Aerospace Corporation to provide continuous written communication with Gulfstream aircraft operators concerning aircraft maintenance and operations. Operator communications are relayed via the following:

The application of operator communications enables operators and maintenance technicians to maintain the aircraft to the latest Gulfstream-suggested configuration, continually improve maintenance practices, and reduce maintenance downtime.

NOTES

• Service News • Maintenance and Operations Letters • Service Bulletins • “Gulfstream Intercom” Service News provides general service news, information on improved parts, and operator experiences. Maintenance and Operations Letters provide general information and items of interest concerning the aircraft. Service Bulletins provide aircraft feature or design changes, incorporating a statement of importance, inspection requirements, maintenance not covered in maintenance manuals, a strict time compliance, and timely information of major importance. Response to a Service Bulletin is accomplished via a Service Reply Card. “Gulfstream Intercom” (provided on website www.Gulfstream.com) provides a weekly communication of topics to Gulfstream owners, operators, and employees and include current events, upcoming events, and technical updates.

NOTE The technical content of operator communications is for information only and is not to be used in the maintenance or service of any Gulfstream aircraft.


2-23



CHAPTER 5–12 AIRCRAFT GENERAL CONTENTS CHAPTER 5 ............................................................................................................................ 5-1 INTRODUCTION ................................................................................................................... 5-1 GENERAL............................................................................................................................... 5-1 DESIGN CHARACTERISTICS ............................................................................................. 5-2 Aircraft Performance..........................................................................................................5-2 G500 Aircraft Weights ..................................................................................................... 5-2 G550 Aircraft Weights ..................................................................................................... 5-2 Cabin Standards ............................................................................................................... 5-3 Differences from Previous Gulfstream Aircraft............................................................... 5-7 AIRCRAFT SYSTEMS........................................................................................................... 5-9 Airframe Innovations ....................................................................................................... 5-9 Engines............................................................................................................................. 5-9 Auxiliary Power Unit..................................................................................................... 5-11 Hydraulics...................................................................................................................... 5-13 Flight Controls....................................................................................................................... 5-13 Landing Gear ................................................................................................................. 5-13 Environmental Control System...................................................................................... 5-15 Fuel System.................................................................................................................... 5-15 Electrical Power ............................................................................................................. 5-17 G500/G550 Avionics Equipment Functions .................................................................. 5-19 AIRCRAFT PLACARDS...................................................................................................... 5-21 General........................................................................................................................... 5-21 Caution and Warning Placards....................................................................................... 5-21


5-i



Instruction and Information Placards............................................................................. 5-21 Locator Placards ............................................................................................................ 5-21 CHAPTER 12—SERVICING............................................................................................... 12-1 INTRODUCTION ................................................................................................................. 12-1 GENERAL ............................................................................................................................ 12-1 JACKING, LIFTING, AND SHORING ............................................................................... 12-3 Operational Requirements ............................................................................................. 12-3 Individual Nose and Main Gear Jacking........................................................................ 12-5 Fuselage Jacking ............................................................................................................ 12-7 Lifting and Shoring........................................................................................................ 12-9 PARKING, STORAGE, AND MOORING ........................................................................ 12-11 Parking Procedures ...................................................................................................... 12-11 Storage ......................................................................................................................... 12-13 Mooring ....................................................................................................................... 12-15 LEVELING AND WEIGHING .......................................................................................... 12-17 Leveling ....................................................................................................................... 12-17 Weighing...................................................................................................................... 12-19 TOWING............................................................................................................................. 12-19 SERVICING........................................................................................................................ 12-25 Fuel System Servicing ................................................................................................. 12-25 Oil System Servicing .................................................................................................. 12-27 Hydraulic System Servicing ........................................................................................ 12-35 Pneumatic System Servicing ....................................................................................... 12-39 Oxygen System Servicing ........................................................................................... 12-43 Anti-Icing/Deicing....................................................................................................... 12-44 Aircraft Washing.......................................................................................................... 12-45

5-ii





Title

Page

5-1

Aircraft Dimensions ................................................................................................. 5-4

5-2

Emergency Exits....................................................................................................... 5-6

5-3

Rolls-Royce Deutschland BR710 Engine ................................................................ 5-8

5-4

Engine Instrument Displays ..................................................................................... 5-9

5-5

Auxiliary Power Unit ............................................................................................. 5-10

5-6

APU Synoptic Page............................................................................................... 5-11

5-7

2/3 Hydraulic Synoptic Page.................................................................................. 5-12

5-8

Flight Control Locations ........................................................................................ 5-12

5-9

Flight Control Synoptic Page ................................................................................. 5-13

5-10

ECS/PRESS Synoptic Page.................................................................................... 5-14

5-11

Fuel System Synoptic Page .................................................................................... 5-14

5-12

Gulfstream G500/G550 Electrical Schematic........................................................ 5-16

5-13

AC and DC Synoptic Pages.................................................................................... 5-17

5-14

PlaneView System Overview ................................................................................. 5-18

5-15

Flight Displays ....................................................................................................... 5-19

5-16

Locator Placard—Emergency Window Exits ........................................................ 5-20

12-1

Aircraft Jacking...................................................................................................... 12-2

12-2

Axle Jack Provisions .............................................................................................. 12-4

12-3

Main Landing Gear With Jack Adapter ................................................................. 12-5

12-4

Fuselage Jack Point Locations ............................................................................... 12-6

12-5

Locations of Pneumatic Bags and Stabilizing Line Attachments .......................... 12-8

12-6

Landing Gear Pins................................................................................................ 12-10

12-7

Protective Covers ................................................................................................. 12-12


5-iii



12-8

Tire Covers........................................................................................................... 12-13

12-9

Battery Disconnect............................................................................................... 12-13

12-10

Mooring Rings ..................................................................................................... 12-14

12-11

Longitudinal Leveling Brackets........................................................................... 12-16

12-12

Lateral Leveling Brackets .................................................................................... 12-17

12-13

Torque Link Disconnect....................................................................................... 12-18

12-14

Nose Strut Extension............................................................................................ 12-20

12-15

Nosewheel Steering Collar................................................................................... 12-21

12-16

Nose Wheel Well Parking Brake Accumulator Gage .......................................... 12-22

12-17

Tow Bar Attachment ............................................................................................ 12-23

12-18

Single-Point Pressure Refueling .......................................................................... 12-24

12-19

Overwing Fueling ................................................................................................ 12-24

12-20

Ground Service Control Panel ............................................................................ 12-26

12-21

Engine Oil Tank Location .................................................................................... 12-26

12-22

Oil Tank Sight Gage............................................................................................. 12-28

12-23

APU Gearbox....................................................................................................... 12-28

12-24

Remote Oil Replenishment System ..................................................................... 12-30

12-25

Air Turbine Starter ............................................................................................... 12-32

12-26

Hydraulic System Schematic ............................................................................... 12-34

12-27

Hydraulic Reservoir ............................................................................................. 12-35

12-28

Remote Hydraulic Replenishing System ............................................................. 12-36

12-29

Landing Gear Strut Filler Valves ......................................................................... 12-38

12-30

Emergency Extension Blowdown Bottles............................................................ 12-40

12-31

Oxygen Service Panel .......................................................................................... 12-42

5-iv




CHAPTER 5 AIRCRAFT GENERAL

INTRODUCTION This training manual presents a description of the major airframe systems and engines installed on the Gulfstream G500/G550 aircraft. The information contained herein is intended only as an instructional aid. This material does not supersede, nor is it meant as a substitute for, any of the manufacturer’s operating manuals. The material presented has been prepared from the basic design data. All subsequent changes in aircraft appearance or system operation will be covered during academic training and subsequent revisions to this manual.

GENERAL The Gulfstream G500/G550 is a low-wing, twin fan-jet, pressurized transport category airplane, specifically designed for all-weather operations and certified to fly at altitudes up to 51,000 feet. The minimum crew required is a pilot and copilot. Many aircraft systems and standards, along with their effect on aircraft performance, are unique to the Gulfstream G500/G550 .

This chapter describes the design characteristics of the Gulfstream G500/G550 aircraft and identifies the major aircraft systems, aircraft placards, interior furnishings and equipment, and outfitting options. It covers material from the following ATA chapters: • 6—Dimensions and Areas • 11—Placards and Markings • 25—Equipment/Furnishings


5-1



DESIGN CHARACTERISTICS

The basic operating weight of the Gulfstream G550 includes the manufacturer’s bare empty weight and the typical operating items listed below:

The mission for which the Gulfstream G550 is designed is to provide executive travelers with the longest-range non-stop capability available.

Flight crew—three (3) at 170 lbs each = 510 Flight attendant—one (1) at 170 lbs = 170 Crew baggage—four (4) at 30 lbs = 120

AIRCRAFT PERFORMANCE

Engine oil = ........................................123

The Gulfstream G550 has been designed to a range specification of 6,750 nm. This criterion dictated fuselage size, wing, engines, fuel capacity, and weight.The normal cruising speed of the Gulfstream G550 is 459 knots true airspeed (KTAS)/0.80 Mach, which is also the long-range cruise airspeed. Initial cruising altitude is 41,000 feet and will be achieved in 21 minutes. Maximum cruising speed is 499 KTAS (maximum operating Mach 0.885).

Unusable/Undrainable fuel ................189 Supplies ................................................688 Maximum payload for the Gulfstream G550 with maximum fuel on board is 1,600 lbs (727 kg), including passengers (8) and baggage. The maximum useable fuel weight for the Gulfstream G550 is 41,300 lbs (18,773 kg).

Maximum cruising altitude is 51,000 feet with a climb rate of 4,188 fpm, and total flight time is approximately 14 hours. This is based on maintaining 0.80 Mach, 99% Boeing standardized winds, and ATC airway routing. With the aircraft at 51,000 feet, the cabin altitude would be equivalent to 6,000 feet (1,829 meters) with 10.17 psid.

G550 AIRCRAFT WEIGHTS • Maximum ramp—91,400 lbs (41,458 kg) (Allows 400 lbs of fuel for engine run-up and taxi) • Maximum (41,277 kg)

takeoff—91,000

lbs

• Maximum landing—75,300 (34,227 kg)

lbs

• Maximum zero fuel—54,500 lbs (24,721 kg) (Maximum allowable weight of a loaded aircraft without fuel).

5-2




CABIN STANDARDS

NOTES

Seats All passenger compartment seats must meet regulatory requirements. The maximum number of passengers the Gulfstream G500/G550 can carry is 19 (13 is typical). All seats are secured to the aircraft structure, have three point restraints, and may be configured as follows: • Single forward facing seats are designed to meet a 16-g requirement. Seats have various levels of foam options (soft to firm) and various seat coverings from which to choose. Side facing seats must also meet a 16-g requirement. Seat belts are lever or push-button release type. • Double seats are available with the same standards as the single seats. • Triple (divan) style seats 37 to 39 inches wide are available and include stowage underneath.


5-3



93.45 ft

35.17 ft

25.86 ft

96.40 ft

Figure 5-1. Aircraft Dimensions

5-4




Interior Dimensions and Capacities

NOTES

Cabin length .................. 43 ft 11 in. (13.4 m) Cabin height ...................... 6 ft 2 in. (1.9 m) Cabin width ........................ 7 ft 4 in. (2.2 m) Cabin volume .......... 1,669 cu ft (47.3 cu m) Baggage compartment length .................................... 6 ft 2 in. (1.9 m) Baggage compartment volume ............................ 226 cu ft (6.4 cu m) Baggage compartment capacity .......................... 2,500 lb (1,134 kg)

Exterior Dimensions Wing span ........................ 93 ft 6 in. (28.5 m) Wing area (each wing) .............. 1,136.5 sq ft Height (top of vertical stabilizer to ground) ...... 25 ft 10 in. (7.9 m) Horizontal stabilizer span .............. 35 ft 2 in. (10.72 m) Horizontal stabilizer area ...... 260.85 sq ft (24.23 sq m) Total length (tip of nose to end of horizontal stabilizer) .... 96 ft 5 in. (29.4 m) Fuselage stations (FS) represent edges of vertical planes perpendicular to the horizontal reference plane and show measurement of length along the longitudinal (X) axis. The stations locate points along the fuselage from FS 0.00 located 4 inches aft of the forward tip of the nose radome (Figure 5-1).


5-5



PRIMARY ESCAPE ROUTES SECONDARY ESCAPE ROUTES TO BE USED ONLY ON INSTRUCTIONS FROM FLIGHT CREW

ENTRANCE DOOR

EXTERNAL BAGGAGE COMPARTMENT DOOR EMERGENCY ESCAPE WINDOWS (2 PER SIDE)

Figure 5-2. Emergency Exits

5-6




DIFFERENCES FROM PREVIOUS GULFSTREAM AIRCRAFT Datum Line The datum line is an imaginary vertical plane from which all horizontal measurements are taken. On the Gulfstream G500/G550 aircraft the datum line is located at FS –4.00. This means that the datum line is located 4 inches in front of the tip of the nose radome, which is FS 0.00.

Water Line (WL) The water lines represent edges of planes parallel to the horizontal reference plane. This plane is parallel to the fuselage centerline and locates points, components, and distances above a theoretical datum line (WL 0.00). For the Gulfstream V, WL 0.00 is 100 inches below the centerline of the fuselage datum.

Windows There are fourteen windows in the passenger cabin (seven per side) (Figure 5-1). They are located 49 inches apart on center and are elliptical in shape (19 x 26 inches), water tight, and electrically heated.

Emergency Exits Of the fourteen windows in the passenger compartment, four are removable—two on the left and two on the right side located over the wing. The removable windows are capable of being opened from either inside or outside for emergency egress. The main entrance door is certified as an emergency exit, and the baggage compartment door may also be used as an auxiliary emergency exit (Figure 5-2).

NOTES

Buttock (Butt) Line (BL) The butt line shows measurement of width to the left and right of the aircraft centerline. Measurements, in inches, to the left of BL 0.00 are designated left buttock line (LBL), and measurements to the right of BL 0.00 are designated right buttock line (RBL). Wing stations represent planes perpendicular to the wing reference plane and parallel to the fuselage centerline and are measured from WS 50.00 (FS 397.513 BL 0.00 WL 39.982) to WS 531.000 (BL 531.000).

Doors The baggage compartment door is a 40 x 36inch plug-type door on the left side of the airplane. The main entrance door is a 36 x 60-inch passenger entrance door. An unpressurized tail compartment houses the auxiliary power unit (APU), the air-conditioning units, and other equipment. The entrance to this compartment is from the ground, through an access (ventral) door with a self-contained folding ladder. There are also several service doors located all over the aircraft; these are discussed in Chapter 51–57, “Structures.”


5-7

5-8

international

FlightSafety GULFSTREAM G500/G550 MAINTENANCE TRAINING MANUAL


Figure 5-3. Rolls-Royce Deutschland BR710 Engine



AIRCRAFT SYSTEMS

NOTES

AIRFRAME INNOVATIONS Innovations in the airframe include an enlarged cockpit, relocated airstair, new trailing-edge components, new wing technology, nacelle and thrust reversers, and newly designed winglets.

ENGINES The Gulfstream G500/G550 has two aft fusel a g e - m o u n t e d R o l l s - R oy c e D e u t s c h l a n d BR700-710C4-11 engines (Figure 5-3) with a static thrust of 15,385 pounds each at ISA +15°C (86°F). Each engine is a high bypass turbofan with a bypass ratio of 4.0:1 and is controlled by a full authority digital engine control (FADEC). Control of the BR710 engine is electronic, via dual-channel engine electronic controllers (EEC’s). Nothing mechanical connects the pilot’s power levers and the engines. The power levers send electronic signals to the EEC’s microprocessors, which command engine power. This system protects against engine overspeed and overtemperature.


5-9



PRIMARY ENGINE

ALTERNATE ENGINE

SECONDARY ENGINE

COMPACTED ENGINE

Figure 5-4. Engine Instrument Displays

5-10




Primary Engine Indications

NOTES

The normal configuration format at initial application of full electrical power places the primary engine display 1/6 window in the top right section of DU No. 2 with the secondary engine display 1/6 window directly below in the bottom right of DU No. 2. In this configuration, the primary engine contains analog dial representations for EPR, TGT and LP rpm and digital indications for HP rpm and fuel flow (FF) with a split arrow icon showing differences in engine FF. The secondary engine display 1/6 window will display only digital indications for the following engine parameters: oil pressure, oil temperature, LP and HP EVM, hydraulic pressures, fuel tank temperature and fuel quantity (a split arrow icon representing differences in tank quantities). If for any reason the flight crew chooses to rearrange the default display format, the following options are available: 1. If the secondary engine 1/6 window is eliminated from display, the primary engine 1/6 window format will change to an alternate engine 1/6 window display where the digital indications of HP rpm and fuel flow (and split arrow difference icon) are replaced with digital readings of fuel quantity including a total fuel indication and a split arrow tank difference icon. 2. A compacted engine 1/6 window may be elected for display during normal engine operations. If only battery power is available for the aircraft electrical system, and engines are operating, the 1/6 compacted engine display is the default presentation. The compacted window contains digital indications for EPR, TGT, LP and HP rpm, FF, oil pressure, oil temperature and fuel quantity.


5-11



Figure 5-5. Auxiliary Power Unit

5-12




AUXILIARY POWER UNIT The auxiliary power unit (APU) is the RE220, developed by Honeywell. The APU is a gas turbine engine that starts with aircraft or ground DC power and then operates on aircraft-supplied fuel. The APU is to provide an alternate source of pneumatic power for the main engine start system and the environmental control system, and provide shaft power to drive the auxiliary AC generator. Engine operation is controlled by four systems: fuel, lubrication, electrical and pneumatic. The control system consists of an electronic control unit (ECU) and sensors that measure APU operating parameters, which the ECU uses to control the engine. The ECU ensures the APU and all of its subsystems operate correctly in response to all environmental and load conditions.

The APU provides pneumatic power in the form of compressed air for operation of aircraft main engine starters and environmental control. The engine has five basic operating modes: ready to load (full rpm with no shaft or bleed load), main engine starting (bleed load), environmental control (bleed load), electrical power generation (shaft load) and combination operation (simultaneous shaft and bleed loads). The APU can be started up to 43,000 feet, although 39,000 feet is guaranteed. The auxiliary AC generator is rated at 40 kVa electrical load up to 45,000 feet. APU operations can be monitored from the APU control panel located on the cockpit overhead panel or using the 1/6 synoptic page APU BLEED (Figure 5-6).

APU

EGT

RPM

495 °C

101.1 %

Open

R

L

45 Psi

45 Psi

Bleed Air Pressure Figure 5-6. APU Synoptic Page


5-13



HYDRAULICS

Left 12 ∞C

Right 24 ∞C

Full

Full

Low

Low 1.5g

4.7g

Aux 3000 psi

Left 3000 psi L T/R

Right psi R T/R

Aileron Elev Flt Spl Gnd Spl Stl Bar YD1

Rudder

YD2

NWS PTU 0 psi

Main Door Flaps Gnd Spl Ctrl

HMG

Brakes Ldg Gear

ACCUM 3000 psi

BOTTLE

3100 psi

Figure 5-7. 2/3 Hydraulic Synoptic Page

ELEVATOR TRIM TAB AILERON

LATERAL AXIS

ELEVATOR ELEVATOR TRIM TAB

FLAP

RUDDER LONGITUDINAL AXIS

SPOILERS AILERON TRIM TAB AILERON

VERTICAL AXIS

Figure 5-8. Flight Control Locations

5-14




HYDRAULICS Hydraulic power is supplied to the aircraft by three independent systems designated as left system, right system, and auxiliary system. All systems operate at 3,000 psi nominal pressure. The left hydraulic system is powered by one variable displacement pump mounted on the left engine. This system powers the flight controls, ground spoiler control, wing flaps, brakes, nose wheel steering, landing gear, and hydraulic motor generator. The right hydraulic system is powered by one variable displacement pump mounted on the right engine. This system powers the flight controls and through the power transfer unit; ground spoiler control, wing flaps, brakes, nose wheel steering, landing gear, and hydraulic motor generator. A DC electrically-driven pump is provided for operation of the main entrance door, alternate power for flaps, brakes, steering, rudder, yaw

damper, and for ground operation of the landing gear. The 2/3 HYDRAULICS synoptic page displays system operation and indicates any abnormal condition (Figure 5-7).

FLIGHT CONTROLS Gulfstream G500/G550 flight controls feature conventional mechanical linkages with dual servo hydraulic boost for all axes. Roll authority is augmented by combined aileron and spoiler action (Figures 5-8 and 5-9). An aileron and elevator disconnect is provided in case of jammed controls and both electric and manual pitch trim is available. Roll and yaw trim are manually operated, and dualchannel yaw dampers are provided. The 2/3 or a 1/6 FLIGHT CONTROL synoptic page depicts system operation and indicates any abnormal condition in the system (Figure 5-9)

1/6 PAGE

1

1

2/3 PAGE

Figure 5-9. Flight Control Synoptic Page


5-15



1/6 PAGE

2/3 PAGE Figure 5-10. ECS/PRESS Synoptic Page

12,400

6200

43

21

6200

38

14

Figure 5-11. Fuel System Synoptic Page

5-16




LANDING GEAR The Gulfstream G500/G550 landing gears are fully retractable, tricycle landing gear with standard dual wheels. Oil-pneumatic shock struts provide support of the aircraft during landing and while on the ground. The gear is electrically controlled and hydraulically actuated and is normally pressurized from the left hydraulic system. The landing gear incorporates an electronically controlled antiskid braking system and a nose wheel steering system. The FLIGHT CONTROL synoptic page can be displayed as 1/6 or 2/3, and depicts landing gear position with weight on wheels information (Figure 5-9).

A fuel return to tank system is also provided to warm the fuel in the tanks for high altitude operation or any time the fuel tank temperatures are very low. The FUEL synoptic page depicts operation of the fuel system and indicates any abnormal conditions within the system (Figure 5-11).

NOTES

ENVIRONMENTAL CONTROL SYSTEM The environmental control system (ECS) can sustain a maximum pressure differential of 10.17 psi. This allows a 6,000-foot cabin altitude at FL 510. The ECS consists of dual air cycle machines, a three-zone temperature control, water separators, and ozone filters. Dual digital controls that are tied into the flight management system normally govern cabin pressurization. The 2/3 or a 1/6 ECS/PRESS synoptic page depicts operation of the pneumatic system and indicates any abnormal conditions within the system (Figure 5-10).

FUEL SYSTEM The fuel system consists of two integral wing tanks with a total usable capacity of 41,300 pounds (Figure 5-11). Cross-flow capability by way of a cross-feed valve allows fuel from either tank manifold to feed the opposite tank manifold. Fuel transfer between tanks is provided through the intertank valve. This provides for fuel balancing. Refueling is accomplished by gravity or pressure fueling. Automatic or manual refueling capability is available in the pressure fueling mode.


5-17

5-18 LAC

AAC

L GCU

APU GCU

CTA 1

R GCU

EAC

CTA 2

CTA 4

L BPCU

LAXC

RIGHT MAIN AC BUS

RAXC


REAC LMTAC

ATAC 1

RMTAC

R BPCU

LEAC

HMG

R MAIN TRU Sw to LAC

ATAC 2 LEFT MAIN TRU

LEFT ESS TRU

AC Aÿ

HE 1 HE 3

LSAC EDC/ ADC

HE 8

E-INV

LMDC

LEDC

AUX TRU

HE 5

APC R BPCU

L BPCU

L STBY AC BUS

RSAC

HE 2

HE 4

RMDC

REDC

R STBY AC BUS

RMDXC

LMDXC L MAIN DC BUS

RIGHT ESS TRU

RIGHT MAIN TRU

ESS AC BUS

L BPCU

R BPCU

R BPCU w/ SEP R ESS Sw OFF

R MAIN DC BUS

R BPCU

LEDXC

REDXC

REDBC

LEDBC LEIDC

GSBC 1

REIDC

L ESS DC BUS

BC 1

R ESS DC BUS

BC 2

HE 7

AUX HYD MOT CONT

APU START CONT

R Batt Sw ON

R BATT CHGR

Figure 5-12. Gulfstream G500/G550 Electrical Schematic

international

L Batt Sw ON

GND SRV BUS

RBCC

LBCC L BATT CHGR

GSBC 2

HE 6

FlightSafety

BATT TIE BUS


LEFT MAIN AC BUS

RAC

034T1



ELECTRICAL POWER The Gulfstream G500/G550 electrical power system (EPS) provides AC and DC power and a means of control, protection, and distribution of electrical power required for ground and inflight operations of the aircraft (Figures 5-12 and 5-13). Primary electrical power is provided by the AC power system, comprised of two engine-driven integrated drive generators (IDGs) and an auxiliary-power-unit-driven generator. DC power is provided by five transformer rectifier units (TRUs) and supplemented by two nickel-cadmium batteries. A DC-power ground service bus (GSB) is provided to allow routine aircraft servicing without powering other aircraft systems. The aircraft is equipped with a standby electrical power system that is powered by a hydraulic motor generator (HMG).

The HMG is able to provide power to the standby AC busses and to the essential DC busses via the AUX TRU during emergency conditions. Emergency power is provided by four 9 amp-hour battery packs. Two battery packs provide power for the emergency lighting system and two battery packs provide power for the avionics backup battery system. The emergency lighting system provides cabin and over-the-wing emergency exit lighting, and the avionics backup battery system provides power for standby flight instruments and inertia reference units (IRU) during emergency conditions.

Figure 5-13. AC and DC Synoptic Pages


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Figure 5-14. PlaneView System Overview



G500/G550 AVIONICS The purpose of the Planeview system is to provide pilots with real-time information about airports, weather, air traffic and terrain, all displayed on four large landscape, flat-screen monitors. These advancements greatly improve situational awareness for pilots and provide instant access to information necessary to make safe flying decisions in adverse or low-visibility conditions.The most significant advantage is improved safety.

The Primus Epic system takes advantage of advancements in flat panel display technology and cursor control devices and couples these with the modular integration of many of the stand-alone utilities functions into the avionics suite. Many control functions that were previously individual line replacement units (LRUs) in older systems are functionally integrated into the modular avionics unit (MAU) and the modular radio cabinets (MRC) of the Primus Epic system.

The Planeview system is based on the Honeywell EPIC architecture and consists of three 16 user slot dual channel MAUs, four large format flat panel displays, dual EGPWS, dual EPIC radios, TCAS, three AV-900 (ACPs) audio panels, various controllers, sensors and servos.

PRIMARY FLIGHT DISPLAY (PFD)

NAVIGATION DISPLAY (INAV)

Figure 5-15. Flight Displays


5-21



OPEN DOOR PULL HANDLE UP PUSH WINDOW IN

ESCAPE WINDOW RELEASE

PUSH

LOCATOR PLACARDS

Figure 5-16. Locator Placards—Emergency Window Exits

5-22




AIRCRAFT PLACARDS

NOTES

GENERAL Aircraft placards are classified as either exterior or interior. In either case, Gulfstream G500/G550 placarding is consistent with FAA requirements for safety and emergency rescue. Both exterior and interior placards fall into one or more of the following categories: • Caution and warning • Instruction and information • Locator

CAUTION AND WARNING PLACARDS Caution and warning placards provide for the prevention of injury to personnel and damage to equipment.

INSTRUCTION AND INFORMATION PLACARDS Instruction and information placards provide information on the operation of controls and equipment.

LOCATOR PLACARDS Locator placards point out locations such as the emergency exits and identify components such as vents, electrical connectors, and access panels (Figure 5-16).

NOTE If placards need replacing, refer to the Illustrated Parts Catalog for ordering information.


5-23



CHAPTER 12 SERVICING

INTRODUCTION This chapter presents an overview of the Gulfstream G500/G550 aircraft ground handling and servicing procedures. All values such as pressures, temperatures, rpm, and power requirements are used for their illustrative meaning. The current manufacturer’s Maintenance Manual must be consulted for all maintenance specifications, tolerances, and the actual values. This data must be determined from approved Gulfstream reference material.

GENERAL Aircraft ground handling and servicing involves the equipment, methods, and procedures used on the ground in the following situations: • To move an aircraft when it is impossible or impractical to move it under its own power

• To recover a disabled aircraft • To determine the exact weight of an aircraft • To properly store an aircraft for an extended period of time • To replenish expended consumables


12-1

12-2

international



Figure 12-1. Aircraft Jacking



This chapter covers material from the following ATA chapters: • 7—Lifting and Shoring • 8—Leveling and Weighing

CAUTION If stairway is to remain extended while aircraft is on jacks, ensure lower step of air stair door is supported.

• 9—Towing and Taxiing

CAUTION

• 10—Parking and Mooring • 12—Servicing

JACKING, LIFTING, AND SHORING Jacking is performed when it is necessary to remove, repair, replace, and functionally check the landing gear and its components. Jacking is also a method used to weigh the aircraft. Lifting and shoring are performed when it is necessary to recover an aircraft that has landed with one or more collapsed landing gear (Figure 12-1).

OPERATIONAL REQUIREMENTS

• Ensure there is sufficient tail clearance when jacking aircraft while inside hangar. • To prevent aircraft from falling on its tail section while being jacked, counterweights can be attached to nose gear of aircraft.

• To prevent damage to nose gear shock strut, ensure that nose wheel steering torque links are connected.

CAUTION

Jacking the aircraft is performed in accordance with the GAC Maintenance Manual (07-10-00) by using one of two methods: individual nose and main gear jacking or threepoint fuselage jacking.

CAUTION • If aircraft is to be jacked out of doors, head aircraft into the wind.

To prevent the possibility of aircraft slipping off jacks while being raised, ensure that parking brake has been released.

CAUTION Prior to applying electrical power to a jacked aircraft, ensure that probe heat and cabin window circuit breakers are pulled and cabin window heat switch is selected to off.

• Do not jack if maximum gusts exceed 20 mph. • Do not jack if wind velocity exceeds 10 mph, unless the aircraft is snubbed at all mooring points.


12-3



11.16"

JACK ADAPTER TO GROUND

NOSE LANDING GEAR RH WHEEL OMITTED FOR CLARITY

FLAT TIRE GROUND LEVEL

16.45" MAIN LANDING GEAR LH WHEEL OMITTED FOR CLARITY

FLAT TIRE GROUND LEVEL

GEAR JACK SPECIFICATIONS JACK POINT NLG MLG

ROLL UNDER FLAT TIRE INCHES 11.61 16.45

EXTENDED INCHES 14.04 20.70

GEAR DIMENSIONS ITEM TIRE DIA NORMAL ROLLING STATUS FLAT TIRES RADIUS DISTANCE BETWEEN NORMAL TIRES DISTANCE BETWEEN FLAT TIRES

MAIN LANDING GEAR INCHES 35.00

NOSE LANDING GEAR INCHES 21.25

15.22

9.15

11.90

7.30

7.06

5.47

5.50

3.62

Figure 12-2. Axle Jacking Provisions

12-4




INDIVIDUAL NOSE AND MAIN GEAR JACKING A fold-down jack pad is provided at the bottom of the nose gear shock strut. The jack pad is spring-loaded to the down or folded position and must be held in place while the jack is placed in position, which requires a 25-ton low-profile axle jack (Figure 12-2).

essary to use the fuselage jack point. If both tires are flat on the main gear, it will be necessary to use the wing jack point. The maximum aircraft weight for single wheel jacking is 90,900 pounds, which is also the maximum gross weight of the aircraft.

WARNING

For individual main gear jacking, a jack pad adapter is inserted into the aft portion of the main landing gear trailing arm between the two wheels. A 25-ton low-profile jack is required with the use of a wheel jack pad for main gear jacking. Because of the close proximity of the brake lines to the jack pad, caution must be used when positioning the jack on the pad.

Do not exceed the jack extension screw beyond the specific limits stenciled on the jack. Serious injury to personnel or damage to the aircraft can result.

A jack with a minimum capacity of 12 tons and with an extension range of 11.61–14.04 inches is required to raise the aircraft at the nose jack point. For jacking individual main landing gear struts, an axle jack with a minimum of 20 tons and with an extension of 16.45–20.70 inches is required with the use of a wheel jack pad. If both nose tires are flat, it will be nec-

Walking on the empennage or outer wing of the aircraft, while on jacks, may cause movement of the aircraft. Caution must be observed at all times and personnel must be limited to those absolutely necessary for proper performance of this operation.

CAUTION

MAIN LANDING GEAR TRAILING ARM

BUSHING JACK ADAPTOR ENVIRONMENTAL CAP

Figure 12-3. Main Landing Gear With Jack Adapter FOR TRAINING PURPOSES ONLY

12-5



WING TRIPOD JACK FS 635.314 RBL 201.605 WL 46.922

NOSE LANDING GEAR TRIPOD JACK FS 120.000 RBL 0.000 WL 53.000

WING TRIPOD JACK FS 635.314 RBL 201.605 WL 46.922

Figure 12-4. Fuselage Jack Point Locations

12-6




FUSELAGE JACKING

CAUTION

There are three fuselage jacking points: one at the nose, located at fuselage station (FS) 120.00°, and one at each wing, located at FS 635.31 (Figure 12-4). A 12-ton jack with a 46- to 75-inch extension range is required for the nose fuselage jacking point. The wing fuselage jacking points require 25- or 30-ton jacks with a 54- to 81-inch extension range. Aircraft gross weight restrictions apply if 25ton tripod jacks are employed. The maximum aircraft weight for three-point jacking is 77,260 pounds using 25-ton jacks and 90,900 pounds using 30-ton jacks. It is imperative that personnel closely follow notes, cautions, and warnings associated with aircraft jacking published in Chapter 7 of the manufacturer’s Maintenance Manual.

Bolts of correct length must be used to install wing jacking pads. If improper bolts are used, damage to dome nuts in wing will result, causing a fuel leak which will require extensive maintenance and repair.

CAUTION To prevent damage to the nosewheel steering unit, do not operate the nosewheel steering while aircraft is on jacks with torque links connected.

NOTES

CAUTION If the nose gear is to be retracted, the fuselage jack must be installed with one leg facing aft to prevent damage to the nose gear fairing door. Ensure that the special nose gear pad has been installed.


12-7



1

REAR VIEW

SEE NOTE 1 3

4

5

6

LIFTING BAGS (REF) JACK POINTS STABILIZING LINE ATTACHMENT POINTS

SEE NOTE 2

3 FT MINIMUM (2 PLACES

2 6

TOP VIEW

3 5

4

STABILIZING LINE (REF) (TYPICAL 6 PLACES)

LIFTING BAGS (REF) 1

PROVIDE CLEARANCE FOR JACK (2 PLACES)

1

FS 133 (REF)

8

SIDE VIEW

SEE NOTE 1 1

2

NOTES: 1. TO AVOID POSSIBLE STRUCTURAL DAMAGE, USE RATIO OF 8 TO 1 ON ALL STABILIZING LINES. NEVER ALLOW RATIO TO EXCEED 4 TO 1 DURING RAISING PROCESS. 2. STAKES ARE SHOWN FOR CLARITY ONLY.

Figure 12-5. Locations of Pneumatic Bags and Stabilizing Line Attachments

12-8




LIFTING AND SHORING

NOTES

Before performing any lifting and shoring of the aircraft, personnel must be familiar with the aircraft structure and its limitations and be proficient in the use of lifting equipment and its limitations (Figure 12-5). Compliance with all notes, cautions, and warnings in Chapter 7 of the manufacturer’s Maintenance Manual is necessary to properly lift or shore the aircraft and prevent injury to personnel or damage to the aircraft. Lifting of the aircraft is accomplished using either pneumatic lifting bags or a crane and sling. Close attention must be given to aircraft attitude during air bag lifting. The aircraft should be kept in equilibrium by controlling the air bag inflation and relying on stabilizing lines to further preclude structural damage. When using the crane and sling method, it should be noted that there are no provisions for hoisting the entire weight of the aircraft. When using mechanical lifting equipment for hoisting the nose of the aircraft, slings should be placed at FS 133.

NOTE Tail support or an air bag should be placed under the aft fuselage to prevent the aircraft from tipping on its tail.

Pneumatic lifting bags are used when there is not sufficient ground clearance available to use standard jacks. Shoring is generally used to support a portion or the entire aircraft during periods of extended, heavy maintenance. Contact Gulfstream Technical Operations for shoring procedures.


12-9



DOWNLOCK PIN

NOSE LANDING GEAR

DOWNLOCK PIN

MAIN LANDING GEAR

Figure 12-6. Landing Gear Pins

12-10




PARKING, STORAGE, AND MOORING

NOTES

The purpose of parking and mooring is to safely secure, protect, and prevent inadvertent movement of the aircraft in all weather conditions. There are two types of parking methods: parking and storage. Outside storage procedures are based on a humid, tropical atmosphere. Based on this somewhat severe climate, an aircraft remaining inoperative for 15 days or more is considered as having aircraft in storage. Requirements for inside storage are the same as outside storage, as specified i n C h a p t e r 1 0 o f t h e m a n u f a c t u r e r ’s Maintenance Manual.

PARKING PROCEDURES Parking procedures are required any time the aircraft remains inoperative for 14 days or less. When the aircraft is parked for a short period of time, the downlock pins must be installed on each gear (Figure 12-6), and the parking brake, located on the center console, must be set until the chocks are in place. Once the chocks are in place, the parking brake should be released in order to prevent undue wear and inadvertent locking of the brakes. A gust lock is provided to lock all primary flight controls without the use of external locking devices. The gust lock lever is located in the cockpit on the center pedestal to the right of the throttle quadrant. With the gust lock on, each flight control surface is protected in winds up to 60 mph.


12-11



(LH AND RH) COWLING AND DOOR VENT COVER

(LH AND RH) ACOC OUTLET COVER (LH AND RH) ENGINE EXHAUST COVER (LH AND RH) AIR CYCLE MACHINE EXHAUST COVER

(LH AND RH) ANGLE OF ATTACK PROBE COVER

(RH ONLY) CABIN OUTFLOW VALVE COVER

(RH ONLY) STATIC PORT/SAFETY VALVE COVER

FUEL VENT SCREEN (WING UNDERSIDE)

(LH AND RH) (RH ONLY) AIR CYCLE MACHINE APU EXHAUST EXHAUST COVER COVER

COVER INLET VENT

PRECOOLER EXHAUST COVER (PYON UNDERSIDE) PYLON EXHAUST COVER (NEW LOCATION)

PYLON INLET COVER ICE DETECTOR PROBE COVER PYLON INLET COVER

COVER INLET VENT

PYLON EXHAUST COVER (NEW LOCATION) PRECOOLER EXHAUST COVER (PYON UNDERSIDE)

FUEL VENT SCREEN (WING UNDERSIDE) (RH ONLY) RAM AIR SCOOP COVER ENGINE INLET COVER TAI OUTLET COVER

PITOT STATIC PROBE COVER

(RH ONLY) RAM AIR SCOOP COVER ENGINE INLET COVER TAI OUTLET COVER

PITOT STATIC PROBE COVER

Figure 12-7. Protective Covers

12-12




If the aircraft is to be left unattended for a period longer than one overnight stay, or if weather conditions make it advisable, install all protective covers (Figure 12-7).

Extended Parking If the aircraft is scheduled to be parked longer than one night, or weather conditions make it advisable, all protective covers should be installed. The gust lock control, located on the center console, should be set to prevent unwanted movement of the control surfaces. Tire covers should be installed to protect the tires from undue weathering and spillage from fuels, solvents, and other fluids (Figure 12-8). The main entrance door, baggage compartment door, and tail compartment door should all be closed. Finally, if the aircraft will be parked for a period of more than three days, disconnect the main batteries located in the tail compartment, via the battery connectors, to prevent depletion of the aircraft battery charge (Figure 12-9).

STORAGE When the aircraft is parked for a period of time greater than 15 days, it will be considered storage. Storage inspection intervals are the same for aircraft parked inside or outside. These inspection intervals include parking and storage checks. For specific storage requirements, consult the GAC Maintenance Manual or the appropriate vendor manual. Unless the engines have been preserved, it is recommended to operate the engines at least once every 7 days to maintain optimum performance. Consult Chapter 10 of the GAC Maintenance Manual and the BR710 Maintenance Manual for specific requirements for flight-ready, short- and long-term storage. The tires must be rotated and checked for adequate pressure. Rotation of the tires can be accomplished by towing or taxiing when combined with running engines. This should be accomplished as a function of the seven-day storage inspection requirement. The electronic equipment must be energized weekly for one hour to ensure proper operation, which can be accomplished using ground power, the APU, or combined with the engine run.

NOTE When the aircraft is parked for extended periods (considered storage), parking requirements must also be satisfied.

Figure 12-8. Tire Covers

Figure 12-9. Battery Disconnect


12-13



45° 45

NOSE GEAR MOORING

45° 45

MAIN GEAR MOORING

Figure 12-10. Mooring Rings

12-14




MOORING

NOTES

The aircraft must be moored if it is unprotected and the wind speed is predicted to be above 30 knots. If the wind is expected to exceed 30 knots due to a severe storm or wind condition, the aircraft should be hangared. If hangaring or flying the aircraft to a safe location is not possible, the aircraft must be moored. Refer to the manufacturer’s Maintenance Manual for the appropriate procedures. The aircraft must be positioned with its nose facing into the prevailing wind and have a total distance equaling the length of the aircraft, plus 15 feet separating it from other aircraft. For mooring purposes, there are two mooring rings located on each nose gear strut and one mooring ring located on each main gear strut. Once the aircraft has been moored, ensure that all the requirements for short-term parking have been completed. Tiedown lines should be attached to the mooring rings on both the main and nose gear (Figure 12-10).

CAUTION After mooring and before flight, ensure that the mooring rings are returned to their stowed position by the torsion springs.


12-15



NOSE WHEEL WELL

SPIRIT LEVEL

STA 44.5 BULKHEAD LEVEL LUG

STA 61.500

JIG POINT WL 70.00

LEVEL LUGS

DETAIL OF JIG POINT

FW

D

STA 71.500

LONGITUDINAL LEVELING POINTS (RIGHT-HAND SIDE)

DIGITAL LEVEL

Figure 12-11. Longitudinal Leveling Brackets

12-16




LEVELING AND WEIGHING LEVELING Leveling the aircraft is required to prevent tipping when jacking the aircraft. It is also required prior to weighing the aircraft in order to obtain an accurate weight.

A spirit level is positioned on these brackets, and the aircraft is level when the bubble is centered. The lateral leveling brackets are located on the face of the nose wheel well bulkhead at FS 44.50 (Figure 12-12).

Leveling the aircraft is accomplished by raising or lowering the individual struts or the individual fuselage jacks. The longitudinal leveling brackets are located on the right side of the nose wheel well at FS 61.5 and FS 71.5 (Figure 12-11).

Before working in any wheel well, ensure that the landing gear and landing gear door ground safety devices are installed.

STA 44.50

SPIRIT LEVEL

FW D

LEVEL LUGS

JIG POINT WL 70.00

LATERAL LEVELING POINTS (FORWARD BULKHEAD) NOSE WHEEL WELL

DIGITAL LEVEL

Figure 12-12. Lateral Leveling Brackets


12-17



REMOVE PIP PIN BEFORE TOWING

NOSE LANDING GEAR

Figure 12-13. Torque Link Disconnect

12-18




WEIGHING

TOWING

Weighing the aircraft is done in order to verify the empty weight of the aircraft and to determine a new empty weight for center of gravity (CG) and takeoff and landing calculations.

Towing is performed when the aircraft must be moved and it is impossible or impractical to move it under its own power.

The aircraft can be weighed using one of two methods: portable platform scales or an electronic weighing kit (PN JP-50K-3), which uses load cells attached to the fuselage and wing jacks. In order to obtain an accurate empty weight, an inventory of equipment for the current aircraft configuration must be obtained, regardless of the method used. Before weighing, the landing gear must be extended, and the flaps and thrust reversers must be retracted. All fluids, such as hydraulic fluid, APU oil, and engine oil, must be serviced to operational levels. The toilet, as well as wash and galley water, should be drained in a c c o r d a n c e w i t h t h e m a n u f a c t u r e r ’s Maintenance Manual. The aircraft should be defueled, and all equipment and protective material that are not part of the basic aircraft inventory should be removed from the interior and exterior of the aircraft. The aircraft must be clean and dry with all protective covers removed. The entry door must be closed. The aircraft must be weighed in a closed hangar with no blowers or ventilating systems impinging on the aircraft.

CAUTION The nosewheel steering unit torque links must be disconnected prior to towing the aircraft. Rotation of the nosewheel beyond its normal limit of 82° can cause serious damage to the nosewheel steering unit. With the steering unit torque links disconnected, the nosewheel can rotate 360° (Figure 12-13 and Figure 12-15).

CAUTION Do not forcibly remove the nose landing gear torque link safety pin.

Refer to Chapter 8 of the manufacturer’s Maintenance Manual and the Weight and Balance Manual for specific details concerning aircraft weighing procedures.


12-19



A

X

B

Figure 12-14. Nose Strut Extension

12-20




A minimum of six personnel is recommended for aircraft towing: two wing walkers, one tail walker, one brake rider, one supervisor, and one tow vehicle operator. Consult the local operational requirements and the manufact u r e r ’s M a i n t e n a n c e M a n u a l f o r t ow i n g specifics.

CAUTION To prevent possible damage to the nosewheel self-centering cams, do not tow aircraft if dimension X, as shown on nose landing gear strut inflation instruction plate, exceeds 13.5 inches (Figure 12-14).

OVERTRAVEL INDICATOR

Figure 12-15. Nosewheel Steering Collar


12-21



PARKING/EMERGENCY BRAKE ACCUMULATOR INDICATOR VALVE

EMERGENCY LANDING GEAR BOTTLE INDICATOR/FILLER VALVE

DUMP

PSIG X 1000

WARNING

3

1 0

1

3 4

0

2

TIGH LOCKNUT PSXIG1000 BEFORE G REMOVIN FILLER TOR CONNEC

4

2

G WARNIN TEN

ACCUM PRESS

NO

ALVE FILLER V

RELEASE PAR BR PRI DUMP

R ACCUMULATO TT LG EMERG BO

LE

Y ENC ERG E M E G PSI R SI GEA VE & GDAOF 3100E TO 70 P G N I L ELOA REAS F D A V N LA ILLER TO PR R DEC W 70° O F LO LE

E GAG & E ALV ER TVOR: ERVED L L I R F ULA OBS ATOE ACCUMVE UNTIL EASE W 70°F L U O UM ARG VAL NCR EL ACC TO CHNLOADER I AT 70°FOIVE OR B E K B U S S A BRA CUM ILIZE 00 P 0°F C B BE 12 CH 1 A TE A STA TUAPRESS AD TOI FOR E ING C A 1. AGE RELO 5 PS RN N G UM P ASE 2 WA OGE C CRE C A R . 2 R DE NIT O

E E BOTT REAS OR B RGE 70°F INC ABOVE A H C 10°F E AT ING ONLY SUR EACH RN PRES FOR WA GEN

RO NIT E S U

LY ON

E US

D TB U O D FW

Figure 12-16. Nose Wheel Well Parking Brake Accumulator Gage

12-22




CAUTION If the torque links are not disconnected and the towing angle exceeds 82°, the steering unit hard stops may be sheared. A red overtravel warning indicator will pop up on the steering unit collar if 82° has been exceeded. Mandatory inspection for sheared stops shall then be performed prior to the next flight (Figure 12-15).

A parking brake accumulator pressure gage is located on the copilot’s flight panel. There is also a direct reading gage on the left side of the nose wheel well. The indicator should read 3,000 psi for a full charge (Figure 12-16). If cockpit gage indicates less than 3,000 psi minimum required pressure, select auxiliary pump to ON to replenish hydraulic brake pressure. Six full applications of parking and emergency brake system can be made with a fully charged accumulator (3,000 psi).

WARNING Do not use tow bars that are not rated for a Gulfstream G500/G550. The possibility of failure could result in serious injury to personnel.

Using an approved tow bar, the aircraft can be towed forward or pushed backward on hard surfaces. Provisions for attaching the tow bar are located on the axle of the nosewheel assembly (Figure 12-17). In order to tow the

TOW BAR PINS

NOSE STRUT

Figure 12-17. Tow Bar Attachment


12-23



FUELING ADAPTER

Figure 12-18. Single-Point Pressure Refueling

SINGLE-POINT PRESSURE REFUELING

FUELING CAP

Figure 12-19. Overwing Fueling

12-24




aircraft at a gross weight of 90,900 pounds in dry conditions, a tow vehicle capable of providing 12,000 pounds of drawbar pull is required. When towing the aircraft from a soft surface, the aircraft must first be defueled, and 3⁄4-inch cables must be secured to the main gear mooring rings. Two towing vehicles are required with a 12,000-pound maximum drawbar pull. The maximum recommended towing speed for all surface conditions is 5 mph. Consult Chapter 9 of the manufacturer’s Maintenance Manual for specific towing procedures

WARNING Before refueling, ensure aircraft is bonded to the fuel source.

WARNING Do not operate radar within 100 yards of a fueling/defueling operation.

CAUTION Fueling operations should be conducted with wings level. A nonlevel condition may result in fuel imbalance.

CAUTION Before proceeding to tow aircraft, ensure there is enough clearance at wingtips in the even a turn is required. Due to wing sweepback, when aircraft is turned, the wing will swing out as it completes an arc. On sharp turns, the tail will require more clearance than wings.

CAUTION Maximum fuel imbalance shall not exceed 2,000 pounds. Aircraft jacking is prohibited with any fuel imbalance.

SERVICING Aircraft servicing is performed to replenish consumables expended during flight operations or ground maintenance. Servicing is a means of maintaining efficiency and reducing the risk of mechanical damage to various aircraft systems. When systems requiring servicing use unusual amounts of fluids, gases, or lubricants, this may be an indicator of leakage, wear, possible component failure or system failure.

FUEL SYSTEM SERVICING Fuel system servicing includes fuel/defuel operations, single-point pressure fueling, overwing gravity fueling, auto fueling function, and ground service operation control panel.

G550


12-25



OIL QTY PINTS BELOW FULL L ENG

– 1.5

R ENG

– 2.0

APU

–1.5

FILL

ON

ON

OFF GND SVC BUS SWITCH TEST

OFF

Figure 12-20. Ground Service Control Panel

Figure 12-21. Engine Oil Tank Location

12-26




Fuel/Defuel Operations The total capacity of the Gulfstream G550 fuel system is 6,118.5 U.S. gallons. This equates to 3,059.25 gallons per wing tank. The fueling nozzle pressure optimum range is 35 to 55 psi, which is the standard for most large commercial and military aircraft. Chapter 12 of the manufacturer’s Maintenance Manual provides complete details of the most acceptable fuel grades found around the world.

Single-Point Pressure Fueling Single-point pressure fueling is accomplished through a fueling adapter located in the forward right wing-to-fuselage fillet area at access point 192CB. The optimum fueling hose pressure should be between 35 and 55 psi (Figure 12-18).

Overwing Gravity Fueling Overwing gravity fueling is accomplished through a fueling cap on the top of each wing, outboard near the winglets (Figure 12-19).

Auto Fueling Function/Control Panel Location The Gulfstream G500/G550 is capable of “auto fueling,” which means that preprogrammed fuel loads can be preselected on the ground service control panel located on the systems monitor test panel in the cockpit. When the preselected fuel load is reached, fueling ceases automatically (Figure 12-20).

NOTE The ground service bus must be energized to accomplish the auto fueling function.

OIL SYSTEM SERVICING Be aware that working around turbine engine equipment requires special equipment, protective clothing, and the use of good judgm e n t . To x i c c h e m i c a l s a n d i n h e r e n t l y hazardous conditions provide the potential for a dangerous work environment.

WARNING Engine components may be hot enough to cause injury to personnel for up to one hour after shutdown. Wear protective equipment and appropriate clothing when working in the vicinity of a hot engine. If burned, flush skin with cold water and seek medical attention.

WARNING Synthetic oils approved for use in the BR710 engine may contain an additive called Tricreasyl Phosphate, which is an asphyxiant. It is highly poisonous and can be absorbed through the skin.


12-27



Figure 12-22. Oil Tank Sight Gage

HOIST MOUNT PAD

OIL COOLER TUBE INTERFACE

PRESSURE FILL AFT OIL SUPPLY AFT OIL RETURN

HOIST MOUNT PAD

GRAVITY OIL FILL

OIL FILL REMOTE

LEVEL SENSOR

APU GROUNDING LUG

GRAV OIL FILL

HARNESS CLAMP BOSS OIL HEATER BOX

COMPRESSOR SEAL PORT DRAIN PLUG WITH MAGNETIC PICKUP COLLECTOR

Figure 12-23. APU Gearbox

12-28


APU MOUNT



CAUTION Oil systems should not be serviced cold. All lubrication systems of the Gulfstream G500/G550 should be checked after operation while still at, or at least near, operating temperatures. Overfilling of oil tank can o c c u r i f e n g i n e o i l l ev e l i s n o t checked between 5 and 30 minutes after engine shutdown. Damage to engine can occur if oil system is overfilled.

The Hispano-Suiza gearbox holds 14.4 quarts of oil. Refer to the GAC Maintenance Manual or the vendor manual for a complete and accurate list of approved oils (Figure 12-22).

NOTE For an accurate quantity reading, ensure that the engine has been shut down for at least 5 minutes, but not longer than 30 minutes, prior to determining servicing requirements.

APU Oil System Servicing Engine Oil System Servicing The oil tank is located on the left side of either engine accessory drive gearbox and is an integral part of that module. The servicing points are different on the left and right engines, due to their location in relation to the cowl access panels. Servicing the left engine requires opening the lower engine cowl door, whereas servicing the right engine is accomplished through an access panel located on the fixed cowl on the inboard side (Figure 12-21).

The auxiliary power unit (APU) is housed within the left side of the tail compartment in a fireproof, titanium, box-like enclosure, directly aft of the primary pressure bulkhead. The oil tank holds a total of 5.25 quarts of either Type I or Type II military-specification oils. The APU oil sump should be filled through the gravity fill cap until oil can be seen at the bottom of the screen in the filler port. The oil level should be checked 5 to 15 minutes after shutdown. It is recommended that the APU be disabled via circuit breakers, etc., prior to performing maintenance or service. This is done to help prevent injury to personnel. The GAC Maintenance Manual cautions against mixing different specifications or brands of oil in the APU gearbox sump (Figure 12-23).


12-29



L

ENG

R E N

SELECTOR VALVE

U

OF F

AP

VENT

OIL REPLENISHER TANK

DETAIL A

ENGINE OIL CAP

SIGHT GAGE 0–14 PINTS QUICK DISCONNECT AUXILIARY FILL ACCESS OIL QUANTITY INDICATOR

SEE DETAIL A SEE DETAIL B

OIL QTY PINTS BELOW FULL L ENG

– 1.5

R ENG

– 2.0

APU

–1.5

FILL

ON

TEST

OFF

DETAIL B

Figure 12-24. Remote Oil Replenishment System

12-30


G



NOTES

Remote Oil Replenishment System The Gulfstream G500/G550 incorporates a convenient servicing feature that allows the replenishment of the engine and APU oil systems via a remote servicing panel. This panel is located in the left side of the tail compartment at the top of the entrance ladder. Its reservoir is located just above the panel and has a capacity of 6 U.S. quarts of oil (1.5 gallons) (Figure 12-24).


12-31



V BAND COUPLING CLAMP STARTER QAD ADAPTER

FRONT STARTER AIR DUCT ACCESSORY GEARBOX

OIL FILL PLUG DRAIN PLUG/CHIP DETECTOR

Figure 12-25. Air Turbine Starter

12-32




NOTES

Starter Servicing The starter is mounted on the right front face of the engine accessory drive gearbox (Figure 12-25). Its servicing port is on the left side of the unit, closest to the engine centerline. It is accessed only by dropping the lower engine cowl door. The air turbine starter holds a maximum of 260 cc of oil in its sump. Mixing of engine oils is not recommended, but approved brands may be mixed if operationally essential. Changes from one approved oil to different approved oil must be made slowly by the usual procedure to fill the oil system during servicing. Refer to vendor manual when changing brand of oil.

NOTE Oils listed below are those oils that are approved for G500/G550 oil-serviced equipment and are commonly used on the engine and engine starter. See Rolls-Royce Deutschland Maintenance Manual for complete list of approved lubricating oils.

• Aeroshell Turbine Oil 500 (ROYCO 500) • Castrol 5000 Gas Turbine Oil • Esso/Exxon 2380 Turbo Oil • Mobil Jet Oil II • Mobil Jet Oil 254


12-33



ELEVATORS STALL BARRIER RUDDER YAW DAMP AILERONS FLT SPOILERS SPEED BRAKES RIGHT T/R

ELEVATORS STALL BARRIER AILERONS FLT SPOILERS SPEED BRAKES LEFT T/R RUDDER/ YAW DAMP 1

GROUND SPOILERS L SYS OR PTU OR AUX PRESS SIGNAL REQ'D FOR R SYS USE

GROUND SPOILERS

GND SPLR SERVO PRESS

STANDBY ELECTRICAL POWER MASTER

WING FLAPS LANDING GEAR NOSE WHEEL STEER BRAKES

ON

HMG MOTOR

PWR XFR UNIT

STBY RUD

OFF/ARM

ON

NOT ARM

ON

ON

AUX PUMP ACCUM

LEFT ENG PUMP

AUX SOV AUX BOOST PUMP

ACCUM

RIGHT ENG PUMP

DISCH

DISCH

1

2

2

L

AUX PUMP OFF/ARM

ON

NOT ARM

ON

1

R R SYS

L SYS

AUX L SYS R SYS PTU AUX

L SYS/PTU L SYS/PTU/AUX NITROGEN ELECT CONT

CHECK VALVE SHUTOFF VALVE

Figure 12-26. Hydraulic System Schematic

12-34


(

FLOW

)



Integrated Drive Generator

Type II oils listed below are those oils that are approved for G500/G550 oil-serviced equipment and are commonly used between all equipment (engine, engine starter, etc.):

T h e i n t eg r a t e d d r ive g e n e r a t o r ( I D G ) i s mounted on the right rear face of the accessory drive gearbox. An oil level sight gage is provided to show when the oil level is low or if there is an overfill condition. The oil capacity of the IDG oil system is 3.4 to 4.15 U.S. quarts. Damage to the IDG will result if it is allowed to operate with insufficient oil quantity in the sump.

• Aeroshell Turbine Oil 500 (ROYCO 500) • Castrol 5000 Gas Turbine Oil • Esso/Exxon 2380 Turbo Oil • Mobil Jet Oil II

NOTE

BLEED BUTTON

R SYS RESERVOIR

FW

D

L SYS/AUX RESERVOIR

INDICATOR

LL FU

REFILL

D

TY

EMP

FW

Figure 12-27. Hydraulic Reservoir FOR TRAINING PURPOSES ONLY

12-35



REPLENISHER TANK

1.50 1.25

SIGHT GAGE

1.00 20 25

OU

TB

D

WD

F

NS CTIO TRU INS G ING DH FAT NOIS R E 'G OP OP

IH HER NO ENIS OIHO EPL HIG H IC R NO AUL OIHHC ON R P D G G HY ON H8P IDH

IH IUH DNFIO 9H JOIO BIU F NP 1. F G FS RG BS F IUO 9P OIU GU HG 2, N IUB P98 NIN ET P IH U 8H 3. S P9H89 N 'G E TH IHF 4. IP DFGP H GIO H ]ND F GP PID C NO

OFF

T

H RIG

T

LEF

SELECTOR VALVE

RESERVOIR QUANTITY INDICATOR

Figure 12-28. Remote Hydraulic Replenishing System

12-36




• Mobil Jet Oil 254 The mixing of oil types when servicing the IDG is not permitted. Refer to the Hamilton Sundstrand Standard Practices Manual for complete list of approved oils.

HYDRAULIC SYSTEM SERVICING As in most large jet aircraft, the Gulfstream G500/G550 has two main hydraulic systems: a backup hydraulic system and an emergency system. The Gulfstream G500/G550 systems are labeled “left system,” “right system,” “power transfer unit” (PTU), and “auxiliary system” (Figure 12-26). The left and right system reservoirs require manual servicing of hydraulic fluid. The auxiliary reservoir is an integral part of the left system reservoir and is full any time the left reservoir contains more than 2 gallons of fluid. The left hydraulic system has a total capacity of 20.6 U.S. gallons. The reservoir itself holds 5.7 gallons and is serviced to 4.8 gallons. The right system has a total capacity of 7.0 U.S. gallons. The reservoir holds 1.8 U.S. gallons and is serviced to 1.5 gallons. Each reservoir has a cable-operated, directreading gage. The hydraulic fluid grade speci f i c a t i o n i s P h o s p h a t e - E s t e r Ty p e I V (SKYDROL) (Figure 12-27).

Ground Hydraulic Service Panel The ground hydraulic service panel is located on the aircraft belly, just forward of the tail compartment access door. It consists of six quickdisconnect fittings. The quick-disconnects provide for connection of a ground hydraulic test rig and external fluid servicing cart.

Remote Hydraulic Replenishing System There is a hydraulic replenishing feature incorporated in the Gulfstream G500/G550 that makes servicing both left and right hydraulic systems convenient through a single replenishing source. The hydraulic replenishing panel is positioned in the right side of the tail compartment, at the top of the ladder, opposite the engine/APU remote servicing panel. The replenishing tank is located directly above the hydraulic replenishing panel and holds 1.5 U.S. gallons. There is a sight gage for di-


12-37



MAIN GEAR STRUT SERVICE VALVE

VALVE CAP

VALVE STEM SWIVEL NUT

VALVE BODY NOSE GEAR STRUT SERVICE VALVE

PIN 0-RING BACKUP RING 0-RING

Figure 12-29. Landing Gear Strut Filler Valves

12-38




rect tank level readings (Figure 12-28).

WARNING SKYDROL Type IV phosphate ester hydraulic fluid is combustible and can be a health hazard. Inhalation of vapor and contact with skin and eyes should be avoided. The fluid should not be exposed to extreme heat or open flames. All material safety data sheet recommendations for health and safety precautions should be followed. To prevent injury to personnel and damage to equipment, protective caps should be installed on all open electrical disconnects, open hoses and ports.

CAUTION SKYDROL Type IV phosphate ester hydraulic fluid can damage paints, rubber and plastic materials. Care must be taken to prevent spillage from remaining on surfaces or damage may result.


NOTES

12-39



T OU OU

T

IN

IN

NOSE WHEEL WELL (RIGHT SIDE)

NOSE WHEEL WELL (LEFT SIDE)

Figure 12-30. Emergency Extension Blowdown Bottles

12-40




PNEUMATIC SYSTEM SERVICING

Landing Gear Strut Servicing

NOTE All pneumatic servicing requirements of the Gulfstream G500/G550 aircraft call for dry nitrogen.

Tire Servicing—Main Landing Gear/Nose Landing Gear Recommended tire pressures will yield optimum tire life and the lowest operating tire temperatures. These recommended pressures are listed in Chapter 32 of the Aircraft Maintenance Manual. The operator is required to maintain the minimum tire pressure at each given weight.

NOTE P r e s s u r e s a r e b a s e d o n t a k e o ff weight, not landing weight.

The procedure for servicing the landing gear struts is found in Chapter 32, “Landing Gear,” of the GAC Maintenance Manual rather than Chapter 12, “Servicing.” There are two methods of pneumatic servicing of landing gear struts: weight on wheels or aircraft on jacks. Using the weight-on-wheels method, the shock absorber “X” dimension may be checked with the aircraft on the ground using a high pressure gage (0 to 6,000 psi)and the information on the data plate for the shock absorber. This method is not as accurate as the following procedure (with the aircraft on jacks) due to shock absorber friction, and should not be used indiscriminately. The procedure for the aircraft-on-jacks method c a n b e f o u n d i n C h a p t e r 3 2 o f t h e G AC Maintenance Manual. It requires the use of a nitrogen source with a standard 3⁄4-inch servicing valve and MIL-H-5606, MIL-H-83282 or MIL-H-87257 hydraulic fluid (Figure 1229).

For weight off wheels condition, subtract 4%.

Tire pressure varies with temperature change. For temperature corrections, add 1% pressure for each 5°F above 70°F, subtract 1% for each 5°F below 70°F. Allow adequate time for tire to cool down after operation prior to checking pressure (approximately 2 hours).

WARNING Do not stand in front of wheel flanges while servicing.


12-41



O

P.

F

RELIE

XY. H.

RELIEF VALVE INDICATOR

0 100 0

200

EN YG OX NDER LI CY EW CR

0

FI

NG

ER

EN YG ALVEED X O R V ROV L L E A P P A RY I N G

PAS

SE

E N G CIS IO AR ER UT CH EX RECAN RE HALL P HE EA S Y S W E AR N, DRE P E EW ST VALV CLEA -FRE CR BE D OIL AN

100 0

0 200

0

EN YG OX NDER I L CY EW CR

Figure 12-31. Oxygen Service Panel

12-42




Landing Gear Blowdown Bottle

NOTES

A standard 3⁄4-inch servicing valve is used to replenish the nitrogen to the blowdown bottle system. The Gulfstream G500/G550 has two blowdown bottles located in the nose wheel well (Figure 12-30). One is on the left forward side, adjacent to the parking/emergency brake accumulator; the other is on the right forward side. Both bottles are serviced by a common manifold, via the servicing valve. The system is serviced to 3,100 ±50 psi at 70°F. Further servicing information for the emerg e n cy b l ow d ow n b o t t l e c a n b e f o u n d i n Chapter 32 of the GAC Maintenance Manual.

WARNING Before working in any wheel well, ensure all landing gear and landing gear door safety devices are installed.

Parking/Emergency Brake Accumulator A standard 3⁄4-inch servicing valve is used to replenish the nitrogen to the brake accumulator. The valve is located in the nose wheel well on the left forward side, adjacent to the landing gear emergency blowdown bottle servicing valve. The accumulator is serviced to 1,200 ±50 psi at 70°F (See Figure 12-16).

WARNING Before working in any wheel well, ensure all landing gear and landing gear door safety devices are installed.

CAUTION Aircraft should be chocked, as brakes will release when the accumulator is unloaded.


12-43



OXYGEN SYSTEM SERVICING The oxygen system of the Gulfstream G500/G550 aircraft comprises two independent systems: one for the flight station crew and one for the passengers. Only the flight station system is installed in production aircraft. The passenger system options are varied, and provisions are covered during aircraft outfitting.

WARNING Do not permit open flame, fire or ignition sources within 50 feet of aircraft during oxygen servicing.

WARNING Do not permit aircraft servicing or maintenance operations during oxygen servicing.

CAUTION If the cylinder is filled too rapidly, excessive heat will develop.

Oxygen System Servicing Port/Valve Prior to servicing, a visual inspection of the relief valve indicator should be accomplished. It is located at FS 254 on the lower right section of the fuselage, adjacent to the oxygen servicing panel (Figure 12-31). Although there are two independent oxygen systems on the Gulfstream G500/G550 aircraft, there is only one servicing port/valve for both systems. The system is configured with a common manifold assembly, which incorporates a check valve for each system to prevent depletion of both bottles in the case of leakage, failure, or isolated system use. The servicing port is located on the lower right fuselage, just forward of the fuselage wing fillet panel (Figure 12-31).

WARNING Ensure hands, tools and clothing are free of grease and oil. These contaminants will ignite upon contact with pure oxygen under pressure.

CAUTION Electrostatically ground the aircraft and electrostatically bond oxygen servicing equipment to the aircraft.

CAUTION Only use aviators breathing oxygen for servicing oxygen system. Do not use oxygen intended for medical purposes, or such industrial uses as welding. Such oxygen may contain excessive moisture that could freeze in valves and lines of the oxygen system.

12-44




ANTI-ICING/DEICING

NOTE

The Federal Aviation Authority prohibits takeoff with snow, ice, or frost adhering to the wings or control surfaces of the aircraft. The standard practice is to inspect and deice the aircraft using the appropriate means available and then depart as quickly as possible. Use of heated solutions for deicing, followed by an anti-icing process using a colder but more concentrated solution, produces a significantly lower freezing point on the aircraft surfaces. There are two types of deicing fluids, which are referred to as freezing point depressant (FPD) fluids. The Society of Automotive Engineers (SAE) or the International Standards Organization (ISO) Type I fluids are generally used in North America. SAE/ISO Type II fluids, also referred to as Association of European Airlines (AEA) fluids, are generally used in Europe.

Both Type I and Type II FPD fluids are recommended for use on Gulfstream aircraft (GAC Operations Manual, Volume 3, Chapter 7).

Anti-icing/Deicing Procedures Gulfstream does not recommend the use of undiluted Type II FPD fluids. The table used for flight crew guidelines in determining antiicing holdover times is also a good guideline for ground personnel in predetermining ratio parameters, concentration levels, and ambient temperatures associated with mixing and applying FPD fluids to the aircraft. Severe weather diminishes protection times for anti-ice fluids. Heavy precipitation rates, high moisture content, high wind velocity, or jet blast will reduce the estimated holdover range. Skin temperature that is lower than ambient temperature will also reduce protection time.

Anti-icing and Deicing Methods of Removal The removal of snow, ice, or frost can be accomplished by mechanical means, chemical means, or a combination of both. Removal by mechanical means involves using brooms, brushes, squeegees, and similar equipment. Removal by the use of chemicals can be a onestep process or a two-step process. The onestep process uses a heated, water-diluted deicing fluid to remove ice, snow, or frost and also to protect the treated surface from further immediate accumulation. The two-step process uses the same heated fluid as does the one-step process but is immediately followed by a colder, more concentrated anti-ice fluid. The two-step method usually involves using Type II FPD fluids. Removal by a combination of methods involves combining mechanical and/or chemical methods to properly deice the aircraft.

CAUTION If a de-icing solution is inadvertently sprayed into the engine or APU inlets or contacts the exhaust when the engines or the APU are operating, a potentially unsafe condition could develop in the cabin. Engine bleeds should be off and doors and outflow valve closed during de-icing operations to minimize the risk of cabin environment contamination. It is recommended that the APU not be running and the APU air inlet door be closed during de-icing operations.

CAUTION Do not use deice fluid on the brakes and/or wheels.


12-45



CHAPTER 21 AIR CONDITIONING CONTENTS Page INTRODUCTION ................................................................................................................. 21-1 GENERAL ............................................................................................................................. 21-1 BLEED-AIR SYSTEM .......................................................................................................... 21-3 General............................................................................................................................ 21-3 Engine and APU Source Control .................................................................................... 21-3 Bleed-Air Flow Control and Regulation System ............................................................ 21-5 Bleed-Air Control System Operation ........................................................................... 21-19 Bleed-Air Temperature Regulation System.................................................................. 21-25 AIR-CONDITIONING SYSTEM........................................................................................ 21-31 General.......................................................................................................................... 21-31 Airflow Control ............................................................................................................ 21-31 Operation and Indications............................................................................................. 21-53 Distribution ................................................................................................................... 21-57 Operation and Indications............................................................................................. 21-62 Cockpit and Cabin Temperature Control...................................................................... 21-65 Controls and Indicators................................................................................................. 21-71 CABIN PRESSURIZATION CONTROL SYSTEM (CPCS) ............................................. 21-75 General.......................................................................................................................... 21-75 System Components ..................................................................................................... 21-77 System Operation ......................................................................................................... 21-83 System Indications........................................................................................................ 21-89


21-i




Title

Page

21-1

Air-Conditioning Pack ........................................................................................... 21-2

21-2

ECS Control Panel ................................................................................................. 21-2

21-3

High-Stage Valve ................................................................................................... 21-4

21-4

Servo Pressure Regulator and Torque Motor ......................................................... 21-6

21-5

Mid-Stage Check Valve.......................................................................................... 21-8

21-6

Precooler Inlet Temperature Sensor ....................................................................... 21-8

21-7

Door Seal Regulating Manifold—5001 ............................................................... 21-10

21-8

Isolation Valve ..................................................................................................... 21-12

21-9

APU Load Control Valve ..................................................................................... 21-12

21-10

External Air Connection ...................................................................................... 21-14

21-11

External Air Check Valve .................................................................................... 21-15

21-12

Bleed-Air Augmentation Valve............................................................................ 21-16

21-13

Bleed-Air Control System Schematic.................................................................. 21-18

21-14

Normal/Descent Mode Trip Points ...................................................................... 21-20

21-15

ECS/PRESS Synoptic Page ................................................................................. 21-22

21-16

Bleed-Air Temperature Regulation System Block Diagram................................ 21-24

21-17

Fan-Air Modulation Valve and Servo Pressure Regulator ................................... 21-26

21-18

Precooler Outlet Temperature Sensor .................................................................. 21-28

21-19

Air Conditioning Flow Diagram .......................................................................... 21-30

21-20

Pack Inlet Valve ................................................................................................... 21-32

21-21

Ozone Converter .................................................................................................. 21-34

21-22

Air-Conditioning Pack Components.................................................................... 21-34

21-23

Air Cycle Machine............................................................................................... 21-36


21-iii



21-24

Temperature Control Flow Schematic ................................................................. 21-38

21-25

ACM Fan Plenum and Bypass Check Valve........................................................ 21-40

21-26

Compressor Bypass Check Valve ........................................................................ 21-42

21-27

ACM Compressor Outlet Temperature Sensor .................................................... 21-42

21-28

Reheater and Condenser Assembly ..................................................................... 21-44

21-29

Turbine Inlet Temperature Control Valve ............................................................ 21-44

21-30

Turbine Inlet Temperature Sensor........................................................................ 21-46

21-31

Low Limit Valve and Servo Pressure Regulator .................................................. 21-46

21-32

Turbine Bypass Valve .......................................................................................... 21-48

21-33

Air-Conditioning Pack Outlet Temperature Sensor ............................................. 21-48

21-34

Ram-Air Check Valve .......................................................................................... 21-50

21-35

Airflow Control Switches .................................................................................... 21-51

21-36

Airflow Control Schematic .................................................................................. 21-52

21-37

ECS/PRESS Synoptic Page—Airflow Control Indications................................. 21-54

21-38

Trim-Air Valves ................................................................................................... 21-56

21-39

Baggage Compartment and Ventilation Valve Reset Switch ............................... 21-58

21-40

Cockpit Airflow Valve ......................................................................................... 21-58

21-41

Forward Electrical Equipment Cooling ............................................................... 21-60

21-42

Forward Cabin Zone Temperature Sensor ........................................................... 21-64

21-43

Cockpit Zone Temperature Sensors (ASC 85A).................................................. 21-66

21-44

Supply Duct Temperature Sensor ........................................................................ 21-68

21-45

Supply Duct Overheat Thermostat....................................................................... 21-68

21-46

Air-Conditioning Control Panel........................................................................... 21-70

21-47

Air Temperature Control Schematic .................................................................... 21-72

21-48

Cabin Pressurization Control System Block Diagram......................................... 21-74

21-iv




21-49

Cabin Pressurization Control System Components—Cockpit ............................ 21-76

21-50

Cabin Pressurization Control System Components—REER............................... 21-80

21-51

Cabin Pressure Control and ECS/Press Synoptic Page........................................ 21-82

21-52

Pressure Relief Valve Cutaway ............................................................................ 21-86

21-53

Cabin Pressure Indicator ...................................................................................... 21-88

TABLE Figure

Title

Page

21-1

Cabin Pressure Controller Semi-Auto Schedule.................................................. 21-83

21-2

Cabin Pressure Controller Auto Schedule ........................................................... 21-84


21-v



CHAPTER 21 AIR CONDITIONING

INTRODUCTION The Gulfstream G500/G550 environmental control system (ECS) controls air pressure and temperature used for temperature control, ventilation, and dehumidification in the aircraft during flight or on the ground. The ECS also provides a breathable atmosphere for the crew and passengers. Compressed bleed air used by the ECS comes primarily from the aircraft engines. The auxiliary power unit (APU) and ground-supplied pneumatic equipment are the two alternate sources.

GENERAL Three major subsystems make up the ECS: the bleed-air (pneumatic) system, air-conditioning system, and cabin pressurization system. The bleed-air system supplies air for the air-conditioning, pressurization, and anti-ice systems. Two air-conditioning packs provide dehumidified temperature-controlled air to the forward and aft cabin zones, cockpit, and electronic equipment (Figure 21-1).

The crew monitors and controls pressurization of the aircraft through the cabin pressure indicator, the selector, and the pressure control panels. A dual-channel cabin pressure controller senses cabin altitude and ambient altitude. The pressurization system regulates the amount of air leaving the airplane through the outflow valve. The dual-channel cabin pressure controller establishes control of the air leaving the airplane through the thrust recovery outflow valve (TROV).


21-1



Figure 21-1. Air-Conditioning Pack

Figure 21-2. ECS Control Panel

21-2




BLEED-AIR SYSTEM

Engine Bleed-Air Selector Switches

GENERAL

Mounted on the overhead BLEED AIR control panel, the left and right engine bleed-air selector switches allow selection of the left or right engine as the bleed-air source (Figure 212). Switch ON selection provides a discrete input to the MAU 1 DGIO 1, for the left system and MAU 2 DGIO 2 for the right system for switch position. The switch ON selection also provides 28 VDC input to the bleed-air control relays through the de-energized contacts of the on-side bleed off relays. With the bleed-air control relays energized, a 28 VDC discrete input is sent to the bleed-air controllers ON command and MAU 1 DGIO 1 for the left side and Mau 2 DGIO 2 for the right side. A 28 VDC input from the on-side bleedair control relay is also sent to the bleed-air pressure regulator/shutoff valve.

The bleed-air system consists of the following subsystems: • Engine and APU source control • Flow control and regulation • Temperature regulation • Pressure indication It provides control, regulation, and monitoring of the bleed-air flow, pressure, and temperature in the left and right bleed-air manifolds. The bleed air is used for cabin air conditioning, wing leading-edge anti-ice, engine start, door sealing, aspiration of total air temperature (TAT) probes, and cabin pressure relief valves, as well as other miscellaneous services.

NOTE Refer to the Maintenance Schematic Manual, Chapter 21, for corresponding schematics.

ENGINE AND APU SOURCE CONTROL General The engine and APU control system provides ON–OFF control. It also provides for control of the automatic switching from APU air to left & right engine bleed-air at 1500 AGL. The engine and APU source control system is controlled by the BLEED AIR control panel, located on the cockpit overhead panel, just above the cabin temperature control panel. The BLEED AIR control panel contains the following controls: • Left and right guarded bleed-air switches • APU bleed-air switch

When turned off, the bleed-air switches illuminate amber OFF legends. The bleed-air control relay is de-energized, which in turn de-energizes the solenoid of the selected bleedair pressure regulating/shutoff valve. This also causes the selected controller command input, along with the MAU switch position inputs, to return to the OFF position.

APU Bleed-Air Selector Switch The APU bleed-air control system switch allows selection of the APU as a source of bleed air while the aircraft is below 1500 ft. AGL. A blue “ON” in the switchlight capsule appears on the BLEED AIR control panel when the switch is turned on. The selection of the switch to the ON position provides a switch position signal to the APU electronic control unit to open the APU air valve if the aircraft is below 1500 ft. AGL, and to the MAU 1 DGIO 1 for switch position status. Power is also applied to the isolation valve solenoid, automatically opening the isolation valve if pneumatic pressure is available. A ground is also applied to the bleed off relays.

• Isolation valve switch


21-3



60-MICRON FILTER ASC 48

REFERENCE PRESSURE REGULATOR

HIGH-STAGE VALVE MODULATED CHAMBER "A" –2 VALVE

HP 8 AIRFLOW

TORQUE MOTOR

OPEN

CNTRL LEVER

DOWNSTREAM PRESSURE

CLOSED 0–100mA B.A.C. 57±3PSI

CNTRL NOZZLE

FEEDBACK SERVO

Figure 21-3. High-Stage Valve

21-4




Isolation Valve Control Switch

Components

The isolation valve control switch is located below the APU bleed-air switch and provides the means to manually select the bleed-air isolation valve open in flight or on the ground. When the valve is closed, the switch face is black. When the isolation valve is open, the switch capsule illuminates as a white horizontal bar. This switch illuminates whenever the valve is open, regardless of switch position.

Bleed-Air Controllers (BACs)

BLEED-AIR FLOW CONTROL AND REGULATION SYSTEM General

The on-side BAC receives sensor data from the precooler inlet temperature sensor, the precooler outlet temperature sensor, and the pressure sensor (transducer). The on-side BAC also controls the high-stage valve via the servo pressure regulator/torque motor assembly to control the minimum bleed-air manifold flow. The fan-air valve is also controlled by the on-side BAC to regulate the maximum temperature in the tail compartment bleed-air manifold duct.

The bleed-air flow control and regulation system provides control, regulation, and monitoring of the bleed-air temperature and pressure in the left and right bleed-air manifolds (Refer to MSM, chapter 21). The system engine-mounted components consist of the following:

The bleed-air controllers are dual-purpose and dual-powered microprocessors with software designed to automatically control and regulate the engine bleed-air manifold pressure and temperature. Located in the baggage compartment electronics equipment rack, they also control the on-side wing anti-ice system.

• Mid-stage check valve

The left or right BAC communicates over the ARINC 429 data bus to the MAU 1 DGIO 1 for the left system and MAU 2 DGIO 2 for the right system. It receives outside air temperature, aircraft altitude, engine N 1 speed, bleed-air switch status, air-conditioning pack switch status, and wing anti-ice switch status data from the MAUs. It also transmits sensor data for display and warning and system fault data to the on-side MAUs and CMC.

• Manifold pressure regulator/shutoff valve

High-Stage Valve

• Fifth- and eighth-stage HP compressor bleed ports • Servo pressure regulator/torque motor assembly • High-stage valve

• Precooler inlet temperature sensor Additional components mounted in the tail compartment are as follows: • Precooler outlet temperature sensor • Manifold pressure sensor • Door seal regulator manifold • Bleed-air isolation valve • APU load (APU enclosure)

control

valve

• APU isolation check valve • External air isolation check valve

The high-stage valve (Figure 21-3) is a pneumatically operated, spring-loaded-closed modulating valve. It is located at the HP eighth-stage bleed ports and is connected to the interservices fairing on each BR710 engine. The function of the high-stage valve is to augment or replace mid-stage (HP5) bleed air when the bleed-air manifold pressure drops below the minimum required by the operating configuration. It has no electrical components and receives its opening servo pressures from the servo pressure regulator/torque motor as commanded by the bleed-air controller. The high-stage valve downstream pressure provides modulation.

• Bleed-air augmentation valve FOR TRAINING PURPOSES ONLY

21-5



Figure 21-4. Servo Pressure Regulator and Torque Motor

21-6




Servo Air Pressure Regulator and Torque Motor Assembly

NOTES

The servo air pressure regulator and torque motor assembly is located on the bottom of each engine near the high-stage valve (Figure 21-8). This unit is a diaphragm-actuated, poppet-type regulator combined with an electromechanically actuated modulating valve (torque motor) at the outlet. Its purpose is to provide operating pressure to control the high-stage bleed valve which is proportional to the torque motor input power level provided by the BAC. The servo air pressure regulator receives high-stage bleed pressure, which it regulates to 57 ±3 psi for the torque motor. The BAC provides an electrical current, ranging from 0 to 100 mA, to the torque motor. A 60 Micron element inline filter is installed in the high (8th) stage duct, upstream of the high stage valve servo pressure regulator/torque motor assembly. The filter eliminates FOD (including soot) found in the servo pressure regulator.


21-7



Figure 21-5. Mid-Stage Check Valve

Figure 21-6. Precooler Inlet Temperature Sensor

21-8




NOTES

Mid-Stage Check Valve The mid-stage check valve (Figure 21-5) is a duct-mounted, pin-located, spring-loaded closed, split-flapper-type check valve which prevents reverse airflow into the fifth-stage bleed port when the high-stage valve is open.

Bleed-Air Pressure Regulator and Shutoff Valve The bleed-air pressure regulator and shutoff valve is a 4-inch-diameter, spring-loaded closed, solenoid-controlled, pressure-regulated valve. The unit is located downstream of the engine bleeds and regulates bleed-air manifold pressure to a maximum of 40.5 ±3.5 psig. When the onside bleed-air switch is selected to the ON position, the bleed-air pressure regulator and shutoff valve solenoid receives 28-VDC power.

NOTE Refer to the Maintenance Schematic Manual, Chapter 21, for corresponding schematics

Precooler Inlet Temperature Sensors A precooler inlet temperature sensor (Figure 21-6) is located in each bleed-air manifold duct just upstream of the exit from the engine nacelle to the pylon. Each sensor (left and right) provides temperature data to its respective bleed-air controller. This data is used by the bleed-air controller to control the highstage flow during wing anti-ice operation. It also provides for a bleed hot fault indication.


21-9



TAT PROBE SOLENOID VALVE

DOOR SEAL REGULATOR

MISC. BLEED LINE

Figure 21-7. Door Seal Regulating Manifold—5001

21-10




Precooler Outlet Temperature Sensor Mounted downstream of the precooler heat exchanger in the tail compartment, each precooler outlet temperature sensor is a single nonrepairable LRU with a two-piece corrosionresistant steel housing, a temperature-sensing element, and a connector. The temperature sensor is a hermetically sealed unit that uses a platinum element to sense the precooler outlet air temperature. The output signal from the temperature sensor controls the fan air valve and precooler outlet temperature. The temperature is also displayed on the ECS/PRESS synoptic.

Manifold pressure is also supplied to the following: • Total air temperature (TAT) probe aspiration solenoid valve, which aspirates both the left and right total air temperature probes when the weight on wheels interface is in the ground mode. The TAT valve also supplies pressure to the cabin pressure relief valve jet pump when the weight on wheels interface is in the ground mode. • Miscellaneous bleed-air pressure, normally used to pressurize the aircraft water system.


NOTES

Bleed-Air Manifold Pressure Sensors (Transducers) The left and right bleed-air manifold pressure sensors (transducers) are located in the tail compartment. They provide manifold pressure readings to their respective bleed-air controllers. This pressure reading is used to control high-stage flow when the pressure drops below the minimum required for the aircraft operating environment and configuration. It also provides high and low pressure warnings for the bleed-air system.

Door Seal Regulating Manifold The door seal regulating manifold is located in the tail compartment and provides bleed-air pressure from either or both the left and right bleed-air manifolds (Figure 21-7). Check valves isolate this flow from the left and right bleed-air manifolds. The manifold supplies bleed air to the door seal regulator, which in turn supplies 18 ±1 psi air through the baggage compartment smoke evacuation valve to the baggage door inflatable seal.


21-11



MANUAL LOCKING COLLAR

Figure 21-8. Isolation Valve

Figure 21-9. APU Load Control Valve

21-12




Isolation Valve The isolation valve (Figure 21-8) is a bidirectional pneumatic shutoff valve; it is 3 inches i n d i a m e t e r, s p r i n g - l o a d e d c l o s e d , a n d solenoid-operated. Located between the left and right bleed-air manifolds, it provides isolation of the left and right bleed-air manifolds when in the normally closed position. The valve incorporates a manual wrenching device. An internal valve position indication switch closes when the valve is in the open position. The position switch provides an OPEN indication in the switchlight as well as generating a blue ISO VALVE OPEN EICAS message. Additionally, the isolation valve position is displayed on the ECS/pressurization synoptic page. Pressure from either the left or right bleed-air manifold can open the valve when the solenoid is energized. The isolation valve is energized open when one of the following switches is selected: • APU air switch, aircraft below 1,500 ft AGL. • Engine master start or crank switch

APU Load Control Valve and APU Check Valve The APU load control valve is located on the APU turbine case and is connected to the bleed-air ducting (Figure 21-9). The valve opens to provide bleed air from the APU while the aircraft is on the ground and up to 1500 feet AGL. The load control valve can also supply bleed air if needed for main engine starting in flight, up to 30,000 feet. The APU check valve is a 3.5-inch-diameter, spring-loaded, split-flapper-type check valve. It is located in the APU bleed-air ducting and opens to allow airflow from the APU load control valve to the right bleed-air manifold ducting.


The check valve closes to prevent reverse flow from the right manifold when it is pressurized from another source and bleed manifold pressure exceeds APU pressure.

• Isolation valve control switch An amber “Bleed Configuration” EICAS message and amber precooler and duct symbols are displayed on the ECS/PRESS synoptic page anytime the isolation valve is in the open position and both the left and right engine bleed air switches are selected on, or the APU air switch and either the left or right engine bleed air switches are selected on. A blue “Bleed Configuration” message is also displayed on EICAS if the APU air switch is selected on and the left and right engine bleed air switches are selected off, or the APU air switch is selected on and the APU is not on speed.


21-13



Figure 21-10. External Air Connection

21-14




External Air Connection and External Air Check Valve The external air connection provision is located through an access door on the external fuselage forward of the tail compartment door (Figure 21-10). It provides an external source of air pressure, used primarily to start the main engines when the APU is not available.

The external air check valve is a duct-mounted, pin-located, spring-loaded, split-flapper-type check valve located in the external air duct (Figure 21-11). It allows flow from an external air source to pressurize the right bleed-air manifold and prevents reverse airflow to the external air connector.

Figure 21-11. External Air Check Valve


21-15



Figure 21-12. Bleed-Air Augmentation Valve

21-16




Bleed-Air Augmentation Valve (BAAV)

NOTES

The bleed-air augmentation valve (BAAV) is a spring-loaded-closed, solenoid-controlled valve located in the left bleed-air manifold (Figure 21-13). It is energized open by the ECU when starting the APU above 35,000 feet.


21-17

21-18

R BLEED AIR CONTROLLER +28 VDC AUTO

TO WING A/I CONT

+28 VDC AUTO RETURN MAU 2

R ESS 28 VDC BUS

BLEED AIR ON/OFF COMMAND CHASSIS GND WEIGHT ON WHEELS

DUAL GENERIC I/O MODULE 2

WING ANTI-ICE

AIR

R BLEED-AIR SW ON (GND/OPN) GND

(PART OF WOW RELAY)

ID PIN 1 ID PIN 3


ANNUN LTS PWR

ANNUNCIATOR LTS DIM & TEST CONTROLLER

TORQUE MOTOR

(OPN/GND)

ON

(OPN/GND) OFF

ON

SOLENOID (ENG MOUNTED)

R BLEED-AIR CONTROL RELAY

OFF

28 VDC FROM APU AIR SWITCH

MAU 2

TO ISOLATION VALVE SOLENOID

R PRECOOLER INLET RESISTIVE ELEMENT

R MANIFOLD PRESS SENSOR IN OUT

PRESSURE SENSOR

DUAL GENERIC I/O MODULE 2 R BLEED-AIR SW ON (GND/OPN)

28VDC FROM ICE DETECT SYS <1500 FT

R SERVO AIR PRESS REG TORQUE MOTOR

OFF

ANNUNCIATOR LTS DIM & TEST CONTROLLER DIM TEST

MAU 1

APU AIR SW ON (28 VDC OPN)

(OPN/GND)

TORQUE MOTOR

28VDC FROM APU AIR SW

TO ISOL VALVE SOLENOID

28VDC FROM L B-AIR SW ON

APU CONTROL 28 VDC FROM R BATT BUS AND/OR L ESS 28 VDC BUS

T/M

R PRECOOLER OUTLET TEMP SENSOR

RESISTIVE ELEMENT

TO BLEED-AIR CONTROL RELAY

(OPN/GND)

OFF

L BLEED-OFF RELAY

SERVO REG TORQUE MTR SIGNAL SERVO REG TORQUE MTR RETURN

PRECOOLER OUTLET TEMP SENSOR SIGNAL PRECOOLER OUTLET TEMP SENSOR RTN CHASSIS GND

MAU 2 DUAL GENERIC I/O MODULE 2 ARINC 429 TRANSMIT R FACEC GENERAL ARINC 429 RECEIVE R BASS FAIL (GND/OPN)

Figure 21-13. Bleed-Air Control System Schematic

BLD AIR SWITCH STATUS, PACK SWITCH STATUS WING ANTI-ICE SWITCH STATUS ALT, OAT, N1 ARINC 429 TRANSMIT CONTROLLER FAIL

international

TO L/R BLEEDOFF RELAYS APU AIR SWITCH (COCKPIT OVERHEAD PANEL)

BLEED AIR ISO VALVE SOLENOID

MANIFOLD PRESSURE SENSOR SUPPLY MANIFOLD PRESSURE SENSOR SIGNAL

FlightSafety

DUAL GENERIC I/O MODULE 1

PRECOOLER INLET TEMP SENSOR SIGNAL PRECOOLER INLET TEMP SENSOR RTN

ON

R BLEED-OFF RELAY

L ESS 28 VDC BUS

FAN AIR VLV TORQUE MTR RETURN

R MANIFOLD PRESS REG/SHUTOFF VALVE

DIM TEST

OFF

FAN AIR VLV TORQUE MTR SIGNAL

T/M


R FAN AIR VALVE (COP) BLEED-AIR SWITCH R ENGINE



BLEED-AIR CONTROL SYSTEM OPERATION

NOTES

BLEED AIR Control Panel The BLEED AIR control panel provides the means to control system operation. Selection of the left and right engine bleed-air control switches to the ON position provides power from the on-side essential 28-VDC bus to the respective BAC relay to initiate bleed-air control. This energizes the bleed-air pressure regulator solenoid (Figure 21-13).


When powered, each bleed-air controller provides sensor display and fault data to MAU 1 DGIO 1 for the left system and MAU 2 DGIO 2 for the right system. Isolating diodes allow dual power through either the bleed-air control or wing anti-ice circuit breakers. With the bleed-air switch selected OFF, the bleed-air control relay is deenergized, and the signal to the BAC relay is removed. The bleed switch position discrete signal to MAU 1 DGIO 1 for the left system and MAU 2 DGIO 2 for the right system is also off, and the pressure regulator and shutoff valve solenoid deenergizes. Selecting the bleed-air switch to the ON position energizes the bleed-air control relay. This provides 28-VDC power to the BAC for bleed air on command to initiate bleed-air control. This selection also provides 28-VDC power to the bleed-air pressure regulator and shutoff valve solenoid and supplies the on-side bleed switch ON position discrete signal to the No. 1 or No. 2 MAU. The mid-stage (HP5) bleed is the normal source of engine bleed air. During normal operation, mid-stage air supplies all bleed flow during takeoff, climb, and cruise. If the bleedair manifold pressure decreases to below 14 ±2 psi, the high-stage bleed valve opens to maintain a minimum of 14 ±2 psi.


21-19

21-20 ENGINE PERCENT N1 VERSUS AIRCRAFT ALTITUDE AND STATIC AIR TEMPERATURE


ENGINE N1 (%)

90

14 PSIG CRUISE REGION

80

70 22 PSIG DESCENT/HOLD REGION

60

50 –1000

20000

25000

30000

35000

40000

41000

43000

45000

ALTITUDE (FT) 10

0

20

40

60

80

49000

51000

international

FlightSafety

TEMP °C

47000


100

Figure 21-14. Normal/Descent Mode Trip Points



At low power on the ground and for descent, VMIN, and hold operation, the BAC receives engine N1, static air temperature, and pressure altitude data from the ARINC 429 data bus. At a predetermined trip point, the BAC resets high-stage regulation to maintain a 22 psi minimum bleed-air manifold pressure. As an example of the predetermined trip point, when the aircraft is at an altitude of 47,000 feet, with a static air temperature of –60°C, and an engine N 1 speed of less than 77%, the bleedair controller will establish a minimum manifold pressure of 22 psi (Figure 21-14).

override opening during ground operation. The isolation valve opens when the engine master start or crank switch is selected to the ON position.

NOTE Only one of these switches is selected at a time. Selection of master start switch and master crank switch simultaneously will result in a START SW CONFIG CAS message regardless of weight on wheels condition.

During single air-conditioning pack operation, the high-stage bleed valve opens to maintain a minimum of 35 psi in the bleed-air manifold.

NOTES

Wing Anti-Ice Operation During wing anti-ice operation, the high-stage valve opens to augment mid-stage bleed-air pressure. The BAC controller holds the highstage valve open until 620°F is achieved at the precooler inlet temperature sensor. Achieving 620°F depends on operation of the engine parameters. At lower engine rpms, 620°F may not be possible.

APU Air Operation The APU bleed-air switch allows selection of the APU as a bleed-air source if the aircraft is below 1,500 ft. AGL. Selecting the APU switch on the BLEED AIR control panel provides a signal to the APU ECU requesting APU air if the aircraft is below 1500 ft. AGL. Selecting the APU air switch also energizes the isolation valve solenoid when the aircraft if the aircraft is below 1,500 ft. AGL and energizes the “Bleed off” relays which deenergizes the left and right engine bleed air control relays.

Isolation Valve Operation When the isolation valve switch is selected to the ON position, the isolation valve automatically opens to connect the left and right bleedair manifolds. If the isolation valve does not open electrically, a manual wrenching fixture at the butterfly shaft end allows emergency


21-21



ECS/PRESSURIZATION Zone Temperatures ∞F

76∞ 78∞

76∞ 78∞

74 ∞F

80∞ 80∞

Zone Temperatures ∞F

78 ∞F

74 ∞F

70∞ 70∞72∞ 7 2 ∞ 75∞ 75∞ 8 5 ∞F 101 ∞F

78 ∞F

35 ∞F

35 ∞F

Air

78 ∞F

35 ∞F 0 psi 370 ∞F

370 ∞F

42 psi

42 psi

Ldg Elev Cab Alt -450 1030 Right Eng

78 ∞F

78 ∞F APU

Cab Alt -450

Rate 0

P 0.25

Mode Auto 1

Figure 21-15. ECS/PRESS Synoptic Pages

21-22

35 ∞F

0 psi

Left Eng

L Elev 1030

9 2 ∞F


Rate 0

Mode Auto1 P .25



Indications and Display

NOTES

The ECS/PRESS 2/3 and 1/6 Synoptic pages provide the primary display for bleed air flow control and regulation. (Figure 21-15) The 2/3 synoptic page displays bleed air, air conditioning and pressurization information. The 1/6 synoptic page displays air conditioning and pressurization information only. The ECS/PRESS synoptic page displays the positions of the left and right engine bleed-air pressure regulating/shutoff valve, APU load control valve, and the isolation valve. It also displays left and right bleed-air manifold pressures as well as precooler inlet and outlet temperatures. When the bleed-air pressure is at 75 psi or higher, an amber “L or R Bleed Pressure Low” message is displayed on the EICAS. When the left or right bleed pressure is 5 psi or less for more than 10 seconds, an amber “L or R Bleed Pressure Low” message appears. Bleed-air pressure is also displayed on the summary, engine start, and APU synoptic pages.


21-23



COWL ANTI-ICE SYSTEM

SERVOPRESSURE REGULATOR 46 ±10 PSI

BLEED PRESSURE REGULATOR VALVE

HP8 BLEED PORT

AIR TURBINE STARTER STARTER AIR VALVE

FAN AIR PORT

RIGHT BAC

ARINC TXR FAN AIR VALVE SERVOPRESSURE

R BLEED-AIR CONTROL RELAY

MAU 2 CUSTOM I/O MODULE 28VDC FROM ICE DETECT SYSTEM < 1500 FT

R BLEED OFF RELAY

PRECOOLER OUTLET TEMP SENSOR

MAU 2 MAU GEN 2 DUAL GENERIC I/O MODULE

MANIFOLD PRESSURE SENSOR AC PACK INLET VALVE

R BLEED AIR SW ON (GND/OPN) R BLEED AIR SW ON (28 VDC OPN)

WING ANTI-ICE CONTROL VALVE

ARINC TXR ARINC REC

APU ELECTRONIC CONTROL UNIT

EXTERNAL AIR CHECK VALVE

MAU 1 MAU GEN 2 DUAL GENERIC I/O MODULE

MISC BLEEDS BAGGAGE DOOR SEAL DOOR SEAL REGULATOR

APU AIR LOAD VALVE

APU AIR CHECK VALVE

TAT AND PRV

APU AIR SW ON (28 VDC/OPN)

ISOLATION VALVE BAAV APU AIR INLET

LEFT ENGINE IDENTICAL ENGINE START PANEL

LEGEND SHUTOFF VALVE

ESSENTIAL 28 VDC PRESSURE/FLOW CONTROL

REGULATING VALVE

TEMPERATURE CONTROL

TEMPERATURE SENSOR PRESSURE SENSOR

ARINC 429 DATA BUS

CHECK VALVE

Figure 21-16. Bleed-Air Temperature Regulation System Block Diagram

21-24

ARINC REC


APU



BLEED-AIR TEMPERATURE REGULATION SYSTEM General The bleed-air temperature regulation system uses engine fan discharge air as a cooling medium to control temperature inside the aircraft tail compartment bleed-air ducting. The units and components within the bleed-air temperature regulation system include the following (Figure 21-16):

Because all engine bleed air flows through the precooler heat exchanger, bleed-air temperatures may vary widely and additional cooling capability is sometimes required, depending on engine power setting and use of fifth- or eighth-stage extraction. Engine fan air can be directed to flow across the precooler heat exchangers for cooling when necessary. The bleed-air controller monitors precooler outlet temperature and modulates the opening of the fan-air valve as required.

NOTES

• Fan-air ports • Precooler heat exchangers • Engine fan-air modulation valves • Servo pressure regulators • Precooler outlet temperature sensors


Components Fan-Air Port and Precooler Heat Exchanger The fan-air port provides for control of engine bleed-air temperature by supplying air via the engine bypass duct to the precooler heat exchanger. Located in each engine pylon, the precooler heat exchanger provides for control of the maximum temperature of engine bleed air, using fan discharge air as a cooling medium. Each engine bleed-air precooler heat exchanger is composed of a single-pass, crossflow plate and fin assembly of brazed and welded construction. The left and right precooler heat exchangers differ only in port configuration and mounting provisions required to adapt to each pylon; however, since there is a difference, they do have different part numbers.


21-25



Figure 21-17. Fan-Air Modulation Valve and Servo Pressure Regulator

21-26




Engine Fan-Air Modulation Valve

NOTES

The engine fan-air modulation valve (Figure 21-17) regulates the engine fan air flowing through the precooler heat exchanger. It is located aft of the precooler in the pylon. The fanair valve is a 4-inch-diameter, spring-loaded open, pneumatically operated, torque-motorcontrolled, butterfly-type modulating and shutoff valve. To modulate the fan-air valve, the BAC must apply a 0 to 100 mA current to the valve’s torque motor. Temperature at the precooler outlet is controlled to 400 ±10°F (204°C) under normal operating conditions, within the capacity of the engine. During single-bleed operation with wing anti-ice on, precooler outlet temperature is controlled to 500 ±10°F (260°C), within the capacity of the engine. This is also true when both bleeds are on and a single wing anti-ice valve is operating.

Servo Pressure Regulator One regulator is installed in each pylon with its air inlet fitting installed upstream of the bleed-air pressure regulator/shutoff valve (40psi valve) (Figure 21-17). The regulated outf l ow i s f u r n i s h e d t o t h e e n g i n e f a n a i r modulation valve and is calibrated to 16 ±1 psi with relief set at 25.5 ±1 psi. If the outlet pressure exceeds 25.5 ±1 psi, the diaphragm force opens the relief valve to vent the excess pressure. This usually occurs during transient reduced-demand conditions.


21-27



Figure 21-18. Precooler Outlet Temperature Sensor

21-28




Precooler Outlet Temperature Sensor Mounted downstream of the precooler heat exchanger in the tail compartment, each precooler outlet temperature sensor is a single nonrepairable LRU with a two-piece corrosionresistant steel housing, a temperature-sensing element, and a connector (Figure 21-18). The temperature sensor is a hermetically sealed unit that uses a platinum element to sense the precooler outlet air temperature. This precooler output temperature is used by the bleed controller for control, display, and fault monitoring.

The specific temperature values are provided from the BAC via the ARINC 429 data bus and are displayed on the ECS/PRESS synoptic page. Overtemperatures (765 ±10°F) at the precooler inlet temperature sensor result in an amber “L or R Bleed Hot” fault warning message. Overtemperatures (550 ±10°F) at the precooler outlet temperature sensor result in an amber “L or R Bleed Hot” fault warning message.

NOTES

Operation and Indications The bleed-air temperature regulation system uses fan air as a cooling medium through a precooler heat exchanger to control bleed-air temperature, monitoring precooler outlet temperature and modulating the engine fan-air modulation valve as required. Engine fan air crossflows through the precooler to cool the bleed air when necessary. Temperature at the precooler outlet is controlled to 400 ±10°F (204°C) under normal operating conditions, regardless of wing anti-ice operation, within the capacity of the engine.


During single-bleed operation, precooler outlet temperature is controlled to 500 ±10°F (260°C) during wing anti-ice operation, within the capacity of the engine. This is also true when both bleeds are on and a single wing anti-ice valve is operating.


21-29



AIR CONDITIONING SYSTEM • COCKPIT CONTROLLED TEMPERATURE • COOLS CABIN AND COCKPIT ZONES • CABIN PRESSURIZATION

Figure 21-19. Air Conditioning Flow Diagram

21-30




AIR-CONDITIONING SYSTEM


GENERAL The air-conditioning system conditions fresh air taken from the bleed-air system and supplies it to the passenger cabin and cockpit to maintain a comfortable environment for the aircraft occupants and crew (Figure 21-19). It also provides ventilation and temperature control for the cockpit and the forward and aft cabin zones and cooling for the electronic equipment using return airflow. The system consists of the following subsystems:

NOTES

• Airflow control • Distribution • Temperature control

AIRFLOW CONTROL General The airflow control system relies on the airconditioning packs and their supporting equipment and controls the cool and dehumidified air supply as well as hot air to maintain selected temperatures and ventilation rates within the cockpit and cabin. The conditioned air also provides aircraft pressurization and cools the avionics. The air-conditioning airflow control system is split into the left and right systems, each with its own air-conditioning controller, pack inlet valve, ozone converter, and air-conditioning pack. the left and right air-conditioning controllers mounted in the baggage compartment control the airflow through the left and right air-conditioning packs. Each pack inlet valve’s torque motor receives a milli-amp current from the on-side controller to regulate the flow of bleed air through the valve.


21-31



Figure 21-20. Pack Inlet Valve

21-32




Components Air-Conditioning Controllers (ACCs) The air-conditioning controllers located in the baggage compartment vertical rack are dual-powered, dual-function microprocessors that provide airflow control through the packs and control pack outlet temperature. They also provide automatic temperature control of the cockpit and the forward and aft cabin zones. Each controller receives hard-wired discrete signals from the temperature sensors, which are used to modulate torque-motor-controlled valves, thus regulating airflow and temperatures. Data is received and transmitted over the ARINC 429 data bus for control, display, and fault indication.

Pack Inlet Valves (PIVs) The pack inlet valves (PIVs) are located in the pack bleed-air supply ducts, at the left and right of the tail compartment (Figure 21-21). They are controlled by a differential pressure (DP) servo valve which receives inputs from a torque motor to influence flow rates. A shutoff solenoid is incorporated to close the valve. The pack inlet valve is a spring-loaded-open, flow-regulating and shutoff valve, which normally regulates the flow of bleed air via torque motor inputs from the air-conditioning controller. It closes to shut off bleed-air flow when its shutoff solenoid is energized from the cockpit.

During single-pack operation, torque motor current is reduced to 0 mA. The pack inlet valve is positioned to its spring-loaded full open position and maximum nominal airflow can be achieved. The pack inlet valve also protects the air cycle machine (ACM) from over-speeding with a current of up to 100 mA from the air-conditioning controller. In this minimum flow condition, a minimum of 25 ppm airflow rate enters the air-conditioning pack and the ACM turbine speed is reduced, preventing the turbine from exceeding 100,000 rpm.

ASC 004 ASC 004 installs upgraded pack inlet valves in the left and right bleed air supply ducts to improve cabin airflow during certain flight conditions. As a result of this service change, the air conditioning/bleed air controllers also will require an upgrade. With this service change, the air controllers will incorporate the following enhancements: • Upgraded pack inlet valve flow schedules • Upgraded default A429 parameters related to pack inlet valve flow schedules • Added pack inlet valve monitoring logic (pack undertemp signal)

NOTES

The air-conditioning controller normally uses aircraft altitude received from the ARINC 429 data bus to schedule airflow through the pack inlet valve and air-conditioning pack. At sea level, the on-side air-conditioning controller provides approximately 82 mA current to the pack inlet valve torque motor. As the aircraft ascends, the torque motor current gradually decreases. When the aircraft is at 51,000 feet, the torque motor current is approximately 38 mA. This increases the airflow through the pack inlet valve and pack to compensate for decreased air density.


21-33



Figure 21-21. Ozone Converter SECONDARY HEAT EXCHANGER COMPRESSOR OUTLET TEMPERATURE SENSOR

PRIMARY HEAT EXCHANGER

WATER EXTRACTOR (HIDDEN)

ACM FAN PLENUM AND BYPASS CHECK VALVE

COMPRESSOR BYPASS CHECK VALVE

ACM TURBINE BYPASS VALVE

PACK OUTLET TEMP SENSOR

AIR CYCLE MACHINE REHEATER/ CONDENSER

LOW LIMIT VALVE

Figure 21-22. Air Conditioning Pack Components

21-34




NOTES

Ozone Converters Ozone converters are installed for the left- and right-side cooling systems (Figure 21-21). Each one is mounted in its respective bleed-air ducting between the pack inlet valve and the inlets to the hot-air manifold and air-conditioning pack. The ozone converters change bleed-air ozone to oxygen before it enters the air-conditioning packs to enhance passenger comfort.

Air-Conditioning Pack (ACP) Assemblies The air-conditioning pack (ACP) assemblies are installed in the tail compartment downstream of the ozone converters. Each ACP consists of the following components (Figure 21-22): • Three-wheel air cycle machine (ACM) • Primary and secondary heat exchanger • ACM fan, plenum and bypass valve • ACM compressor bypass check valve • ACM compressor outlet temperature sensor • Reheater and condenser assembly • Water extractor and spray nozzle • Turbine inlet temperature control valve • Turbine inlet temperature sensor • ACM turbine bypass valve • Low-limit valve • Servo pressure regulator • Pack outlet temperature sensor


21-35



COMPRESSOR INLET TURBINE INLET

TURBINE OUTLET

COMPRESSOR OUTLET (HIDDEN IN PHOTO)

TURBINE INLET TEMPERATURE SENSOR CONNECTION (HIDDEN IN PHOTO)

FAN OUTLET (HIDDEN IN PHOTO)

FAN INLET

Figure 21-23. Air Cycle Machine

21-36




NOTES

Air Cycle Machine (ACM) The air cycle machine is the heart of the airconditioning pack assembly. Each ACM consists of a compressor, a turbine, and fan components (Figure 21-23). The ACM operates in a “bootstrap” fashion to cool pack air at its outlet. The ACM maximum operating speed is approximately 95,000 to 100,000 rpm. Each ACM also incorporates two journal air bearings to support axial shaft rotation and a set of thrust air bearings to support fore and aft movement of the assembly.


21-37

21-38

RAM AIR ENGINE START

TEMP CONTROL L PACK

RAM AIR

MASTER

R PACK

CRANK

START START

L ENG

R ENG

TO AC PACK OUTLET

FROM BLEED-AIR SYSTEM

LEFT AC PACK

PACK INLET VALVE


HOT AIR MANIFOLD

REGULATED PRESSURE

ACM COMPRESSOR BYPASS CHECK VALVE

400-500°F RIGHT BLEED

GASPER AIR ACM FAN OVERBOARD

35°F COLD AIR MANIFOLD RIGHT PACK GASPER AIR

TURBINE BYPASS VALVE

Figure 21-24. Temperature Control Flow Schematic

international

PACK OUTLET TEMP SENSOR

FlightSafety

AC CONTROLLER (ACC)


LOW LIMIT VALVE



Primary and Secondary Heat Exchanger

NOTES

The primary and secondary heat exchanger is a single assembly with two different bleed-air passages. It provides the first and second stages of bleed-air cooling and uses ram air in flight or induced air on the ground as a cooling medium. The bleed-air temperature is lowered by this assembly without significantly reducing pressure. The left and right heat exchangers are installed in the tail compartment between the ram-air inlet duct and the fan plenum chambers (Figure 21-30). The primary heat exchanger provides for cooling the air exiting the ozone converter, while the secondary heat exchanger cools air from the ACM compressor outlet (Figure 21-24).


21-39



BYPASS CHECK VALVE RAM/FAN AIR INLET

RAM/FAN AIR OUTLET

ACM FAN MOUNT

RAM/FAN AIR INLET

BYPASS CHECK VALVE

RAM AIR (FLIGHT)

FAN AIR (INDUCED ON THE GROUND)

RAM/FAN AIR OUTLET

WATER DRAIN PORT

Figure 21-25. ACM Fan Plenum and Bypass Check Valve

21-40




ACM Fan Plenum and Bypass Check Valve

NOTES

The ACM fan plenum and bypass check valve are part of the air-conditioning pack assembly and are attached to the bottom of the primary and secondary heat exchanger (Figure 21-25). The plenum directs the flow of the ram air out of the primary and secondary heat exchanger and overboard. It also provides drainage of water that condenses in the primary and secondary heat exchanger as the ram air is exhausted overboard through a vent mounted in the fuselage. The bypass check valve opens in flight to allow most of the airflow to bypass the fan and be directly ported to the overboard vent.


21-41



Figure 21-26. Compressor Bypass Check Valve

Figure 21-27. ACM Compressor Outlet Temperature Sensor

21-42




Compressor Bypass Check Valve

NOTES

The compressor bypass check valve is mounted in-line between the primary heat exchanger outlet and the secondary heat exchanger inlet ducting (Figure 21-26). The ACM compressor bypass check valve allows a portion of the airflow to bypass the compressor during the initial startup of the pack. This airflow passes through the downstream components to the ACM turbine, providing the power to drive the compressor and fan. When compressor discharge pressure exceeds inlet pressure, the check valve closes.

ACM Compressor Outlet Temperature Sensor The compressor outlet temperature sensor is located downstream of the associated compressor outlet and before the secondary heat exchanger inlet (Figure 21-27). The sensor provides air cycle machine compressor outlet temperature as a discrete electrical signal to the air-conditioning controller. This signal is used to provide the feedback logic for monitoring and controlling the ACM maximum speed. The compressor outlet temperature sensor is a sealed, nonrepairable unit with a single platinum element. It has a positive temperature coefficient and a nominal resistance of 500 ohms at 32°F (0°C). The air-conditioning controller monitors compressor outlet temperature. If the temperature reaches 425°F (217°C), the controller drives the pack inlet valve into a minimum flow (25 ppm) condition to protect the air cycle machine from overspeed. Reducing the airflow through the pack should reduce the temperature at this sensor. Should the resistance of this sensor input to the air-conditioning controller indicate greater than 450°F (238°C), an amber L or R COOL TURB HOT message is generated on the EICAS.


21-43



Figure 21-28. Reheater and Condenser Assembly

Figure 21-29. Turbine Inlet Temperature Control Valve

21-44




Reheater and Condenser Assembly The reheater and condenser assembly is mounted on the forward end of the air-conditioning pack assembly (Figure 21-29). This is a single welded, dual-flow heat exchanger assembly. Its purpose is to cool the air to condense moisture content for water extraction, then reheat the air prior to its entering the turbine. It also provides mounting provisions for the servo pressure regulator, water extractor, and air cycle machine turbine bypass valve. The reheater/condenser assembly includes a provision for anti-icing the condenser, sense ports for deicing the condenser, and a bleed port for air supply to the servo pressure regulator.

Water Extractor and Water Spray Nozzle One water extractor is mounted on each pack between the condenser outlet (first pass) and reheater inlet (second pass). It is an in-line, integral duct-type device that incorporates a helix (swirl vanes) and water shave-off collector. The flow entering the water extractor from the condenser unit is set into circular motion by the swirl vanes. Condensed moisture is centrifuged by the swirl vanes and shaved off to the water drain port. This port is connected to the water spray nozzle on each ram-air inlet duct. The air pressure flowing through the water extractor forces the water to flow to and spray from the nozzle.

A metal strainer is installed at the water spray line connector on the water extractor. The strainer prevents water from being trapped in the line between the extractor and the water spray nozzle, reducing the possibility of freezing the water spray line, causing leaks.

Turbine Inlet Temperature Control Valve The turbine inlet temperature control valve is mounted in-line between the reheater outlet and the turbine inlet (Figure 21-30). Its purpose is to direct flow from the reheater to the turbine and ensure that ice crystals do not enter the turbine. The valve is operated by a thermal sensing element, which is both sensor and actuator. It contains a wax eutectic that contracts when chilled. The valve opens when reheater outlet air temperature drops below 75° F. This allows hot air from the air cycle machine compressor outlet duct to enter as needed to restore the temperature level to a minimum of 75° F. Should this valve fail to control the minimum temperature and allow cold airflow, it will be detected by the turbine inlet temperature sensor and annunciated via the airconditioning controller.

The water spray nozzle is a machined restrictor fitting mounted in the ram-air duct above the primary and secondary heat exchanger. The water that is removed by the water extractor is forced through tubing to the spray nozzle by air pressure in the water extractor. When the water reaches the restrictor orifice under pressure, it is atomized and sprayed into the ram air entering the heat exchanger. Evaporation of this water spray cools the ram air and increases the efficiency of the heat exchanger.


21-45



Figure 21-30. Turbine Inlet Temperature Sensor

Figure 21-31. Low-Limit Valve and Servo Pressure Regulator

21-46




Turbine Inlet Temperature Sensor The turbine inlet temperature sensor is mounted on the top side of the associated air cycle machine turbine housing (Figure 2130). The sensor detects turbine inlet air temperature that is used by the air-conditioning controller to monitor operation of the turbine inlet temperature control valve.

operate the low-limit valve. The servo pressure regulator also provides regulated pressure to operate the trim air valves.


The turbine inlet temperature sensor is a sealed, nonrepairable unit comprising two thermistortype elements connected in a series. The sensing elements are embedded in gold-filled epoxy, making contact with the shield for fast response. The unit has a negative temperature coefficient and nominal resistance of 5,648 ohms at 32°F (0°C) and 413 ohms at 160°F (71°C). The turbine inlet air temperature is normally limited to no lower than 75°F (25°C) by the turbine inlet temperature control valve. Should the input of this sensor indicate less than 50°F (10°C) for two minutes, the air-conditioning controller generates a blue “L or R ACS Maint Required” message on the EICAS and records an “L-R Turb Under Temp” fault parameter in the CMC.

NOTES

Low-Limit Valve The low-limit valve is a spring-loaded-closed pneumatically operated modulating valve, mounted on the air cycle machine turbine outlet duct (Figure 21-31). Both a torque motor and a differential pressure servo control this valve. The low-limit valve has two modes of operation. It provides hot air to maintain a minimum of 35°F pack outlet temperature via the torque motor. The valve also provides hot air to deice the condenser via the differential pressure servo.

Servo Pressure Regulator The servo pressure regulator is a spring-loadedopen, pneumatically operated, pressure regulating and relief valve. It is mounted on the air cycle machine condenser as part of the airconditioning pack assembly (Figure 21-31). The regulator taps turbine inlet air from the reheater outlet to provide a regulated supply pressure to


21-47



Figure 21-32. Turbine Bypass Valve

Figure 21-33. Air-Conditioning Pack Outlet Temperature Sensor

21-48




NOTES

Turbine Bypass Valve The turbine bypass valve is a spring-loadedclosed, solenoid-controlled, and pneumatically operated valve mounted on the condenser outlet (Figure 21-32). It allows a portion of the secondary heat exchanger outlet air to bypass the air-conditioning machine turbine. The valve is held open when the solenoid is energized to increase the airflow at altitudes above 35,000 feet.

Air-Conditioning Pack Outlet Temperature Sensor The air-conditioning pack outlet temperature sensor is located in the air ducting downstream of the air-conditioning pack outlet (Figure 21-33). Each sensor detects the temperature of the air exiting the associated air-conditioning pack. This temperature is sent as a discrete signal to the air-conditioning controller for control and indication functions. The air conditioning controller limits the minimum pack outlet temperature indications on the ECS/PRESS synoptic page. An amber “L or R Cool Turb Hot” message is displays on the EICAS if the pack outlet temperature exceeds 160°F. The sensor is a thermistor-type element with a nominal resistance of 15,000 ohms at 77°F (25°C). The sensor has a negative temperature coefficient; therefore, when sensed air temperature increases, the unit resistance decreases.


21-49



Figure 21-34. Ram-Air Check Valve

21-50




Ram-Air Check Valve The ram-air check valve is a 3-inch-diameter, split-flapper-type valve mounted in the left pack ram-air duct (Figure 21-34). During normal air-conditioning operation this valve remains closed, preventing loss of conditioned air to the ram-air duct. If the cold-air manifold becomes unpressurized (both packs off), ram air enters the left pack outlet duct through the ram-air check valve. It continues through the left supply air check valve to the coldair manifold and into the three zone delivery ducts. This supply of ram air can be used to ventilate the aircraft if required and provides ventilation for smoke evacuation in flight.

Controls and Indicators Controls and indicators consist of the left and right pack and ram-air switches on the cockpit overhead panel (Figure 21-35). Since the normal position of the left and right pack switches is ON, they illuminate their amber OFF legends only when selected OFF or when the pack inlet valve is closed either by the ram switch or during the engine start sequence on the ground. The function of the left and right pack switches is to provide ON and OFF control of the pack inlet valves. They also provide the ON and OFF command to the air-conditioning controllers and to the on-side DAU. The ram-air switch is normally selected off and illuminates amber RAM only when selected on. This also illuminates the left and right pack switches OFF. When ram air is selected, the left and right pack inlet valves are turned off, leaving ram air as the only means of cabin and cockpit ventilation.

Figure 21-35. Airflow Control Switches


21-51

21-52 R AIR COND CONTROLLER (ACC)

R PACK CONT R ESS 28 VDC BUS

R ESS 28 VDC BUS

R PACK LOW LIMIT VALVE TO CKPT TEMP CONT


TO EERR COOLING FANS & VALVE (4 POLES)

<35K FT

28 VDC FROM R PACK SW (R PACK CONT RELAY DE-ENERGIZED) FROM L PACK 25K RELAY

TORQUE MOTOR

E

G

PACK OUTLET TEMP SENSOR X1 X2

R PACK TURBINE BYPASS VALVE CONT RLY

28 VDC BELOW 1500 FT TRIP

R TURBINE BYPASS VALVE SOLENOID

LOW LIMIT VALVE T/M SIGNAL

–

R TURBINE INLET TEMP SENSOR TURBINE INLET TEMP SENSOR

R PACK OUTLET TEMP SENSOR

>35K FT

+28 VDC AUTO RETURN +

X1 X2 A/I INHIBIT

+ –

(ENERGIZE L/R COWL TO CLOSE) AUTO A/I 28 VDC L/R WING FROM F AUTO A/I R PACK (4 POLES) CONT ICE DETECT RELAY AUTO ALT INHIBIT RELAY

R COMPRESSOR OUTLET TEMP SENSOR R PACK INLET VALVE + –

SOLENOID (ENGINE TO CLOSE) + TORQUE MOTOR –

COMPRESSOR OUTLET TEMP SENSOR

+ –

(TAIL COMPARTMENT)

LOW LIMIT VALVE RETURN ID PIN 2 ID PIN 3 TURBINE INLET TEMP SENSOR SIGNAL TURBINE INLET TEMP SENSOR RETURN PACK OUTLET TEMP SENSOR SIGNAL PACK OUTLET TEMP SENSOR RETURN COMPRESSOR OUTLET TEMP SENSOR SIGNAL COMPRESSOR OUTLET TEMP SENSOR RETURN PACK INLET VALVE T/M SIGNAL PACK INLET VALVE T/M RETURN

+

AIR

–

ARINC 429

WEIGHT ON WHEELS

(TAIL COMPARTMENT)

MAU 2 DUAL GENERIC I/O MODULE (2) ARINC 429 TX

GND

CHASSIS GND

(P/O C WOW RLY) ARINC 429

ARINC 429 RX

CKPT TEMP DATA

R ACS FAIL (A)

CONTROLLER FAILED

international

FlightSafety

CKPT/CABIN TEMP IND

BLEED AIR SW STATUS, PACK SW STATUS, WING A/I SW STATUS, ALT, OAT, N1 ARINC 429 TX (REF 3-16)

ARINC 429


D CKPT AUTO TEMP

+28 VDC AUTO TURBINE BYPASS VALVE CIRCUIT MONITOR PACK ON/OFF COMMAND TURBINE BYPASS VALVE OPEN COMMAND

28 VDC TO R PACK SW & R PACK CONT RELAY

Figure 21-36. Airflow Control Schematic



OPERATION AND INDICATIONS The air-conditioning system uses bleed air to supply the left and right air-conditioning packs that provide a nominal 35°F airflow to a common cold-air manifold. Hot air is supplied to the hot-air manifold by both air-conditioning packs. The 28-VDC essential bus powers the air-conditioning controller via the L-R PACK CONTROL circuit breaker and the CABIN or COCKPIT AUTO TEMP circuit breakers. The air-conditioning controller monitors the pack outlet temperature sensor, the compressor outlet temperature sensor, and the turbine inlet temperature sensor. It also controls the pack inlet valve torque motor, the low-limit valve torque motor, and the turbine bypass valve control relay (Figure 21-36). The air-conditioning controller provides a ground to energize the turbine bypass valve control relay, causing the air cycle machine turbine bypass valve to open above 35,000 feet. The turbine bypass valve control relay also has input to the electrical equipment cooling system and wing/cowl automatic anti-ice control. During engine start on the ground, both packs shut off to provide the greatest volume of air available to start the engine. Selecting the master start or master crank switch energ i z e s t h e r i g h t p a c k i n l e t va l ve s h u t o ff solenoid. The right pack remains off as long as the switch is selected ON. When the left or right engine start switch is selected ON, the starter air valve relay is energized to shut the left pack off during the start period. When engine start is accomplished, left pack operation is reinstated.


NOTE The air-conditioning packs will not shut off during in-flight starts.

The air-conditioning packs are powered when the left and right pack switches on the TEMP CONTROL panel are selected ON. During maximum cooling operation, most of the bleed air from the pack inlet valves flow into the airconditioning packs at their primary heat exchanger inlet. The bleed-air temperature is reduced by the primary heat exchanger to a level sufficiently low for the compressor of the air cycle machine. This transfers the heat into the heat sink cooling air. The fan bypass check valve in the fan/plenum allows additional ram air to bypass the fan during flight operations. This increases the cooling flow through the heat exchangers to achieve higher efficiency. The fan and bypass air is exhausted overboard after passing through the primary and secondary heat exchanger. During ground operations, the only cooling air heat sink for the primary and secondary heat exchanger is provided by fan-induced airflow. This is accomplished through the ram-air scoop and the air cycle machine fan, which is mounted to the same shaft as the rotating turbine and compressor assembly. The bleed air from the primary heat exchanger first enters the compressor stage of the air cycle machine. This compression increases the pressure and temperature of the air. Some air passes through the compressor bypass check valve. This allows the air cycle machine to reach operating speed quickly. The secondary heat exchanger transfers the heat of compression to the heat sink provided by the ram air, thereby reducing secondary heat exchanger bleed-air outlet temperature. The air cycle machine compressor outlet temperature sensor protects the air cycle machine in the event of a ram-air passage blockage or an air cycle machine overspeed. When a temperature of 425° F is reached, the air-conditioning controller sends up to 100 mA to the pack inlet valve torque motor, and the pack inlet valve flow can be reduced to a minimum 25 ppm.


21-53




76∞ 78∞

76∞ 78∞

74 ∞F

78 ∞F

80∞ 80∞ 78 ∞F

74 ∞F

35 ∞F

35 ∞F

78 ∞F

0 psi

0 psi 370 ∞F

370 ∞F

42 psi

42 psi

Left Eng

Right Eng 78 ∞F

78 ∞F APU

L Elev 1030

Cab Alt -450

Rate 0

P 0.25

Mode Auto 1

Figure 21-38. ECS/PRESS Synoptic Page—Airflow Control Indications

21-54




Bleed air exiting the secondary heat exchanger is ducted to the reheater inlet, where it is initially cooled. The bleed air next passes through the condenser, where the temperature is further reduced by the subfreezing turbine outlet air. This causes the water vapor present in the high-pressure bleed air to condense substantially. The high-energy water extractor removes up to 92% of the condensed water coming from the condenser outlet air. The extracted water is blown through tubing to the ram-air duct and is sprayed above the secondary heat exchanger through the water spray nozzle. Hot air from the air cycle machine compressor outlet duct is bypassed through hollow cores on the face of the condenser. This hot air prevents the cold side of the condenser from icing up. From the water extractor, the airflow returns to the reheater. The reheater increases air temperature exiting the water extractor prior to entering the air cycle machine turbine inlet. This decreases the possibility of any turbine icing and increases the energy to the turbine. The dehumidified bleed air entering the air cycle machine turbine generates power to drive the compressor and fan. As a byproduct, this air rapidly expands, resulting in subfreezing temperatures at the turbine outlet. A thermally controlled turbine inlet control valve keeps the turbine inlet temperature at a minimum 75°F, protecting the air cycle machine turbine from failures due to icing. Hot air is ducted to the turbine inlet temperature control valve from the air cycle machine compressor outlet duct. The turbine inlet temperature sensor monitors the turbine inlet temperature for the airconditioning controller. Temperatures below 50°F for more than two minutes cause the airconditioning controller to generate a blue “LR ACS Maint Required” CAS message and

the CMC to display an “L-R Turb Under Temp” fault parameter. This message indicates a possible fault of the turbine inlet temperature control valve. The low-limit valve is commanded open by the air-conditioning controller whenever the air-conditioning pack outlet temperature drops below 35°F as monitored by the pack outlet temperature sensor. The hot air consequently mixes with air cycle machine turbine outlet air. The low-limit valve incorporates a differential pressure servo that receives a condenser inlet (cold side) and outlet pressure. A 3.5-psi differential causes the low-limit valve to open to melt ice accumulation on the condenser inlet. The air cycle machine turbine bypass valve is commanded open when the airplane altitude is above 35,000 feet. The air-conditioning controller provides a ground for the turbine bypass control valve relay, and the turbine bypass valve solenoid energizes. This bypasses a portion of the secondary heat exchanger outlet air around the air-conditioning machine turbine, resulting in increased volume flow to the air-conditioning distribution system. The turbine bypass control valve relay deenergizes below 33,000 feet. The indications for the air-conditioning airflow control system are found on the EICAS display, accessed through the cockpit display controller. The air-conditioning controller sends the information provided by the air-conditioning pack outlet temperature sensor to the EICAS for display. The air-conditioning packs are represented as fan symbols in the center of the ECS/PRESS synoptic page (Figure 21-37). An amber “CAS L or R Cool Turb Hot” message is displayed when the air-conditioning controller input from the compressor outlet temperature sensor exceeds 450°F. The same message is displayed if the pack outlet temperature sensor input exceeds 160°F.


21-55



FORWARD CABIN

COCKPIT

AFT CABIN

Figure 21-38. Trim-Air Valves

21-56




DISTRIBUTION

Ram-Air Ducting and Ram-Air Check Valve

General Air-conditioning distribution provides the necessary paths for conditioned air to maintain temperatures and cabin pressure for maximum comfort. Air-conditioning distribution components include the following: • Cold-air manifold • Ram-air ducting

Ram-air ducting connects to the left airconditioning pack outlet ducting. The ram-air duct inlet and its check valve are connected just above the left air-conditioning pack primary and secondary heat exchanger, and the outlet is joined to the air-conditioning pack outlet ducting. The check valve allows ram air to be fed directly into the left air-conditioning pack duct, supplying the cold-air manifold when ramair pressure exceeds cold-air manifold pressure.

• Supply-air check valves • Hot-air manifold

Supply-Air Check Valves

• Trim-air valves

• Baggage compartment ventilation valve

The supply-air check valves are mounted in the ducting beneath the baggage compartment floor, between the cold-air supply duct from the airconditioning packs and the cold-air manifold. These check valves allow ram air into the coldair manifold and prevent cold air from escaping during single air-conditioning pack operation.

• Personal service units

Hot-Air Manifold

• Cockpit airflow valve

The hot-air manifold is located beneath the baggage compartment floor. Its purpose is to provide distribution for hot air tapped from the air-conditioning pack inlet ducting aft of the pack inlet valve and ozone converter.

• Hot trim-air check valves • Cabin silencers • Supply duct

These components supply conditioned air to the cockpit, baggage compartment, and two cabin zones.

Trim-Air Valves


Components Cold-Air Manifold The cold-air manifold is located under the floor of the baggage compartment. Its purpose is to collect conditioned bleed air from the airconditioning packs or from ambient ram air. It becomes an expansion chamber and a distribution point for cold-air supply or ventilation as needed throughout the aircraft.

The trim-air valves are spring-loaded-closed, pneumatically operated, torque-motor-controlled valves. They are mounted below the baggage compartment floor, on the hot-air manifold (Figure 21-38). There is one valve for each of the temperature control zones: cockpit, forward cabin, and aft cabin. Each valve provides temperature control by mixing hot air with the cold air from the cold-air manifold. The zone trim air control valves provide the crew with the ability to adjust temperatures by adding heat to the cold air flowing into the cabin and cockpit zones.


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Figure 21-39. Baggage Compartment Ventilation Valve Reset Switch

Figure 21-40. Cockpit Airflow Valve

21-58




Hot Trim-Air Check Valves The hot trim-air check valves are mounted in the ducting beneath the baggage compartment floor, between the pack hot-air supply and the hot-air manifold. These check valves allow hot bleed air from the air-conditioning pack inlet ducting into the hot-air manifold and prevent loss of hot air during operation of a single air-conditioning pack.

Cabin/Cockpit Silencers The cabin silencers are located beneath the floor in the aft cabin (zone 2). They provide each zone supply duct with a muffler to keep airconditioning noises out of the cabin areas. A separate silencer for the cockpit zone is located in the cockpit air supply duct, underneath the floor in the forward cabin, to minimize airconditioning noises in the cockpit.

Supply Duct Foot-level ducting is provided along the sides of the aircraft to distribute the conditioned air to the two cabin zones. A separate air-supply duct furnishes conditioned air to the aft baggage compartment. The baggage compartment supply duct is tapped off the aft cabin distribution duct under the floor on the right side.

Baggage Compartment Ventilation Valve and Return Air Check Valve

valve to open the valve and re-establish airflow to the baggage compartment. The ventilation valve cabin pressure sense port and the smoke evacuation valve are also located on this panel. The secondary pressure bulkhead separating the baggage compartment from the cabin has a sliding door that is required to be closed and locked above 40,000 feet. The return air check valve provides ventilation airflow back to the cabin under normal conditions. When the baggage compartment is depressurized, the check valve closes to prevent loss of cabin pressure.

Personal Service Units (Gaspers) The personal service units (PSUs), known as “gaspers,” are installed in both zones on each side of the cabin to supply cold air directly from the cold-air manifold. This allows selection of cold air for individual comfort.

Cockpit Airflow Valve The cockpit airflow valve (Figure 21-40) is a DC motor-driven butterfly valve with an orifice section. It is located in the cockpit conditioned-air supply duct aft of the cockpit silencer. The valve restricts cockpit airflow by approximately 50% when closed.

The baggage compartment ventilation valve is installed in the baggage compartment supply air duct, beneath the right side of the aft cabin floor. This valve is a differential pressure shutoff valve and incorporates a cabin pressure port and baggage compartment pressure port. The valve shuts off the flow of air to the baggage compartment if pressure is lost due to a rupture or smoke evacuation valve use. A solenoid on the valve can be energized to reopen the valve and repressurize the baggage compartment. An aft baggage compartment ventilation valve guarded reset switch is installed on a panel (Figure 21-39) above the secondary pressure bulkhead door. Actuating this momentary switch energizes the solenoid on the ventilation


21-59



CABIN PRESSURE RELIEF VALVE

THRUST RECOVERY OUTFLOW VALVE

EVS COOLING VALVE

RIGHT EER DU NO. 4

RIGHT EER FAN

EMERGENCY INVERTER TRUs

DU NO. 3 DU NO. 2

PSU FAN

DU NO. 1

LEFT EER FAN

L CABIN RETURN AIR

LEFT EER

Figure 21-41. Forward Electronic Equipment Cooling

21-60




NOTES

EER Fans The electronic equipment rack cooling is accomplished by utilizing fans that distribute ambient cabin air throughout the racks and across the forward underfloor electrical and avionics equipment. The left and right electronic equipment rack DC fans are positioned at or near the bottom of each rack. These fans change from high speed to low speed above 35,000 feet. When operating on a single air-conditioning pack, the fan speed remains at high. The left PSU DC fan is located under the cabin floor on the left side, just aft of the entrance door. It pulls air down from the forward cabin and blows it forward over the entrance compartment underfloor electronic equipment. This fan changes from low speed to high speed above 35,000 feet.

NOTE All electronic equipment cooling fans incorporate failure detection devices.

The aft electronic equipment rack, located in the baggage compartment, also incorporates a constant-speed DC fan. The aft cabin air is ducted to the electronic equipment rack, and the fan pulls the air through and exhausts it to the baggage compartment. The baggage compartment ventilation airflow is provided back to the cabin through a check valve on the secondary pressure bulkhead.


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OPERATION AND INDICATIONS The cold-air manifold normally receives 35°F air under pressure from the air-conditioning packs, which are powered by the bleed-air manifold. The hot-air manifold receives hot bleed air from the bleed-air manifold after it has passed through the pack inlet valve and the ozone converter. It distributes this air to the trim-air valves. The trim-air valves, via cockpit control or air-conditioning controller control, allow hot air to mix with cold air in the cabin and the cockpit zone supply ducts to control temperatures throughout the aircraft. The hot and cold mixed air is distributed through the forward and aft cabin zones, through the cabin silencers to the foot level ducting and the baggage compartment. It is also distributed through ductwork and the cockpit silencer to the cockpit zone. Additionally, unmixed air directly from the cold-air manifold is distributed to the PSUs for both cabin zones. For aircraft ventilation and smoke evacuation at low altitude, the ram-air switch can be selected to shut off both air-conditioning packs. The ram-air pressure will then exceed cold-air manifold pressure and provide ram air through the ram-air check valve to the air-conditioning system for distribution. Additional cooling fans are located in areas where electronic equipment is installed to ensure steady cooling airflow and prevent overheating. The left electronic equipment rack cooling fan is mounted at the base of the equipment rack to pull air down through the electronics and disperse it to the underfloor equipment area. The right electronic equipment rack cooling fan is mounted to the bott o m o f t h e l ow e r s h e l f o f t h e e l e c t r o n i c equipment rack. The right and left electronic equipment rack fans change to low speed above 35,000 feet when the air-conditioning controller energizes the turbine bypass valve control relay. The fans operate at high speed below 35,000 feet and at low speed above 35,000 feet when both air-conditioning packs are operating.

21-62


NOTES



The left PSU fan changes from low speed to high speed above 35,000 feet and pulls air from the forward cabin under the floor. It blows this air forward to cool the electronic equipment in the entrance compartment underfloor area.

NOTES

The status for the air-conditioning distribution system equipment cooling fans is indicated by crew alert messages. When the left PSU, left or right electronic equipment rack, aft equipment fan fails, a blue message illuminates.



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FAN

TEMP SENSOR

LOW SPEED WARNING DEVICE

Figure 21-42. Forward Cabin Zone Temperature Sensor

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COCKPIT AND CABIN TEMPERATURE CONTROL General Air-conditioning temperature control provides the means to control and display the cockpit and cabin area temperatures. Temperature control components include the following: • Air-conditioning controllers • Zone temperature sensors • Supply duct temperature sensors • Overheat thermostats • Display switches • Temperature indicators • AU T O - M A N t e m p e r a t u r e c o n t r o l switches • Auto/manual dual potentiometer

Components Air-Conditioning Controllers Two air-conditioning controllers (ACCs) are located in the baggage compartment electronic equipment rack. Each air-conditioning controller accommodates both automatic and manual modes of temperature control.

NOTE In the manual mode, control of the trim-air valves is based on crew input to the ACC via direct control from a dual potentiometer.

Zone Temperature Sensors There are two cabin zone temperature sensors. one sensor is located in each of the two cabin zones. The forward cabin zone temperature sensor is located between the No. 1 and No. 2 windows, just above the fuselage centerline on the right side of the fuselage (Figure 21-42). The aft cabin zone temperature sensor is located aft of window 7, just above the foot level duct on the left side of the fuselage. These are production installation locations and sensors may be relocated during the completion phase per customer outfitting requirements. Each sensor assembly incorporates a fan to draw ambient cabin air across the sensor for accurate temperature readings for its zone. The assembly also incorporates a low-speed warning device (LSWD) to alert the crew should the fan assembly fail. The temperature sensors provide temperature information to the left airconditioning controller for automatic temperature control and display.

The left air-conditioning controller controls the temperature in the two cabin zones, and the right one controls the cockpit zone. The air-conditioning controllers receive crew input for desired temperature in each zone. They also receive actual temperature data from the zone and duct temperature sensors. They provide control of their associated trim air valve torque motors based on a selected and actual zone and duct temperature. Selected temperature range is from 60 to 90°F in the AUTO mode. The air-conditioning controller also limits supply duct temperature to 160°F for the two cabin zones (180°F for the cockpit) and transmits temperature data to the cockpit temperature indicator and CAS.


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LOCATED AT LOWER LEFT SIDE OF CENTER PEDESTAL

LOCATED ABOVE 328 PANEL

Figure 21-43. Cockpit Zone Temperature Sensors

21-66




NOTES

Cockpit Zone Temperature Sensors The cockpit zone temperature sensors are located behind the pilots seat on the forward side of the LEER, and on the left side of the center pedestal. The sensors provide temperature information to the right air-conditioning controller.


21-67



FORWARD CABIN DUCT TEMPERATURE SENSOR

AFT CABIN DUCT TEMPERATURE SENSOR

Figure 21-44. Supply Duct Temperature Sensor

Figure 21-45. Supply Duct Overheat Thermostat

21-68




Supply Duct Temperature Sensors

NOTES

There are three supply duct temperature sensors: one is in the cockpit air supply duct, and the other two are in each of the cabin zone supply ducts. All three sensors are located under the aft cabin floor, just forward of the secondary pressure bulkhead (Figure 21-44). The supply duct temperature sensors provide duct temperature information to their respective airconditioning controllers for automatic temperature control and display.

Supply Duct Overheat Thermostats Three supply duct overheat thermostats are mounted to their respective air supply duct, just below the cabin floor and downstream of the trim-air valves (Figure 21-45). One is located in the cockpit air supply duct, and the other two are in each of the cabin zone supply ducts. All three sensors are located under the aft cabin floor, just forward of the secondary pressure bulkhead. The supply duct overheat thermostats provide backup control against overheating by causing the trim-air valves to close when excessive temperatures are detected. Thermally controlled pneumatic valves open at 215 ±15°F and vent trim-air valve torque motor pressure to ambient. The spring-loaded trim-air valves then close.



21-69



Figure 21-4 6. Air-Conditioning Control Panel

21-70




CONTROLS AND INDICATORS

AUTO–MAN Switches

The temperature controls and indications are contained on the ECS panel, which is located on the cockpit overhead panel (Figure 21-46). These consist of the AUTO TEMP SELECT, TEMP DISPLAY, COCKPIT AIRFLOW, and AUTO–MAN switches, the TEMP DISPLAY °F indicators, and the cockpit/cabin dualpotentiometer control knobs.

The AUTO–MAN switches and dual-potentiometer control knobs allow the crew to manually select automatic temperature control or manually control the temperature for each z o n e . W h e n i n t h e AU T O p o s i t i o n , t h e AUTO–MAN switch is illuminated blue, and the air-conditioning system automatically regulates the temperature for each zone between 60 and 90° F, depending on the setting of the control knobs. The power for cabin automatic temperature is from the left essential 28-VDC bus. The auto/manual control relays are energized in the AUTO mode, and the dual-potentiometer signals to the air-conditioning controllers are active for automatic temperature control.

AUTO TEMP SELECT Switch The AUTO TEMP SELECT switch, when selected ON, causes the temperature windows to indicate the selected temperature for their respective zones The selection of this switch extinguishes the TEMP DISPLAY legends. The automatic temperature select range is from 60° F in full cold to 90° F in full hot. In a manual temperature control mode, selection of the AUTO TEMP SELECT switch to ON causes the dashes to be displayed in that zone. Deselecting the AUTO TEMP SELECT switch extinguishes the blue ON legend and causes the TEMP DISPLAY switch to illuminate blue ZONE or green DUCT, depending on the last selection. The actual zone or duct temperature is displayed on the temperature indicators. Selecting manual temperature control will not affect this display.

COCKPIT AIRFLOW Switch

NOTE Provisions are made for the outfitter to install VIP remote potentiometers, which provide remote ±6°F input to the air-conditioning controllers.

The selection of MAN allows direct control of each zone’s trim-air valve through its respective air-conditioning controllers, which convert main DC bus 28-VDC power to 0 to 100 mA for direct torque motor control.

The COCKPIT AIRFLOW switch allows the crew to control the amount of airflow into the cockpit area by controlling the cockpit airflow valve. The selection of LOW decreases airflow into the cockpit by approximately 50%. The temperature display indicators provide a digital readout for each of the cockpit and cabin zones and display selected auto temperature, actual zone, or duct temperature. The respective air-conditioning controllers transmit this data over the ARINC 429 data bus.


21-71

21-72 OVERHEAD OVERHEAT THERMOSTAT

FOOT LEVEL OUTLET

FWD CABIN COCKPIT ZONE TEMP SENSOR

COCKPIT DUCT TEMP SENSOR

ZONE TEMP SENSOR

FWD CABIN SUPPLY TEMP SENSOR FOOT LEVEL OUTLET

AFT CABIN ZONE TEMP SENSOR GASPER AIR

HOT AIR MANIFOLD


FOOT LEVEL OUTLET

AFT CABIN SUPPLY TEMP SENSOR TRIM AIR VALVES

MAU

COCKPIT INPUT AFT CABIN FWD CABIN

RIGHT AIR CONDITIONING CONTROLLER

CKPT IND

CMC

LEFT AIR CONDITIONING CONTROLLER MAU

COCKPIT TEMPERATURE INDICATOR

international

FlightSafety


COLD AIR MANIFOLD

FOOT LEVEL OUTLET

COCKPIT AIRFLOW VALVE

Figure 21-47. Air Temperature Control Schematic



Operation and Indications In the auto temperature control mode, the airconditioning controller monitors the temperature in each zone and its supply ducts (Figure 21-47). The air-conditioning controller compares this data with the crew settings on the cockpit temperature control panel to control the temperature throughout the aircraft. The TEMP DISPLAY switch, when indicating ZONE, causes the TEMP DISPLAY °F indicators to display the actual temperature for their respective cabin and cockpit areas, using inputs from the zone temperature sensors. When DUCT is selected, the TEMP DISPLAY °F indicators display the actual temperature of the respective cabin or cockpit supply air duct, taken from the duct temperature sensors.

The ECS/PRESS synoptic page indicates actual temperatures of each cabin zone and the cockpit using information from the zone temperature sensors. The actual supply duct temperature is also displayed on the ECS/PRESS synoptic page. When in the automatic temperature control mode, the ECS/PRESS synoptic page displays the selected temperature for each zone in blue digits. In the manual temperature control mode, the selected display is amber dashes. The crew alert messages pertaining to the air-conditioning system are also displayed on DU4, to the left of the ECS synoptic page.

NOTES

When any of the three zone AUTO–MAN switches is in the AUTO position, the air-conditioning controller controls the mixing of hot and cold air via the respective trim-air valve. When in the amber MAN position, the COLD–HOT control knobs allow the crew to manually control the temperature for their respective zones between full cold of 35°F and full hot maximum of 230°F.


The AUTO TEMP SELECT switch controls the TEMP DISPLAY °F indicators. When the AUTO TEMP SELECT switch is selected ON, the TEMP DISPLAY °F indicators indicate the selected temperature for their respective zones as determined by the position of the C O L D – H OT d u a l p o t e n t i o m e t e r c o n t r o l knobs. The selection of this switch extinguishes the ZONE–DUCT switch. The COCKPIT AIRFLOW switch allows the crew to control the amount of airflow into the cockpit area. The selection of LOW causes the cockpit airflow valve to restrict airflow into the cockpit area by approximately 50%.


21-73



CPCS 1

0–5 VDC

CPCS 2 ARINC 429

POWER DATA

ADM NO. 1 ADM NO. 2 ADM NO. 3

SELECTOR PANEL

TO DISPLAY SYSTEM

CPAM

CPI

ARINC 429 DATA IN

DISCRETES TO DAUs

FMSNO. 1 FMSNO. 2

0–5 VDC

WEIGHT ON WHEELS (WOW) DOORS CLOSED

0–28 VDC

THROTTLE LEVERS

+28 VDC CAB IN P CXR RESS GNL URE J CON TRO L

DISCRETES

FTLN GB FTLN GB FTLN GB

FTLN GB

MANUAL CONTROL

FTLN GB FTLN GB

115 VAC NO. 1 115 VAC NO. 2

CABIN PRESSURE CONTROL PANEL

+28 VDC FOR PRV

SERVO CONTROL NO. 1 SERVO CONTROL NO. 2

DUAL CHANNEL CONTROLLER

CABIN PRESSURE SENSE

HIGH PRESSURE SOURCE (20 PSIG)

PRESSURE RELIEF VALVE

STATIC ATMOSPHERE SENSE

OUTFLOW VALVE

Figure 21-48. Cabin Pressure Control System Block Diagram

21-74




CABIN PRESSURIZATION CONTROL SYSTEM (CPCS) GENERAL The Gulfstream V cabin pressurization control system (Figure 21-48) controls cabin pressurization by regulating the amount of air leaving the cabin through the outflow valve. The system is designed to provide cabin altitude equal to 6,000 feet when the aircraft is at 51,000 feet. The system also monitors the cabin pressure to ensure maximum passenger comfort and safety. Some of the essential elements of the system include the following:

ferential pressure based on a predetermined schedule, controls the rate limiting for climb and descent, and controls depressurization on landing. In the semiautomatic mode, the dual-channel controller provides control based on crew inputs. Pressurization control is essentially the same as for automatic mode except that FMS inputs are not used. In the manual mode, the dual-channel controller is inoperative and has no function. The crew must manually control the outflow valve.

NOTES

• Cabin pressurization limit control • Cabin rate limit control • Cabin pressure indication • High-altitude warning The cabin altitude and ambient altitude are sensed through electronic devices to establish control over the outflow valve. A dual-channel controller provides dual automatic control channels to perform cabin pressure limit control, cabin rate limit control, pressure indication, and high cabin altitude warning. The Gulfstream G500/G550 cabin pressure limit control consists of a dual redundant system that has three modes of control: automatic, semiautomatic, and manual. In the automatic mode, the dual-channel controller provides hands-off pressurization control. The system receives digital data inputs from the flight management system (FMS), air data computers (ADM), and cabin pressurization acquisition module (CPAM). It also receives discrete inputs from the weight-on-wheels (WOW), doors-closed signals, and throttle quadrants. The cabin pressure control panel on the cockpit overhead panel provides crew inputs to the system. This mode automatically prepressurizes the aircraft before takeoff, controls the cabin dif-


21-75



1

2

3

A

4

5

6

7

5

5

1

L LWR PROBE HTR

L LFR PROBE HTR

W RADAR CONT #1

C

ANTI-ICE

10

9

1

8

MADC #2

7

6

5

4

3

1

1

5

5

71/ 2

CLOCK #2

W RADAR CONT #2

R UPR PITOT HTR

R LWR PITOT HTR

#2 AOA HTR

1

MADC #3

10

10

10

10

10

10

DISPLAY UNIT #3 SEC

DISPLAY UNIT #5

DISPLAY UNIT #6

DISPLAY UNIT #4 SEC

DISPLAY UNIT #4 PRI

DISPLAY UNIT #2

2

A

B

20 R FRONT WSHLD

C

5

2

5

5

5

1

5

5

5

5

5

1

5

5

5

2

5

2

COMBINED WOW

WHE EL SPE ED

LEFT AIL HYD S/ O

LEFT ELEV/ HYD S/ O

FLAP/ STAB LEFT DC

LEFT SD D

FWC #1

SYM GEN #1

SYM GEN #3

SYM GEN #2

FWC #2

RIGHT SD D

FLAP/ STAB RIGHT DC

RIGHT ELEV/ HYD S/ O

RIGHT AIL HYD S/ O

RUD DER HYD S/ O

TOA

RIGHT WOW

5

4

71/ 2

1

5

3

71/ 2

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5

IAC #1

STBY ALT/ AIRSPD

CDU #2

YAW DAMP #2

71/ 2

FLAP/ STAB L STBY AC

A/ P SERVO #1

YAW DAMP #1

HUD OVHD

STBY HORIZON

2

2

2

2

5

10

2

2

2

2

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5

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SHAKER #1

DAU #1A

DAU #1B

STBY RMI

HUD COMP

W RADAR R/ T

DAU #2B

DAU #2A

SHAKER #2

STAL L BAR R VALVE #2

A/ T SERVO

SPLR FLT PWR S/ O

GND SPLR

5

2

TCAS

NAV RCVR #1

2 GPS #1

ATC #1

FLT INSTRUMENTS

2

71/ 2

ADF #1

VHF COMM #1

5

1

71/ 2 L/ R IGN 1

2

3

DME #1

L ES S DC CONT #1

SYSTEM TEST

FLT INSTRUMENTS

5

5

5

5

5

L/ R START A

L TR CONT

L TR MAN STOW

R TR MAN STOW

R TR CONT

ENG SYS/ CNTRLS

2 RFMU #1

ICE DET

SMOKE DET

AN N LIGHTS

DO OR CLOSE

DO OR SAFETY

LDG GR DUMP V

FD R/ MDAU EVENT

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TEST

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ON

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ENG VIB MON TEST PRI/ SEC

EQPT OVHT

TEST

TEST

SEC

R ENG LO OP A LO OP B

LO OP B

FAULT TEST

FAULT OF F

TEST

FAULT

FAULT

OF F

OF F

OF F

ARM

ALT OF F

ON

ALT

97.5 99.8 — —

R ENG

ON

OF F

ISOLATION

20 SECONDS 20 SECONDS

R/ H LP .22 .14 .12 .10 .04 .03 .02

TEMP CONTROL L PACK

RAM AIR

OF F

RAM

OF F

AUTO TEMP SELECT

TEMP DISPLAY

COCKPIT AIRFLOW

ZONE

NORM

DUCT

LOW

FIRE EXT DISCHD

RPM %

ON

FIRE

OF F

ISLN

888 STOP

ON

OF F

MAN AFT CABIN

COWL OPEN COLD

REMOTE FUELING

OF F/ AUTO

L SHUTOF F

R SHUTOF F

OF F

CLSD

CLSD

HOT

COLD

2 3 4

1 UP

INTER TANK

R TANK

FPM X 1000

R PUMPS MAIN ALT

4 2 3

COLD

HOT

OF F

OF F

OF F

L ENG PWR XFR UNIT OF F/ ARM ON

L ENG

R ENG

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ON

ON

NOT ARM

D E S C E N D

ON

OF F

AUTO

FLIGHT

AUTO

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EXTERIOR LTS

AUTO

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ICE

ON

ON

ON

ON

R LDG

PULSE

TAXI

WINGTIP

WHL WEL L

NO SMOKE

ON

ON

ON

ON

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ON

ON

ON

AUTO SEMI

SEMI

MAN HOLD

CAUTION DFRN PRES SURE SHOULD NOT EXCE ED 0.3 PSI DURING TAXI TAKEOF F OR LANDING

OVHD PNL

CB PANELS MASTER CONTROL

WINDSHIELD HEAT

OF F

ON

ON OF F PNL

CABIN WDO-HT

WSHLD BLWR

COCKPIT LIGHTS

SEAT BELT

CABIN PRES SURE CONTROL

AUTO

MANUAL

MANUAL

ON

ON

PAS S WARN

NAV

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VALVE

FAULT

LANDING

LANDING

R WING ON

AUX PUMP OF F/ ARM ON

NOT ARM

E

FAULT

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C L I M B

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L COWL

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88888 8888

6

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ANTI ICE L WING

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MAN HOLD

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88888 8888

6

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4 2 3

EN

RIGHT

0088.8 088

5

6

DN 1

5

6

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OP

LEFT

HOT

CABIN ALT FT

RC

FUEL SYSTEM L TANK

L PUMPS ALT MAIN

MAIN BAT TERIES

MASTER

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AUTO

MAN

OF F

VOLTS/ AMPS

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FWD CABIN

E-INV

AUTO

ON OF F

ON

888

FWD CAB

MAN COCKPIT

R GEN

ON

888

COCKPIT

AUTO

START

READY

FUEL RETURN

AUTO

ON AVAIL EXT PWR

ON GND SVC BUS

STROBE

R PACK

GAL LEY

OF F

R BUS TIE

ISLN

ON

APU CONTROL MASTER

AUTO

HP .20 .21 .21 .22 .21 .41 .21

TEMP DISPLAY °F

APU

FIRE EXT

2 3 4

1 UP

20 SECONDS 20 SECONDS

— — 900°C 905°C

— —

HP .26 .26 .10 .22 .17 .23 .13

8888 888

L AC

MASTERS

ELECTRICAL POWER CONTROL L BUS TIE

L/ H LP .08 .08 .04 .03 .03 .01 .01

EGT C

R MAIN

R AC

CABIN

0088.8 008

101.3 101.5

S/ N %LP MAX 85 65 55 45 35 IDLE

TRU

R ES S

ON

APU GEN

APU

OF F

MOMENTARY MOMENTARY

COPILOT

L MAIN

LIGHTS AV PWR

L ES S

ON

UNRESTRICTED UNRESTRICTED

* 10 MINUTES IN THE EVENT OF AN ENGINE FAILURE * 10 MINUTES IN THE EVENT OF AN ENGINE FAILURE

EICAS

OF F

LIGHTS AV PWR

ON

ON

MOMENTARY MOMENTARY 5 MINUTES* 5 MINUTES*

850°C 860°C — —

UNRESTRICTED UNRESTRICTED MIN -40°C MINUNRESTRICTED -4090 °C°C MAX FUEL TEMP (ENG) MAX 140 °C°C FUEL TEMP (ENG) MAX UNRESTRICTED (15 MIN) TRANSIENT 120 MAX (15 MIN) TRANSIENT 165°C OIL TEMPERATURE MINIMUM ENGINE OIL PRES SURE OIL TEMPERATURE MINIMUM ENGINE OIL PRES SURE FOR COMPLETE MIN. FOR START -40°C FOR COMPLETE MIN. FOR START -40°C T/ O FLIGHT MIN. BEFORE T/ O FLIGHT MIN. BEFORE 30 PSI INCREASING +10°C BELOW 72 3% N2 40 PSI 72 0% 3% N2 N2 70 35 25 PSI INCREASING +10°C BELOW ABOVE 95 PSI 50 PSI PSI POWER ABOVE 95 0% N2 35 PSI 45 PSI POWER MAXIMUM +160°C STRAIGHT LINE VARIATION IN BETWE EN MAXIMUM +160°C STRAIGHT LINE VARIATION IN BETWE EN

PFD

NORM

ALT OF F

EMERGENCY POWER LIGHTS

OF F

850°C 900°C 900°C

96.5 98.9

78.0 66.0 — —

L ENG TIME LIMIT TIME LIMIT

800°C 700°C

97.5 99.6

98.8 101.0

COPILOT

NORM

STANDBY ELECTRICAL POWER

L GEN

RC BLE ED AIR

MAX MAX TGT TGT

— —

101.3 101.1

MAX CONTINUOUS MAX CONTINUOUS MAX OVERSPE ED MAX OVERSPE ED MAX REVERSE MAX REVERSE

1159F50290-11 1159F50290-11

MASTER

AC

HP % HP % (N2) (N2)

— —

MAX TAKE-OF F MAX TAKE-OF F

MAX OVERTEMPERATURE MAX OVERTEMPERATURE

AV PWR

DC

LP % LP % (N1) (N1)

CABIN ALT FT

G

2 GPS #2

COM/ NAV

LOW IDLE (MIN) LOW IDLE (MIN)

SYS/ EICAS

DISPLAY SYSTEM CONTROL

PILOT

RESET

F

2 NAV RCVR #2

2 RFMU #2

BR700-710A1-10 OPERATING LIMITATIONS BR700-710A1-10 OPERATING LIMITATIONS

NORM ENG/ EICAS

PILOT

NORM

RIGHT LO OP A LO OP B

ON

5 ATC #2

2

MAX AIR START MAX AIR START

NAV DISPLAY SWITCHING NORM

FIRE DETECTION

LEFT

OF F

FLT CTRL/ HYD

2 ADF #2

DME #2

CONDITION CONDITION

PFD APU

TEST

71/ 2 VHF COMM #2

3 L ES S DC CONT #2

MAX GROUND START MAX GROUND START

FIRE TEST L ENG

71/ 2 L/ R IGN 2

3 R ES S DC CONT

ELECTRICAL #1

TEST

LO OP A LO OP B

5 L/ R START B

ENG SYS/ CNTRLS

3 R ES S DC CONT #2 ELECTRICAL

FUEL

FAULT

E

2 STAL L BAR R VALVE #1

GPWS

LO OP A

FLAP/ STAB R STBY AC

5 LDG GEAR CONT

TEST

TEST

A/ P SERVO #2

5

COM/ NAV

SELCAL

IAC #2

D

5

DO OR CONT / WARN

FLT CTRL/ HYD

F

G

CDU #1

71/ 2

CLOS

E

BCN

1

5 L SIDE WSHLD

ANTI-ICE

2

5

TEST

10

1 MADC #1

10 DISPLAY UNIT #3 PRI

LEFT WOW

D

PAS S OXYGEN

9

1 CLOCK #1

10 DISPLAY UNIT #1

20 L FRONT WSHLD

8

71/ 2 #1 AOA HTR

5 R SIDE WSHLD

B

ANTI ICE HTR

ADA

LEFT

RIGHT

UP PER

OF F

OF F

PROBE

SIDE

LOWER

LF/ RS

RF/ LS

OF F

OF F

OF F

OF F

OF F

ORIDE OF F PNL

OF F LTS

LIM AIR KN ITATIOSPEE OT S & NS D V MA -C.A.S A CH . V NO LE . 20 V 250K 6K LO 225K/.70 /.70 10° V FE 20° FLAP 39° FL AP 250 FL AP 220 170

NO GND SPLRS

PLAN

ON

ON

TRS

SYSTEM

ON

ON

ON

DISP

ON

ON

BRT

8.88

BARO

NAV

ON FLT REF

ON TEST

PUSH STD

BG O OF F GPWS BG VOR2 CHECKLIST ID WAYPT VERT PROP ID NAVAID TUNED TCAS ID AIRPT WIND XY VECT

CRS

ON

888

V NAV

CHG

ON

MAN

ON

SYNC

BC

ON

ON

ON

1.55

EPR

ST

CTL

10

DH

MILEN

10 00

G

10

SVO IGN

TIL LS SAV 12

10

NBC

80.0

HP

80.0

20

20

DH 150

360 CRS

359

N 33 30 W

330

11RW36

FF

ON

SVO IGN

PUL L TO CAGE

15500

FMS2

1

0 9 8

ALT

6

SAV 12

IN HG

4

5

ON

GS 20

20

10

10

10

NBC

10

140

STBY HDG

11RW36

060/ 040

HDG MODE

MDAU

D

33

G

A D F

VOR

FAULT

AUTO

3

A D F

+1 -1

20 DH 150

360 CRS

359 0 10

N

-3 20 -5 39 STB FLP

HDG

360

NAV1 13.7 N M

29.92 IN +3000

3

6 3 2 1

+1500 ADF1

1 2 3 6

20:34 FT

ST

SELECT

CONTROL

SEL

CTL

D E S C E N D

FLIGHT

FAULT

AUTO

LANDING

MANUAL

SEMI

LANDING

MANUAL

SEMI

VOR AERONETICS

FAULT / MAINTENANCE PARAMETERS COMPONENT OPERATION MEMORY USED MDAU SW PART NUMBERS END OF FLIGHT REPORT DISABLED RECOR DINGS:NONE

Compas s No.

Plane No. Date Lat. Magnetic Heading 0 15 30 45 60 75 90 105120135150165 Compas s Course Magnetic Heading 180195210225240255270285300315330345 Compas s Course

GA 252 5/ 83

Compas s Correction Card

2 1 BRAKE PRES S 3 0

4

N L

MDAU SCRE ENS ARE FOR ADVISORY PURPOSES ONLY

0.58M AOA 0.60

33

SAT -21 TAS 160 GSPD 185

HDG TEST

0

H

MDAU TOP MENU

MAV13

WX

STBY FAIL

100

120

50

15

HDG 330 ALT HDG

CHRONOMETER

GNT LT

20 50

8

KNOTS

IAS

10 00

G

2 9 9 2

0 60

50 BARO

2500 1500

DADC3

DH

ISE

9

TIL LS

3

7 1 0 1 3

IF39 1

5000

12

DISENG

STBY RUD

OF F

RAD ALT

160

2

9

DISENG

TER RAIN DISPLAY

AP1 AT1

180

MILEN

6 0 0

6

PITCH TRIM ENG/ DISENG

LOC

160 200

TRK CHG OF FSET

MB

15000

N

ALD

0

9 8

30

PRI HDG

L-R ENG MAINT STD EVENT RECOR D EXCE EDANCE RECOR D L-R ENG MAINT150 L-R ENG MAINT LTD MAINT REQUIRED MDAU FAIL MDAU MFM 90% FUL L YAW DAMP ENG/ DISENG

GPWS O'RIDE

BELOW G/ S G/ S INH

LOC

SAV

FL390

VOICE O'RIDE

INHIBIT

C

30

FUEL TANK TEMP 30

SAT -21 TAS 160 GSPD 185

FL410

WARN INHIBIT

W

20

B/ CRS

VOR2

FL410

MASTER WARN SET

10

ILS

ILS

AUX 0

30500

MAV13

WX

PUSH STD

BG O OF F GPWS BG VOR2 CHECKLIST ID WAYPT VERT PROP ID NAVAID TUNED TCAS ID AIRPT WIND XY VECT

BRT

10

0.57

060/ 040

SAV HDG 330

ON SYSTEM

ON

ON

E

HDG

E

-3 20 -5 39 STB FLP

4900

50

15

29.92 IN +3000 6 3 2 1 +1500 1 2 ADF1 3 6

NAV1 13.7 N M 3

BARO

NAV

ON TRS

ON

6

0 10

6

+1 -1

50

100

120

PLAN

ON

DISP

ON

30

0.27

RIGHT 3000

PTU 0

FUEL QTY

140

0.58M AOA 0.60

HP

HYD PRES S LEFT 2900

30

FT

CONTROL

SEL

10

AT T HDG ALT

COMP

FLT REF

ON TEST

20

EVM LP

0.56

OF F

90.0 REV

21

GNT LT

SELECT

ISE

9

6

20:34

IF39 1

160

ON SENSOR

90

0.26

ALT LP

CRS SYNC

OF F

74 OIL TEMP

90 712

ALT 90.0 REV

3

ALD

MAP

74

A/ I TGT

FMS1 VERT ALERT TRK CHG OF FSET

N 33 30

180

CHRONOMETER

360

2500 1500

MADC3

20

ON

OIL PRES S

1.50

SYNC

GS 20

ALT HLD

G/ S

FLEX TO/ GA

A/ I 702

AP1 AT1

88888

LR

ON

FLIGHT

ALTITUDE

ON

PFD-CMD

CHG

APR

1.55

1.50

LOC

160 200

A/ P

8888.8

LOW

SYNC

FLCH

SET

30

COMP

ON SENSOR

W

C

27

W

BELOW G/ S G/ S INHIBIT

VS-FPA

BANK

24

MASTER WARN

INHIBIT

RAD ALT

HEADING

18

WARN INHIBIT

GPWS O'RIDE

VALVE

NO GND SPLRS

L NAV

15

MAP

VOICE O'RIDE

E

EN OP

V 270 TUR .75 K BUL M TO ENC AB 32, E OV 000 BE V E 32, FT LO MO 000 . W /M FT . 8,0 8,000 MO 27, 00 TH FT 860 RU .30 0K 27, AB FT 860 OV 300 30, 930 FT E K . 32, 160 FT. .850M 43, 500 FT .875M 45, . 51, 000 FT. .885M 000 FT .88 FT . .885M . .86 0M 0M

E ENC FT. . BUL 000 FT V TUR 32, 000 K TO E 32, 270ABOV MO 0K / M FT.30 M .75 V MO 00 8,0 RU W TH 300K LO 00 BE 8,0 FT E 0M 860 27, ABOV. .85 5M FT .87 . 5M 860 FT .88 5M 27, 930 . FT .88 30, . 0M 160 FT .88 0M . 32, 500 FT .86 . 43, 000 FT 45, 000 51,

SPE ED

C L I M B

CLOS

. D EE A.S SP -C. NO. AIR TIONSCH ITA MA 6K LIM S & 20 OT 0 KN /.7 A 0K 25 /.70 V 5K LE 22 V LO 0 25 V VFE 0 AP 22 FL AP 170 ° 10 FL AP ° 20 FL ° 39

R

HORN SILENCE

LOCK RELEASE

DO NOT OPERATE WITH L/ G EXTENDED ABOVE 2 0 ,0 0 0 FT

AIRCRAFT/ CABIN

LANDING ELEV

BARO/ COR R

1000

IN HG

FT

100

Figure 21-49. Cabin Pressure Control System Components—Cockpit

21-76




SYSTEM COMPONENTS

NOTES

General The purpose of the cabin pressure control system is to provide control, regulation, and monitoring of cabin pressure to ensure maximum passenger comfort and safety. It includes the following components located in the cockpit (Figure 21-49): • Cabin pressure indicator • Cabin pressure control panel • Cabin pressure selector panel In addition, the cabin pressure control system also consists of four other components located in the right electronic equipment rack: • Cabin pressure controller • Thrust recovery outflow valve (TROV) • Pressure relief valve • Cabin pressure acquisition module


21-77



Cabin Pressure Indicator and Control Panel The cabin pressure indicator and control panel are both located on the cockpit overhead panel. The cabin pressure indicator provides the crew with a visual presentation of the cabin altitude, cabin differential pressure, and cabin pressure rate of change. The control panel provides the crew with the means for selection and control of the cabin pressurization control system via three switchlights: FLIGHT–LANDING, FAULT–MANUAL, and AUTO–SEMI.

frequent flashing, which is proportional to valve opening or closing speed. The valve position meter provides a visual representation of the TROV position regardless of mode of operation.

The FLIGHT–LANDING momentary-position switch provides the means to manually prepressurize or depressurize the aircraft on the ground or to select in flight a descent schedule for landing. In the automatic or semiautomatic mode, the cabin pressurization controller automatically switches from LANDING to FLIGHT if the doors are closed and either of the throttles are advanced to 35.5° of throttle lever angle or the aircraft taxi speed exceeds 9 knots. The FAULT–MANUAL alternate-position switch alerts the crew to select the manual mode if a cabin pressure controller failure causes the FAULT annunciator to illuminate. The AUTO–SEMI momentary-position switch provides selection of the automatic or semiautomatic mode of operation. The MAN HOLD control knob provides the means to manually open and close the outflow valve via the DC motor drive. The DESCEND legend equates to valve closed, and the CLIMB legend equates to valve open. When in the manual mode, the outflow valve motor indicator lamp illuminates amber. This provides a visual indication of the outflow valve’s rate of movement. As the MAN HOLD control knob is rotated, a varying amount of power is applied to the DC motor on the outflow valve. The motor indicator light pulses as power is applied. More power causes more

21-78


NOTES



Cabin Pressure Selector Panel The cabin pressure selector panel is located on the center pedestal. It consists of the following: • Aircraft cabin indicator • Barometric correction indicator • Landing field elevation indicator • Rate-of-change indicator The aircraft/cabin indicator provides the means to display the actual aircraft altitude and the computed cabin altitude in 1,000-foot increments. The barometric correction indicator provides the means to display the barometric correction in inches of mercury. The landing field elevation (LFE) indicator provides the means to display the landing field elevation in feet. The cabin rate-of-change (climb or descent) indicator displays nominal rate of change in 100-feet-per-minute increments.

controller and the selector panel from the FMS. In the semiautomatic mode, the crew enters the landing field elevation via the landing field elevation selector knob. In the automatic mode, the cabin pressure controller schedules cabin rate of change to 500/300 fpm. In the semiautomatic mode, the cabin rate of change, normally 500/300 fpm, is entered by the crew, using the cabin rate-ofclimb or -descent selector knob.

NOTE The selector panel displays dashes in manual control.

NOTES

In the automatic mode, the aircraft/cabin indicator displays the actual aircraft altitude and cabin altitude. The cabin pressure controller generates this data. In the semiautomatic mode, the aircraft/cabin indicator displays the selected altitude (aircraft or cabin). The aircraft/cabin selector knob provides the means to input data in the semiautomatic mode for aircraft cruise and cabin altitude. In the automatic mode, the barometric pressure corrections are input at the pilot’s BARO SET knob on the display controller and transmitted to the cabin pressure controller through the ADMs. The cabin pressure controller then transmits the barometric pressure correction to the BARO COR window on the CPSP. In the semi-automatic mode, the crew inputs the barometric pressure corrections, in inches of mercury, directly to the cabin pressure controller via the BARO COR selector knob on the CPSP. In the automatic mode, the landing field elevation data is an input to the cabin pressure


21-79



CP CONTROLLER

TEST OFV SEL PRV

CABIN PRESSURE CONTROLLER

PCP

CH 1 CH 2

R I G H T P D B

PRESSURE RELIEF VALVE (PRV)

CABIN PRESSURE ACQUISITION MODLE (CPAM)

OUTFLOW VALVE (OFV)

Figure 21-50. Cabin Pressure Control System Components—REER

21-80




Cabin Pressure Controller (CPC)

Thrust Recovery Outflow Valve

The cabin pressure controller is located in the right electronic equipment rack (REER) (Figure 21-50). It is a dual-channel automatic microprocessor-based electronic unit that uses digital control technology. It senses cabin pressure directly and static pressure electronically. The cabin pressure controller compares actual to programmed pressure schedules, and it controls the cabin pressure outflow valve electronically to maintain the correct pressure and rate.

The thrust recovery outflow valve (TROV) is located on the aircraft fuselage under the lower shelf of the right electronic equipment rack. The valve consists of a frame, 2 outflow doors, and an actuator/gearbox assembly along with connecting links. Its purpose is to provide an exhaust point for pressurized cabin air.


Only one channel is actively controlling the cabin pressure system at a time. The channel control alternates after each flight. Failure of the controlling channel causes an automatic switch to the alternate cabin pressure controller channel. Each channel has its own power source and receives digital and discrete data from the various data inputs independently. Each channel also controls its own dedicated AC motor on the TROV valve. The controlling channel provides outputs to the cabin pressure indicator and the CAS for display and fault indication. The cabin pressure controller provides continuous and manually initiated built-in test (BIT) at system power-up. The BIT detects degraded system performance and isolates faults to the line replaceable unit (LRU) level. Four red light emitting diode (LED) fault indicators (monitoring the outflow valve, selector panel, pressure relief valve, and cabin pressure control panel) are located on the controller’s face to provide fault isolation information during flight-line troubleshooting. Two additional LEDs indicate which channel is faulty. Faults can be cleared by selecting and then deselecting the MANUAL–FAULT switch on the cabin pressure control panel. This action also changes the active channel.

The valve also employs an irreversible geartrain that prevents outflow door motion, except when commanded by one of three drive motors, the AUTO–SEMI mode channel No. 1 and No. 2 AC drive motors, or the manual mode DC motor. The valve also incorporates a valve position potentiometer.

Pressure Relief Valve The pressure relief valve (PRV) is located under the lower shelf of the right electronic controller and is mounted with a marmon flange and a V-band clamp forward of the outflow valve (Figure 21-50). The pressure relief valve provides positive differential pressure relief, negative differential pressure relief, and an additional outflow area in the ground mode. The pressure relief valve incorporates two metering sections for maximum positive pressure differential relief. The first metering section relieves at 10.28 to 10.48 psi. The second metering section relieves at 10.48 to 10.68 psi. It also incorporates a jet ejector pump and a solenoid to actuate the valve open when the aircraft is on the ground. The valve has its own cabin pressure sense port and two dedicated static sense ports. The static pressure sense ports are located on the right fuselage, forward of the wing.


21-81




CP CONTROLLER

TEST

76∞ 78∞

OFV SEL

76∞ 78∞

74 ∞F

80∞ 80∞ 78 ∞F

74 ∞F

PRV PCP

78 ∞F

CH 1

35 ∞F

35 ∞F

78 ∞F

CH 2

0 psi

0 psi 370 ∞F

370 ∞F

42 psi

42 psi

Left Eng

Right Eng 78 ∞F

78 ∞F APU

L Elev 1030

Cab Alt -450

Rate 0

P 0.25

Mode Auto 1

Figure 21-51. Cabin Pressure Controller and ECS/PRESS Synoptic Page

21-82




Cabin Pressure Acquisition Module The cabin pressure acquisition module is located on the bottom of the right electronic equipment rack (Figure 21-50). It is a self-contained microprocessor unit, with a dedicated cabin pressure sensing port. The module uses the aircraft pitot-static system for static pressure sensing and backs up the cabin pressure control system controller for display, essential warning, and caution messages.

SYSTEM OPERATION Cabin Pressure Control Panel Operation The cabin pressure control panel selectable modes of operation are automatic (AUTO), semiautomatic (SEMI), and manual (MAN). The AUTO mode is the normal mode of operation for this system. In the AUTO mode, with all the FMS data valid and available, the cabin pressure controller provides all processing to deliver automatic operation. The cabin pressure selector panel displays information from the FMS and ADMs via the cabin pressure controller (see Figure 21-48).

Table 21-1. CABIN PRESSURE CONTROLLER SEMI-AUTO SCHEDULE SELECTED ALTITUDE

CABIN ALTITUDE

0 50000 10000 15000 20000 21000 22000 23000 24000 25000 26000 27000 28000 29000 30000 31000 32000 33000 34000 35000 36000 37000 38000 39000 40000 41000 42000 43000 44000 45000 46000 47000 48000 49000 50000 51000

–1900 –1900 –1900 –1900 –1900 –1900 –1900 –1900 –1900 –1577 –1155 –741 –337 56 440 814 1178 1533 1877 2213 2538 2854 3159 3450 3730 4001 4258 4507 4747 4975 5194 5406 5608 5808 5989 6000

SELECTOR DISPLAY AIRCRAFT CABIN 0.0 5.0 10.0 15.0 20.0 21.0 22.0 23.0 24.0 25.0 26.0 27.0 28.0 29.0 30.0 31.0 32.0 33.0 34.0 35.0 36.0 37.0 38.0 39.0 40.0 41.0 42.0 43.0 44.0 45.0 46.0 47.0 48.0 49.0 50.0 51.0


–2 –2 –2 –2 –2 –2 –2 –2 –2 –1 –1 –1 –0.3 0.1 0.4 0.8 1.2 1.5 1.9 2.2 2.5 2.9 3.2 3.4 3.7 4.0 4.3 4.5 4.7 5.0 5.2 5.4 5.6 5.8 6.0 6.0

21-83



Table 21-2. CABIN PRESSURE CONTROLLER AUTO SCHEDULE AIRCRAFT ALTITUDE (FEET) 0 3,000 6,000 9,000 12,000 15,000 18,000 21,000 24,000 27,000 30,000 33,000 36,000 39,000 42,000 45,000 48,000 51,000

21-84

AMBIENT AUTO CABIN DELTA PRESSURE SCHEDULE PRESSURE PRESSURE (PSIA) (FEET) (PSIA) (PSID) 14.70 13.17 11.78 10.50 9.35 8.29 7.34 6.48 5.70 4.99 4.36 3.80 3.30 2.85 2.47 2.14 1.85 1.60

–600 –450 –289 –114 214 448 712 1,008 1,347 1,743 2,238 2,645 3,047 3,574 4,249 4,916 5,496 6,000

15.02 14.94 14.85 14.76 14.58 14.46 14.32 14.117 13.99 13.79 13.55 13.34 13.15 12.89 12.58 12.27 12.00 11.78


0.32 1.77 3.07 4.25 5.24 6.17 6.98 7.69 8.30 8.80 9.18 9.54 9.85 10.04 10.10 10.13 10.15 10.17



NOTE

NOTE

Refer to the Maintenance Schematic Manual, Chapter 21, for corresponding schematics.

The FLIGHT–LANDING switch changes from LANDING to FLIGHT automatically if the airplane begins to move at greater than 9 knots or if the throttles are advanced to takeoff power setting.

The SEMI mode of operation is used when FMS inputs are invalid or the landing field elevation is missing. If FMS inputs are lost, the selector panel freezes and displays the last valid FMS inputs. The cabin pressure selector panel knobs are functional, and the crew must provide input. When both channels of the cabin pressure control fail, the FAULT switch legend illuminates. Crew action is required to select (depress) the switch, which places the system in manual. An amber “CPCS Fail-Sel Man” message also appears on the CAS. Selection of the FAULT–MANUAL switch during normal operation places the system into the manual mode. The MAN HOLD knob controls the position of the TROV. Once the MAN HOLD knob is released and returns to its spring-loaded center position, the outflow valve remains in its last position until the knob rotates again. In manual control, the valve speed is proportional to the MAN HOLD knob position. During the AUTO and SEMI modes of operation, the manual TROV DC driver is disabled. The active channel control alternates after each power-down sequence. In the manual mode, the channel 1 and channel 2 outflow valve AC drivers are disabled (Figure 21-48). Automatic or manual selection of the FLIGHT–LANDING switch to FLIGHT prepressurizes the cabin pressure control system to 500 feet (0.25 psid) below takeoff field elev a t i o n p r i o r t o a c t u a l t a k e o f f . Prepressurization is initiated if the doors are closed and any of the following occur: • Switch is selected to the FLIGHT position.

When LANDING is selected on the FLIGHT–LANDING switch, the cabin pressure control system automatically switches to the descent depressurization schedule (300 fpm). If the switch is not manually selected and the aircraft descends from cruise at least 1,000 feet, the FLIGHT–LANDING switch automatically switches from FLIGHT to LANDING. A climb of 1,000 feet or a new level-off at a lower altitude (after one minute) causes the CPCS to return to the FLIGHT mode. The descent pressurization schedule (300 fpm) is maintained until the cabin altitude is 250 feet below landing field elevation. Once weight is on wheels, the cabin is depressurized at a rate of 500 fpm for one minute. Therefore, 30 seconds after landing the cabin should be at landing field elevation. When the TROV is fully open and one minute has elapsed, the CPCS enters ground operation, energizing the pressure relief valve jet ejector pump solenoid valve and transferring from the active channel to the inactive channel in preparation for the next flight.

NOTE When landing at high field elevation airfields, the cabin may be at an altitude lower than landing field elevation. In these scenarios, during the aircraft descent, the CPCS will depressurize the cabin at a rate of 500 fpm to a value that is 250 feet below landing field elevation.

• Airplane begins to move at greater than 9 knots. • Throttles are advanced to the takeoff power setting (35.5° or greater throttle lever angle).


21-85



CPC +28 VDC PC

Pa

Pa 10.28 TO 10.48 PSID

10.48 TO 10.68 PSID

PC

PC

TAT SOLENOID VALVE BLEED AIR

CABIN

ATMOSPHERE

LEGEND AMBIENT PRESSURE (PA) REF CHAMBER PRESSURE CABIN PRESSURE

Figure 21-52. Pressure Relief Valve Cutaway

21-86




Cabin Pressure Selector Panel Operation In automatic mode, the cabin pressure selector panel LCDs provide viewing of FMS- and ADM-produced information while the selector knobs are disabled. The LCDs are backl i g h t e d a n d c a n b e d i m m e d d ow n t o t h e predetermined panel lighting threshold voltage. If the FMS information becomes invalid, the selector freezes the last valid data. The selector knobs become active when the SEMI mode is selected on the cabin pressure control panel. In the SEMI mode, selector setting knobs for aircraft cruise/cabin altitude, barometric correction, landing field elevation, and cabin rate of climb and descent are used for data input. If new landing altitude data are not provided via the selector knobs, the landing field elevation retained in memory is used for computation. A barometric correction to the landing field pressure is also required.

Cabin Pressure Controller (CPC) Operation In the AUTO or SEMI mode of control, the active controller senses the cabin pressure. The outflow valve is commanded to open or close as required to control the cabin pressure by regulating airflow out of the fuselage. Both channels produce individual discrete outputs, but only the active channel drives the outflow valve via a dedicated motor. Active control alternates between channels on every new flight. The essential 115-VAC phase A bus powers channel 1 of the cabin pressure controller. The right main 115-VAC phase A bus powers channel 2. The built-in test equipment (BITE) monitors the micro air data inputs, FMS inputs, shared discrete inputs, cabin pressure selector panel, cabin pressure control panel, pressure relief valve, and outflow valve. It also provides a cross-check of the measured cabin pressure between channels.

The cabin pressure controller limits the cabin pressure change rate during climb and descent. During the AUTO or SEMI mode of operation, the differential pressure across the fuselage is limited by the CPCS. The cabin pressure controller transmits cabin pressure data to the symbol generators and the cabin pressure indicator for display. Crew alert messages, advisory messages, component failures, and other fault data are transmitted via the ARINC 429 data bus to the EICAS and the CMC.

Outflow Valve Operation Each channel of the cabin pressure controller drives its own dedicated TROV AC motor and provides tachometer feedback into the cabin pressure controller (Figure 21-48). The cabin pressure control panel MAN HOLD knob drives the DC motor when in the manual mode of operation. The actuator provides door position feedback from the TROV position potentiometer to the control panel indicator in all modes of operation.

Pressure Relief Valve (PRV) Operation The pressure relief valve is designed to prevent damage to the aircraft fuselage from excessive positive or negative pressures (Figure 21-53). The maximum positive pressure relief occurs when cabin pressure differential reaches 10.28 to 10.48 psi. If the first metering section fails, a second metering section relieves maximum positive pressure when the differential pressure increases to 10.48 to 10.68 psi. Negative pressure relief occurs when ambient pressure exceeds cabin pressure by 0.25 psi.


21-87




76∞ 78∞

76∞ 78∞

74 ∞F

78 ∞F

74 ∞F

35 ∞F

78 ∞F

80∞ 80∞

35 ∞F

78 ∞F

0 psi

0 psi

CPI

370 ∞F

370 ∞F

42 psi

42 psi

Left Eng

Right Eng 78 ∞F

78 ∞F APU

L Elev 1030

Cab Alt -450

Rate 0

P 0.25

Mode Auto 1

CAS DISPLAY ARINC 429

CPAM

ARINC 429

CONTROLLER

Figure 21-53. Cabin Pressure Indicator

21-88


ADVANCED GRAPHICS MODULE



Due to the large cabin volume, the valve is commanded open by the cabin pressure controller when in the AUTO or SEMI mode of operation one minute after the aircraft has landed and the TROV is fully open. The cabin pressure controller energizes a solenoid on the pressure relief valve. Bleed air flowing through a jet pump ejector decreases the reference chamber pressure, causing the valve to open.

NOTE Refer to the G500/G550 Maintenance Schematic Manual, Chapter 21, for corresponding schematics.

Cabin Pressure Acquisition Module (CPAM) The cabin pressure acquisition module is located under the lower shelf of the right electronic equipment rack (see Figure 21-50). It is a self-contained microprocessor unit with its own cabin pressure sense port.

SYSTEM INDICATIONS General An indication system provides the flight crew with current data from the cabin pressure control system. The system indications provide a readout of the following: • Cabin altitude in feet • Cabin differential pressure in psi • Cabin rate of climb and descent • Mode of operation The indications are available to the crew on the CAS and the synoptic pages.

Cabin Pressure Indicator (CPI) The cabin pressure indicator is located adjacent to the cabin pressure control panel in the cockpit overhead panel (Figure 21-53). The indicator receives indication data from the controlling channel of the cabin pressure controller when in the AUTO or SEMI mode of operation. When in the manual mode, indication data is received from the cabin pressure acquisition module (CPAM). It provides numeric readouts for CABIN ALT FT (cabin altitude in feet), DFRN PRESS (cabin differential pressure in psi), and RC (a rate-of-change analog meter which indicates the cabin rate of change in increments of one thousand feet per minute).

It receives static pressure from the aircraft pitot-static system. It then computes cabin and static pressure and communicates cabin altitude in feet, differential pressure in psi, and rate of climb or descent to both channels of the cabin pressure controller and to the cabin pressure indication over the ARINC 429 data bus. The module also provides fault logic discrete signals to the MAUs.

Indication System Operation The cabin pressure indicator displays cabin pressure information. This information is transmitted from the active channel of the cabin pressure controller via the ARINC 429 data bus when in the AUTO or SEMI mode of operation. If both channels of the cabin pressure controller become inoperative or the data becomes invalid, the cabin pressure indicator displays the cabin pressure information computed by the cabin pressure acquisition module via its dedicated ARINC 429 data bus. This data is also displayed whenever the cabin pressure control panel is selected to MANUAL. The cabin pressure acquisition module processes the information internally. The module has one ARINC 429 bus to the cabin pressure indicator. The cabin pressure acquisition module outputs four discrete signals, three to the MAU 1 DGIO 10 and one to the MAU 2 DGIO 8.


21-89



The discrete signals sent to the MAU 1 DGIO slot 10 are: • CABIN PRESSURE LOW • CABIN DFRN ≥ 10.48 (actual CAS message is red CABIN DFRN - 10.48) • CPAM FAIL The discrete signal sent to the MAU 2 DGIO slot 8 is CABIN DFRN ≥ 10.28 (actual CAS message is amber CABIN DFRN - 10.28). The CAS ECS/PRESS synoptic page displays landing field elevation in feet, cabin altitude in feet, rate of climb or descent in feet, differential pressure in psi, and mode of operat i o n ( AU TO o r S E M I a n d C P C c h a n n e l controlling). The CAS SUMMARY synoptic page displays landing field elevation in feet, cabin altitude in feet, rate of climb or descent in feet, and differential pressure in psi.

21-90


Please refer to Chapter 34, “Navigation” for information regarding Autoflight information.



CHAPTER 23 COMMUNICATIONS CONTENTS Page INTRODUCTION ................................................................................................................. 23-1 GENERAL ............................................................................................................................ 23-1 RADIOS ................................................................................................................................ 23-3 AUDIO CONTROL PANELS ............................................................................................... 23-5 ACP Operation............................................................................................................... 23-7 MULTI-FUNCTION CONTROL DISPLAY UNITS (MCDUs) .......................................... 23-9 MCDU (Radio Tuning) Operation............................................................................... 23-11 MCDU Icons................................................................................................................ 23-11 Tuning COM 1 With MCDU....................................................................................... 23-13


23-i




Title

Page

23-1

Communications Subsystem Components............................................................. 23-2

23-2

Audio Control Panels ............................................................................................. 23-4

23-3

Audio Control Panel Function Selection................................................................ 23-6

23-4

Multi-Function Control Display Units (MCDUs).................................................. 23-8

23-5

MDCU (Radio Tuning) Operation ....................................................................... 23-10

23-6

MCDU Line Select 1L ......................................................................................... 23-12

23-7

MCDU Line Select 2L ......................................................................................... 23-12


23-iii



CHAPTER 23 COMMUNICATIONS

MODE

OPR

FREQ/CHAN

PWR

II H F

VOL

SQL

FREQ/LD

DSBL

CHAN

CURSOR

VALUE

INTRODUCTION The communications system contains the subsystems necessary for communication within the aircraft, between different aircraft, and between the aircraft and ground stations. These subsystems include components which supply voice and data communication

GENERAL The communications system is divided into the following subsystems: • Voice/Navigation Radio Systems • Audio Control Panels • MCDU (Multifunction Control Display Unit) Radio Tuning


23-1



Figure 23-1. Communications Subsystem Components

23-2




RADIOS

NOTES

The PlaneView radio system consists of Honeywell digital radios contained in two modular radio cabinets, a third stand-alone VHF digital NAV/COM and two HF radios. Each modular radio cabinet (MRC) consists of an ADF, DME, Mode S transponder (XPDR), network interface module (NIM), VHF data radio (VDR), and VOR/ILS/VDL (VIDL). Each of the functions is contained in a line relaceable module that is self-contained within its own housing and has its own power supply. MRC No. 1 is located in the left electronic equipment rack (LEER), and the No. 2 MRC is located in the right electronic equipment rack (REER). The two HF radios are located in the tail compartment. The HF communication system provides long range voice and data communication capabilities to and from the aircraft. It is further divided into the HF No. 1 system and the HF No. 2 system. Each system consists of an HF transceiver and an associated coupler unit utilizing a shared antenna. The third NAV/COM radio is located in the REER. It is a stand-alone unit which contains the functionality of the VDR and VIDL modules. Functionality of being a VDR or a VIDL is based on the frequency to which the radio is tuned.


23-3



H'MIC

COM 1: 72

COM 1: 72

COM 1: 72 COM 1: 72

Figure 23-2. Audio Control Panels

23-4




AUDIO CONTROL PANELS

NOTES

Three AV-900 audio control panels (ACPs) are installed in the cockpit: one at each pilot seat and one at the observer station. All panels are identical. All three audio panels are linked over redundant digital busses. Digital busses carry microphone transmission and audio receive selection data to and from the network interface modules (NIMs) in the modular radio cabinets (MRCs).

NOTE The NIMs translate voice analog audio into digital data.

The MRCs contain the communication and navigation radios (COM, NAV, ADF, DME and XPDR), a non-volatile memory that stores radio and audio configuration data, and an internally mounted cooling fan. Both the analog t h i r d d u a l f u n c t i o n NAV / C O M r a d i o (NAV/COM 3) and the high frequency (HF) radios are located outside of the MRCs and are interfaced to the audio control panels through ARINC-429 connections to the MAUs. The ACPs provide the following functions to the flight crew: • Ten rectangular transmit button selections • Twenty round audio reception button selections • Three square function button selections, including selected calling (SELCAL) • Volume control • Emergency volume control • LCD display


23-5



H'MIC

COM 1: 72

H'MIC

COM 2: 72 Figure 23-3. Audio Control Panel Function Selection

23-6




ACP OPERATION

NOTES

Cockpit audio system functions shall be activated by selecting the button on the ACP associated with the function when the function is turned OFF (associated LED OFF). Similarly, cockpit audio system functions shall be de-activated by selecting the button on the ACP associated with the function when the function is turned ON (associated LED ON).

NOTE Only one transmit selection may be made at a time, and selection of the transmit button will automatically activate the audio reception button for the associated radio, if not previously selected.

When a function is first selected, the volume may be adjusted by using the volume button and observing the volume level in the display portion of the ACP. The audio panels at the pilot and copilot stations provide an output of eight (8) watts to the respective cockpit speaker on the overhead. Transmit audio output from each panel is also provided to the cockpit voice recorder (CVR) system.


23-7



Figure 23-4. Multi-Function Control Display Units (MCDUs)

23-8




MULTI-FUNCTION CONTROL DISPLAY UNITS (MCDU)

NOTES

The MCDU is a keyboard control/display device that provides general purpose data entry, control, and display for the flight crew. The MCDU provides keyboard entry for the flight management system (FMS), radio system tuning, and SATCOM dialing. The MCDU is the primary interface to the radio management and control system for the flight crew. It allows the pilot to preset and activate frequencies for the aircraft radios. The three MCDUs are installed in the center pedestal (Figure 23-4)


23-9



Figure 23-5. MCDU (Radio Tuning) Operation

23-10




MCDU (RADIO TUNING) OPERATION The radio tuning function is accessed via the RADIO function key on the MCDU, which displays the RADIO 1/2 page. All other pages are accessed from RADIO 1/2 using the line select keys or the NEXT and PREV function keys. The tuning knob is used to dial in frequencies or other numeric values. The knob uses an outer knob and upper post for tuning. M ove m e n t b e t w e e n t h e R A D I O 1 / 2 a n d RADIO 2/2 pages is done with the NEXT and PREV function keys. Access to the COM DETAIL, TCAS/XPDR, and NAV DETAIL pages is via line select keys from RADIO 1/2. Access to the HF DETAIL, COM/NAV, and ADF DETAIL pages are via line select keys from RADIO 2/2. All of the pages that correspond to a particular radio type (e.g., VHF communications or ADF) typically consist of a details page along with two memory pages. The details page allows the specific features of each radio to be configured and provide access to the memory pages associated with each radio type. In keeping with the titling and numbering conventions used by MCDU functions, each MCDU page is arranged with a centered title at the top and a page number in the upperright corner. Page numbers are formatted as the current page number (among those with the same page a slash/, and the number of pages with the same title. For example, there are two pages entitled RADIO, the first is labeled RADIO 1/2 and the second is RADIO 2/2.

MCDU ICONS

Page Indicator When this icon is displayed, pushing the adjacent LSK changes the display to another page. The page to be displayed is either labeled explicitly or is a detail page for the radio in the associated field.

Exclusive Selection This icon is displayed next to a list of mutually exclusive options. Each time the adjacent LSK is pushed, the next options become selected, wrapping around to the first when the last option is reached. The selected value is displayed in the active color (green) and large font, while the other selections are displayed as smaller characters.

Immediate Function This icon indicates that the function identified in the field will be carried out immediately when the adjacent LSK key is pushed.

Copy Value This icon is used on the memory pages to indicate that the frequency highlighted by the format cursor will be copied into the active frequency for the associated radio.

Tuning Curl this icon indicates that the data value highlighted by the format cursor can be changed by turning the MCDU tuning knob.

Swap Frequencies This symbol indicates that exchanges between the active and preset frequencies for the associated radio can be made. This has the effect of saving the currently active frequency in the preset memory, and tuning the radio to the frequency previously stored as the present.

Format Cursor The cursor box highlights the value in the currently preset field.


23-11



Figure 23-6. MCDU Line Select 1L

Figure 23-7. MCDU Line Select 2L

23-12




TUNING COM 1 WITH MCDU

NOTES

All radios can be tuned in the same manner as described herein. To tune a radio, the operator must first press t h e R A D I O f u n c t i o n k ey o n t h e M C D U (Figure 23-6).

Line Select 1L Line 1L is displayed in green and represents the current active frequency. Pressing LSK 1L exchanges the active and preset frequencies, (the white frequency shown at 2L), for VHF COM 1. If a valid COM frequency had been entered in the scratchpad entry, then pressing 1L replaces the active frequency with the frequency in the scratchpad.

Line Select 2L Line 2L is displayed in white and represents the current COM1 preset frequency. This is the default field for the format cursor when the RADIO function key is pressed. Pressing LSK 2L when the format cursor (white box) is already in the field displays the COM 1 page. When the format cursor (white box) is around the preset frequency, then the operator can change the preset frequency by using the tuning knobs. The outer tuning knob changes the frequency left of the decimal point, and the inner knob changes the frequency right of the decimal point. Once the desired frequency is selected, then pressing 1L will swap the newly tuned preset frequency with the active (green) frequency at 1L.


23-13



CHAPTER 24 ELECTRICAL POWER CONTENTS Page INTRODUCTION ................................................................................................................. 24-1 Buses and Relays ........................................................................................................... 24-7 Power Sources................................................................................................................ 24-8 Controls and Indications.............................................................................................. 24-11 AC POWER SYSTEM........................................................................................................ 24-25 Introduction ................................................................................................................. 24-25 Engine-Driven AC Generation System........................................................................ 24-27 Generator Control Units (GCUs)................................................................................. 24-36 Bus Power Control Units (BPCUs) ............................................................................. 24-43 APU-Driven AC Generation System ........................................................................... 24-53 Emergency AC Inverter System................................................................................... 24-59 External AC Power System ......................................................................................... 24-61 Generated AC Power Distribution ............................................................................... 24-67 DC POWER SYSTEM........................................................................................................ 24-77 Introduction ................................................................................................................. 24-77 Primary DC Power....................................................................................................... 24-79 Transformer Rectifier Units......................................................................................... 24-79 Primary DC Contactors ............................................................................................... 24-87 AC Contactors ............................................................................................................. 24-89 DC Power Source Distribution .................................................................................... 24-93


24-i



No-Break Power Transfer ............................................................................................ 24-93 Operation ..................................................................................................................... 24-93 BATTERY POWER ............................................................................................................ 24-95 Battery Power System Components ............................................................................ 24-95 APU Start Control ..................................................................................................... 24-101 Auxiliary Hydraulic Pump Circuit ............................................................................ 24-101 Battery Chargers..........................................................................................................24-105 Charge Modes............................................................................................................ 24-107 DC External Power.................................................................................................... 24-111 SUPPLEMENTAL POWER ............................................................................................. 24-113 Introduction ............................................................................................................... 24-113 Standby Electrical Power System.............................................................................. 24-113 Emergency Power System ......................................................................................... 24-125 60-Hz Power.............................................................................................................. 24-133 Cabin/Galley System ................................................................................................. 24-139 MDAU MESSAGES ......................................................................................................... 24-145 CIRCUIT-BREAKER PANELS ....................................................................................... 24-145 ELECTRICAL POWER SYSTEM REVIEW .................................................................. 24-145

24-ii





Title

Page

24-1

Equipment Locator................................................................................................. 24-2

24-2

Left and Right Electronic Equipment Racks.......................................................... 24-4

24-3

Left and Right Power Distribution Boxes .............................................................. 24-5

24-4

Main and Auxiliary Power Contacts ...................................................................... 24-7

24-5

Electrical Power Control Panel ............................................................................ 24-11

24-6

AC POWER Synoptic Page ................................................................................. 24-12

24-7

DC POWER Synoptic Page ................................................................................. 24-16

24-8

SUMMARY Synoptic Page ................................................................................. 24-20

24-9

Bus Diagram ........................................................................................................ 24-22

24-10

AC Power System Overview................................................................................ 24-24

24-11

Integrated Drive Generator (IDG)........................................................................ 24-26

24-12

IDG Assembly ..................................................................................................... 24-28

24-13

IDG Servicing ...................................................................................................... 24-32

24-14

IDG Sight Gage and Standpipe............................................................................ 24-33

24-15

IDG Oil Cooler .................................................................................................... 24-34

24-16

RPBD .....................................................................................................................24-48

24-17

Current Transformers........................................................................................... 24-48

24-18

IDG GCU Diagram .............................................................................................. 24-50

24-19

APU Generator..................................................................................................... 24-52

24-20

APU Generator—GCU Block Diagram, APU Generator ON............................. 24-54

24-21

Underfloor Equipment Locator—Emergency Inverter ........................................ 24-56

24-22

Emergency Inverter Operation ............................................................................. 24-56

24-23

External AC Power Receptacle ............................................................................ 24-60


24-iii



24-24

EXT AC Block Diagram...................................................................................... 24-60

24-25

Ground Power Carts—Type 1 Block Diagram .................................................... 24-62

24-26


24-27


24-28

LAXC Control ..................................................................................................... 24-66

24-29

NBPT Contactor Control Diagram ...................................................................... 24-68

24-30

PDB Circuit-Breaker Panel—LEFT MAIN AC .................................................. 24-72

24-31

PDB Circuit-Breaker Panel—RIGHT MAIN AC................................................ 24-74

24-32

DC Power System ................................................................................................ 24-76

24-33

TRU Location ...................................................................................................... 24-78

24-34

TRU Description .................................................................................................. 24-80

24-35

Hall-Effect DC Current Sensor ............................................................................ 24-82

24-36

BPCU Protective Trips......................................................................................... 24-84

24-37

Primary DC Contactor Control ............................................................................ 24-86

24-38

DC Crosstie Contactor Control............................................................................ 24-86

24-39

AC Contactor Location ........................................................................................ 24-88

24-40

Standby Electrical Power Junction and Relay Panel ........................................... 24-89

24-41

Left Main TRU AC Control ................................................................................. 24-90

24-42

DC Power System Diagram ................................................................................ 24-92

24-43

Main Aircraft Batteries ........................................................................................ 24-94

24-44

Battery Indication................................................................................................. 24-96

24-45

FWD External Switch Panel ................................................................................ 24-98

24-46

Battery Control .................................................................................................. 24-100

24-47

REDBC Control................................................................................................. 24-102

24-48

Battery Chargers ................................................................................................ 24-104

24-iv




24-49

Battery and Charger Diagram ............................................................................ 24-106

24-50

Battery Charger TR Mode Control .................................................................... 24-108

24-51

DC External Power Circuit ................................................................................ 24-110

24-52

Standby Power Circuit—HMG ON................................................................... 24-112

24-53

Hydraulic Motor Generator ............................................................................... 24-114

24-54

Underfloor Equipment Locator—Hydraulic Motor Generator Control Unit .... 24-116

24-55

Standby Junction and Relay Panel..................................................................... 24-118

24-56

HMG System Diagram ...................................................................................... 24-120

24-57

System Monitor Test Panel ................................................................................ 24-122

24-58

Emergency Power Battery Pack......................................................................... 24-124

24-59

Electrical Power Control Panel.......................................................................... 24-128

24-60

Emergency Power Battery Schematic................................................................ 24-130

24-61

60-Hz Frequency Converter............................................................................... 24-132

24-62

60-Hz Receptacle............................................................................................... 24-134

24-63

Baggage Equipment Rack.................................................................................. 24-136

24-64

Cabin/Galley Master Switches .......................................................................... 24-136

24-65

60-Hz Power System Diagram .......................................................................... 24-138

24-66

Cabin/Galley Master Control............................................................................. 24-140

24-67

Ground Service Bus Diagram............................................................................ 24-144

24-68

Pilot Circuit-Breaker Panel................................................................................ 24-146

24-69

Copilot Circuit-Breaker Panel ........................................................................... 24-147

24-70

Left Electronic Equipment Rack Circuit-Breaker Panel.................................... 24-148

24-71

Right Electronic Equipment Rack Circuit-Breaker Panel ................................. 24-149


24-v



TABLES Table

Title

Page

24-1

Aircraft Electrical Component Locations .............................................................. 24-3

24-2

Electronic Equipment Rack Component Locations ............................................... 24-4

24-3

Power Distribution Box Component Locations ..................................................... 24-6

24-4

Main Electrical Power Control and Indication .................................................... 24-10

24-5

AC Source Range, Resolution, and Display Color Information .......................... 24-13

24-6

AC Source Status Box Colors .............................................................................. 24-14

24-7

DC Source Range, Resolution, and Display Color Information.......................... 24-17

24-8

DC Source Status Box Colors.............................................................................. 24-18

24-9

Left Bus Power Control Unit Power Inputs ......................................................... 24-42

24-10

Right Bus Power Control Unit Power Inputs....................................................... 24-42

24-11

Main and Essential Buses Priority Logic............................................................. 24-69

24-12

Component Locations .......................................................................................... 24-77


24-vii



CHAPTER 24 ELECTRICAL POWER

G EN PL #1 IL O DC #1 EN G O RV M T SE TE S 1 # SY HO T T BA

T BA

FF

O

ACEN G

INTRODUCTION The Gulfstream G500/G550 electrical power system (EPS) provides AC and DC power and a means of control, protection, and distribution of electrical power required for ground and in-flight operations of the aircraft. Primary electrical power is provided by the AC power system, comprised of two engine-driven integrated drive generators (IDG) and an auxiliary-power-unit-driven generator. DC power is provided by five transformer rectifier units (TRUs) and supplemented by two nickel-cadmium batteries. A DC-power ground service bus (GSB) is provided to allow routine aircraft servicing without powering other aircraft systems. The aircraft is equipped with a standby electrical power system that is powered by a hydraulic motor generator (HMG). The HMG can provide AC power if the APU generator and both main AC generators are not available. If all the previously mentioned power sources fail, emergency power battery packs provide power to equipment essential for safety of flight. There is also a 60-Hz converter to provide power to outfitter-installed equipment. See Figure 24-1 for component location.


24-1

24-2 MAIN WHEEL WELL Hydraulic Motor Generator

RIGHT ELECTRONIC EQUIPMENT RACK (REER)

Left Fire Handle Switch Right Fire Handle Switch Left Fuel Shutoff Switch Right Fuel Shutoff Switch

RIGHT ENGINE Right Integrated Drive Generator

Annun Lights Dim & Test Box Modular Avionics Unit #2 Dual Generic I/O 2 Module APU Generator Control Unit Right Generator Control Unit Right Bus Power Control Unit Right Power Distribution Box

RIGHT IDG OIL COOLER

Left Battery Right Battery AUX Power Relay Box Left Battery Charger Right Battery Charger 60 Hz Converter


14K

14J

14H 1A

COCKPIT OVERHEAD PANEL (COP) APU Generator Switch Right Generator Switch Left Generator Switch

LEFT IDG OIL COOLER LEFT ELECTRONIC EQUIPMENT RACK (LEER) Left Bus Power Control Unit Left Generator Control Unit Modular Avionics Unit #1 Dual Generic I/O 1 Module Left Power Distribution Box APU Ready to Load Relay

UNDERFLOOR COMONENTS

international

Figure 24-1. Equipment Locator

LEFT ENGINE Left Integrated Drive Generator Left IDG Oil Cooler

FlightSafety

Standby Junction Relay Panel Left Essential TRU Left Main TRU Aux TRU Emergency Inverter Right Main TRU Right Essential TrU HMG Generator Control Unit External AC CTA TRU Hall Effect Sensors (5) LMTAC—Left Main Tru Contactor RTAC—Right Main Tru Contactor

APU GENERATOR


PEDESTAL AREA

TAIL COMPARTMENT COMPONENTS



Table 24-1. AIRCRAFT ELECTRICAL COMPONENT LOCATIONS Standby junctions and relay panel (underfloor)

• AUX TRU AC contactors 1 and 2 • Left standby AC contactor • Right standby AC contactor

Underfloor LRUs

• • • • • • • • • •

Main wheel well LRU

• Hydraulic motor generator

Tail compartment LRUs

• • • • • • •

Auxiliary power relay box APU generator Left battery Right battery Left battery charger Right battery charger 60-Hz converter

Engine-mounted LRUs

• • • •

Left IDG Left oil cooler (not shown) Right IDG Right oil cooler (not shown)

Cockpit overhead panel

• Electrical power control panel • Controls switches • Indicators

External power receptacles

• AC • DC

Auxiliary TRU Left main TRU Left essential TRU Right main TRU Right essential TRU Emergency inverter Hydraulic motor generator control unit Hall effect sensors (5) LMTAC—Right main TRU contactor RMTAC—Righ main TRU contactor


24-3



Figure 24-2. Left and Right Electronic Equipment Racks

Table 24-2. ELECTRONIC EQUIPMENT RACK COMPONENT LOCATIONS

24-4

LEER LRUs

L EBATT L GCU L BPCU L PDB

REER LRUs

R E BATT R GCU R BPCU APU GCU R PDB




LEFT SIDE OF LEFT PDB—AC POWER

RIGHT SIDE OF LEFT PDB—DC POWER

LEFT POWER DISTRIBUTION BOX

LEFT SIDE OF RIGHT PDB—AC POWER

RIGHT SIDE OF RIGHT PDB—DC POWER

RIGHT POWER DISTRIBUTION BOX

Figure 24-3. Left and Right Power Distribution Boxes


24-5



Table 24-3. POWER DISTRIBUTION BOX COMPONENT LOCATIONS Left PDB

• • • • • • • • •

Left AC contactor Left AC crosstie contactor APU AC contactor Left main DC contactor Left ESS DC contactor Left emergency inverter DC contactor Left main DC crosstie contactor Left ESS DC crosstie contactor Left ESS DC battery contactor

Right PDB

• • • • • • • • • • • • •

Right AC contactor Right AC crosstie contactor External AC contactor Right main DC contactor Right ESS DC contactor Right emergency inverter DC contactor Right main DC crosstie contactor Right ESS DC crosstie contactor Right ESS DC battery contactor Galley DC contactor External DC contactor Auxiliary power contactor Hall effect sensor 5

24-6




BUSES AND RELAYS

tain three main power contacts. These main contacts carry the power from the AC sources or buses to the bus or component requiring power. Built into the contactor are auxiliary sets of contacts that are used to control other relays and provide relay position information to the component controlling the contactor (Figure 24-4).

The Gulfstream G500/G550 electrical power system diagram used in this lesson is a single-line drawing. This means that single lines are used for simplification. In the case of AC buses, the single bus shown actually represents three buses (phases A, B, and C). In addition, the contactors in the AC power system con-

035A2P5 P

AC OUTPUT PHASE C AC OUTPUT PHASE B AC OUTPUT PHASE A AC OUTPUT NEUTRAL

T3 T2 T1 N

3

3X162-1C

2

3X161-1B

1

3X160-1A

3X162-0C 3X161-0B 3X160-0A

A B C D T3

TB6 TB5 TB4

034TB3 3X200-1N

APU GENERATOR 031G3 LOC: APU COMP’T

N M L U T S R

E4131B

C2

C1

B2

B1

A2

A1

NC 7 NC 25

12

NO 24

13

T2 T1 035A1P1

CTA-2

EE

4

DD CC BB NN MM

1+

035A2-TB6

TO SHEET 3: EXTERNAL AC POWER FROM EXTERNAL AC CONTACTOR (R PDB)

035A2-TB5 035A2-TB4

C B A

X206-0C X205-0B X204-0A

AAC

19–

C2

C1

B2

B1

A2

A1

TB12 TB11 TB10

NC 7 NC 25

LEGEND RELAY COIL MAIN CONTACTS

GG HH

LH MAIN AC Ø C LH MAIN AC Ø B LH MAIN AC Ø A 035A1P2 J K T U

4 12 1+

AUXILIARY CONTACTS

LAXC

BB

19–

AA

LEFT POWER DISTRIBUTION BOX 035A1

LOC: L EER

Figure 24-4. Main and Auxiliary Power Contacts


24-7



POWER SOURCES Primary AC Power Primary AC power is provided by the enginedriven integrated drive generators (IDGs). An integrated drive generator (IDG) is mounted on, and driven by, each engine’s gearbox. Each IDG consists of a hydromechanical constantspeed drive (CSD) and an oil-cooled generator. The CSD portion of the IDG converts variable-input speed from the engine-driven gearbox to a constant speed to drive the generator portion of the IDG. Each IDG is controlled by a dedicated generator control unit (GCU), which provides voltage regulation, over/undervoltage protection, over/underfrequency protection, and speed control. The IDGs are rated at 40 KVA and outputs threephase, 400 Hz, 115 VAC power.

Auxiliary AC Power An APU-driven oil-cooled generator provides 115 VAC, 400 Hz, three-phase power at a capacity of 40 KVA and is able to power the same buses as the engine-driven IDGs. The APU generator GCU provides voltage regulation, over/undervoltage protection, and over/underfrequency protection.

to a 28 VDC nominal output. The output voltage of TRUs is unregulated and therefore dependent on loads. Each TRU operates in a normal range of 29 to 26 VDC with loads in the range of 20 to 250 amps, respectively.

Main Battery Power Main battery power is provided by two 21-cell, 24 VDC nickel-cadmium batteries rated at 53 amp-hours. These batteries provide power for APU starting, running the auxiliary hydraulic pump, and powering the essential DC buses, when required.

Battery Chargers Two battery chargers are provided to recharge the main batteries. Each battery charger provides temperature-compensated charging in the charge mode, or an output of constant 28.75 VDC at up to 50 amps in the transformer-rectifier mode of operation.

Emergency Inverter An E-inverter provides 115 VAC, 400 Hz, 1 KVA, single-phase power for phase A of the ESS AC bus. The E-inverter can be powered from the left or right essential (L ESS or R ESS) DC bus.

External AC Power External AC power provides three-phase AC power at 115 VAC, 400 Hz to power main AC buses via the AC crosstie bus. External AC power protection is provided by the left bus power control unit. This unit provides overcurrent protection, over/undervoltage protection, over/underfrequency protection, phase sequence protection, AC interlock protection, and failed current transformer assembly protection.

DC Power Five transformer rectifier unit (TRUs), left and right main, left and right essential, and auxiliary TRU, provide DC power. These TRUs convert three-phase, 115 VAC, 400 Hz power

24-8

External DC Power External DC power provides 28 +4, –7 VDC through an external DC power receptacle, located just forward of the external AC power receptacle. This power source can power the main and essential DC buses through the external DC contactor (EDC), the battery tie bus through the auxiliary power contactor (APC), and the ground service bus through ground service bus contactors (GSBCs 1 and 2). External DC power can also be used for APU starts and to run the auxiliary hydraulic pump when battery power is not available. The right bus power control unit provides overcurrent protection, over/undervoltage protection, polarity protection, and DC interlock protection.




Ground Service Bus (GSB)

60-Hz Power System

The ground service bus is provided for normal aircraft servicing without having to needlessly power avionics equipment. Power for the GSB can be provided by the right main DC bus, an external DC power supply, or from the right main aircraft battery.

A 60-Hz power system provides 115 VAC, 60 Hz power for outfitter-installed equipment. A 3.5 KVA, 400 to 60 Hz frequency converter uses main AC bus power to produce singlephase, 115 VAC, 60 Hz power. On the ground, power for the 60 Hz system can also be obtained from an external power source, using an external 60 Hz power receptacle located in the tail compartment.

Standby Electrical Power System (SEPS) In the event primary or auxiliary power is not available, the Gulfstream G 5 0 0 / G 5 5 0 i s equipped with a standby electrical power system. This system is powered by a hydraulic motor generator (HMG), which utilizes left hydraulic system fluid and pressure or power transfer unit (PTU) pressure to provide AC power to the left and right standby AC buses and the auxiliary TRU. The auxiliary TRU provides DC power for the L ESS and R ESS DC buses when selected.

NOTES

The HMG is rated at 10 KVA and provides three-phase, 400 Hz, 115 VAC power. It is controlled by a dedicated generator control unit (GCU). (HMG GCU is not interchangeable with the APU or IDG GCUs.) Voltage, frequency, and loads are monitored and controlled by the GCU.

Emergency Power Battery Packs If all sources of power are lost, two emergency power battery packs, rated at 9 amp hours, provide up to 45 minutes of electrical power to the left and right emergency (L/R EMER) DC buses, essential flight instruments (ESS FLT INST) bus, and inertial reference units to support equipment essential for safety of flight. The left emergency power battery pack also provides power for the shutdown cycle of the central maintenance computer module.


24-9



Table 24-4. MAIN ELECTRICAL POWER CONTROL AND INDICATION Left bus-tie switch

AUTO (extended position): Left main AC bus is powered as a function of priority. ISLN (depressed position): Left main AC bus is isolated from lower priority AC power sources by inhibiting the closing of the LAXC.

Right bus-tie switch

AUTO (extended position): Right main AC bus is powered as a function of priority. ISLN (depressed position): Right main AC bus is isolated from lower priority AC power sources by inhibiting the closing of the RAXC.

Left generator switch ON (depressed position): Allows left IDG to excite and automatically close LAC if power quality is acceptable. OFF: Illuminates amber when switch is depressed but IDG is not producing acceptable power. Note: Switch will only illuminate OFF if it is in the depressed position. If the switch is extended, it will remain blank.

Right generator switch

ON (depressed position): Allows right IDG to excite and automatically close RAC if power quality is acceptable. OFF: Illuminates amber when switch is depressed but IDG is not producing acceptable power. Note: Switch will only illuminate OFF if it is in the depressed position. If the switch is extended, it will remain blank.

APU generator switch

ON (depressed position): Allows APU generator to excite and automatically close AAC if power quality is acceptable. Note: Switch will illuminate only if it is in the depressed position and good APU power is available. If the switch is extended, it will remain blank.

External power switch

AVAIL (extended position): Illuminates if AC or DC external power is available and power quality is acceptable and battery power is selected ON. ON (depressed position): Illuminates when switch is depressed while AVAIL is illuminated. Causes AVAIL indication to extinguish if another source of acceptable external power is not available.

AC/DC reset switch

AC: Illuminates when an inadvertent parallel trip is detected or when L or RAC or L or RAXC are locked out due to a protective trip. Cycling the switch resets the lockout and allows the tripped contactor to reclose. DC: Illuminates when a DC contactor is opened due to a protective trip. Cycling the switch resets the lockout and allows the tripped contactor to reclose if conditions are normal.

Emergency inverter switch

AUTO (extended position): Inverter is powered by the left or right essential DC bus as a function of priority. OFF (depressed position): LEIDC and REIDC are prevented from energizing, which will inhibit the inverter from receiving power.

Ground service bus indicator

24-10

Illuminates ON when the ground service bus is powered by either the right battery or an external DC power cart.




CONTROLS AND INDICATIONS Electrical Power Control Panel (EPCP) Normal control of the EPS is accomplished through switches located on the electrical power control panel (EPCP) (Figure 24-5) in the cockpit overhead panel. The switches provide manual and automatic control of the EPS. The switches have color-coded legends to indicate system status to the aircrew. The colors are blue for advisory indications, green for normal in-flight configurations, and amber for abnormal in-flight configurations. Main battery volt/amp meters are also located on the EPCP. The emergency power control panel section of the EPCP provides ON, ARM, and OFF controls and indicators for the emergency batteries.

Main electrical power control and indication are provided through the left bus-tie switch, right bus-tie switch, left generator switch, right generator switch, APU generator switch, external power switch, E-inverter switch (alternate action switches), and AC/DC reset switch (momentary switch). Refer to Table 24-4 for details on the individual switch operation. There is also a ground service bus indicator. The main batteries control panel is located just below the electrical power control panel (See Figure 24-5). The left and right main batteries ON–OFF switches provide control of the m a i n b a t t e r y p o w e r. L e f t a n d r i g h t VOLTS/AMPS liquid crystal displays (LCDs) indicate the charge level of each battery. A negative indication “–” (minus sign) displayed in the amp meter indicates the main batteries are online and discharging.

The standby electrical power system (SEPS) control panel section of the EPCP provides control of the HMG system through MASTER, L ESS, and R ESS switches.

Figure 24-5. Electrical Power Control Panel


24-11



PFD

MAP

1/6

ON

ON

ON

SYS

ON

SENSOR

FLT REF

TRS

NAV

ON

ON

ON

ON

TEST

CHKLST

HUD

ON

ON

ON

BARO

2/3

BRT

BRAKES AC POWER DC POWER FUEL TCAS

SUMMARY ECS/PRESS HYDRAULICS WAYPT LIST NEXT

PUSH STD

SET

Figure 24-6. AC POWER Synoptic Page (Sheet 1 of 2)

24-12




The TRU control panel is located in the cockpit overhead panel (COP) on the upper right side of the EPCP (See Figure 24-5). The guarded switches are not illuminated when they are in their normal extended position. The TRU L MAIN switch, when depressed, energizes the LMTAC contactor. This allows the left main TRU to be powered by the right main AC bus. The switch will illuminate amber R AC to annunciate the TRUs input power source. The TRU R MAIN switch, when depressed, energizes the RMTAC contactor. This allows the right main TRU to be powered by the left main AC bus. The switch will illuminate amber L AC to annunciate the TRUs input power source.

The electrical power system can be displayed graphically on the display units to provide a synopsis of the system. These synoptic pages are accessed through the display controller by selecting “SYS” 1/6, 2/3 or using the CCD. The AC POWER synoptic page is a 2/3 page and is accessed by selecting the line-select key (LSK) marked “AC POWER.” With the AC POWER page displayed, AC power source information is available including left IDG (L GEN), APU generator (APU GEN), external AC (EXT AC), right IDG (R GEN), and hydraulic motor generator (HMG) (Figure 24-6). See Table 24-5 for individual AC power source range, resolution, and display color information and Table 24-6 for AC power source status box colors.

AC POWER Synoptic Page Table 24-5. AC SOURCE RANGE, RESOLUTION, AND DISPLAY COLOR INFORMATION POWER SOURCE AC power source voltage readouts

RANGE RESOLUTION 0 to 150 VAC 1 volt • • • • • •

READOUT COLORS Amber < 105 VAC White ≥ 105 through ≤ 122 VAC Amber > 122 VAC Amber dashes—invalid source White dashes—respective switch off and no DAU failure White dashes—take priority over amber

AC frequency readouts 0 to 512 Hz

1 Hz

• • • • • •

Amber < 380 Hz White ≥ 380 Hz through ≤ 420 Hz Amber > 420 Hz Amber dashes—invalid sensor White dashes—respective switch off and no DAU failure White dashes take priority over amber

AC percentage load readouts

1%

• • • • •

White ≤ 100% Amber > 100% Amber dashes—invalid sensor White dashes—respective switch off and no DAU failure White dashes take priority over amber

Auxiliary TRU 0.0 – 50.0 • Only shown VDC powered when it is powered by the HMG • Only DC power source displayed on AC synoptic page

0.1 VDC

• • • • • •

Amber < 22.0 VDC White ≥ 22.0 VDC through ≤ 32.5 VDC Amber > 32.5 VDC Amber dashes—invalid sensor White dashes—HMG switch off and no DAU failure White dashes take priority over amber AC

Auxiliary TRU percentage load readouts

1%

• Amber dashes—invalid AUX TRU load sensor • White dashes—HMG switch off and no DAU failure • White dashes take priority over amber

0 to 200%

0 to 200%


24-13



PFD

MAP

1/6

ON

ON

ON

ON

SENSOR

FLT REF

TRS

NAV

ON

ON

ON

ON

TEST

CHKLST

HUD

ON

ON

ON

SYS

BARO

2/3


BRT

0

PUSH STD

SET

0

0 0


0 40

0

Figure 24-6. AC POWER Synoptic Page (Sheet 2 of 2)

24-14




Table 24-6. AC SOURCE STATUS BOX COLORS SOURCE BOX

STATUS BOX

AC generators

• White when respective switch is off, or switch on and engine not running • Green when respective switch is on and normal voltage • Amber for other conditions

External AC power

• White when respective switch is off and external power not applied • Green when respective switch is on or off and normal voltage connected • Amber for other conditions

HMG power

• White when HMG switch is off • Green when HMG switch is on with normal voltage and load • Amber for all other conditions

Left and right main AC bus

• • • •

White bus name label White when bus is not selected to be powered Green when the bus is powered and voltages are correct Amber for all other conditions

Essential AC bus

• • • •

White bus name label White when bus is not selected to be powered Green when the bus is powered and voltage is correct Amber for all other conditions

Left and right standby AC bus

• • • •


Left and right essential DC bus

• • • •

White bus name label White when bus is not selected to be powered Green when the bus is powered and voltages are correct Amber for all other conditions

E-inverter power

• White when the bus is not selected to be powered • Green when the bus is powered

Feeder Lines

• • • •

Always displays 22 lines White when not selected to be powered Green when powered and voltages are correct Amber for any other condition


24-15



PFD

MAP

1/6

ON

ON

ON

ON

SENSOR

FLT REF

TRS

NAV

ON

ON

ON

ON

TEST

CHKLST

HUD

ON

ON

ON

SYS

BARO

2/3

BRT



PUSH STD

SET

Figure 24-7. DC POWER Synoptic Page (Sheet 1 of 2)

24-16




DC POWER Synoptic Page The DC POWER synoptic page is a 2/3 page and is accessed through the display controller 2/3 menu and then selecting the DC POWER LSK or with the CCD. With this page displayed, left essential TRU, left main TRU, right main TRU, right essential TRU, auxiliary TRU, left battery, external DC, and right battery power source information is available (Figure 24-7). Tables 24-7 and 24-8 provide range information, resolution, and readout colors for the various DC power sources. Table 24-7. DC SOURCE RANGE, RESOLUTION, AND DISPLAY COLOR INFORMATION POWER SOURCE

RANGE

RESOLUTION

READOUT COLORS

TRUs and external DC voltage readouts

0.0 to 50.0 VDC

0.1 VDC

• Amber < 22.0 VDC • White ≥ 22.0 VDC through ≤ 30.0 VDC (32.5 for AUX TRU) • Amber > 30.0 VDC (32.5 for AUX TRU) • Amber dashes—invalid sensor

Main batteries voltage readouts

0.0 to 50.0 VDC

0.1 VDC

• Amber < 22.0 VDC • White ≥ 22.0 VDC through ≤ 36.0 VDC • Amber > 36.0 VDC • Amber dashes—invalid sensor

TRUs and external DC percentage load readouts

0 to 200%

1%

• White ≤ 100% • Amber > 100% • Amber dashes—invalid sensor

Batteries DC current readouts

–400 to +400 amps 1 amp

Note: There is a 30-second delay changing from white to amber when < 22.0 VDC.

• Digital readout is always white • Amber dashes—invalid sensor


24-17



PFD

MAP

1/6

ON

ON

ON

ON

SENSOR

FLT REF

TRS

NAV

ON

ON

ON

ON

TEST

CHKLST

HUD

ON

ON

ON

SYS

BARO

2/3

BRT



PUSH STD

SET

Figure 24-7. DC POWER Synoptic Page (Sheet 2 of 2)

24-18




Table 24-8. DC SOURCE STATUS BOX COLORS SOURCE BOX

STATUS BOX

TRU power

• White when input power is not available • Green when respective input power is available and the TRU provides normal voltage and load • Amber for all other conditions

External DC power

• White when external DC power is not available • Green when external DC power is available and voltage is normal • Amber for other conditions

Main battery power

• Green when its respective battery control switch is on or off and battery voltage is normal • Amber for other conditions with a 30-second delay in change to amber for < 22.0 VDC

L-R main DC bus

• • • •


L-R essential DC bus

• • • •


Ground service bus

• White when not powered • Green when powered with acceptable voltages • Amber for all other conditions

Left and right main AC bus

• White when bus is not selected to be powered • Green when the bus is powered and all phrase voltages are correct • Amber for all other conditions

Feeder lines

• • • •

Always displays 27 feeder lines White when not selected to be powered Green when powered and voltages are correct Amber for any other condition


24-19



PFD

MAP

1/6

ON

ON

ON

ON

SENSOR

FLT REF

TRS

NAV

ON

ON

ON

ON

TEST

CHKLST

HUD

ON

ON

ON

SYS

BARO

2/3

BRT



Figure 24-8. Summary Synoptic Page

24-20


PUSH STD

SET



SYSTEM Synoptic Page 1/6

NOTES

The 1/6 AC/DC system synoptic page displays AC and DC power information on the selected DU window (Figure 24-8) when accessed by the display controller or the cursor control device. Selecting the AC/DC PWR line select key on the display controller opens the 1/6 AC/DC synoptic window on the selected area of the DU. With this page displayed, GEN, HMG EXT power BATT and TRU information is available.

CAS Messages The monitor warning function monitors electrical power system information provided by the bus power control units and displays various advisory and caution messages to the crew (Figure 24-8).


24-21




MAU No.1 Duel Generic I/O Slot 9/10

MAU No.2 Duel Generic I/O Slot 7/8

ARINC 429 No.1

ARINC 429 No.1

ARINC 429 No.2

ARINC 429 No.2

LEFT GENERATOR CONTROL UNIT

APU GENERATOR CONTROL UNIT EPS Serial Data Bus A EPS Serial Data Bus B ARINC Data Bus

Figure 24-9. Bus Diagram

24-22


RIGHT BUS POWER CONTROL UNIT

APU POWER READY SIGNAL (APU ECU)

APU GEN SWITCH (FOP)

RIGHT GENERATOR CONTROL UNIT



Inter-LRU Serial Link Communications Bus

NOTES

The GCUs and BPCUs communicate with each other over redundant inter-LRU serial link data buses. Both buses are capable of transmitting identical data. Only one of the two buses is active at any given time. Data bus No. 1 is the primary, and data bus No. 2 is the backup. If bus No. 1 fails, bus No. 2 will become active and take over data transmission. The electrical and protocol aspects of these communications buses is compatible with MIL-STD-1553B requirements. Both BPCUs serve as controllers of the redundant buses. GCU CAS data and data to be used to do AC power system no-break power transfers are transmitted over these buses.

L BPCU The left BPCU transmits information via ARINC 429 to MAU No. 1, DGIO 1, slot 9/10. The DGIO places this information on the virtual backplane bus to be utilized internally and is also shared over ASCB with MAUs No. 2 and No. 3.

R BPCU The right BPCU transmits information via ARINC 429 to MAU No. 2, DGIO 2, slot 7/8. The DGIO places this information on the virtual backplane bus to be utilized internally and is shared over ASCB with MAUs No. 1 and No. 3.

Parameters and Faults The CMC receives parameter and fault information from the L BPCU via the virtual backplane bus in MAU No. 1. The CMC receives R BPCU parameter and fault information from MAU No. 2 via ASCB to MAU No. 1’s virtual backplane bus.


24-23



LEFT IDG

(LEFT PDB)

PILOT CB PANEL

LEFT MAIN AC BUS E-INV ESSENTIAL AC BUS

TO AC LOADS LEER CB PANEL

APU GEN

COPILOT CB PANEL EXT AC

(RIGHT PDB) RIGHT MAIN AC BUS

TO AC LOADS

RIGHT IDG

REER CB PANEL

Figure 24-10. AC Power System Overview

24-24




AC POWER SYSTEM

NOTES

INTRODUCTION The AC power system provides source, control, protection, and distribution of 115 volt, three-phase, 400-Hz electrical power (Figure 24-10). This section discusses the main enginedriven generators system, APU-driven generator system, emergency inverter system, external AC power, AC source distribution, and AC power distribution. The first system covered is the main engine generator system, the integrated drive generators (IDGs).


24-25



ENGINEMOUNTED HP GEARBOX

CSD HYDROMECHANICAL TRANSMISSION

OILLUBRICATED SPLINE

SERVO VALVE

GENERATOR PMG

EXCITER ARMATURE

RDA

GEN ROTOR

GEN STATOR

T1

EXCITER FIELD

T2

T3

N

POR

140–160 VDC 70 to 100 VAC L-L

CTA GCU

Figure 24-11. Integrated Drive Generator

RR072

24-26




ENGINE-DRIVEN AC GENERATION SYSTEM The components, controls, and indicators of the main AC power system are the left and right generator control switches, left and right AC contactors, left and right bus power control units, current transformer assemblies, left and right generator control units, integrated drive generators, and IDG oil coolers.

Integrated Drive Generator System The integrated drive generators (IDGs) are the engine-driven power sources (Figure 2411). One IDG is mounted on the rear of each engine reduction gearbox by means of a bracketed V-band clamp. Each IDG consists of a hydromechanical constant-speed drive (CSD), mounted in line with an oil-cooled generator. The CSD portion of the IDG converts variableinput speed from the engine reduction gearbox to a constant speed to drive the generator portion of the IDG. The generator is rated at 40 KVA and provides three-phase, 400-Hz, 115 VAC power at the point of regulation.

The hydraulic servo valve is controlled electronically by the speed servocontrol in the generator control unit (GCU) to maintain a generator speed of 12,000 revolutions per minute. The hydraulic servo valve is also used during the no-break power transfer operation between an IDG power source and non-IDG power source such as the APU generator or external power. The GCU will send a signal to the servo valve to change the position of the variable displacement hydraulic unit. The variable displacement hydraulic unit will either increase or decrease the speed of the generator so that it operates at the same frequency/speed as the non-IDG power source. The speed summing is done in the differential. The differential adds or subtracts the trim speed. The output of the differential is a constant speed. The differential is connected to the generator rotor. The generator rotor then turns at a constant speed.

The IDG converts the varying engine reduction gearbox speeds of 4,666 to 8,438 revolutions per minute to a constant generator speed of 12,000 revolutions per minute. The generator operates at a constant output frequency of 400 ±4 Hz. The variable-speed shaft power is provided to the IDG input shaft and coupled to the carrier shaft of the CSDs differential. The variable displacement hydraulic unit is hydraulically coup l e d t o a fi xe d h y d r a u l i c u n i t , w h i c h i s mechanically connected to the input ring gear of the differential. The speed and direction of the input ring gear are controlled by the displacement of the variable hydraulic unit. The displacement of the variable hydraulic unit is controlled by a hydraulic servo valve. The hydraulic servo valve acts on a control cylinder connected to the variable hydraulic unit.


24-27



VENT VALVE

ELECTRICAL CONNECTOR

OIL IN PORT ELECTRICAL TERMINAL COVER

PRESSURE FILL FITTING WITH CAP

NOT A HANDLE

OVERFULL

ADD OIL

ALTERNATE PRESSURE FILL FITTING LOCATION

SCAVENGE FILTER COVER

OVERFLOW DRAIN PLUG CASE DRAIN PLUG

OIL OUT PORT

Figure 24-12. IDG Assembly

24-28


DIFFERENTIAL PRESSURE INDICATOR



Vent Valve

CAUTION This is a single LRU. Do not attempt to separate CSD from GEN.

As the permanent magnet generator (PMG) rotor turns, it induces an AC voltage in the three-phase windings of the permanent magnet generator stator. This AC voltage is supplied through a connector on the IDG housing and aircraft wiring to the generator control unit (GCU) where it is rectified into DC voltage. The rectified DC voltage from the PMG is the primary power source for the GCU. The GCU voltage regulator uses this DC voltage to control the current to the generator exciter field stator.

A vent valve is provided to depressurize the IDG prior to oil servicing or other IDG maintenance. This will prevent oil spray when the pressurized system is opened. The vent valve is a spring-actuated valve located on the IDG input housing that is designed to maintain internal pressure in the housing. A biasing spring, which acts on a piston, closes the vent valve. This action lets the air pressure build up in the IDG case. The built-up internal pressure is relieved when the vent valve is pushed.

NOTES

The magnetic field produced by the DC in the windings of the exciter stator field induces an AC voltage in the rotating windings of the exciter armature (rotor). This three-phase AC voltage is converted to DC voltage by the rotating diode rectifier assembly on the armature. The resultant DC voltage is supplied directly to the windings of the main field to produce a rotating magnetic field. As the magnetic field rotates, it induces an AC voltage in the windings of the main generator stator. The windings of the main generator stator are connected to the terminal block located on the IDG housing. Aircraft wiring supply IDG power to the AC load buses in the power distribution boxes (PDBs).

IDG Assembly The AC generator contained in the IDG (Figure 24-12) has a three-stage, brushless, rotating spray oil-cooled generator. The generator rotor assembly contains an exciter rotor, main field rotor, and diode rectifier assembly. The two rotors are mounted on a common shaft supported by a roller bearing at the drive end and ball bearing at the nondriving end. The permanent magnet generator (PMG) stator and rotor are located in the CSD housing, and the exciter stator, main field stator, and generator current transformer assembly are located in the generator housing.


24-29



Thermal Disconnect A thermal disconnect will decouple the engine reduction gearbox rotational power from the IDG in the event of a high oil temperature. The thermal disconnect components consist of a “dog-tooth” splined clutch (which has a splined clutch and an input shaft), a solder ring, and a disengagement spring. The solder ring acts as a spacer to keep the dog teeth of the splined clutch engaged with the dog teeth on the input shaft. This transmits the input shaft power to the IDG. An overtemperature condition in the IDG or external oil cooler circuit resulting in a rise of IDG internal temperature to 354 ±3°F (179 ±2°C) will cause the solder ring to melt. This liquified solder flows into a cavity, where it stays.

yond the face of the indicator to provide visual indication of a clogged filter. A temperature lockout feature is incorporated to prevent a false pressure differential indication during startup with cold oil. This is done by a bimetal element that locks the pop-out button in place during low-temperature operation but expands to permit actuation during the normal operating temperature range if the filter becomes clogged.

When the solder ring melts, the spring force will separate the input clutch from the input shaft. The spring force is sufficient to keep the input clutch from reengaging. When the dog teeth are fully moved apart, the input shaft continues to turn freely and the IDG will slow to a stop. This type of disconnect is automatic and requires no crewmember involvement.

Input Shaft Shear Section The IDG has an input shaft section with a reduced diameter that will shear to protect the IDG when the input torque reaches 3,400 +250 pound-inches. After the shaft shears, there is no damage to the gearbox output spline and no contamination of the gearbox oil system. The gearbox can continue to operate.

Differential Pressure Indicator A differential pressure indicator (DPI) is provided to indicate a clogged scavenge filter. The DPI is located in the IDG scavenge oil system and senses pressure upstream and downstream of the filter element. The DPI housing contains a spring-loaded, pop-out button, which is attracted to a magnet during normal operation. When the pressure drop across the filter is more than the bias force of the spring (60 psid), the magnet is forced away from the indicator button, which breaks the magnetic attraction. The pop-out button then extends be-

24-30


NOTES



NOTES

IDG Oil Supply System The purpose of the oil supply system is to provide IDG with pressurized lubrication and cooling oil during operation. The integrated drive generator oil supply system has a rotating deaerator assembly, charge pump, charge relief valve, generator scavenge pump, scavenge pump, scavenge relief valve, inversion scavenge pump, scavenge filter, and a mechanical differential pressure indicator. These components are located in the cored housing of the IDG. Normal charge pressure is regulated by the charge relief valve. The scavenge relief valve limits the discharge pressure of the scavenge pump due to blockage in the scavenge filter or external circuit. The externally cooled lubrication oil enters the IDG through the oil-in port and flows through the rotating deaerator to the charge pump. Non-deaerated oil lubricates the input seal. The gear-driven rotating deaerator separates the captured air from the oil as it turns. The oil then moves through the g e a r - d r iv e n , positive-displacement, vane charge pump to lubricate the internal IDG components. The charge relief valve is a spring-loaded valve that regulates operating pressure of the charge oil circuit. Normal charge pressure is regulated at 240 to 280 pounds per square inch. The pressurized oil flows to the generator, pump, motors, servo valve, control piston, and lubricating system. The gear-driven, positive-displacement vane generator and scavenge pump has two chambers. One chamber scavenges oil from the generator, and the other chamber pump scavenges oil from the IDG sump to the scavenge filter. Also, a gear-driven, positive-displacement vane inversion pump scavenges oil from the IDG sump during negative “G” maneuvers. The scavenge relief valve limits the discharge pressure of the scavenge pump. If blockage in the scavenge filter or external circuit causes the oil pressure to rise to a nominal 350 pounds per square inch, the scavenge relief valve opens and ports oil directly to the charge pump inlet. The scavenge pump moves the oil through the scavenge filter and through the oil-out boss out into the external system for cooling. FOR TRAINING PURPOSES ONLY

24-31



VENT VALVE

STEP ONE

PRESSURE FILL FITTING DUST CAP

• RELIEVE CASE PRESSURE BY DEPRESSING VENT VALVE FOR APPROXIMATELY 5 SECONDS. • REMOVE OVERFLOW DRAIN PLUG. • SOME OIL MAY COME OUT OF THE OVERFLOW DRAIN WHEN THE PLUG IS REMOVED.

NOT A HANDLE

OVERFULL

ADD OIL

OVERFLOW DRAIN PORT CONTAINER

PRESSURE FILL HOSE

STEP TWO

PRESSURE FILL FITTING

ADD OIL

NOT A HANDLE

OVERFULL

• • • •

ATTACH A SHORT OVERFLOW DRAIN HOSE. REMOVE PRESSURE FILL DUST CAP. ATTACH PRESSURE FILL HOSE. PUMP OIL INTO THE IDG UNTIL ONE QUART OF OIL FLOWS FROM THE OVERFLOW DRAIN HOSE.

OVERFLOW DRAIN HOSE

CONTAINER

STEP THREE

PRESSURE FILL FITTING DUST CAP

• DISCONNECT PRESSURE FILL HOSE. • INSTALL PRESSURE FILL DUST CAP. • WHEN DRAINAGE SLOWS TO DROPS, DISCONNECT THE OVERFLOW DRAIN HOSE. • INSTALL OVERFLOW DRAIN PLUG WITH NEW O-RING INSTALLED. • CHECK SIGHT GLASS FOR CORRECT OIL LEVEL.

ADD OIL

OVERFLOW DRAIN PORT

NOT A HANDLE

OVERFULL

OVERFLOW DRAIN HOSE CONTAINER

NOTE: THIS IS A MAINTENANCE ADVISORY ONLY. REFER TO THE APPLICABLE GULFSTREAM MAINTENANCE MANUAL FOR SPECIFIC INSTRUCTIONS.

Figure 24-13. IDG Servicing

24-32




IDG Servicing Features

Overflow Standpipe

The IDG has several features to achieve the proper oil level (Figure 24-13). These enable a simple go/no-go approach to servicing. When properly serviced, the system holds a minimum of 3.4 quarts and a maximum of 4.15 quarts.

An internal overflow standpipe (Figure 2414) drains out the overflow drain port and establishes the proper oil level for the system. This is done without the need for judgment by the mechanic based upon the appearance of oil level is a sight glass. The concept is similar to the standard fill-overflow port for most automotive differential gearboxes. The automotive differential oil level is correct when oil stabilizes at the overflow port. This also applies to the IDG.

Oil Level Sight Glass An oil level sight glass (Figure 24-14) is provided to show when the oil level is below the normal IDG oil level or if there is an overfill condition. When the oil level is above or drops below the designated normal IDG oil level area (gray band), oil servicing is required. When the oil level is within the gray band, oil servicing is not required.

OVERFULL MAX ACCEPTABLE OIL LEVEL

SIGHT GLASS STANDPIPE OVERFULL

GRAY BAND

MINIMUM ACCEPTABLE OIL LEVEL

ADD OIL

ADD OIL

Figure 24-14. IDG Sight Gage and Standpipe


24-33



BYPASS CASE

OIL INLET

BYPASS VALVE

OIL OUTLET

A

IDG OIL COOLER

AIRSIDE COOLING FINS

A

Figure 24-15. IDG Oil Cooler

24-34




NOTES

IDG Oil Cooler The purpose of the oil cooling system is to provide externally cooled oil for the operation of the IDG. Cooling system components include the heat exchanger, bypass valves, interconnecting tubing, and fittings. The oil in the IDG is kept in the specified operating temperature range by a surface-fin-type air-oil heat exchanger (oil cooler) and is filtered before the oil enters the heat exchanger (Figure 24-15). The cooler is located in the engine bypass duct and is connected to the IDG oilin and oil-out ports with oil lines. The oil cooler contains a pressure-relief valve that protects the cooler from overpressurization and is set at a nominal 50 psid. During bypass operation, two-thirds of the cooler is bypassed. The external cooling system provides a pressure drop of 90 psid maximum during normal operation and will decongeal without causing an IDG overtemperature during 13,000 centistokes starting conditions. The volume of the system is approximately 1,200 cubic centimeters. The normal external system limits for the normal IDG oil-in temperature is 150 to 200°F (65.5 to 93°C) and normal oil-out temperature is 168 to 221°F (75.5 to 105°C). The oil that is heated by IDG operation is pumped from the IDG by the internal supply pump through a scavenge filter to the external oil circuit. The heat is removed from the oil circuit as it flows through the oil coolers. The cooled oil is then returned to the IDG through the oil-in port and circulated through the IDG hydraulic circuit by the charge pump inside the IDG. The IDG and external oil circuit are initially filled with oil pumped into the IDG through the pressure fill port. The oil pumped into the IDG flows through the scavenge filter and then through the external oil circuit and into the IDG case. Air in the oil circuit is forced out ahead of the oil and escapes out the opened overflow drain port. When the oil level in the IDG sump stabilizes at the top of the overflow standpipe, the IDG system is correctly filled.


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GENERATOR CONTROL UNITS (GCUs) There are three generator control units (GCUs) located in the electronic equipment racks (EERs). Each generator control unit operates in conjunction with its respective IDG, APU generator, line CTA, and cockpit overhead panel to supply system protection and control functions, and built-in test capability for the AC electrical generating equipment. A GCU is provided for the left IDG, right IDG, and APU generator. The GCUs are physically identical and interchangeable. The GCUs are pin programmable; therefore they know which position they are in and which generator they are controlling because of the wiring in their mounting rack. Input power sources to each of the GCUs are as follows:

GCU can be reset from a fail-safe condition when power is removed from the GCU or when the generator control switch on the cockpit overhead panel (COP) is operated from OFF to ON. This reset causes the software to reinitialize. If the conditions that caused the microprocessor fail-safe are still present, the microprocessor will fail-safe again. The GCU performs the following control functions: • Generator main contactor control • DC control power • POR voltage control • Frequency control (IDG only) • Generator excitation

Contactor Control

• L GCU—Left IDG PMG voltage and left essential DC bus

The GCU controls the contactors and relays in the aircraft system as follows:

• R GCU—Right IDG PMG voltage and right essential DC bus

• Main generator control relay (GCR/AGCR) control (internal to the GCU)

• APU GCU—APU generator PMG voltage and left essential DC bus The GCUs provide voltage regulation, frequency regulation (IDGs only), control, system protection, and monitor the frequency reference unit (FRU) signals that are transmitted from the BPCU to the GCU. It does this to provide IDG frequency and phase angle control for no-break power transfers (NBPTs) to and from the IDG power sources. In addition, built-in test (BIT) capability is included for on-aircraft fault isolation via the CMC.

Control Functions The GCUs work with the aircraft electrical system to provide quality power at the point of regulation (POR). If an inoperative GCU creates an unpowered load bus, it does not affect the control and protective functions of the operative power sources. The system continues to operate without loss of protective or control functions of the operative power sources, even if any combination of one or two inoperative power sources exists. The

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• Generator AC line contactors (LAC, RAC, and AAC) • Crosstie contactors (LAXC, RAXC) lockout capability • Essential AC contactor (LEAC, REAC) lockout capability Control commands for the contactors are started from a switch on the COP. A hardwired switch status line supplies +28 VDC when the associated contactor switch is depressed. The auxiliary contacts are used to provide two complimentary inputs to the status switch position.

DC Control Power Inputs Each GCU receives power from two independent sources. 28 VDC essential DC bus power is routed to the GCUs from circuit breakers on the EERs. This voltage “wakes up” the GCU and starts communications with the BPCU over the data bus. It also supplies power to operate the GCU and control the IDG, with the




exception of the exciter field and servo valve current. The primary power source for the GCUs, however, is rectified PMG voltage from the generator. The three-phase, 1,200 Hz, AC (70 to 100 VAC L–L) PMG voltage is rectified by the GCU to produce the primary GCU power. The GCU will not excite the generator without the PMG voltage input. The GCU provides the power to control the relay that illuminates the AC reset light. In addition, the GCU provides the DC control output for the following control switches: • Left and right IDG switch (L/R GEN) • APU switch (APU GEN)

The GCU, with a frequency reference unit selected by the left BPCU (LBPCU) or an internal 400 Hz reference, provides the IDG frequency and phase angle control. If a frequency reference unit is unavailable, the GCU uses an internally generated reference for frequency control and phase control. The IDG frequency will stay within 0.3 Hz of the reference averaged over one minute. The APU GCU monitors the frequency of the APU generator, which is driven at a constant 12,000 rpm by the APU. If the APU generator speed is out of limits, this condition is seen by the APU GCU as a frequency fault.

Generator Excitation Control

• Fire switch • Engine fuel control switch The GCU also provides the DC control output to power the L/RAC and AAC line contactors.

During fault conditions, the excitation level to the generators is automatically controlled to limit current and/or power out of the associated generator.

POR Voltage Control The voltage regulation portion of the GCU controls the IDG and APU generator voltage at the point of regulation (POR). The voltage regulator senses the voltage at the point of regulation and compares this against a reference voltage. If a difference exists, it will adj u s t t h e g e n e r a t o r e x c i t e r fi e l d c u r r e n t accordingly to maintain a constant voltage at the point of regulation.

Frequency Control The frequency regulation circuit in the IDG GCU regulates the IDG output frequency by sensing the PMG frequency and comparing it to a reference frequency. If a difference exists, the IDG GCU adjusts the servo valve current as required to maintain a constant generator speed of 12,000 revolutions per minute (rpm). The frequency regulator function is designed to maintain generator output frequency when the IDG input speed range is 4,666 to 8,438 rpm with input speed changes of ±1,800 rpm per second. (This acceleration rate is for rated input speed only; maximum acceleration rate during startup is 240 rpm per second.)


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System Protection The protective functions are implemented by a microprocessor internal to the GCU that provides control of the aircraft electrical power generating system. The GCU provides the following protective functions:

NOTE In the cases where the protective function of the GCU drops the generator off line, it is a result of the GCU removing excitation from the generator. There is a generator control relay (GCR) internal to the GCU. The GCR is energized by the microprocessor to allow rectified PMG voltage to power the voltage rectifier. The voltage regulator provides the generator excitation voltage. The GCR is deenergized during a protective trip to shut down the generator by removing the excitation voltage.

Overvoltage Protection The GCU overvoltage (OV) protection circuitry senses each of the phase voltages at the input side, or point of regulation (POR), of the respective line contactors. If the voltage of the highest of the three phases exceeds the overvoltage threshold of 125.5 ±1.5 V rms, the GCU overvoltage protection function disables the voltage regulator and opens the GCR and the L/R/AAC after an inverse time delay. Undervoltage Protection The GCU undervoltage (UV) protection circuitry senses each of the phase voltages at the input side, or point of regulation (POR), of the respective line contactors. If the lowest of the three POR phases is less than 102.5 ±2.5 V rms, the undervoltage protection function disables the voltage regulator and opens the GCR and the L/R/AAC after a time delay of 4.5 ±0.25 seconds. If the lowest of the three POR phases is less than 70 ±1.5 V rms, the undervoltage protection function disables the voltage regulator, opens the GCR and the L/R/AAC after

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a time delay of 160 milliseconds maximum. The GCR and the L/R/AAC may be manually reset by cycling the appropriate GEN switch.

Overfrequency Protection The overfrequency (OF) protection receives input from the PMG, which has a 3:1 frequency ratio to the main aircraft generator. There are three levels of overfrequency protection (OF1, OF2, OF3). If the IDG or APU generator frequency sensed by the GCU is more than 420 +0/–5 Hz, the OF1 protective function disables the voltage regulator, opens the GCR, servo valve switch (SVS) (IDG only), and the L/R/AAC after a time delay of 4.0 ±0.25 seconds. If the APU generator frequency sensed by the GCU is more than 440 ±5 Hz, the OF2 protective function opens the GCR and AAC after a 160-millisecond delay maximum. If the IDG frequency sensed by the GCU is more than 454 ±8 Hz, the OF3 protective function disables the voltage regulator and opens the GCR, SVS, and L/RAC after a 90-millisecond delay maximum.

Underfrequency Protection The underfrequency (UF) protection receives input from the PMG, which has a 3:1 frequency ratio to the IDG. If the IDG or APU generator frequency sensed by the GCU is equal to or less than 380 Hz, the UF1 protective function disables the voltage regulator and opens the GCR, SVS (IDG only), and the L/R/AAC after a time delay of 4.0 ±0.25 seconds. If the IDG or APU generator frequency sensed by the GCU is less than 300 ±10 Hz, the U F 2 p r o t e c t iv e f u n c t i o n d i s a b l e s t h e voltage regulator and opens the GCR, SVS (IDG only), and the L/R/AAC after a 30 +20/–0-millisecond delay.




Overcurrent Protection The GCU overcurrent (OC) protection circuitry receives input from the generator current transformer. An overcurrent condition exists when the highest phase current exceeds the limit of 122 ±5 amperes rms. The GCU will clear the fault after a time delay. The GCU OC1 protective function will first lock out the crosstie contactor (L/RAXC). If this clears the fault, no further action is necessary. If the overcurrent is still present for 200 milliseconds after the crosstie lockout, OC2 protective function will lock out the left or right essential AC contactor (L/REAC). If the overcurrent fault is cleared after the L/REAC lockout, no further action is necessary. If the overcurrent is still present for 200 milliseconds after the L/REAC is locked out, the OC3 protective function will disable the voltage regulator and open the GCR and L/RAC. Because the APU GCU does not have a crosstie lockout function, the sequence of events during an overcurrent condition for the APU GCU is different than it is for the L/R GCU defined above. The APU GCU sends a request to the RBPCU to open the RAXC. If the fault is cleared, then no further action is necessary. If the overcurrent fault is still present for 200 milliseconds after the request was sent to the RBPCU, then the RAXC is reclosed and the APU GCU sends a request to the LBPCU to lock out (open) the LAXC. If the fault is cleared, no further action is necessary. If the overcurrent is still present for 200 milliseconds after the LAXC was locked out, then the RBPCU will lock out (open) the RAXC and the APU GCU will disable the voltage regulator and open the GCR and AAC.

Differential Fault Protection/Feeder Fault Protection The GCU uses the line current transformer assembly (CTA) and generator current transformers to provide differential fault protection (DP) against generator, generator feeder, and A/L/RAC faults. Phase A, phase B, and phase C are monitored by the generator CT, located in the generator, and the line CTA, located at the point of regulation (POR). The GCU compares each phase of the generator current with the load bus current. If the comparison of the individual phase current sensed by the generator CT and the line CT differ by more than 20 ±5 amperes, the appropriate generator line contactor (A/L/RAC) will trip in 90 milliseconds maximum. The GCR will trip 130 milliseconds maximum after DP detection to allow the fault to be isolated. If the fault was cleared after tripping the generator line contactor, then the adjacent crosstie contactor will be locked out. If the fault remained after tripping the generator line contactor then no crosstie lockout will occur. The APU GCU will send a request to lock out both crossties (L/RAXC) should the fault clear after tripping the AAC.


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Open Phase Protection The GCU monitors load current through the generator CTs to protect it against an open phase. If the GCU detects a low-phase current of less than 4.5 ±.5 amperes and the next lowest phase current is greater than 30 ±5 amperes, the open-phase protective function disables the voltage regulator and trips the GCR and the A/L/RAC after a time delay of 2.5 ±0.25 seconds. Shorted Rotating Diode Protection The GCU monitors the generator exciter field current and compares it to the load current to detect a shorted rotating diode. Excessive current for a specific amount of load indicates a shorted rotating diode. The shorted rotating diode protective function disables the voltage regulator and opens the GCR and the A/L/RAC after a time delay of 3.0 ±0.25 seconds. Shorted Permanent Magnet Generator Warning A shorted permanent magnet generator (PMG) is detected by a drop in the rectified PMG voltage applied to the positive side of the exciter field. This voltage will also vary with applied load. Therefore, shorted PMG protection is performed by monitoring the rectified PMG voltage and comparing its value to a predetermined line or exciter field current value. When the rectified PMG voltage is less than the expected amount over a time delay of 2.0 ±0.25 seconds, a shorted PMG condition exists. A shorted PMG is only detectable when the generator is excited and the A/L/RAC is open, which occurs during powerup or shutdown. Underspeed Protection The magnetic pickup/monopole unit (MPU) provides the signal for the IDG input speed. The IDG underspeed protective function will trip the L/RAC after a time delay of 100 milliseconds if the IDG input speed goes below 4,500 ±45 rpm. When the aircraft has completed a flight and the engines are shut down, each IDG is automatically removed from its bus through the operation of the GCU’s under-

24-40

speed protection. The AAC will trip within 540 milliseconds maximum if there is no APU READY TO LOAD signal.

Phase Sequence Protection The GCU checks the three power phases at the POR during powerup to verify they are in proper sequence. If the GCU detects that the phase sequence is not A–B–C, the GCU prevents the closure of the AAC or L/RAC and the GCR will be tripped after a time delay of 60 milliseconds maximum. Inadvertent Parallel Trip Protection An inadvertent parallel trip (IPT) occurs when two AC power sources remain in parallel. Three levels of protection are provided. The BPCU provides the first level of protection for IPT. Once an IPT is detected by the BPCU, the BPCU will command open its crosstie contactor after a time delay of 150 milliseconds in an attempt to clear the fault. If the fault condition is cleared, no other protective action is required. The GCUs will provide the second and third level of protection for IPT. If a GCU senses an inadvertent paralleling condition for 200 milliseconds, the crosstie contactor will be locked out by the L/R GCU. On the APU generator, the APU GCU will command the AAC to trip in 250 milliseconds. If the fault is not cleared, then the L/R GCU will command the L/RAC to trip, 250-millisecond time delay for the RAC and 300-millisecond time delay for the LAC. Total time required to clear the fault will be less than 360 milliseconds.

Fire Switch (IDG Only) The GCU receives a discrete input signal from the fire handle in the flight deck. The GCU generates a signal to trip the GCR, SVS, and the LAC/RAC 0.8 ±0.2 milliseconds after the fire handle switch is sensed.




Central Processing Unit (CPU) Failure Protection If the CPU of a GCU should fail, the GCU will go into a fail-safe configuration, initiated by the GCU when any of the following conditions takes place: • The CPU power supply fails. • The CPU cycle (software cycle) exceeds its time allotment. • Three software errors occur in the CPU within a 1-second time limit. • The watchdog circuit is detected as failed during a powerup reset. The fail-safe circuitry is powered by a separate supply derived from the integral control power. The fail-safe condition will cause the inter-LRU serial communications bus to go quiet and will open the SVS, GCR, and A/L/RAC.

Servo Valve Protection (IDG) Only The servo valve gets input from the PMG, which has a 3:1 frequency ratio to the IDG. The PMG frequency is monitored for modulation that exceeds ±15 Hz (±5 Hz for the generator main field) above the reference frequency at a modulation rate greater than or equal to 4 Hz. The servo valve protective function disables the voltage regulator and trips the GCR, SVS, and L/RAC after a time delay of 2.0 ±0.25 seconds. Indicator Functions The GCU senses the following signals and controls indicators in the aircraft as follows: • Generator frequency, voltage, and load sensing

Generator Frequency, Voltage, and Load Sensing The IDG/APU generator GCU monitors the IDG/APU generator frequency, voltage, and load. The GCU processes this data and transmits the information over a bidirectional serial data bus to the BPCU. The L BPCU transmits information via ARINC 429 to MAU No. 1, DGIO 1 slot 9/10 for display. The R BPCU transmits information via ARINC 429 to MAU No. 2, DGIO 2 slot 7/8 for display. • Generator frequency—Sensed on phases A and C of the PMG and phase A of the POR • POR voltage—Sensed on phase A • Generator load—High phase sensing (individual highest phase)

Contactor Status The GCU monitors the electrical system contactors for control and protection and to provide the CAS (with data sent through the BPCU) with system information. The system functions are monitored as follows: • IPT • Anticycling and lockout • Status for the CAS • Detection of contactor failure

AC Reset Light The left and right GCUs (L/R GCUs) control the lockout for the L/RAXC and L/REAC. If either is locked out, the L/R GCUs will provide a 28-VDC signal to control the illumination the AC reset light on the COP.

• Contactor status • AC reset light


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Table 24-9. LEFT BUS POWER CONTROL UNIT POWER INPUTS ORIGIN

CB/LOCATION

NOMENCLATURE

FUNCTION

External AC power cart ΦA

EXT AC SENSE/RPDB

External AC sense ΦA

Monitors voltage and frequency for power quality and is rectified to 28 VDC to power the BPCU and provide E/F interlock voltage when external AC is the only power available

External AC power cart ΦΒ

EXT AC SENSE/RPDB

External AC sense ΦB

Monitors voltage and frequency for power quality

External AC power cart ΦC

EXT AC SENSE/RPDB

External AC sense ΦC

Monitors voltage and frequency for power quality

Left essential TRU (two inputs)

L ESS TRU SENSE/ PWR/LPDB

Left essential TRU sense Voltage

System protection and indication

Left main TRU

L MAIN TRU SENSE/ PWR/LPDB

Left main TRU sense voltage


Auxiliary TRU

AUX TRU L SENSE/ PWR/RPDB

Auxiliary TRU sense voltage


External DC power cart (+ pin)

EXT DC–L BPCU/RPDB

External power input power

Input power to left BPCU

Right essential DC bus

L BPCU PWR/REER

Right essential 28-VDC input power

Input power to left BPCU

Table 24-10. RIGHT BUS POWER CONTROL UNIT POWER INPUTS ORIGIN

CB/LOCATION

NOMENCLATURE

FUNCTION

Essential AC bus ΦA

ESS AC–RBPCU/LPDB

Essential AC sense ΦA

Monitors voltage to make sure it is >102 VAC

Essential AC bus ΦΒ

ESS AC–RBPCU/LPDB

Essential AC sense ΦB


Essential AC bus ΦC

ESS AC–RBPCU/LPDB

Essential AC sense ΦC


Right essential TRU (two inputs)

R ESS TRU SENSE/ PWR/RPDB

Right essential TRU sense voltage

Monitors voltage for system protection and indication

Right main TRU

R MAIN TRU SENSE/ PWR/RPDB

Right main TRU sense voltage


Auxiliary TRU

AUX TRU SENSE/ PWR/RPDB

Auxiliary TRU right sense voltage


External DC cart (+ pin)

EXT DC–R BPCU/RPDB

External power input power Polarity protection and input power to right BPCU

External DC cart (sense pin) INTERLOCK SENSE/RPDB 28-VDC interlock sense

Input power to right BPCU, provides interlock protection, and monitors power quality

L essential DC bus

Input power to right BPCU

24-42

R BPCU PWR/LEER

Left essential 28-VDC input power




BUS POWER CONTROL UNIT (BPCUs) There are two bus power control units (BPCUs) located in the left and right electronic equipment racks (LEER/REER). The BPCUs monitor and protect the AC and DC external power sources, control and protect the DC buses, control the master frequency and NBPT (LBPCU only), and provide protection for the essential AC bus. The BPCUs interface with the crew alerting system (CAS) for built-in test (BIT) functions and system status information. The BPCUs also supply power and control logic for most of the AC and DC contactors. The input power sources to each of the BPCUs are listed in Table 24-9 and 24-10.

The RBPCU controls the AC contactors in the aircraft system as follows (in order of priority): • Right AC crosstie contactor (RAXC) • Right essential AC contactor (REAC) • Auxiliary TRU AC contactor (ATAC1) If a BPCU has the requirement to close both an AC and DC contactor at the same time, the AC contactor will have priority over the DC contactor.

DC Contactor Control The LBPCU controls the DC contactors in the aircraft system as follows (in order of priority): • Left essential DC contactor (LEDC)

The left BPCU (LBPCU) and right BPCU (RBPCU) are physically identical and completely interchangeable.

• Right essential DC crosstie contactor (REDXC)

The major BPCU system functions include system control and protection, indicator functions, and BIT capability.

• Right main DC crosstie contactor (RMDXC)

Control Functions Control functions handled by the BPCU include:

• Left main DC contactor (LMDC)

The RBPCU controls the DC contactors in the aircraft system as follows (in order of priority): • Right essential DC contactor (REDC) • Left essential DC crosstie contactor (LEDXC)

• AC contactor control

• Right main DC contactor (RMDC)

• DC contactor control

• Left main DC crosstie contactor (LMDXC)

• Frequency reference • Air–ground mode • Power transfer control

AC Contactor Control The LBPCU controls the AC contactors in the aircraft system as follows (in order of priority): • Left AC crosstie contactor (LAXC) • External AC contactor (EAC) • Left essential AC contactor (LEAC)


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Frequency Reference The LBPCU selects a frequency reference unit (FRU) signal for the left and right generator control units (L/R GCUs) to control frequency and phase angle of the GCU’s integrated drive generator (IDG). The FRU is used to provide frequency control between an IDG power source and a non-IDG power source, which lets the system do the NBPT. The sources for the frequency reference are as follows (in order of priority): • APU generator • AC external power • BPCU internal • GCU internal The reference source must be 400 ±10 Hz and 115 ±5 VAC (power-ready parameters satisfied) before it can be used as a reference. When bringing an IDG online, if a frequency reference of good quality is not found, then the GCU will switch to the internal 400-Hz reference and a NBPT will not occur. In this case, the IDG must be isolated before it can be brought online. When taking an IDG offline, if the reference source is out of limits, then a break transfer will occur with that source.

external power. The BPCUs provide the following protective functions:

Central Processing Unit Failure Protection If the CPU of the BPCU should fail, the BPCU will go into a fail-safe configuration. The EAC, EDC and IMR will be opened, the L/RAXC will remain in their prefault condition, provided their control switches on the COP remain selected to the AUTO position, the inter-LRU serial data bus will go quiet, and all ARINC 429 transmitter outputs will go offline. Opening the IMR consists of removing 28 VDC from pin F of the E/F interlock. Fail-safe is initiated by the BPCU when any of the following conditions occur: • The CPU power supply fails. • The CPU cycle (software cycle) exceeds its time allotment. • Three software errors occur in the CPU within a 1-second time limit. • The watchdog circuit is detected as failed during a powerup reset.

Air–Ground Mode The BPCU can sense whether the aircraft is in the air or on the ground. The BPCU receives this information through “gear down” and weight-on-wheels (WOW) signals from the landing gear. This information is provided for the BPCU’s built-in test equipment (BITE). System Protection The BPCUs have circuitry for monitoring external power parameters, including voltage, frequency, and current. They also have the capability to remove the source of external power in the event of a system fault or when external power quality is out of limits. This protects the aircraft equipment connected to the bus. The LBPCU has control of the AC external power, and the RBPCU has control of the DC

24-44




BPCU AC Bus Protective Functions Inadvertent Parallel Trip Protection (IPT) An inadvertent parallel trip (IPT) occurs when two AC sources remain in parallel. Three levels of protection are provided. The BPCU provides the first level of protection for IPT. Once an IPT is detected by the BPCU, the BPCU will command open its crosstie contactor (LAXC/RAXC) in 150 milliseconds in an attempt to clear the fault. If the fault condition is cleared, no other protective action is required.

deenergize the LEAC before the REAC can be energized, since auxiliary contacts of the LEAC do not allow both contactors to be energized at the same time. The LBPCU deenergizes the LEAC, sends a command to the L GCU to lock out the LEAC, and the AC reset light is illuminated on the COP. The LEAC and REAC are now configured to allow the right main AC bus to power the essential AC bus. If a fault now occurs on the right main AC bus, the RBPCU deenergizes the REAC and sends a request to the R GCU to lock out REAC. Power for the essential AC bus is now provided by the E-inverter.

If the fault is not cleared, the GCUs will provide the second and third level of protection for IPT. If a GCU senses an inadvertent paralleling condition for 200 milliseconds, the crosstie (L/RAXC) will be locked out by the L/R GCU. On the APU generator, the APU GCU will command the AAC to trip in 250 milliseconds. If the fault is not cleared, then the L/R GCU will command the L/RAC to trip, 250-millisecond time delay for the RAC and 300-millisecond time delay for the LAC. Total time required to clear the fault will be less than 360 milliseconds.

NOTES

AC Essential Bus Undervoltage Protection The RBPCU undervoltage (UV) protection function monitors the voltage of the essential AC bus. The left main AC bus supplies power to the essential AC bus during normal operation. Once power is supplied to the left main AC bus, the LBPCU energizes the LEAC. At this time the RBPCU outputs power to close the REAC when aircraft ladder logic permits it (when the LEAC deenergizes). This configures the LEAC and REAC to allow the left main AC bus to supply power to the essential AC bus. The RBPCU monitors each phase of the essential AC bus for an undervoltage (UV) protection fault. If the essential AC bus voltage is less than 102 ±2.5 VAC for a maximum of 6.0 seconds, then the RBPCU sends a request to the LBPCU to deenergize the LEAC. The LBPCU must


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Indicator Functions The BPCU is designed to operate with the control switches and indicators on the cockpit overhead panel (COP). The BPCU provides 28 VDC through the control switches to provide reset capability. The BPCU also provides 28 VDC to relays to illuminate the AC/DC reset light. The BPCU provides current limiting to protect aircraft wiring between the COP and the BPCU. The control switches and indicators are as follows: • WOW (LBPCU only) • Geardown switch (RBPCU only) • AC–DC reset switch • DC reset switch light • External power switch • External power switchlight

DC faults whenever L/RMDC, L/REDC, or the EDC/ADC is tripped offline because of a fault. The BPCUs monitor the output of the system TRUs and the DC ground cart. The first DC contactor downstream from either of these sources will trip offline if the limits of a protective function are exceeded and illuminate the DC reset light. Pressing the AC/DC reset switch resets the tripped contactor when the fault clears and the power quality is restored. ATAC1 is included as part of the AC reset function. The RBPCU energizes ATAC1 if an undervoltage is detected on the left main AC bus and power quality exists on the right main AC bus. If the AC/DC reset switch is selected after ATAC1 has been energized and power quality has been restored on the left main AC bus, then ATAC1 will deenergize. This will allow the left main AC bus to provide power to the input of the auxiliary TRU.

WOW The LBPCU senses either a WOW active (closed circuit), when the aircraft is assumed to be on the ground, or a WOW false (open circuit). This signal is used in system control, protection, and BIT coordination. Gear Down Switch The RBPCU senses either an active (closed circuit) when the landing gear is down or a false (open circuit) when the landing gear is up. This signal is used in system control, protection, and BIT coordination. AC/DC Reset Switch The AC/DC reset switch is a momentary switch located in the COP. When selected, it closes a circuit, routing 28 VDC, sent by one LRU (GCU for AC reset and BPCU for DC reset) back to the cross-side LRU. The AC/DC reset light illuminates for AC faults whenever L/RAXC or L/REAC is locked out. Lockouts occur for the purpose of clearing a fault. Pressing the AC/DC reset switch resets the lockout and allows the tripped contactor to reclose. The AC/DC reset light illuminates for

24-46


NOTES



External Power On/Available (ON/AVAIL) Switch The external power on/available switch is a latching switch located on the COP. If external power, AC or DC, satisfies the power-ready requirements, the available (AVAIL) light will illuminate. (It is not necessary to have the main aircraft batteries selected on the bring external power onto the aircraft, but the AVAIL light will illuminate only if the batteries have been selected on.) The BPCUs provide the power and the logic for controlling the external power AVAIL switchlight. The ON light is controlled by the auxiliary contacts in the external power contactors, EAC, and EDC. Cycling the ON/AVAIL switch will close the external power contactor (EDC/ADC), as a function of the system priority of the external source satisfying the power-ready requirements. Once the EDC/ADC closes, the AVAIL light will go off and the ON light will illuminate. If both external sources are on the aircraft, external AC shall have priority over the external DC power. In this case, when the external power switch is selected, the ON illuminates to indicate external AC is powering the aircraft, and the blue AVAIL remains illuminated to indicate external DC power is available.


NOTES

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Figure 24-16. RPBD

Figure 24-17. Current Transformer

24-48




Current Transformer Assemblies (CTA)

NOTES

Board mounted current transformers within the PDBs and internal current transformers in the IDG generators are provided for AC load monitoring. The line current transformers are located in the left and right power distribution boxes (Figure 24-16). These bus bar feedthrough units provide load current information to the GCU. AC load monitoring is done by the three-phase line CTs that have a transformer ratio of 500/1. There is one line CT for each phase lead to sense the current. The line CT works with the integrated drive generator and auxiliary generator CTs to provide sensing information to the GCU. The IDG and auxiliary generator line CT give current information for current limit operation. This current is also compared between line CTs and generator CTs for differential fault protection (feeder fault). A line CTA is used to give current sensing signals to the LBPCU for overcurrent protection for the external AC ground power cart (Figure 24-17).


24-49



HYDRO MECH TRANSMISSION

PERMANENT MAGNET GENERATOR

EXCITER ARMATURE

ROT DIODE ASSY

GEN GEN ROTOR STATOR

LAC N

L MAIN AC BUS CTA

T3

T2

T1

SERVO CONT

L GEN

ANN LTS 28 VDC POWER

DIM

ON

TEST

OFF

DIM TEST

CMD ON GCS OFF CMD CMD RTN

GCS ON CMD GCS 28 VDC

CTA PHASE A, B, C

POR PHASE A, B, C GEN CT PHASE A, B, C

FIRE HANDLE

FIRE SW FIRE SW RTN

EXCITATION FLD PMG PHASE A, B, C SERVO CONT

DATA BUS TO BPCU

FUEL CUTOFF FUEL CUTOFF RTN

HP SPEED PROBE

Figure 24-18. IDG GCU Diagram

24-50


ENGINE FUEL S/O SWICH



AC Contactors

CAUTION

The LAC and RAC contactors are located in the left and right PDBs and provide connection from the IDGs to the main AC buses. They are energized by their respective GCUs (Figure 24-18).

Do not remove or reseat GCU while engines are running as PMG voltage is 70 to 100 VAC (line to line) and could result in damage to personnel or equipment.

Generator Control The generator control units, located in the left and right electronic equipment racks, provide voltage regulation, frequency regulation, and system protection and monitor the frequency reference unit (FRU) signals that are transmitted from the BPCU to the GCU. The GCUs also provide automatic and manual control of the IDGs. L/R GEN switches located on the electrical power control panel (EPCP) provide manual control of the IDGs by providing on or off requests to their respective GCU. With the L/R GEN switches selected to the depressed position, the GCUs automatically energize the LAC/RAC when the IDG input shaft speed is above 4,666 (55% hp). In the depressed position, the generator switches illuminate an amber OFF when LAC/RAC is not energized, indicating that the IDG is not online. The depressed switch will illuminate a green ON when LAC/RAC is energized, indicating that the IDG is online (Figure 24-18). The engine fire and fuel shutoff switches (Figure 24-18) provide inputs to the GCU that disabled the generator operation at engine shutdown or engine fire handle activation. Input from the engine HP speed probe is provided to the GCU to ensure that the IDG is rotating at a proper speed to deliver power.

The rectified voltage from the PMG is the primary power source for the GCU. With the generator switch selected ON, the GCU voltage regulator uses this DC voltage to excite the generator exciter field stator. The exciter armature windings develop a voltage that is then rectified by the rotating diode assembly that provides direct current to the main field rotor. The rotating electromagnetic field created in the main field rotor induces voltage into the main field stator. The resulting threephase AC voltage is now available to be routed to the aircraft buses through the LAC and RAC contactors.

CAUTION Exciter voltage is 140 to 160 VDC.

The GCUs monitor the output of the generators at a location between the generators and the LAC and RAC contactors. This is referred to as the point of regulation (POR). The GCUs monitor the POR voltage, CTA current inputs, and on–off requests from the generator switches and control the LAC and RAC contactors. The GCUs also output data to the BPCUs for display of synoptic page information and messages on the CAS display unit.

IDG Operation As the hydromechanical transmission rotates the generator armature shaft, the PMG rotor turns and induces an AC voltage into the windings of the PMG stator (Figure 24-18). This AC voltage is supplied to the GCU where it is rectified to a DC voltage. The GCU will not operate without this PMG input.


24-51



UP D

XO

X OX

R O TA TIO

SN 2 1 4X3O5X X0O0X0O X3O4X5 O 89

FW

XO

N

Figure 24-19. APU Generator

24-52




APU-DRIVEN AC GENERATION SYSTEM The APU generator provides an auxiliary source of AC power both in flight and on ground. The APU generator system consists of an APU generator, APU generator control unit (GCU), APU current transformer assembly, APU generator contactor, and APU generator control switch.

AC Contactor The AAC is located in the left power distribution box. It provides connection of the APU generator output to the main AC tie bus and is controlled by the APU GCU.

NOTES

APU Generator The APU generator is mounted to and mechanically driven by the APU (Figure 24-19). Although physically different, the output of the APU generator is electrically identical to the IDG generator output. It is rated at 40 KVA, 115 volt, three phase, 400 Hz. The APU generator differs physically from the IDG in that the APU generator has an internal PMG, and its output frequency is a function of APU speed. Cooling and lubrication are provided by the APU oil system.

Generator Control Unit (GCU) The APU GCU is located in the right electronic equipment rack. The APU GCU provides voltage regulation, over/undervoltage protection, over/underfrequency protection, feeder fault protection, phase sequence protection, open phase protection, control interface, and controls contactor AAC. Details of the APU GCU are discussed in the previous GCU section.

Current Transformer Assembly (CTA) The APU line current transformer assembly (CTA) is located in the left power distribution box. The CTA provides the APU GCU with line current information. Details of CTA functions are discussed in the previous CTA section.


24-53



PERMANENT MAGNET GENERATOR

EXCITER ROT GEN GEN ARMATURE DIODE ROTOR STATOR ASSY

AAC N

AC X-TIE BUS CTA

T3

T2

T1

APU ANN LTS 28 VDC POWER

DIM TEST

ON

CMD ON GCS OFF CMD GCS ON CMD GCS 28 VDC

CMD RTN CTA PHASE A, B, C

POR PHASE A, B, C GEN CT PHASE A, B, CAPU READY TO LOAD READY TO LOAD RETURN EXCITATION FLD PMG PHASE A, B, C

APU READY TO LOAD RELAY (OPENS ABOVE 99% RPM)

DATA BUS TO BPCU

Figure 24-20. APU Generator—GCU Block Diagram, APU Generator ON

24-54




APU Generator Control The generator control unit, located in the right electronic equipment rack, provides voltage regulation and system protection. The GCU also provides automatic and manual control of the APU generator. The APU GEN switch, located on the electrical power control panel (EPCP), provides manual control of the APU generator by providing on or off requests to the GCU. With the APU GEN switch selected to the depressed position, the GCU automatically energizes the AAC when the APU speed is greater than 99%. In the depressed position, the generator switch is not illuminated when the AAC is not energized, indicating that the APU generator is not online. The depressed switch will illuminate an amber ON when AAC is energized, indicating that the APU generator is online (Figure 24-20).

APU Generator Operation As the APU rotates the generator armature shaft, the PMG rotor turns and induces an AC voltage into the windings of the PMG stator. This AC voltage is supplied to the GCU where it is rectified to a DC voltage. The GCU will not operate without this PMG input.

CAUTION Do not remove or reseat GCU while the APU is running because PMG voltage is 70 to 100 VAC (line to line) and could result in damage to personnel or equipment.

CAUTION Exciter voltage is 140 to 160 VDC.

Three-phase, 115 VAC, 400 Hz power is produced by the APU generator and is supplied to the open contacts of the AAC. The GCU monitors the output of the generator at a location between the generators and the AAC. This is referred to as the point of regulation (POR) (Figure 24-20). Point of regulation (POR) voltage is sampled by the GCU, and, if power quality is good, the GCU commands the AAC to close when the APU READY TO LOAD relay energizes (APU speed greater than 99%). The AAC then closes, the generator switch illuminates ON, and AC power is routed to the main AC tie bus. The GCU continues to monitor the POR voltage, CTA current inputs, and on–off requests from the generator switch. If voltage, current, or frequency fall outside limits, the AAC contactor is opened and excitation is removed from the APU generator. The GCU also sends data to the RBPCU for display of synoptic page information and messages on the CAS display unit. The RBPCU shares this information with the LBPCU so they can close the LAXC and RAXC, providing APU generator power a path to the main AC buses. The AC POWER synoptic page provides the indication for the APU generator system. CAS messages related to APU generator operation are displayed on the CAS display unit.

The rectified voltage from the PMG is the primary power source for the GCU. With the generator switch selected to ON, the GCU voltage regulator uses this DC voltage to excite the generator exciter field stator. The exciter armature windings develop a voltage that is then rectified by the rotating diode assembly that provides direct current to the main field rotor. The rotating electromagnetic field created in the main field rotor induces voltage into the main field stator.


24-55



Figure 24-21. Underfloor Equipment Locator—Emergency Inverter LEFT MAIN AC BUS RIGHT MAIN AC BUS REAC (LOC: RPDP)

LEAC (LOC: LPDB)

ESS AC BUS PHASE A E-INVERTER (LOC: UNDERFLOOR)

NOTE: SHOWN WITH L ESS DC BUS POWERING E-INVERTER. 115 VAC AØ 400-HZ OUTPUT 28-VDC INPUT

L ESS DC

R ESS DC LEIDC (LOC: LPDB)

REIDC (LOC: RPDB)

Figure 24-22. Emergency Inverter Operation (Sheet 1 of 3)

24-56




L MAIN AC BUS

R MAIN AC BUS

ØA ØB ØC

ØA

ØB

ØC

L ESS AC CONTACTOR ON COMMAND L ESS AC ON CMD RETURN L BUS POWER CONTROL UNIT

FROM E INVERTER L ESS AC LOCKOUT RETURN

REAC +

-

L ESS AC CONTACTOR LOCKOUT L GENERATOR CONTROL UNIT (GCU)

ESSENTIAL AC PHASE A, B, C R ESS AC CONTACTOR ON CMD +

-

LEAC

R ESS AC ON CMD RETURN R BUS POWER CONTROL UNIT

R ESS AC LOCKOUT RETURN R ESS AC CONTACTOR LOCKOUT

ESS AC BUS

R GENERATOR CONTORL UNIT ØA

ØB

ØC

ESSENTIAL AC POWER SOURCE

Figure 24-22. Emergency Inverter Operation (Sheet 2 of 3) (LEAC/REAC Control)


24-57



R MAIN AC BUS ØA

ØB

ØC

E INVERTER

AC 1Ø E-INV REAC

L MAIN AC BUS ØA ØB ØC

LEAC

ESS AC BUS

ØA

ØB

ØC

L ESS DC

R ESS DC

LEIDC

AUTO

REIDC

OFF

E-INV

DC POWER SOURCE

Figure 24-22. Emergency Inverter Operation (Sheet 3 of 3) (LEIDC/REIDC Control)

24-58




EMERGENCY AC INVERTER SYSTEM In the event power to the ESS AC bus is not available from the main AC buses, the emergency AC inverter provides 1-KVA, singlephase, 115 VAC, 400 Hz power to the AC essential bus. The components, controls, and indicators for the emergency AC inverter system are the E-inverter, L/R EAC contactors, L/R EIDC contactors, and the E-inverter switch. The E-inverter is located in the underfloor equipment area between the LEER and the REER (Figure 24-21) and supplies single-phase, 115 VAC power to the A phase of the essential AC bus through REAC and LEAC (Figure 24-22).

should fail, REIDC would close and provide DC power to the inverter. The AC POWER synoptic page provides indication for the Einverter system.

NOTES

Emergency Inverter Operation Input power to the E-inverter (Figure 24-22) is supplied by either the left or right essential DC bus through the contacts of the LEIDC or REIDC. LEAC and LEIDC are located in the left PDB. REAC and REIDC are located in the right PDB. The E-inverter switch is located on the electrical power control panel and illuminates a blue AUTO when the switch is extended for normal automatic operation. The E-inverter switch will illuminate an amber OFF when depressed. If the EINV–AUTO/OFF switch is in OFF position, LEIDC and REIDC are not permitted to close, and the E-inverter will not receive input power. Wi t h no input p o w e r, the Einverter has no output power and is unable to power the essential AC bus. The essential AC bus is normally powered by the left main AC bus through LEAC. If LEAC is deenergized for any reason, the next priority power source for the essential AC bus is the right main AC bus through REAC and LEAC. If this power source is lost for any reason, then system control logic will energize LEIDC, which routes DC power to the E-inverter, causing the essential AC bus (phase A) to be powered by the E-inverter. If for any reason the left essential DC bus of LEIDC


24-59



Figure 24-23. External AC Power Receptacle EAC

CTA A

AIRCRAFT LOADS (MAIN AC TIE BUS)

B C

N

EXT AC SENSE CB (RPDB)

E F

LBPCU EXT AC PHASE A

EXT PWR RECPT

EXT AC PHASE B EXT AC PHASE C

CMD ON CMD ON RTN

CT PHASE A, B, C INTERLOCK (E) INTERLOCK (F) PWR AVAIL LT PWR AVAIL LT RTN

PWR ON CMD PWR ON CMD RTN

DIM DIM

TEST

ON AVAIL

TEST

EXT PWR

EXT AC AVAIL PWR RELAY

Figure 24-24. EXT AC Block Diagram

24-60


EXT PWR SWITCH

28-VDC ANN LTS PWR



EXTERNAL AC POWER SYSTEM

NOTES

The external AC power system consists of the external AC receptacle, external AC current transformer assembly, left bus power control unit, external AC contactor (EAC), and external power switch. External power is supplied to the aircraft through the external AC power receptacle, located on the lower right side of the fuselage exterior (Figure 24-23). The external AC power receptacle is the connection point between the aircraft and the external power source and is designed to ensure positive connection through monitoring of the E–F interlock. The E–F interlock ensures pins A, B, C, and D are fully engaged before power can be transferred from the external AC source to the aircraft. The left BPCU senses when external power is properly connected (interlocked) by monitoring pin E of the AC power receptacle for a 28-VDC signal. This 28 VDC can be supplied from the power cart or the BPCU (BPCU rectified phase A from the power cart (Figure 24-24).


24-61



THREEPHASE GENERATOR

TO EXTERNAL AC CONTACTOR (EAC)

N

N

A

A

B

B

ØB

C

C

ØC

E

E

60

F

F

54

ØA

EXTERNAL AC CONNECTOR

GROUND POWER CART

28-VDC INPUT 28-VDC SUPPLY LEFT BUS POWER CONTROL UNIT

Figure 24-25. Ground Power Carts—Type 1 Block Diagram

THREEPHASE GENERATOR 28-VDC SUPPLY


N

N

A

A

B

B

ØB

C

C

ØC

E

E

60

F

F

54

ØA


GROUND POWER CART



GROUND POWER CART THREEPHASE GENERATOR 28-VDC SUPPLY CONTROL PUSHBUTTON


N

N

A

A

B

B

ØB

C

C

ØC

E

E

60

F

F

54


ØA



24-62




External AC Ground Power Carts

NOTES

There are several types of ground power carts that will work with the G-Vs external AC connector. Three different types are shown here to demonstrate ways in which the BPCU monitors the E–F interlock. The first power cart is Type 1. Phase A of the ground power cart’s output is rectified to DC voltage by the BPCU and used to power the BPCU. 28 VDC is sent from the BPCU’s power supply to pin F of the external AC connector. Pin F of the connector is shorted to pin E, allowing the BPCU to provide its own DC power signal (Figure 24-25). The second type isolates pin F of the external connector, and pin E is powered from the power cart’s 28-volt power supply (Figure 24-26). The third type of power cart contains a momentary control switch. When the switch is depressed, the ground power cart’s output contactor is energized by the ground cart’s internal 28 VDC power supply. Phase A of the ground power cart’s output is rectified to DC voltage by the BPCU and used to power the BPCU. 28 VDC is sent from the BPCU’s power supply through pin F of the external AC connector to the coil of the power cart’s output contactor. When the momentary control switch is released, the cart’s output contactor is held closed by this 28 VDC from the BPCU’s power supply. Power is then sent to the BPCU through pin E of the external AC connector to complete the interlock (Figure 24-27).

External Power Current Transformer Assembly (CTA) The CTA is located below the floor between the external AC power receptacle and the right power distribution box. The external power CTA senses the current flow from the external power receptacle and provides load information to the left bus power control unit for system protection and indication (see Figure 24-24).


24-63



External AC Power Protective Functions (LBPCU) The left bus power control unit is located in the left electronic equipment rack. The LBPCU provides external AC power over/undervoltage, over/underfrequency, overcurr e n t , p h a s e s e q u e n c e , o p e n / s h o r t C TA monitoring, and protection. It also interfaces with the CAS via the DAU. The LBPCU controls the application of external AC power by controlling the external power contactor (EAC) located in the right PDB. The AC external power protective functions are provided by the LBPCU as follows:

AC Interlock Protection The LBPCU provides a 28 VDC output on pin F, which is either applied to pin E for the interlock or the AC external power source supplies a 28 VDC input to pin E when 28 VDC is received from pin F for the interlock.

Overvoltage Protection The LBPCU overvoltage (OV) protection circuitry senses each of the phase voltages at the input side, or point of regulation (POR), of the external AC power contactor (EAC). If the voltage of the highest of the three phases exceeds the overvoltage threshold of 125.5 ±1.5V rms, the LBPCU opens the interlock monitoring relay (IMR) and the EAC. Opening the IMR consists of removing 28 VDC from pin F of the E–F interlock.

Undervoltage Protection The LBPCU undervoltage (UV) protection circuitry senses the three-phase voltages at the input side, or POR, of the EAC. If the lowest of the three POR phases is equal to or less than 102.5 ±2.5 VAC, the undervoltage protection function opens the IMR and the EAC after a time delay of 4.5 ±0.25 seconds. Opening the IMR consists or removing 28 VDC from pin F of the E–F interlock.

24-64

Overfrequency Protection If the external AC power frequency sensed by the LBPCU is more than 420 +0/–5 Hz, the overfrequency protection function opens the IMR and EAC after a time delay of 4.0 ±0.25 seconds. Opening the IMR consists of removing 28 VDC from pin F of the E–F interlock.

Underfrequency Protection If the external AC power frequency sensed by the LBPCU is equal to or less than 380 +5/–0 Hz, the underfrequency protection function opens the IMR and EAC after a time delay of 4.0 ±0.25 seconds. Opening the IMR consists of removing 28 VDC from pin F of the E–F interlock.

Overcurrent Protection The LBPCU overcurrent protection circuitry monitors the three phases of the external power using the CTA. An overcurrent condition exists when the highest of the line currents exceeds the limit of 122 ±5 amperes rms for an inverse time delay. If an overcurrent condition exists, the LBPCU sends a request to the RBPCU to open RAXC. If the fault is cleared, no further action is required. If the overcurrent fault is still present for 200 milliseconds after the request was sent to the RBPCU, then RAXC is reclosed and the LBPCU opens the LAXC. If the fault is cleared, no further action is required. If the fault remains 200 milliseconds later, then RAXC is opened. Power is left on the tie-bus long enough to determine if the fault is located there. Then the LBPCU opens the IMR and EAC. Opening the IMR consists of removing 28 VDC from pin F of the E–F interlock.

Phase Sequence Protection The LBPCU checks the three power phases at the POR during powerup to verify they are in proper sequence. If the LBPCU detects that the phase sequence is not A–B–C, the LBPCU prevents closure of the EAC and the IMR will be opened in 80 milliseconds maximum. Opening the IMR consists or removing 28 VDC from pin F of the E–F interlock.




Open-Short CTA Protection The LBPCU uses the line CTA to sense threephase current for open-short CTA protection. When the lowest phase is less than 4.5 ±4.5 amperes rms, while the next lowest phase is more than 30 ±5 amperes rms, the LBPCU openshort CTA protective function will open the IMR and EAC in 2.5 ±0.25 seconds. Opening the IMR consists of removing 28 VDC from pin F of the E–F interlock.

Control and Indication External power control and indicator is provided by a split legend, alternate action EXT PWR switch located on the electrical power control panel. The external power switch illuminates a blue AVAIL in the lower legend when acceptable external power is connected but not selected. The external power switch illuminates an amber ON in the upper legend when the switch is depressed and the EAC is closed. If only external AC power is applied to the aircraft, the blue AVAIL will extinguish when the amber ON illuminates. If both AC and DC external power is available, external AC power has priority. In this case, the amber ON will illuminate, indicating that external AC power is online, and the blue AVAIL will remain illuminated indicating that acceptable external DC power is available. If external DC is powering the aircraft and acceptable external AC power is connected to the aircraft, the system configuration will not change. To bring external AC onto the aircraft in this situation would require cycling the external power switch. Cycling the switch to the extended position will disconnect the external DC power by opening the external DC contactor (EDC). The AVAIL light will illuminate at this time. Cycling the external power switch to the ON position again will bring the external AC power onto the aircraft. Should a fault occur with the AC external power after it has been brought onto the aircraft, and DC power quality is still acceptable, then DC external power will automatically be brought onto the aircraft after the AC external power is tripped offline.

External power is brought on to the aircraft through the closure of EAC. The EAC will be energized when AC external power quality is within acceptable limits, E–F interlock conditions are satisfied, and the external power switch is selected to the depressed position. The EAC will remain energized until the external power plug is removed from the aircraft, the external power quality is no longer within acceptable limits, a system protection trip occurs, or the external power switch is selected to the extended position. Indication and system advisories are provided for on the CAS and the AC POWER synoptic page.


NOTES

24-65



AC X-TIE BUS

LEFT MAIN AC BUS

RIGHT MAIN AC BUS

LGCU

RAXC

LEFT AC X-TIE LOCKOUT

+ LBPCU

LEFT AC - TIE LOCKOUT RTN

LAXC

AC RESET LIGHT RTN LEFT AC X-TIE CMD ON RTN

AC RESET LIGHT AC RESET CMD

LEFT AC XTIE CMD ON

AC RESET CMD RTN

L BUS TIE AUTO ISLN

RESET 28 VDC ANN LTS PWR

AC DC

L TIE BUS SW ON MAU 1 DG I/O MODULE SLOT 9/10

L AC RESET RELAY

Figure 24-28. LAXC Control

24-66




GENERATED AC POWER DISTRIBUTION

NOTES

The AC power source distribution system enables both main AC buses to be powered by any of four available AC power sources. It also provides for uninterrupted bus power during source switching by the use of the no-break power transfer (NBPT) feature of the EPS. The components for the AC power source distribution system include the left and right bus power control unit, left and right generator control unit, APU generator control unit, left and right AC crosstie contactors, AC reset button, and the left and right bus-tie switches. The APU GCU is located in the right electronic equipment rack. The LAXC and RAXC contactors are located in the left and right power distribution boxes, respectively. The control for source distribution is primarily an automatic function of the GCUs and BPCUs. Manually input controls consist of the left and right bus-tie switches (L BUS TIE and R BUS TIE) and the AC reset switch, located on the cockpit overhead control panel (see Figure 24-7). The reset switch will illuminate an amber AC when the left or right AC crosstie contactor (LAXC or RAXC) is locked out due to a protective fault condition. Depressing the momentary switch sends a signal to the left and right generator control units to clear the lockout and reset the logic. If the fault condition is no longer present, this also extinguishes the AC legend and allows LAXC and RAXC to operate (Figure 24-28). The left and right bus-tie isolation switches provide cockpit control of isolation of a main bus from a source other than its respective IDG. When depressed, the bus-tie isolation switches will illuminate ISLN, and will prevent automatic source switching by preventing the LAXC and RAXC from closing.


24-67



R BPCU

L BPCU BUS POWER CONTROL UNIT

BUS POWER CONTROL UNIT

115/ 200 VOLT-400Hz GAC PN 11558GCAVS021 MFG PN 150930 MODEL NO. 40EG514G MFG SERVO MFG DATE MODIFICATION 7

8

9

1

2

115/ 200 VOLT-400Hz

3

4

5

GAC PN 11558GCAVS021 MFG PN 150930 MODEL NO. 40EG514G MFG SERVO MFG DATE MODIFICATION

6

10 11 12 13 14 15 16 17 18 19 20

7

21 22 23 24 25 26 27 28 29 30 31 32 33 34

8

9

1

2

3

4

5

6

10 11 12 13 14 15 16 17 18 19 20

21 22 23 24 25 26 27 28 29 30 31 32 33 34

CAGE CODE 99167

CAGE CODE 99167

SUNSTRAND - ROCKFORD, IL US


ATTENTION

ATTENTION

SENSITIVE ELECTRONIC DEVICE


OBSERVE HANDLING PROCEDURES INSTALL DUST CAP WHEN CONNECTOR IS NOT IN USE


BUS POWER CONTROL UNIT 115/ 200 VOLT-400Hz

APU GCU

BUS POWER CONTROL UNIT 115/ 200 VOLT-400Hz

SERVO CONTROL

GAC PN 11558GCAVS021 MFG PN 150930 MODEL NO. 40EG514G MFG SERVO MFG DATE MODIFICATION 7

8

9

1

2

3

4

5

6

10 11 12 13 14 15 16 17 18 19 20

21 22 23 24 25 26 27 28 29 30 31 32 33 34

GAC PN 11558GCAVS021 MFG PN 150930 MODEL NO. 40EG514G MFG SERVO MFG DATE MODIFICATION 7

8

9

1

2

3

4

5

6

10 11 12 13 14 15 16 17 18 19 20

21 22 23 24 25 26 27 28 29 30 31 32 33 34

BUS POWER CONTROL UNIT

CAGE CODE 99167


115/ 200 VOLT-400Hz

L GCU

GAC PN 11558GCAVS021 MFG PN 150930 MODEL NO. 40EG514G MFG SERVO MFG DATE MODIFICATION

R GCU

7

8

9

1

2

3

4

5

6

10 11 12 13 14 15 16 17 18 19 20

21 22 23 24 25 26 27 28 29 30 31 32 33 34

CAGE CODE 99167

CAGE CODE 99167



SERVO CONTROL

ATTENTION

SENSITIVE ELECTRONIC DEVICE OBSERVE HANDLING PROCEDURES INSTALL DUST CAP WHEN CONNECTOR IS NOT IN USE

ATTENTION

ATTENTION





FREQ REF

L IDG

APU

LAC

AAC

LAXC

EXT

R IDG

EAC

RAC

RAXC

L MAIN AC BUS

R MAIN AC BUS

AC TIE BUS

Figure 24-29. NBPT Contactor Control Diagram

Table 24-11. MAIN ESSENTIAL BUSES PRIORITY LOGIC

24-68

Left main AC bus priority logic

1) 2) 3) 4)

Left IDG APU generator External AC Right IDG

Right main AC bus priority logic

1) 2) 3) 4)

Right IDG APU generator External AC Left IDG

Essential AC bus priority logic

1) Left main AC bus 2) Right main AC bus 3) Emergency inverter




AC Bus Source Priority Logic

NOTES

Both main AC buses can be powered by any of the four AC power sources listed in Table 24-11. AC bus source priority determines which source(s) will power which bus(es). If more than one source is available, source priority logic establishes which source will power that particular bus. Priority logic is primarily a function of the GCUs and BPCUs. Refer to Table 24-11 for priority logic for the main and essential buses. The bus power control unit and generator control unit communicate via inter-LRU data buses when controlling AC source power transfers. This high-speed communication enables the accomplishment of no-break power transfers (NBPTs) (Figure 24-29). Control of NBPTs is a function of the LBPCU. A nobreak power transfer means that there is no break or interruption of power on the affected bus during source switching. An NBPT can only be accomplished when at least one of the two sources involved in the transfer is an IDG (which is frequency controlled). Frequency control allows momentary paralleling of two sources on a bus before breaking connection of the deselected or lower priority source.


24-69



Frequency Reference Unit The LBPCU selects a frequency reference unit (FRU) for the left and right generator control units (L/R GCUs) to control frequency and phase angle of the GCUs integrated drive generator (IDG). The FRU is used to provide frequency control between an IDG power source and a non-IDG power source, which allows the system to do the NBPT. APU generator and external AC power are not frequency controlled by the GCU and BPCU; therefore power transfer between these two sources are break power transfers and result in momentary power loss during transfer. The reference source must be 400 ±10 Hz and 115 ±5 VAC, and phase errors less than 15° (power ready parameters satisfied) before it can be used as a FRU. If the reference source is out of limits or transfer does not happen within three seconds, then a break transfer will occur with that source (if requested) and the next priority source will be used for a frequency reference. If a frequency reference of good quality is not found, then the GCU will switch to the internal 400-Hz reference, and an NBPT will not occur. The sources for the frequency reference are as follows (in order of priority): • APU generator • AC external power • BPCU internal 400-Hz reference • GCU internal 400-Hz reference (When down to GCU, there will be a break power transfer.)

Power Transfer Control All nonfailure-related AC power transfers are NBPTs, except between external power and auxiliary power unit (APU) generator sources. An IDG must be involved in the power transfer to have an NBPTs occur. In addition, the BPCUs have dedicated input and output control signals to perform NBPT.

24-70

NBPT Sequence of Events 1.

Start APU.

2.

APU GCU reports APU generator information to the R BPCU over the 1553 inter-LRU data bus. The R BPCU relays this information to the R GCU and the LBPCU. The L BPCU passes the information along to the L GCU.

3.

The IDG GCUs now have the information needed to synchronize the IDGs to the APU generator available to them.

4.

Start the right engine (generator switch depressed). a. IDG starts to turn with input from the gearbox. b. PMG voltage is supplied to the R GCU (primary power to GCU). c. GCU rectifies the PMG voltage and powers the whole GCU. d. Engine speed is supplied to the GCU byt he HP speed probe. When the GCU determines that input shaft speed to the IDG is 4185 rpm, it sends a servo valve current to the CSD portion of the IDG. the servo valve current will adjust the CSD (which is a hydro-mechanical transmission) which will convert the variable input shaft speed to a steady speed of 12,000 rpm (=400Hz). e. When the engine comes up to speed, the internal generator control relay energizes and power is supplied to the GCU’s voltage regulator. f. The voltage regulator applies excitement voltage to the IDG. g. GCU uses the PMG to determine the IDG’s frequency, and checks the POR to determine the output voltage.




h. When these checks are completed and IDG output is good, an “On Line” signal is sent to the L BPCU (in charge of NBPTs). This is a request to the system to perform a NBPT. 5.

If the APU generator’s output is 115 VAC ±5 VAC and 400 Hz ±10 Hz, the L BPCU tells the R BPCU to tell the R GCU to use the APU generator as the frequency reference unit in order to perform the NBPT.

6.

The R GCU adjusts the servo valve current and the excitation voltage to synchronize the output of the IDG with that of the APU generator.

7.

If the output of the two generators are within .5 Hz and the phase error is within ±15°, the R GCU sends a “Sync True” signal to the L BPCU.

8.

the L BPCU sends a “Close AC: command (RAC) to the R GCU.

9.

The R GCU commands the RAC closed. The R IDG and the APU generator are now paralleled on the X-tie bus.

NOTES

10. When the parallel is confirmed by the L BPCU, it initiates a “Trip AC” command. The R BPCU receives the command and opens the RAXC by removing power from the coil.


24-71



LEER

PILOT OVERHEAD PANEL

Figure 24-30. PDB Circuit-Breaker Panel—LEFT MAIN AC

24-72




AC Power Distribution

NOTES

The purpose of the AC power distribution system is to protect and distribute 115-volt, threephase, 400-Hz power from the bus to the aircraft AC loads. Distribution of AC power is accomplished using five AC buses. Three are classified as primary AC buses: left main (L MN), right main (R MN), and essential (ESS). Two are classified as standby AC buses: left standby (L STBY) and right standby (R STBY). The left main AC bus and essential AC bus are located in the left power distribution box (PDB). The right main AC bus is located in the right PDB. The power distribution boxes provide distribution of main and essential AC bus power to the LEER circuit-breaker panel, REER circuit-breaker panel, pilot’s overhead circuit-breaker panel, copilot’s overhead circuit-breaker panel, and AC power electrical equipment. The left and right power distribution boxes are in the left and right electronic equipment racks.

Left Power Distribution AC power from the left main AC bus is routed from the left PDB to the pilot’s overhead circuit-breaker panel (Figure 24-30). The PILOT 1 and PILOT 2 circuit breakers, located on the left main AC section of the PDB’s circuitbreaker panel, protect the circuit. AC power from the left main AC bus is also routed from the left PDB to the left electronic equipment rack circuit-breaker panel. This circuit is protected by the PILOT 3 circuit breaker, located on the left main AC section of the PDB’s circuitbreaker panel. Additionally, AC power protected by the circuit breakers located in the left main AC section of the left PDB is distributed from the left main AC bus to the labeled ACpowered equipment. Essential AC bus power is available only from the left PDB. Essential AC bus power is routed to all four circuit-breaker panels. This wiring is protected by 10-amp circuit breakers located on the essential AC section of the PDB.


24-73



REER

COPILOT OVERHEAD PANEL

Figure 24-31. PDB Circuit-Breaker Panel—RIGHT MAIN AC

24-74




Right Power Distribution

NOTES

AC power from the right main AC bus is routed from the right PDB to the copilot’s circuitbreaker panel (Figure 24-31). The circuit is protected by the COPILOT 1 and COPILOT 2 circuit breakers, located on the right main AC section of the PDB. AC power from the right main AC bus is also routed from the right PDB to the REER circuit-breaker panels. The circuit is protected by the COPILOT 3 and COPILOT 4 circuit breakers, located on the right main AC section of the PDB. Additionally, AC power protected by the circuit breakers located in the right main AC section of the right PDB is distributed from the right main AC bus to the labeled AC-powered equipment.


24-75



RIGHT MAIN AC BUS

LEFT MAIN AC BUS

LMTAC

RMTAC

ATAC 1 HMG ATAC 2

LEFT MAIN TRU

LEFT ESS TRU

EDC/ ADC

APC

LMDC

LEDC

RIGHT ESS TRU

RIGHT MAIN TRU

AUX TRU

RMDC

REDC

RMDXC

LMDXC

R MAIN DC BUS

L MAIN DC BUS

LEDXC

REDXC REDBC

LEDBC

GSBC 1 L ESS DC BUS

BC 1

R ESS DC BUS

BC 2 BATT TIE BUS

GSBC 2

LBCC

AUX HYD PUMP CONT

APU START CONT

GND SRV BUS

RBCC R BATT CHGR

L BATT CHGR

Figure 24-32. DC Power System

24-76




DC POWER SYSTEM

NOTES

INTRODUCTION The purpose of the DC power system (Figure 24-32) is to provide source, control, protection, and distribution of DC power, as required for operation of the aircraft systems. This chapter discusses primary DC power, battery power, and external DC power.

Table 24-12. COMPONENT LOCATIONS

Left electronic equipment rack (LEER)

• Left bus power control unit (LBPCU) • LEER circuit breaker panel • Left power distribution box (PDB) • Contactors • Bus distribution circuit breakers • TRU AC input circuit breakers

Right electronic equipment rack (REER)

• Right bus power control unit (RBPCU) • REER circuit breaker panel • Right power distribution box • Contactors • Bus distribution circuit breakers • TRU AC input circuit breakers

Under the floorboard

• • • • • • •

Cockpit overhead panel

• Electrical power control panel

Auxiliary TRU Left essential TRU Left main TRU Right main TRU Right essential TRU Hall effect sensors Standby junction relay panel


24-77



14K

14J

14H 1A

Figure 24-33. TRU Location

24-78




PRIMARY DC POWER

NOTES

Primary DC power is provided to meet the power requirements for onboard DC-powered systems. The locations (Figure 24-32) of the components, controls, and indicators of the primary DC power subsystem are shown in Table 24-12.

TRANSFORMER RECTIFIER UNITS Transformer rectifier units provide the primary source of DC power for the main and essential DC-powered systems in the aircraft, with the exception of the APU and auxiliary hydraulic pump. There are five transformer rectifier units (TRUs) located in the G550. All five are located under the floor (See Figure 24-37) between the LEER and REER and are labeled “LEFT ESS TRU,” “LEFT MAIN TRU,” and “AUX TRU,” “RIGHT MAIN” and “RIGHT ESS TRU.”


24-79



115 VAC, 400 Hz, 3 PHASE MAIN AC BUS POWER Aø

Bø

DC BUS

Cø

26–29 VDC 250 AMPS = 100% LOAD

Figure 24-34. TRU Description

24-80


TRANSFORMER RECTIFIER UNIT (TRU)



Transformer Rectifier Unit Description

NOTES

Each TRU converts 115 VAC, three-phase, 400 Hz input power to 28 VDC, 250 ampere continuous output power. The TRUs output voltage is unregulated and therefore dependent on load but operates in a normal range of 29 to 26 VDC with loads in the range of 20 to 250 amps, respectively. TRU input power is supplied by any of the following power sources: IDGs, APU generator, or external AC power. The AUX TRU can also be powered by the HMG. Each of the main and essential TRUs supply power to their respective DC buses exclusively. The system can be configured (through the L and R ESS standby electrical power switches) so that the AUX TRU, when powered by the HMG, can supply power to either or both of the essential DC buses. When the input power is supplied by a source other than the HMG, the AUX TRU supplies power to only one DC bus at a time, based on system configuration and bus priority. In the event of a sensed overtemperature condition or when there is a detected fan failure, a discrete signal provided by three internal temperature switches is output from the TRU. This discrete signal is sent to the DGIOs in MAU 1 and MAU 2.and displayed as a “TRU HOT” CAS message. The switches close at 310 ±8°F and open at 280 ±8°F.


24-81



Figure 24-35. Hall-Effect DC Current Sensor

24-82




Hall-Effect Sensors

NOTES

Hall-effect sensors (HES) monitor the DC current at the output of the TRUs, external DC ground power source, and main aircraft batteries (Figure 24-35). The DC current sensors monitor the TRU current (left main TRU, left essential TRU, right main TRU, right essential TRU, and the auxiliary TRU) and provide inputs to the BPCUs for overcurrent protection and percent load information. The DC current sensor monitoring the external DC ground power source provides input signals to the RBPCU for overcurrent protection and percent load information. The DC current sensors that monitor the main aircraft batteries provide input signals directly to the battery volt/amp meters in the COP and to DGIO 1 in MAU 1 and DGIO 2 in MAU 2 for display on the CAS.


24-83



BUS POWER CONTROL UNIT PROTECTION

LEFT ESS TRU

1. OVERVOLTAGE 32.2 VDC 2. UNDERVOLTAGE 21 VDC 3. OVERCURRENT 270 AMPS

LEFT ESS DC BUS

HE NO. 3

RESET LEDC

LEFT BPCU

DC RESET LIGHT RELAY

Figure 24-36. BPCU Protective Trips

24-84


DC



System Protection

NOTES

The left and right BPCUs provide control and protective functions for the DC transformer rectifier units (TRUs), DC buses, and contactors, and are listed as follows (Figure 24-36):

TRU Overvoltage Protection The BPCU overvoltage (OV) protection function monitors the DC voltage supplied by the TRUs. If the voltage is more than 32.2 ±1 VDC, the BPCU will trip the L/RMDC, L/REDC, or ADC.

TRU Undervoltage Protection The BPCU undervoltage (UV) protection function monitors the DC voltage supplied by the TRUs. If the voltage is less than 21 ±1 VDC, the BPCU will trip the L/RMDC, L/REDC, or ADC in 6.0 ±0.25 seconds.

TRU Overcurrent Protection A Hall-effect sensor (HES) is installed on the source side of the TRU contactors to monitor the current. If the current is more than 270 ±20 amps, the L/RMDC, L/REDC, or ADC is tripped. When a DC contactor is tripped offline because of a protective fault, the DC reset light illuminates and a “L/R DC RESET” CAS message is generated.


24-85



LEFT MAIN AC

RIGHT MAIN AC

RMTAC

LMTAC


RIGHT BUS POWER CONTROL UNIT RMDC

LMDC

LEFT MAIN DC

RIGHT MAIN DC REDC

LEDC

LEFT ESS DC

RIGHT ESS DC

Figure 24-36. Primary DC Contactor Control

LEFT MAIN AC

RIGHT MAIN AC ATAC 1 EXT DC ATAC 2

EDC/ADC LMDXC

RMDXC

LEFT MAIN DC

RIGHT MAIN DC

RBPCU LBPCU

LEDXC

REDXC

LEFT ESS DC

RIGHT ESS DC

Figure 24-37. DC Crosstie Contactor Control

24-86




PRIMARY DC CONTACTORS

NOTES

The primary contactors (LMDC, LEDC, RMDC, REDC) route power from the TRU to its dedicated bus and are controlled by their onside BPCUs (Figure 24-37).

Crosstie Contactors The crosstie contactors (LMDXC, LEDXC, RMDXC, and REDXC) are controlled by the opposite side BPCU and route power from the AUX TRU or external DC power receptacle to t h e bu s ( F i g u r e 2 4 - 3 8 ) . L M D C , L E D C , LMDXC, and LEDXC are located in the left PDB; and RMDC, REDC, RMDXC, and REDXC are located in the right PDB.


24-87



Figure 24-39. AC Contactor Location

24-88




AC CONTACTORS

ATAC1 is controlled by the RBPCU and determines whether the AUX TRU is powered by the left main or right main AC bus. Normal source is the left main AC. If the left main AC bus drops below 102 VAC, the RBPCU will command ATAC1 to energize, routing power to the AUX TRU from the right main AC bus. ATAC2 (Figure 24-40) is normally deenergized and routes power from ATAC1 to the AUX TRU. If the hydraulic motor generator system is in use, ATAC2 is energized to route HMG AC power to the AUX TRU.

T h e AC c o n t a c t o r s c o n s i s t o f R M TAC , LMTAC, ATAC1, and ATAC2. LMTAC and ATAC1 are located in the left PDB (Figure 24-39), RMTAC is located in the right PDB (Figure 24-39), and ATAC2 is located in the standby electrical power junction and relay panel (Figure 24-40).

ATAC 2

038K2 082K1

038XK2

036K4 038K1

036K2 036K3

036XK4 038XK1 036XK3

082XK1

E1A

036K1

034K1

E2E

036S2 036S1 034K8 TJ2

K

L

M

N

M

N

P

04A

E

166

L

D

166

K

C

02C

B

01D

A

04A

E3B 329A1

TJ1

B

03D

A

R

S

T

U

Y

036S3

ATAC 1

Figure 24-40. Standby Electrical Power Junction and Relay Panel


24-89



LEFT MAIN AC BUS

RIGHT MAIN AC BUS

To EICAS

115-VOLT AC 400 HZ

TRU

DIM TEST

L MAIN

R MAIN

R AC

L AC

LMTAC

28-VDC TRU/ E-INV CONT NO. 1 CB LEER

LEFT MAIN TRU

Figure 24-41. Left Main TRU AC Control

24-90




TRU Switches

NOTES

LMTAC controls the AC power source to the left main TRU (Figure 24-41). Normally deenergized, it routes power from the left main AC bus to the left main TRU. When energized by selecting the L MAIN TRU switch on COP, it routes the right main AC bus power to the left main TRU. A blue L TRU–R AC message on the CAS and an amber R AC illuminated in the switch are displayed in this configuration. RMTAC controls the AC power source to the right main TRU. Normally deenergized, it routes power from the right main AC bus to the right main TRU. When energized by selecting the R MAIN TRU switch on COP, it routes the left main AC bus power to the right main TRU. A blue R TRU–L AC message on the CAS and an amber L AC illuminated in the switch are displayed in this configuration.


24-91



RIGHT MAIN AC BUS

LEFT MAIN AC BUS

LMTAC

RMTAC

ATAC 1 HMG ATAC 2

LEFT MAIN TRU

LEFT ESS TRU

EDC/ ADC

APC

LMDC

LEDC

RIGHT ESS TRU

RIGHT MAIN TRU

AUX TRU

RMDC

REDC

RMDXC

LMDXC

R MAIN DC BUS

L MAIN DC BUS

LEDXC

REDXC REDBC

LEDBC

GSBC 1 L ESS DC BUS

BC 1

R ESS DC BUS

BC 2 BATT TIE BUS

GSBC 2

LBCC

AUX HYD PUMP CONT

APU START CONT

GND SRV BUS

RBCC R BATT CHGR

L BATT CHGR

Figure 24-42. DC Power System Diagram

24-92




DC POWER SOURCE DISTRIBUTION DC source distribution is an automatic control function of the bus power control units based on bus source priority logic. Understanding this logic is essential to understanding system operation and isolation of faults during troubleshooting (Figure 24-42). DC bus priority logic for the main DC buses are as follows: • Dedicated Main TRU • AUX TRU • External DC power DC bus priority logic for the essential DC buses are as follows: • Dedicated essential TRU • AUX TRU

The AUX TRU can be also receive power from the hydraulic motor generator (HMG) if necessary, and, in turn, supply power to the left or right or both essential DC buses through the standby electrical power system.

NO-BREAK POWER TRANSFER The DC power system design provides for NBPT to eliminate power interrupts during normal operations. Minimization of power interrupts on the DC buses is primarily a result of AC power NBPTs. NBPT will occur when power on a bus is transferred to a higher priority source, such as, the essential DC buses transfer from battery power to TRU power. An unlikely exception to this rule is when transferring from external DC to AUX TRU. This is because EDC/ADC is a common contactor and must break before make. Power transfers from a higher priority to a lower priority and failure-related transfers result in break-before-make transfers.

• External DC power

OPERATION

• Batteries Other rules of BPCU logic are as follows: • The essential buses have priority over the main buses. • The left bus has priority over the right bus. The AUX TRU normally receives power from the left main AC bus, via ATAC1, but if power is lost on that bus, automatic switching to the right main AC bus will occur. It should be noted that when the AUX TRU is receiving power from a source other than the HMG, it can only supply output power to one DC bus at a time. Priority for this logic is as follows: • Left essential DC bus

A TRU is dedicated for each of the four major DC buses. Each TRU receives 115 VAC, threephase, 400 Hz power from the integrated drive generators, APU generators, or AC external power. The power is then converted to 28 VDC nominal. Should a main AC bus fail, the left and right main TRUs have redundancy provided by relays, which, when energized, connect that TRU to the opposite main AC bus. The right main TRU and right essential TRU are located above the floor, below the right electronic equipment rack (REER). The left main, left essential, and auxiliary (AUX) TRU are located below the floor. The AUX TRU operates in a standby status to automatically provide power to any one of the four major DC buses if its dedicated TRU fails.

• Right essential DC bus • Left main DC bus • Right main DC bus


24-93



Figure 24-43. Main Aircraft Batteries

24-94




BATTERY POWER

NOTES

The main batteries provide basic DC power to perform the following primary functions: • Start the APU. • Operate the auxiliary hydraulic system (AUX) pump. • Operate the ESS DC buses if the MAIN BATTERIES switches are selected to ON and the ESS DC buses have no other power source. The right main battery has an additional function on the ground. It is used to power the ground service bus if it is not powered by right main DC bus or external DC power. The main batteries are located aft of APU in the tail compartment, left and right of centerline (Figure 24-43). These batteries are manufactured by SAFT, rated at 24 VDC with a nominal capacity of 53 amp-hours produced by 21 nickel-cadmium cells. The batteries contain a thermal switch and a temperature sensor (output to charger).

BATTERY POWER SYSTEM COMPONENTS The components, controls, and indicators of the battery power system are the main batteries, contactors, Hall-effect sensors, voltmeter/ammeter, battery switches, external battery switches, and battery chargers.

Battery Contactors The main battery contactors, BC1 and BC2, are located in the auxiliary power relay box and route power from the batteries to the battery tie bus.


24-95



L MAIN AC

R MAIN AC MAIN BATTERIES

24.0 -50 VOLTS/AMPS

ON

ON

LEFT

RIGHT

24.0 -50 VOLTS/AMPS

- +

- +

NO. 1 BATTERY CHARGER


TO EICAS

TO EICAS

RBCC

LBCC TIE BUS BC 1

BC 2

HE 6

HE 7

AUX POWER RELAY BOX + ñ

+ ñ

NO. 1 BATTERY

NO. 2 BATTERY

Figure 24-44. Battery Indication

24-96




Main Battery HallEffect Sensors

NOTES

Hall-effect sensors (Figure 24-44) are located in the auxiliary power relay box and sense the current flow in and out of battery as well as providing input to the battery ammeter and data acquisition unit (DAU).

Voltmeters/Ammeters The voltmeters/ammeters are located on EPCP just outboard of the battery switches. Both indicators are installed as a single LRU. Their liquid crystal display (LCD) indicates individual battery voltage and current charge/discharge and receive inputs from the left and right battery bus A and Hall-effect sensors.


24-97



MAIN BATTERIES

24.0 –35

ON

ON

LEFT

RIGHT

VOLTS/AMPS

VOLTS/AMPS

COCKPIT OVERHEAD PANEL

ICS JACK

GND SVC BUS CLOSE PUSH TO TEST

ON

OFF OUTSIDE DOOR SWITCH

OFF GND SVC BUS SWITCH

EXTERNAL BATTERY

Figure 24-44. FWD External Switch Panel

24-98

24.0 –35




Battery Switches

NOTES

The battery switches are located at the bottom of the EPCP. These switches illuminate amber ON when selected and the associated battery contactor is energized closed. They are not illuminated when the associated battery contactor is open or the switch is selected off. The external battery switch is located on the forward external switch panel (Figure 24-45). Turning the switch on connects both batteries to the battery tie and essential DC buses. The external door control switch is also located in the forward external switch panel. Selecting the switch to ON connects the right battery only to the battery tie bus and essential DC bus for door control and auxiliary hydraulic pump operation.


24-99

24-100

EXT BATT SWITCH

B A T T

RIGHT BATTERY BUS A

ON RIGHT BATTERY BUS B

RIGHT R BATT CONTROL


T I E B U S

RIGHT BATTERY

R CHARGER CONTROL

BC 2

TO RIGHT POWER DISTRIBUTION 80X

RIGHT MAIN BATTERY SWITCH

LEDC

R BATT CHARGER CNTL RELAY

LEDXC L PDB

REDC

P/O APU PILOT RELAY

REDXC R PDB

1

R BATT CNTL RELAY

AUX HYD PUMP CONT RELAY NOTE: TDODE 100MS = TIME DELAY ON DEENERGIZE

Figure 24-46. Battery Control

international

RIGHT BATTERY SYSTEM SHOWN, OPERATION IS THE SAME FOR THE LEFT SYSTEM. THE LEFT SYSTEM IS NOT AFFECTED BY THE DOOR CONTROL SYSTEM.

FlightSafety

1


OUTSIDE DOOR SWITCH

TO BATTERY BUS LOADS

R BATT BUS B RELAY



Battery Switches Operation

APU START CONTROL

Selecting the battery switch to the depressed, ON position routes power from the battery bus A to the battery bus B relay (energizing battery bus B), the battery contactor control circuit, and the battery charger control circuit to energize the battery charger contactor. The power on battery bus B will be used to energize REDBC. When the battery switch is selected to ON, power is also routed through t h e d e e n e rg i z e d c o n t a c t s o f R E D C a n d REDXC, which energizes the coil of the battery control relay (Figure 24-46). When the battery control relay energizes, battery power is routed to battery bus A, through the contacts of the battery control relay, to the coil of the battery contactor. With power to the coil, the battery contactor will energize, routing battery power to the battery tie bus and on to the essential DC bus through the essential DC battery contactor. The battery switch on the COP will illuminate an amber ON when its associated battery contactor is energized.

On APU start, when the APU starter is engaged, the APU pilot relay is energized. With the battery switches depressed, power is routed through the APU pilot relay to the battery control relay closing the battery contactors, illuminating the battery switches, and powering the battery tie bus.

AUXILIARY HYDRAULIC PUMP CIRCUIT When the auxiliary hydraulic pump is selected on, the auxiliary hydraulic control relay routes power directly to the battery contactors, illuminating the battery switches, and powering the battery tie bus.

NOTES

When a higher priority source becomes available for the essential DC bus, the BPCU will route that power by energizing REDC or REDXC. This causes the battery control relay and battery contactor to deenergize, causing the switch to go out and removing battery power from the battery tie bus. An outside battery switch is used to provide power for ground maintenance without entering the aircraft. The switch, when selected to ON, delivers power to the battery tie bus from both batteries through battery contactors BC1 and BC2. An outside door control switch is used to provide power from the right battery for the hydraulic system for the purpose of closing the main entrance door from outside the aircraft.


24-101



LEDC

EDC/ADC

LEDXC

REDC

REDXC

R ESS DC

L ESS DC

LEDBC

REDBC

BATT TIE BUS + –

+ ñ

LEFT BATT

RIGHT BATT

L BATT BUS B C/B

R BATT BUS B C/B

BC 2

Figure 24-47. REDBC Control

24-102




DC Bus Power Priority

NOTES

A function of the main batteries is to power the essential buses when a higher priority source is not available. When the battery switch is selected to ON, battery bus B becomes powered. Assume that the battery switches are on, and only battery power is available. Battery contactors close because REDC and REDXC are in the relaxed position. Power from battery bus B is routed through the contacts of battery contactors to REDBC (via REDXC and REDC). REDBC energizes and routes battery tie bus power to the ESS DC bus. When a higher priority source of power becomes available through the energized REDC or REDXC, the contacts break, which deenergizes the coil of REDBC, disconnecting the battery tie bus from the ESS DC bus (Figure 24-47).


24-103



CHARGER BATTERY

UP D

FW

Figure 24-48. Battery Chargers

24-104




BATTERY CHARGERS

NOTE

Two battery chargers (Figure 24-48) are located in the tail compartment just forward of the tail compartment doorway. The top and bottom of each battery charger have cooling holes, and the sides are fitted with external finned heat sinks to provide convection cooling. Each charger has two external connections located on the front panel. There is a DC power output connection (5/16” negative and 3/8” positive terminals) and an input connector (for AC input power, control, and sense). The front panel also has two health status lightemitting diodes (LED). The chargers operate on 115 VAC, 400 Hz, three-phase input power. The left main AC bus powers the left charger, and the right main AC bus powers the right charger. Two LEDs on the face of each battery charger show the status of the battery and charger. The charger LED illuminates when acceptable power is supplied to the charger. The charger LED will go out when acceptable input power is no longer available or if the charger has an internal fault. These faults include gate drive failure, IGBT (insulated gate bipolar transistor) circuit failure, output rectifier failure or a shorted output. The batter LED is illuminated when the battery is charged and operational. The battery LED will go out when there is no input power to the charge, the battery temperature is too high, the battery sensor is open or shorted, the charging time is too long, or if the battery sense connector becomes disconnected. The charger provides temperature compensated charging in the charge mode or an output of a constant 28.75 VDC in the transformer-rectifier mode of operation.


24-105



BATTERY CHARGER

BATTERY P1

P1

INPUT POWER ØA

4

INPUT POWER ØB

7

INPUT POWER ØC

10

4 115 VAC 400 HZ, 3Ø

JUMPER

1

6 8

3

9

BATTERY TEMP SENSE

11

11

BATTERY TEMP SENSE

12

12

FAULT INDICATION

9

THERMOSTATIC SWITCH

CHARGER INTERLOCK

2 THERMISTOR

TO CAS

R1 1

2 CHARGE MODE CONTROL GROUND=CHARGE MODE OPEN=TR MODE

8

BATTERY CHARGER CONTACTOR

P2

1

32.4K, % OHMS. R1 AND R2 ARE FOR TYPE ONE BATTERIES ONLY. THE THERMOSTATIC SWITCH FORMS A LOOP BETWEEN PINS 8 AND 9 ON TYPE TWO BATTERIES.

2

2460 OHMS AT 73.4°.

Figure 24-49. Battery and Charger Diagram

24-106


R2



CHARGE MODES

NOTES

In the charge mode, the battery charger initiates a charge cycle in a constant current mode. The output voltage will vary with battery state of charge and battery temperature. Initially it will be about 28 VDC, and the current will be about 38 amps. As the battery state of charge increases, the output voltage increases to a 12% overcharge voltage (32.2 volts), then drops to 28.75 volts. After this drop, the charger is in the constant voltage mode (Figure 24-49). The battery charger will initiate a new charge cycle when one of the following occurs: • AC input power is applied or has been interrupted for more than 500 milliseconds and reapplied. • Sensed battery voltage drops below 23 volts. • The charger has operated for more than 500 milliseconds in the TR mode and is switched to the charge mode.


24-107



L MAIN AC

R MAIN AC MAIN BATTERIES

88.8 -888 VOLTS/AMPS

ON

ON

LEFT

RIGHT

88.8 -888 VOLTS/AMPS

- +

- +



TIE BUS

AUX POWER RELAY BOX + ñ

+ ñ

NO. 1 BATTERY

NO. 2 BATTERY

Figure 24-50. Battery Charger TR Mode Control

24-108




Battery Charger TR Mode In the transformer-rectifier (TR) mode of operation, (which is a function of the battery contactor position providing a ground to pin 8 of charger), the charger output is 28.75 volts, constant potential, at up to 50 amps. This voltage will maintain the battery charge without unnecessary electrolyte loss. The TR mode is initiated if the battery switches are selected to ON, the main AC buses are powered, and the following occur: • The AUX pump is turned on • The APU start switch is selected to ON, or • The essential DC buses require power from the batteries

NOTE The BATT switchlights will illuminate to ON when the charger is in the TR mode.

The battery chargers will shut down and an amber “L-R BATT CHGR FAIL” CAS message will illuminate for the affected battery circuit if any of the following conditions occur: • Sense connector (small connector) disconnected—When the interconnecting sense and control cable between the battery charger and the battery is not connected to the battery, an open in the battery interlock circuit will cause the charger to shut down. The system will return to normal operation when the cable is reconnected. Therefore, no resetting is necessary unless the battery power connector is also disconnected.

• Overcurrent—Approximately 65 amps will cause the charger to shut down. An internally caused overcurrent condition or inverter circuit imbalance causes a cyclic shutdown/softstart operation. The overcurrent shutdown signal has a fixed time duration, and when it clears, a soft start will be applied to the inverter. The cyclic shutdown/soft start will repeat as long as an overcurrent condition or inverter circuit imbalance is present. • Overtemperature—A battery temperature of 145°F or above will cause the battery charger to shut down. When the battery cools to below 135°F, the system will return to normal operation. • Loss of 115 VAC input power will cause a shutdown, although no failure may be occurring in the battery/battery charger system.

NOTE The normally discharged battery (above 7 ±1.5 VDC) recharges within approximately 90 minutes. If the charge command continues beyond this maximum recharge time and approaches 100 minutes, the unit will switch into the constant potential mode of operation, topping the constant current charge. If the battery charger detects that the battery is in this discharged condition (<7 ±1.5 VDC), it will not attempt to charge the battery.

• Overvoltage (134 VAC) or undervoltage (94 VAC)—Although the battery charger will shut down when the AC input is over or undervoltage, the system will return to normal operation 10 to 25 seconds after an overvoltage condition clears. It will return to normal immediately when an undervoltage condition clears.


24-109



RIGHT BUS POWER CONTROL UNIT

EXTERNAL DC POWER

28 VDC INTERLOCK SENSE EXT PWR 28 VDC INPUT EXT DC HES

ANN LTS PWR 28VDC ON

DIM

EXT DC PWR AVAIL LIGHT

AUX TRU

TEST

AVAIL

EDC/ADC ON EDC/ADC ON EXT DC PWR EXT DC PWR

DIM

EXT PWR

TEST

HE 5 EDC / ADC

DC EXT AVAIL LT RELAY

RIGHT PDB

LEFT MAIN DC

LEFT ESS DC LEDXC

LMDXC

RIGHT MAIN DC RMDXC

REDXC

Figure 24-51. DC External Power Circuit

24-110

COMMAND COMMAND RTN ON COMMAND ON RETURN


RIGHT ESS DC



DC EXTERNAL POWER External DC power can be used to power the aircraft’s main and essential DC buses (Figure 24-51). It can also be used to power the battery tie bus for APU starts and auxiliary hydraulic pump operation (Figure 24-55). Additionally, the DC external power can be used to power the ground service bus when power is not available from the right main DC bus. Power from an external DC power cart can also power the main and essential DC buses. The external DC power receptacle is located on the forward right side of the fuselage. The external power switch is located on EPCP. The right bus power control unit is located in the REER, the EDC/ADC contactor is located in the RPDB, and the hall effect sensor for the DC external power is located in the RPDB. When an external DC power cart is connected to the aircraft and turned on (Figure 24-52), the voltage and interlock are checked by RBPCU. If voltage and interlock is good, the RBPCU energizes the available relay, the AVAIL light will illuminate (assuming the battery switches are on). When the external power switch is depressed, an ON command is sent to the RBPCU. The RBPCU commands the EDC/ADC to close. With the EDC/ADC closed, power from the external source is supplied to the main and essential buses, and the external power ON light is illuminated. With external power applied, the RBPCU monitors voltage, polarity, current, and interlock. The small pin on the external power connector provides interlock protection. This ensures that the plug is fully engaged before current flows through it. A BPCU lockout caused by any of the protective functions will result in illumination of the DC reset switch and opening of EDC/ADC. External DC power can be used to start the APU in the event of a dead battery as well as running the auxiliary hydraulic pump. This is accomplished through control of the auxiliary power contactor (APC). Certain conditions must be present before the APC can be

energized. Both batteries must be switched off and the APU start circuit or the auxiliary hydraulic circuit must be powered. When the APC is energized, power is routed from the external DC source to the battery tie bus. External power can also be used to power the ground service bus.

DC External Power Protective Functions (RBPCU) The DC external power protective functions are provided by the RBPCU as follows:

DC Interlock Protection The DC external power plug has three pins; +28-VDC/BPCU, +28 VDC sense/interlock, and the +28 VDC return. The +28 VDC sense/interlock pin is shorter than the other two. The 28 VDC signal on this pin is monitored by the RBPCU for power quality, in addition to supplying the BPCU power and interlock. This pin is shorter so connection to the DC external power supply is established prior to the RBPCU sensing the 28 VDC. The BPCU commands the EDC to open in 20 milliseconds if the interconnect is lost.

Polarity Protection When external DC power is first applied through the external DC power plug, the RBPCU verifies the power quality before it allows the EDC to close. One of the power quality checks is for correct polarity. The RBPCU is able to sense incorrect polarity for a voltage of more than 5 ±1 VDC and respond within 60 milliseconds maximum. If incorrect polarity is sensed, the RBPCU will inhibit EDC closure, and a fault message will be generated as long as a source of power is available to the BPCU.

Overvoltage Protection The RBPCU overvoltage (OV) protection function monitors the external DC power voltage. If the voltage is more than 32.2 ±1 VDC, the overvoltage protection function will trip the external DC contactor (EDC) after an inverse


24-111



STANDBY ELECTRICAL POWER

MASTER

L ESS

R ESS

ON

ON

ON

ATAC 2 REAC

HMG

HYD

LEAC ESS AC BUS ØA

AUX TRU

EMERG INVERTER (ØA ONLY)

LH ESS DC BUS

LSAC

HE 8 LH STBY AC BUS

LEDXC

REDXC

LEDBC

REDBC

LEIDC

RH STBY AC BUS

RH ESS DC BUS

REIDC BC 1

BATTERY TIE BUS

BC 2

HE 7

HE 6

LEFT BATTERY

RIGHT BATTERY

Figure 24-52. Standby Power Circuit—HMG ON

24-112

GCU

RSAC




time delay.

Undervoltage Protection The RBPCU undervoltage (UV) protection function monitors the external DC power voltage. If the voltage is less than 21 ±1 VDC, the undervoltage protection function will trip the EDC after 5.0 ±0.25-seconds time delay.

Overcurrent Protection A Hall-effect sensor (HES) is located on the source side of the EDC to monitor external DC power. An overcurrent (OC) condition exists when the current is more than 320 ±20 amperes for an inverse time delay. The RBPCU overcurrent protective function sends a request to the LBPCU to open RMDXC. If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds after the request was sent to the LBPCU, then the RBPCU opens LMDXC. If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds later, then the RBPCU sends a request to the LBPCU to open REDXC. If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds after the request was sent to the LBPCU, then the RBPCU opens LEDXC. If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds later, then the RBPCU will open the EDC side of the EDC/ADC.

SUPPLEMENTAL POWER INTRODUCTION The supplemental power systems are power sources that are not provided by the main AC and DC power systems. These systems provide standby electrical power, emergency power, 60 Hz power, cabin power, galley power, and power for the ground service bus.


24-113



Figure 24-53. Hydraulic Motor Generator

24-114




STANDBY ELECTRICAL POWER SYSTEM The purpose of the standby electrical power system (SEPS) is to provide a supplemental source of electrical power in order to power the essential DC buses, the left and right standby AC buses, and the essential AC bus ØA. This is desirable in order to extend operational time if all main generating power sources have failed. The standby electrical power system (SEPS) consists of a 10 KVA, 115 VAC, three-phase hydraulically-driven motor generator (HMG) and an associated GCU (Figure 24-52).

Hydraulic Motor Generator The HMG (Figure 24-53) uses a hydraulicpowered variable displacement motor to drive a 115 VAC, 3-phase, 400 Hz, AC generator. The HMG, located in the main wheel well, is powered by left hydraulic system fluid and pressure from the left system or the power transfer unit (PTU). The hydraulic motor generator is controlled by the HMG GCU.

The components, controls, and indicators for the standby electrical power system consist of the hydraulic motor generator (HMG), HMG GCU, ATAC 2, L/R SACs, HMG ON relay, HMG overload relay, HMG overload sensors, current transformer assembly, and standby electrical power control panel.


24-115



Figure 24-54. Underfloor Equipment Locator—Hydraulic Motor Generator Control Unit

24-116




NOTES

HMG Generator Control Unit The HMG generator control unit (GCU) (Figure 24-54) is located in the underfloor electronics equipment area. The generator control unit provides voltage regulation, frequency regulation, control, and fault protection including under/overfrequency, under/overvoltage, and overspeed. The GCU controls a solenoid valve for on/off control, and a servo valve to regulate motor speed at 8,000 rpm in order to produce a 400 Hz AC output from the generator. Electrical power from a permanent magnet generator in the HMG supplies primary electrical power for the GCU during normal operation. The PMG also provides feedback to the GCU for speed control.

GCU Protective Functions The GCU monitors the generator output voltage at a point just prior to the ATAC2 contactor, referred to as the point of regulation (POR) and controls the magnetic field in the generator stator so the output voltage at the POR is regulated at 115 ±2 VAC. When the GCU determines that good power is available from the generator, it provides a 28 VDC powerready signal that is used to place the HMG online with the AUX TRU. The GCU also provides fault protection for under/overfrequency, under/overvoltage, and overspeed (fast overfrequency). Delays from 0-6 seconds prevent nuisance trips from occ u r r i n g d u r i n g n o r m a l t r a n s i e n t eve n t s . Anytime a fault (except overspeed) is detected, the GCU removes the power-ready signal, which takes the HMG offline and initially allows the HMG to continue running. During the next 10 seconds, the following two events may occur: • Power quality is restored at the POR, in which the power-ready signal is reapplied and system operation continues as normal. If a second fault is detected within 60 seconds of the onset of the first, the GCU performs an immediate shutdown to prevent cycling.


24-117



HMG OVERLOAD RELAY (REF: STBY ELEC PWR)

038K2 082K1

038XK2

036K4 038K1

AUX TRU AC CONTACTOR (ATAC 2) (REF: STBY ELEC PWR)

HMG ON RELAY (REF: STBY ELEC PWR)

RSAC

LSAC

036K2 036K3

036XK4 038XK1 036XK3

082XK1

E1A

036K1

034K1

E2E

036S2 036S1 034K8 TJ2

K

L

M

N

M

N

P

R

036S3

ATAC 1

HMG OVERLOAD SENSORS

Figure 24-55. Standby Junction and Relay Panel

24-118


04A

E

166

L

D

166

K

C

02C

E3B 329A1

B

01D

A

04A

B

03D

A

TJ1

S

T

U

Y



• Power quality is not restored at the POR, in which the GCU performs an immediate shutdown. In the case of overspeed, which could occur rapidly in the event of speed control failure, the GCU will shut down with no delay to prevent mechanical damage to the HMG. If at anytime the HMG speed falls too low for the PMG to provide adequate voltage to properly power the GCU, the GCU will attempt to temporarily revert to aircraft power until the PMG can supply adequate power. If no aircraft power is available, the GCU will shut down. To enter the “shutdown” state the GCU closes the solenoid valve, thus stopping the HMG. Once shutdown has occurred, the GCU will not allow the HMG to be started until it has been reset. Resetting the shutdown state can be achieved by momentarily removing 28-VDC power to the GCU for at least a second. This requires the SEPS MASTER switch be cycled off then back on.

Overvoltage Protection The HMG GCU overvoltage (OV) protection circuitry senses each of the phase voltages at the point of regulation (POR). If the voltage on any phase exceeds 128.5 ±3 VAC, the HMG GCU will de-excite the generator after an inverse time delay.

Undervoltage Protection The HMG GCU undervoltage (UV) protection circuitry senses each of the phase voltages at the point of regulation (POR). If the voltage on any phase is less than 104.5 ±3 VAC, the HMG GCU will de-excite the generator after an inverse time delay.

Overfrequency Protection The HMG GCU overfrequency (OF) protection circuitry senses each of the phase frequency at the point of regulation (POR). If the frequency on any phase exceeds 425 ±3 Hz, the HMG GCU will de-excite the generator after an inverse time delay.

Underfrequency Protection The HMG GCU underfrequency (UF) protection circuitry senses each of the phase frequency at the point of regulation (POR). If the frequency on any phase is less than 375 ±3 Hz, the HMG GCU will de-excite the generator after an inverse time delay.

Overcurrent Protection Overcurrent protection is provided by three 30amp overload sensors located in the standby electrical power junction and relay panel (Figure 24-55). The overload sensors will cause the HMG to shut down if the load on any of the three phases at the output of the HMG exceeds thirty amps.

Overload Sensors Thirty-amp overload sensors provide protection of each phase of the output of the HMG. Each sensor has a SPDT set of contacts that operate when an overload condition exists. The sensors are wired such that when any one of the sensors detects an overload, the HMG OVERLOAD relay will energize and latch itself through a holding circuit. This interrupts the power-ready signal from the GCU, which takes the HMG offline (HMG ON relay deenergizes, LEDXC and REDXCs deenergize, ATAC 2 deenergizes, LSAC and RSAC deenergize) until the circuit is reset by switching the HMG MASTER switch off and back on. An amber “HMG OVERLOAD” CAS message also appears, based on a discrete received at custom I/02 in MAU 2.

Current Transformer Assembly A current transformer assembly (CTA) is provided to monitor the AC load for display on the AC POWER and SUMMARY synoptic pages. Phase A of the POR is directed to custom I/02 in MAU 2; to monitor phase and voltage for display on the AC POWER and SUMMARY synoptic pages.


24-119

24-120

HMG CONT #2 L ESS DC BUS HMG CONT #1

ANN LTS TEST RELAY

DIM

28 VDC ANN LTS POWER

ON 28 VDC ANN LTS POWER

TEST

BYTE OUTPUT BYTE INPUT

RETURN TO ATAC 1

HMG OVRLD RELAY

TO ATAC 1

POWER READY


SERVO A SERVO B

TO ATAC 1

STBY ELECT PWR "L ESS" SWITCH

SOLENOID POWER

L MAIN BUS DIM ATAC 2

ØA POR ØB POR ØC POR

A

TO LEDXC

B

SOLENOID +

PMG ØA PMG ØB PMG ØC

ØA PMG ØB PMG ØC PMG

FIELD + FIELD -

FIELD + FIELD -

CHASSIS GND

C

L STBY BUS

R MAIN BUS

HMG GEN CONTROL UNIT ØA

STBY ELECT PWR "R ESS" SWITCH

ØB ØC NEUTRAL

L SAC DIM

ON

AC PWR OUT

N ØA ØB ØC TEST

L ESS HMG ON C

% LOAD HMG OVERLOAD

B

HMG A PHASE MAU 2

A

R STBY BUS R SAC

CUSTOM I/O MODULE SLOT 11 R ESS

MAU 1 DG I/O (1) SLOT 9/10

HMG ON R ESS HMG ON MAU 2

Figure 24-56. HMG System Diagram

international

DG I/O (2) SLOT 7/8

FlightSafety

TO REDXC

SERVO A SERVO B

SOLENOID CHASSIS

ON TEST

HYDRAULIC MOTOR GEN


START/STOP

TO AUX TRU

28 VDC ANN LTS POWER

TEST GCU OK

HMG ON RELAY

STBY ELECT PWR MASTER SWITCH

HMG GCU TEST SWITCH



Relays

HMG Test Switch

ATAC 2

The HMG test switch, located on the LEER system monitor panel, provides a way to test the HMG GCU’s protection circuits and power supplies without running the HMG.

The ATAC 2 switching contactor connects the output of the HMG to the AUX TRU for powering the essential DC buses and is located in the standby power junction and relay panel (Figure 24-55).

LSAC and RSAC The LSAC and RSAC connect the HMG output to the L and R standby AC buses. The LSAC and RSAC are also located in the standby power junction and relay panel (Figure 24-55).

HMG ON Relay The HMG ON relay located in the standby junction and relay panel routes the HMG POWER READY signal to energize the HMG ON relay.

Controls The controls and indicators for the standby power system are located on the EPCP section of the cockpit overhead panel. The SEPS is controlled by three guarded switches on the EPCP. These three switches are the MASTER, L ESS, and R ESS. The MASTER switch activates the HMG and allows the HMG GCU to energize the HMG ON relay. The MASTER switch illuminates an amber ON when selected. The L ESS and R ESS switches provide cockpit control for powering the essential DC buses and illuminate the amber ON legends when selected (Figure 24-56).

Operation and Indications

HMG Overload Relay The HMG overload relay removes power from the HMG ON relay during an overload condition (>30 amps).

Before the SEPS is selected on, 28 VDC control power from the left essential DC bus or the right battery bus B is provided to sets of contacts of the HMG ON relay and the MASTER, L ESS, and R ESS switches. When the GCU is selected on via the MASTER switch, it uses 28 VDC control power for startup, until the PMG can supply adequate power. It is also used to illuminate the MASTER switch on. The MASTER switch also provides an OPEN/GRD logic to DGIO 2 in MAU 2 for a blue “HMG SW ON” CAS message. Upon receiving the 28 VDC startup command, the GCU opens a solenoid valve to start the operation of the HMG. At this point, the PMG allows the GCU to provides its own power to operate and uses the 28 VDC control power as an on–off discrete. The GCU monitors the generator output voltage at a point just prior to the ATAC2 contactor, referred to as the point of regulation (POR) and controls the magnetic field in the generator stator so the output voltage at the POR is regulated at 115 ±2 VAC. When the GCU determines that good power is available from the generator, it provides a 28 VDC power-


24-121



Figure 24-57. System Monitor Test Panel

24-122




ready signal to energize the HMG ON relay. When the HMG ON relay is energized, power is routed to contacts of the L and R ESS switches; ATAC2 will energize, directing 115 VAC, three-phase, 400 Hz power from the HMG to the AUX TRU. Once the HMG ON relay has energized ATAC 2, LSAC, and RSAC will automatically be energized to route threephase AC power to the standby AC buses. Once the HMG is supplying power to the AUX TRU, the R ESS and L ESS switches on the standby electrical power panel can be used to direct AUX TRU power to the R ESS and/or L ESS DC buses, through the REDXC and LEDXC contactors, respectively. Selecting the R ESS switch causes the switch to illuminate an amber ON and sends an open/ground logic to DGIO 2 in MAU 2 to display the blue “R ESS HMG SW ON” CAS message. Selecting the L ESS switch causes the switch to illuminate an amber ON and sends an open/ground logic to DGIO 1 IN MAU 1 to display the blue “L ESS HMG SW ON” CAS message.

HMG GCU BITE Test An alternate action GCU TEST switch, located on the LEER system monitor panel (Figure 24-57), can be used to cause the HMG GCU to start its self-test of the protection circuits and power supplies without starting the HMG. When the GCU TEST switch is selected to the depressed position, a blue TEST legend will illuminate on the upper half of the switch while the GCU starts its self-test of the protection circuits and power supplies. After 5 to 6 seconds, if the test is successful, a green GCU OK legend will illuminate on the lower half of the switch. The legend will stay illuminated until the GCU TEST switch is switched off (not depressed).


24-123



XL249

ENERGY LEVEL NOTE: LVL TEST VALID WHEN AIRCRAFT POWER HAS BEEN OFF LONGER THAN ONE HOUR.

LEVEL TEST

LOW MID FULL ENERGY LEVEL FAULT BITE TEST BATT

CHRG SWITCH HTR

FAULT RESET

CB1

CB2

CB3

20

5

5

NOTE: PULL CB1 WHEN SHIPPING OR STORING

XL249

TUCSON, AZ

XL249 P.N. 100-0302-01

S.N.

MOD

DOM

GULFSTREAM AEROSPACE CORP. GAC P/N 1150SCAV524-3

DO-190C

ENV.

CAT.

A1-CA8XXXXXFXXAAAZRZ(A3C3)XX

Figure 24-58. Emergency Power Battery Pack

24-124




NOTES

EMERGENCY POWER SYSTEM Purpose The emergency power system provides power to the left and right emergency DC buses, essential flight instrument bus and inertial reference units (IRUs) to support equipment essential for safety of flight in the event of a loss of normal electrical power.

Emergency Power Battery Packs Emergency power is provided by two Securaplane 9-ampere-hour battery packs (Figure 24-58) located in the LEER and REER. Each emergency power battery pack (EPBP) consists of a separate, field replaceable battery unit, and a separate charger/controller unit. Each 24 VDC battery pack uses two 12 VDC sealed, rechargeable lead-acid batteries. The batteries are packaged so that they can easily be changed in the field. The EPBP consists of the battery case, batteries, heater blanket, overcurrent protection device, and heater blanket/battery charger temperature control sensors. The emergency power battery packs are identical to and interchangeable with the emergency lighting battery packs (ELBPs). The EPBPs are designed for minimum maintenance. An EPBP does not need to be removed from the aircraft for any reason other than outright failure, which can be determined while installed on the aircraft. When a fault occurs, an amber “L/R Emergency Battery Fail” CAS message is generated.

Battery Chargers Each EPBP contains a battery charger that provides constant voltage, temperature-compensated charging in the charge mode, or an output of a constant 28 VDC in the transformer-rectifier mode. Battery charging begins automatically when 115 VAC power is applied to the unit. The EPBP charger is powered by single-phase 115 VAC 400 Hz input power.


24-125



The left EPBP is powered by ØB of the left standby AC bus. The right EPBP is powered by ØC of the right standby AC bus. The charger provides 400 watts of charge power. This equates to 14 amps at 28 VDC. The charging system is designed to replace 80 to 90% of charge in an hour, and full charge in approximately 1.5 hours when the EPBP control switches are in the ARM or OFF mode.

Heater Pads

In the transformer rectifier (TR) mode, the charger system will provide 14 amps at 28 VDC to the emergency power loads in the event power to the essential DC bus is lost and the EPBP still has a power input to the charger. In this case, the battery gets only a trickle charge, if needed, since most of the 400 watts would be directed to user loads.

There are three circuit breakers on the lower face of the EPBPs. CB-1 is a 20-amp circuit breaker for overcurrent protection of the battery pack. Opening CB-1 will terminate all power from the battery pack. CB-2 and CB-3 are 5-amp circuit breakers that protect the IRU outputs (Figure 24-57). There is also one 5-amp and one 15-amp internal electronic circuit breaker, referred to as switches, to provide current protected output to the essential flight instrument bus and the left or right emergency DC bus, respectively.

The IRUs will inhibit the battery charger from operating for 15 seconds every time the IRU is turned on. This check of the battery is initiated by the IRU to determine if the battery status is acceptable to power the IRU.

NOTES

Each emergency power battery pack contains two 28 VDC, 40-watt heater pads that are powered from an internal power supply, not battery power. The heater pads are activated when battery temperature is below 60°F and are controlled by an electronic thermostat circuit.

Protective Circuitry

Battery Level Test The EPBPs energy status can be checked with the battery pack on or off the aircraft. LEDs provide battery level indication for the EPBPs’ batteries. This is accomplished by pressing a momentary button on the front of the EPBP labeled “LEVEL TEST.” When the test switch is activated, the EPBP terminates battery output for 16 seconds, connects the battery to the heater blanket (which acts as a load), and measures the battery voltage over an 8-second period. Three color-coded LEDs, mounted on front of the unit, indicate the energy level. The three LEDs flash for 8 seconds during the test, then the LED indicating current EPBP energy status will remain illuminated for 8 seconds. After the test, the EPBP returns to the previous logic state. The level test will be the most accurate when the aircraft has been powered down for at least an hour and the temperature is around 70°F. The energy values are as follows: • Red—0 to 50% • Amber—51 to 75 % • Green—76 to 100%

24-126




Built-In Test Equipment (BITE)

EPBP Storage

The EPBP self-test circuitry determines if a failure has occurred in the battery, charger, output switches, or heaters. When a fault occurs, an amber “L/R Emergency Battery Fail” CAS message is generated, and the respective fault light emitting diode (LED) will be illuminated on the front cover of the EPBP. Detected faults are latched into fault registers in the control circuit. The fault register is cleared when the RESET switch on the front cover is engaged (Figure 24-57). If an EPBP is indicating a fault, it is not prevented from providing backup power until the battery is depleted.

A fully charged battery pack may be placed in storage (with CB-1 pulled) for approximately nine months to a year with being recharged, if the storage temperature is between –40°F (–40°C) and 72°F (+22°C).

BATT A battery fault indicates that the two batteries are not or were momentarily not balanced. The unit will still function and attempt to charge the pack to correct the imbalance. If the problem persists, replace the batteries. CHGR A charger fault means the unit can not generate proper voltage for charging. The unit will still provide backup power until the battery is expended. SWITCH A switch fault indicates one or more of the seven (five lighting and two avionics) output switches are not operating due to an internal fault or overcurrent situation on the aircraft side. HTR A heater fault means loss of heater function. It is not critical as long as the ambient temperature is greater than 32°F. A BITE TEST switch provides a means to check for proper operation of the EPBPs builtin test equipment (BITE) on the aircraft by stimulating the fault detection circuitry. When the switch is depressed, the BATT, CHGR, SWITCH, and HTR LEDs will illuminate, indicating proper operation of the BITE.

NOTE Do not store the battery packs in ambient temperatures above 80°F (27°C). Doing so will accelerate battery self discharge. A battery’s life is cut by half for every 10°C rise above room temperature (73°F, 23°C).

While in storage, it is recommended that a battery level test be performed every 90 days. The stored battery packs battery level may be tested by warming the battery pack to above 60°F (18°C) and engaging CB-1. The level test will automatically start when CB-1 is pushed in. A bench recharge is required if the amber or red LED illuminates. If the EPBP should require charging, it can be done in the shop with an adapter harness utilizing 115 VAC, 50 to 400 Hz power.

NOTE Immediately following an emergency power total discharge, or when battery power is otherwise determined l o w, t h e b a t t e r y p a c k m u s t b e recharged. If the battery pack will not accept a full charge after being left discharged, the batteries may have become sulfided and need to be replaced. A fully discharged battery may become sulfided in as little as 48 hours. Allow two hours of recharge for a totally discharged battery.


24-127



Figure 24-59. Electrical Power Control Panel

24-128




EPBP Bench Charging

Emergency Power Panel

The battery packs can be charged on the aircraft with the internal battery charger using airc r a f t p ow e r, o r o ff t h e a i r c r a f t u s i n g a Securaplane supplied bench charging harness (PN TE-0014-01). Off aircraft, the EPBP may be charged using a 115 to 120 VAC, 50 to 400 Hz power supply able to deliver a recommended minimum of 1,000 volt-amps. An ordinary North American 120 VAC/60-Hz wall outlet is a sufficient charge source.

Three guarded momentary switches in the EPCP section of the cockpit overhead panel control the EPBPs (Figure 24-59). These switches provide on, arm, and off control of the emergency power system. Each switch has a split legend and control both emergency power and emergency lighting systems. The lower legend of each switch reads “AV PWR” and the upper legend reads “LIGHTS.” Both are amber in color when illuminated.

NOTES

OFF Switch The OFF switch allows the crewmembers to turn off EPBP power only if there is more than 20 VDC on the essential DC bus. When the off function is activated, the OFF and ARM switches illuminate.

ARM Switch The ARM switch arms the EPBP by allowing automatic switching to EPBP power when the e s s e n t i a l D C bu s d r o p s b e l ow 2 0 V D C . Depressing the ARM switch will extinguish the ARM and OFF amber light indication. Depressing the ARM switch with less than 20 VDC on the essential DC bus will turn the EPBPs on.

ON Switch The EMERGENCY POWER ON switch turns on the EPBPs when it is depressed. Illumination of the ON switch indicates that the EPBPs are supplying power to the IRUs and the emergency and essential flight instrument buses.


24-129

24-130

L ESS DC BUS 28V

L E BATT FAIL L E BATT ON MAU 1 DG I/O MODULE SLOT 9/10

LEFT EMERGENCY BATTERY PACK

DIM TEST

L EMERG DC BUS

FAIL OUTPUT


ARM

OFF

LIGHTS

LIGHTS

LIGHTS

AV PWR

AV PWR

AV PWR

MCDU # 3 VHF COM # 1

OFF COMMAND ON

COMBINED WOW

IRS #1

IRS OUTPUT #1

BATT OFF IND ARM COMMAND

MRC # 1 NIM

CMC

NOT ARMED IND IRS OUTPUT #2 ON COMMAND

ELT

CHGR INHIBIT

PPILOT ACP

OUTPUT AMP

BATTERY ON IND OUTPUT AMP

L ESS DC BUS CHGR/TR

BATTERY

CLOCK # 1

L STBY AC BUS

ESS FLT INSTRUMENT BUS

RIGHT EMERGENCY BATTERY PACK

CLOCK # 2 STBY EBDI STBY INST PWR MAGNETOMETER SFD

OFF COMMAND BATT OFF IND ARM COMMAND NOT ARMED IND ON COMMAND

IRS #2 ATC 1

CHGR INHIBIT IRS OUTPUT #2

IRS #3

R EMERG DC BUS

VHF NAV # 1 FUEL IND

OUTPUT AMP

LAND GEAR IND OUTPUT AMP

MCDU # 1 R RNG BATT FAIL R ENG BATT ON

BATTERY

R STBY AC BUS

DG I/O MODULE SLOT 7/8

Figure 24-60. Emergency Power Battery Schematic

R ESS DC BUS

international

MAU 2

CHGR/TR

FlightSafety

BATT ON IND FAIL OUTPUT

IRS OUTPUT #1


DIM TEST

NOTES

DIM TEST




24-131



Figure 24-61. 60-Hz Frequency Converter

24-132




Operation and Indications Each battery has two output ports for supplying up to 80 watts of electrical power from each port to an IRU. The left battery powers IRUs No. 1 and CMC. The right battery powers IRUs No. 2 and No. 3. Each battery also provides one 5-amp and one 10-amp current protected output to the essential flight instrument bus for avionics support. The batteries will always pass essential DC bus power through to loads connected to emergency and essential flight instrument buses, regardless of control switch status when essential DC bus power is available. When battery is in ARM mode, it functions as a ready-to-switch (RTS) power supply to connected emergency loads, providing power during momentary power interruptions or loss of essential DC power.

The operating condition of the EPBPs is provided to the crew on the cockpit CAS display. A blue “L/R Emergency Battery ON” CAS message is displayed when the EPBP is in the on state. DGIO 1 in MAU 1 receives the “L Emergency Battery ON” discrete, while DGIO 2 in MAU 2 receives the “R Emergency Battery ON” discrete. An amber “L/R Emergency Battery Fail” CAS message is displayed when the EPBP has registered an internal fault. DGIO 1 in MAU 1 receives the “L Emergency Battery Fail” discrete, while DGIO 2 in MAU 2 receives the “R Emergency Battery Fail” discrete. The fault CAS message will be displayed when a failure of a battery, charger, output switch, or heater blanket is detected.

NOTES

The three guarded switches located in the EMERGENCY POWER section of the COP operate the emergency power battery packs. Selecting the ON switch will send a ground logic on command to the EPBPs. A BATT ON signal from the EPBPs will illuminate both legends (AV PWR and LIGHTS) of the ON switch. This illumination indicates that the EPBP is supplying power to the IRUs, emergency DC buses and essential flight instrument buses (Figure 24-60). Selection of ARM switch sends an arm command to the EPBPs. This will allow automatic switching to EPBP power should the essential DC bus drops below 20 VDC. When the ARM switch is selected, all emergency power s w i t c h l eg e n d s w i l l ex t i n g u i s h ( n o r m a l operating condition). Selection of the OFF switch sends an off command to the EPBPs. This allows the crewmembers to turn off EPBP power, but only if more than 20 VDC exists on the essential DC bus. When the off function is activated, the EPBP sends a not armed indication and a batt off indication command to illuminate the OFF and ARM switches.


24-133



60HZ POWER

ON 115V/60HZ GND PWR

MAINT ICS JACK

OFF UTILITY LIGHTS ON

OFF GND SVR BUS SW

PUSH TO TEST GND SVC BUS

GND SVR BUS IND

115V/60HZ GND PWR RCPT

Figure 24-62. 60 Hz Receptacle

24-134




60-HZ POWER

NOTES

Purpose The purpose of the 60 Hz power system is to provide 115 V, 60 Hz AC to outfitter-installed equipment requiring. The components, controls, and indicators for the 60 Hz AC power system are the 400 to 60 Hz frequency converter, 115 VAC, 60 Hz receptacle, ground fault circuit interrupter, 60 Hz ground power contactor, and MASTERS control panel.

Converter The 400 to 60 Hz frequency converter is located in the tail compartment (Figure 24-61).


24-135



Figure 24-63. 60 Hz Components

Figure 24-64. Cabin/Galley Master Switches

24-136




60 Hz Receptacle

NOTES

The external 60 Hz receptacle, located on the electrical ground service panel in the tail compartment, provides for 115 VAC 60 Hz connection of the aircraft circuitry to a ground source (Figure 24-62).


24-137



GALLEY MASTER SW CABIN

OVHD ANN LTS GALLEY MASTER POWER NO. 2 RH MAIN 28 VDC

DIM TEST

OFF

400HZ TO 60HZ CONVERTER

CABIN MASTER SW

REMOTE ON/OFF

GALLEY

CABIN MASTER LH MAIN 28 VDC

OVHD ANN LTS POWER NO. 2

CONVERTER FAULT

DIM TEST

OFF

60 HZ REMOTE ON/OFF RELAY

FAULT RETURN CHASSIS GROUND 60 HZ AC OUTPUT

A R MAIN AC BUS

MAU 2 DG I/O MODULE (2) SLOT 7/8

PHASE A INPUT

B

PHASE B INPUT

C

PHASE C INPUT NEUTRAL CHASSIS GROUND

60 HZ GROUND POWER RECEPTACLE

60 HZ GROUND PWR CIRCUIT BREAKER

LINE IN

LINE OUT

LINE IN

LINE OUT

60 HZ POWER TERMINAL BLOC

LINE IN

GROUND FAULT INTERRUPTER

Figure 24-65. 60 Hz Power System Diagram

24-138

R 60 HZ CONVERTER FAIL (OPEN/GND)


60 HZ GROUND POWER RELAY



Ground Fault Interrupter

NOTES

A ground fault circuit interrupter, located on the aft bulkhead of the baggage compartment equipment rack, is provided for protection of the external 60 Hz AC circuitry (Figure 24-63).

60-Hz Ground Power Relay A 60 Hz ground power relay, also located on the aft bulkhead of the baggage compartment equipment rack, is provided to connect available 60 Hz power to the 60 Hz terminal block. The 60 Hz ground power relay is energized w h e n ex t e r n a l 1 1 5 VAC 6 0 H z p ow e r i s connected to the ground power receptacle (Figure 24-63).

60-Hz Terminal A 60 Hz power terminal, located on the aft bulkhead of the baggage compartment equipment rack, is provided as a distribution point for the 60 Hz power when outfitter options are installed (Figure 24-63).

Control Switches The GALLEY or CABIN MASTER switches, located on the MASTERS section of the EPCP, provide cockpit on–off control of the frequency converter. The switches illuminate blue “OFF” when not depressed (Figure 24-64).


24-139

24-140

CABIN MASTER SWITCH CABIN 28-VDC ANN LTS PWR

L MAIN DC BUS

OFF CABIN DC CONTACTOR

OUTFITTER PROVIDED

L MAIN AC BUS A

CABIN MASTER LH MAIN DC 28V

OUTFITTER PROVIDED

ANN LTS CONTROLLER

DIM

CABIN AC CONTACTOR

TEST


GALLEY MASTER SWITCH GALLEY 28-VDC ANN LTS PWR

ANN LTS CONTROLLER

OFF

R MAIN AC BUS A

DIM

B

TEST

OUTFITTER PROVIDED

C GALLEY AC CONTACTOR

GALLEY MASTER RH MAIN DC 28V

R MAIN DC BUS

OUTFITTER PROVIDED

OFF OFF

ON

ON OFF

WOW RELAY

LOAD SHED RELAY

ON

LEFT AC CONTACTOR OFF ON OFF

Figure 24-66. Cabin/Galley Master Control

international

ON

RIGHT AC CONTACTOR

FlightSafety

AUX AC CONTACTOR

GALLEY DC CONTACTOR


B C



Operations and Indications

NOTES

When power is being provided to the main AC buses, the frequency converter is selected on or off by the CABIN or GALLEY MASTER switch. The frequency converter takes 115 volt, 400 Hz three-phase power from the right main AC bus and converts it to 115VAC, 60 Hz, single-phase AC. This single phase of AC power is routed through the contacts of the ground power relay (Figure 24-65). During ground operations, when onboard 400 Hz power is not available, the receptacle on the electrical ground service panel permits connection of an external 115 V, 60 Hz ground power source to the aircraft. When connected, power passes through the ground fault interrupter and energizes the 60 Hz ground power relay. The relay’s contacts pull down, creating a path for power to the terminal block. The external 60 Hz power circuit is protected from overcurrent by a 20-amp circuit breaker labeled “115 VAC/60 Hz GND PWR,” located in the upper left corner of the electrical ground service panel. A fault in the 60 Hz converter sends a discrete signal to MAU 2, DGIO 2, which causes a blue “R 60 Hz PWR FAIL” message on the CAS display.

CABIN/GALLEY SYSTEM Purpose The purpose of the cabin/galley power system is to provide cockpit-controlled 115 VAC, 400 Hz AC and 28 VDC power for cabin and galley loads installed during aircraft outfitting. The primary components for cabin/galley power consist of switching contactors, load shed relay, and control panel switches.


24-141



Contactors The cabin AC contactor (CAC) and cabin DC contactor (CDC) are located behind the 394 panel. The CAC connects left main AC bus power to outfitter provisions when energized, and the CDC contactor connects left main DC bus power to outfitter provisions when energized.

In the single generator, airborne scenario, the energized load-shed relay interrupts power to the coils of the galley AC and DC contactors and sends a 28 VDC discrete to the annunciator control box to illuminate the blue OFF legend on the GALLEY MASTER switch face.

The galley AC contactor (GAC) is located behind the 394 panel. The galley DC contactor GDC) is located in the R PDB. The GAC connects right main AC bus power to outfitter provisions when energized, and the GDC connects right main DC bus power to outfitter provisions when energized. The load-shed function will automatically shed the galley load in flight if the aircraft is down to one source of primary power (Figure 24-66).

Cabin Master Operation When CABIN MASTER is selected on (switch depressed), the switch goes black. The switch provides 28 VDC from the CABIN MASTER circuit breaker to the coils of the CAC and the CDC. The energized contactors connect power from the cabin AC and cabin DC breakers to output connectors on the LPDB (Figure 24-66).

Galley Master Operation When GALLEY MASTER is selected on (switch depressed), the switch provides 28 VDC from the GALLEY MASTER circuit breaker through the deenergized load-shed relay contacts to the coils of the GAC and the GDC. The energized contactors connect power from the galley AC and galley DC circuit breakers to output receptacles on the RPDB. The load-shed function provides automatic shedding of the galley power when the electrical power system is being fed by a single source. The galley load-shed relay will energize whenever the aircraft is weight off wheels and is operating on only one of its primary power sources (L GEN, R GEN, APU GEN).

24-142


NOTES



Ground Service Bus The purpose of the ground service bus (GSB) is to provide DC power to equipment required to perform normal servicing of the aircraft while on the ground. This way, the unnecessary powering of expensive avionics equipment can be avoided.

GSB Power Sources When the aircraft is on the ground and the right Main DC bus is not energized, the GSB can be powered from a 50-amp feed from the right main battery or external DC power when commanded “on” through the GSB control switches. The priority of power sources for the GSB is as follows:

The control relay is located in the standby electrical power junction and relay panel.

GSB Door Switches When the right main DC bus is not powered, the GSB incorporates three door interlock switches to deenergize the GSB and prevent it from being energized when all these doors are in the closed position. The GSB will not be powered when it is no longer needed as indicated by all the access doors being closed. This feature also prevents the depletion of the right battery when the control switches are not used to deactivate the GSB. These switches are located in the following locations: • Tail compartment service door

• Right main DC bus

• Main entry door

• External DC power

• Forward external switch panel door

• Right main battery

NOTES

GSB Contactors During normal operation when the aircraft electrical power system is powered from the IDGs, the APU generator, or an external power source, the GSB is powered from the right main DC bus through the normally closed contacts of GSBC1 and GSBC2, which are located in the 394 panel.

Right DC Bus Sense Relay During normal operation, the right DC bus sense relay, located in the standby electrical power junction and relay panel, will be energized when the right main DC bus is powered. This prevents the GSB from being powered from external DC power through GSBC1 or from the right main battery through GSBC2.

GSB Control Relay During ground operations, the GSB control relay, energized by the GSB control switches, provides a means to power GSBC1 or GSBC2. This allows the GSB to be powered from the external DC power cart through GSBC1 or from the right main battery through GSBC2.


24-143

24-144

GSBC 2 STATUS GND SVC BUS SW ON

COP DIM

MAU 2

GND SVC BUS

TEST

DG I/O MODULE SLOT 7/8

GSBC 1 STATUS EXT DC

FWD

EXT

SW

PNL

GND SVC BUS

GND SVC BUS

RIGHT MAIN DC BUS


GSBC1

OFF

OFF

OFF

AIR GND GSBC2 WOW SWITCH ON

ON

ON

RIGHT DC BUS GSB CONT

MAIN

RELAY

DC BUS

RELAY FWD EXT SW PNL

MAIN DOOR

TAIL COMP DOOR

RIGHT BATT

GND SERV BUS BC 1

BC 2

international

FlightSafety

LEFT BATT

RIGHT

SENSE


MAU 1 DG I/O MODULE SLOT 9/10

TAIL COMP

Figure 24-67. Ground Service Bus Diagram



Weight-On-Wheels (WOW) Switch

NOTES

The right main landing gear WOW switch, when in the “air” mode, prevents the GSB from being powered from any power source other than the right main DC bus. When in the “ground” mode, the GSB may be powered from the external DC power cart through GSBC1 or from the right main battery through GSBC2.

Controls and Indicators The GSB can be manually selected on and off by activating momentary center off toggle switches. When the right main DC bus is not powered and the aircraft is in the ground mode, the GSB can be powered from an external DC power source via the external DC power receptacle, or the right main battery when any one of the three control switches are commanded to the ON position. The GSB can also be deenergized when any one of these switches are commanded to the OFF position. These switches are located in the following locations: • Electrical ground service panel in the tail compartment • Forward external switch panel • LEER system monitor panel GND SVS BUS annunciator lights are located in the ELECTRICAL POWER CONTROL section of the COP, on the electrical ground service panel in the tail, and on the forward external switch panel. These GND SVS BUS “ON” annunciator lights will illuminate when the GSB is powered by the right aircraft battery of external DC power. The status of the GSB is also displayed on the CAS and the DC POWER synoptic page when the CAS display is available.


24-145



Figure 24-68. Pilot Circuit-Breaker Panel

24-146




Figure 24-69. Copilot Circuit-Breaker Panel


24-147



Figure 24-70. Left Power Distribution Box

24-148




Figure 24-71. Right Power Distribution Box


24-149



CHAPTER 25 EQUIPMENT AND FURNISHINGS CONTENTS Page INTRODUCTION ................................................................................................................. 25-1 Flight Compartment....................................................................................................... 25-3 Glareshield ..................................................................................................................... 25-5 Flight Compartment Interior Coverings......................................................................... 25-7 Overhead Console.......................................................................................................... 25-9 Flightcrew Seating ....................................................................................................... 25-11 Equipment Racks......................................................................................................... 25-13 Passenger Compartment .............................................................................................. 25-15 Storage ......................................................................................................................... 25-17 Emergency Equipment................................................................................................. 25-19


25-i




Title

Page

25-1

Flight Compartment Instrument Panel ................................................................... 25-2

25-2

Instrument Panel Glareshield ................................................................................. 25-4

25-3

Cockpit Lighting Power Supply Shelf ................................................................... 25-5

25-4

Pedestal and Side Console Access Panels.............................................................. 25-6

25-5

Cockpit Overhead and Circuit Breaker Panels....................................................... 25-8

25-6

Flightcrew Seating ............................................................................................... 25-10

25-7

Equipment Racks ................................................................................................. 25-12

25-8

Passenger Compartment (Typical) ....................................................................... 25-14

25-9

Aircraft Storage Areas ......................................................................................... 25-16

25-10

Aircraft Emergency Equipment ........................................................................... 25-18


25-iii



CHAPTER 25 EQUIPMENT AND FURNISHINGS

INTRODUCTION The Gulfstream G500/G550 aircraft interior is configured into several compartments. The flight compartment is located forward and provides equipment and furnishings required by the flight crew for aircraft operation. The main cabin is aft of the flight compartment and consists of a passenger compartment, galley and lavatories. Baggage and accessory compartments are located aft of the cabin.The entire aircraft interior is insulated for environmental purposes and to provide noise abatement. Required emergency equipment is positioned throughout the aircraft The current manufacturer’s Maintenance Manual must be consulted for all maintenance specifications, tolerances and actual values. ATA Chapter 5, Time Limits/Maintenance Checks, must be consulted to determine time intervals for components or system inspections, checks, tests, overhaul, or specific life limits.


25-1



A

A—Cockpit Overhead Panel B—Instrument Panel C—Side Console Panels D—Center Pedestal

B

C

C

D

Figure 25-1. Flight Compartment Instrument Panel

25-2




FLIGHT COMPARTMENT

NOTES

The Gulfstream G500/G550 flight compartment, commonly referred to as the cockpit, provides crew seating, instrument panel with attached glare shield, overhead console panel with circuit breaker panels, left and right consoles, center console, pedestal, storage compartments, crew accessories and left, right and under floor electronic equipment racks (Figure 25-1).

Instrument Panel The Gulfstream G500/G550 provides an advanced flight deck instrument panel mounted in front of the flight crew. Of metal alloy construction the panel provides mounting trays for aircraft display units and other flight instruments. The instrument panel contains caution/warning panels for pilot and copilot, flight guidance panel and display controllers. The instrument panel base is secured to aircraft floor structure and left and right sides are secured to longerons. Constructed of machined and stamped metal alloy the panel provides a stable platform for mounting of electronic system components. Mounted in the instrument panel are display units, standby flight instruments, caution / warning panels, display controllers, flight guidance panel and landing gear handle. Two panels are affixed to front of instrument panel and contain aircraft clocks, Emergency Locator Transmitter (ELT) remote switch assembly, brake pressure indicator and pitch trim / yaw damper control panel.


25-3



LEFT GLARESHIELD SCREEN

RIGHT GLARESHIELD SCREEN

LEFT GLARESHIELD EXTENSION

RIGHT GLARESHIELD EXTENSION

Figure 25-2. Instrument Panel Glareshield

25-4




GLARESHIELD The Gulfstream G500/G550 glareshield is mounted on top of the instrument panel. It provides a cover for instrument panel components and as the name implies reduces glare and improves visibility. Stowable extensions are attached to each side of the glareshield (Figure 25-2). The glareshield cover is formed from woven fiberglass / epoxy resin and the bottom is lined with a thin layer of copper mesh for electromagnetic interference protection. The glareshield top is covered with non-reflective material. The cover is mounted to top of instrument panel eyebrow panel. It extends full width of instrument panel.

Glareshield Extensions Glareshield extensions are mounted on right and left aft edge of glareshield cover assembly. An extension cover is riveted to the cover assembly providing a pocket for extension storage. The extension is attached to the glareshield cover with a screw and nut, which

provides a pivot point around which the extension is rotated. A clip attached to the aft outboard edge aids in deploying and stowing the extension. A stop is also incorporated to prevent over extension.

Glareshield Screens A screen is mounted in openings in glareshield cover forward of left and right glareshield extensions. The screen allows for airflow while preventing foreign objects dropping behind the instrument panel.

Cockpit Lighting Power Supply Shelf Unique to the Gulfstream G500/G550 is the cockpit lighting power supply shelf mounted to instrument panel base forward of display unit mountings trays. The cockpit lighting power supply shelf is an 8-inch by 48-inch rectangular panel. The shelf is an aluminum alloy honey combed core bonded sheet metal shelf with attaching brackets and provides mounting slots for 13 individual cockpit lighting power supplies (Figure 25-3).

Figure 25-3. Cockpit Lighting Power Supply Shelf


25-5



LEFT PANEL RIGHT PANEL IDENTICAL, NOT SHOWN) LEFT AFT ACCESS PANEL RIGHT PANEL IDENTICAL, NOT SHOWN)

COCKPIT AIR TEMPERATURE SENSOR

LEFT SIDE CONSOLE ACCESS PANELS (RIGHT SIDE PANELS SIMILAR)

Figure 25-4. Pedestal and Side Console Access Panels

25-6




FLIGHT COMPARTMENT INTERIOR COVERINGS

NOTES

In the Gulfstream G500/G550 cockpit area panels and fairings are used extensively to cover exposed electrical wiring, components and structures. Kickpanels are attached to the pilot and copilot side consoles. Fairings are secured between side consoles and window frames. Panels cover the cockpit ceiling and walls (Figure 25-4).

Pedestal The pedestal assembly provides a platform for the mounting of various electronic and flight control system components. Removable side panels provide access to interior components. The pedestal is located between pilot and copilot seats and is mounted on aircraft centerline.

Pedestal Access Panels Detachable side panels provide access to pedestal interior for maintenance and inspection. A forward and aft access panel is mounted to each side of the pedestal and secured with screws and grommets.

Side Console The side console panels are mounted outboard of the pilot and copilot seats. The panels contain electrical and mechanical devices that display indications from and provide control of various aircraft systems.

Side Console Access Panels The side console panel assembly, commonly referred to as a kickpanel, extends from the cockpit floor to top of side console. The kickpanel is installed on both left and right consoles. It is a three-piece assembly consisting of a forward, aft and closeout panel. The forward and aft panels are attached to the console with a combination of stud assemblies secured to the panels with retaining rings and screws. The closeout panel is attached with screws.


25-7



Figure 25-5. Cockpit Overhead and Circuit Breaker Panels

25-8




OVERHEAD CONSOLE

NOTES

The Gulfstream G500/G550 cockpit overhead console panel is located along aircraft centerline above flightcrew seats. The panel contains controls for operation and monitoring of various aircraft systems to include; engine starting, cockpit lighting, emergency power, auxiliary power, fuel system and cabin pressurization (Figure 25-5). The panel is constructed of aluminum alloy sheet metal. It is attached to aircraft longerons with hinge brackets, which allows for panel lowering to change individual components. Edge light panels are affixed to front of the overhead panel to provide visibility and labeling of mounted components.

Cockpit Circuit Breaker Panel The Gulfstream G500/G550 cockpit circuit breaker panel is located along aircraft centerline aft of the cockpit overhead console. The panel consists of two sections, commonly referred to as the pilot circuit breaker panel and copilot circuit breaker panel. The circuit breaker panel acts as a conduit for AC and DC power supplied to various systems throughout the aircraft. Individual circuit breakers act as protective devices for various system components. Each panel is constructed of aluminum alloy with holes drilled to accommodate a maximum of 70 circuit breakers. They are secured by mounting brackets attached to aircraft longerons. This allows the panels to be individually lowered for maintenance and circuit breaker replacement. Edge lighting panels are secured to circuit breaker panel face to provide component identification and visibility in low light conditions.


25-9



ADJUSTABLE HEADREST

FLIGHT DECK SEATING

ADJUSTABLE ARMRESTS

RESTRAINT SYSTEM

LIFE VEST STOWAGE

BACK CUSHION LUMBAR ADJUSTMENT (RIGHT SIDE) FORE/AFT ADJUSTMENT INBOARD

VERTICAL ADJUSTMENT (OUTBOARD)

THIGH PAD ADJUSTMENT (OUTBOARD)

BACK CUSHION UP/DOWN ADJUSTMENT (LEFT SIDE) RECLINE ADJUSTMENT (OUTBOARD)

HEADREST ADJUSTABLE ARMREST

FOUR-POINT HARNESS HEADREST ADJUSTMENT KNOB

SEATPAN INBOARD/ OUTBOARD HANDLE SEAT BASE

HEADREST ADJUSTMENT KNOB

ARMREST ADJUSTMENT CONTROL

REAR LEG

OBSERVER SEAT TRACK LOCK PIN

Figure 25-6. Flightcrew Seating

25-10




FLIGHTCREW SEATING The Gulfstream G500/G550 flight compartment has seating for a pilot, copilot and observer. Seats for pilot and copilot are identical except for location of seat adjustment controls. Pilot and copilot seats incorporate a five-point harness for the safe restraint of the occupant. The observer seat is mounted on tracks and is stowed when not in use. A fourpoint restraint harness is incorporated into the observer seat. Crew seats are designed to provide maximum long-term comfort while not restricting free movement of the occupant (Figure 25-6).

Pilot and Copilot Seat The pilot and copilot seat is of advanced ergonomic design, providing maximum comfort and ease of movement. The seat is of lightweight construction and the complete assembly is comprised of the upper and base sections. The upper structure consists of lumbar support, recline, thigh pad pressure, seat height (vertical) and seat track lock adjustments. The base structure contains the height and track lock mechanisms. The crew seat is designed to provide maximum long-term comfort while not restricting the free movement of the occupant. The crew seat is of advanced ergonomic design, providing maximum comfort and ease of movement. The seat is of lightweight construction and the complete assembly is comprised of the upper and base sections. The upper structure consists of lumbar support, recline, thigh pad pressure, seat height (vertical) and seat track lock adjustments. The base structure contains the height and track lock mechanisms.

movement. A positive spring-loaded track lock mechanism prevents forward and aft movement on the seat tracks unless adjustment lever is pulled allowing forward and aft movement. The seat assembly is designed to give long service life with a minimum of maintenance and overhaul. The seat is for either pilot or copilot, depending on the configuration of the forward and aft control lever, vertical control lever and recline control lever. A fixed height, adjustable angle headrest is attached to the covered backboard. Life vest stowage is provided on the lower rear of the seat.

Observer Seat The observer seat is of advanced ergonomic design, providing maximum comfort and ease of movement. The track and leg mounted observer seat rolls out from its stowage area on a track mounted to the aircraft floor structure then is pivoted aft to lock its rear leg one of two floor fittings. The seat has a backrest recline of 23°. The armrests pull out from their stowed positions and rotate down for use. The headrest has a locked use position and a stowed position allowing about 85° of forward rotation thereby reducing stowed envelope. The seat is of metal alloy construction using lightweight alloy panels.

The upper seat is of lightweight alloy construction using lightweight alloy panels. The armrests are padded and individually adjustable and may be folded back towards the seat spine when not required for use. The seat base consists of a light alloy box structure. Four claw plate assemblies retain the seat to the aircraft seat tracks and prevent lateral


25-11



LEFT EER

RIGHT EER BAGGAGE EER

UNDER FLOOR EER

Figure 25-7. Equipment Racks

25-12




EQUIPMENT RACKS Left Electronic Equipment Rack The Left Electronic Equipment Rack (LEER) is located aft of the cockpit and contains electronic installations necessary for the aircraft’s safe and efficient operation. The rack is a fiveshelf unit with a closeout panel to hide installed components. Each shelf has mounting trays installed to hold electrical and avionics units. These units are primarily digital line replaceable units. The LEER contains avionics, system components and circuit breaker panels. The LEER also contains a fuel quantity gage and switches for aircraft refueling and for powering the ground service electrical bus (Figure 25-7). The LEER is comprised of a base assembly, forward and aft bulkheads, five equipment shelves, back panel and closeout door. The bulkheads, shelves and closeout door are constructed of honeycomb core material bonded between sheets of metal alloy. The back panel is metal alloy sheeting. Each equipment shelf has mounting trays installed to hold electrical and avionics units. These units are primarily digital Line Replaceable Units (LRUs). A relay and junction panel assembly is mounted to the outside of the forward bulkhead. Power and signal distribution is provided by wiring harnesses routed from connector panels located on forward bulkhead to installed components.

The REER also contains a power distribution box with associated circuit breaker panel and a system test / monitor panel with controls used for system maintenance, equipment diagnostic tests and fault isolation. A cooling fan is installed to prevent overheating of installed components.

Baggage Compartment Equipment Racks An equipment rack is installed on forward right side of baggage compartment. Enclosed in cabinetry, the rack contains multiple shelves equipped with mounting trays for installation of electronic components. A fan is installed to provide cooling air. A shelf assembly for mounting of the conformal water tank is located on the aft right side of the compartment. The baggage compartment EER consists of five equipment shelves secured to the forward baggage compartment bulkhead. A panel is mounted aft of the shelves to provide support and enclose the equipment rack. Closeout doors are mounted to the inboard side of the EER. Two box assemblies with closeout doors are mounted into the forward baggage compartment bulkhead and provides for additional system component mounting. An equipment cooling fan is installed in the top of the EER to provide cooling airflow over installed components.

Under Floor Equipment Rack Right Electronic Equipment Rack The Right Electronic Equipment Rack (REER) is located aft of the cockpit and contains electronic installations necessary for the aircraft’s safe and efficient operation. The rack is a fiveshelf unit with a closeout panel to hide installed components. Each shelf has mounting trays installed to hold electrical and avionics units. These units are primarily digital line replaceable units.

The under floor equipment rack is a series of metal alloy brackets, covers, shelves and trays mounted to the aircraft structure and located under floor of the center aisle. Mounted to the rack are aircraft electronic system components and include inertial reference units, inverters, relay panels, radio altimeter, control units and transformer-rectifier units.


25-13



FORWARD LAVATORY

CREW REST

SINGLE SEAT/CONFERENCE TABLE DIVAN (2 SHOWN)

DOUBLE SEATS/CONSOLE TABLE CREDENZA

AFT LAVATORY GALLEY

VANITY

BAGGAGE COMPARTMENT

Figure 25-8. Passenger Compartment (Typical Layout)

25-14




PASSENGER COMPARTMENT

Lavatories

The Gulfstream G500/G550 passenger compartment provides seating and various accommodations for passengers. Seating accommodations may consist of single seats, double seats and divans. The passenger compartment may also be equipped with tables, credenzas and other amenities which will vary according to individual aircraft configuration (Figure 25-8).

The aircraft is equipped with the provisions for installation of forward and aft lavatories. Depending on aircraft configuration, either one or both lavatories may be installed. The basic lavatory consists of a vacuum toilet. Under floor plumbing connects the toilet to the waste water system located in the aircraft tail compartment. Additional lavatory accessories including wash basins, showers, storage cabinetry and doors are installed after production and information can be found on these items in the individual aircraft completion center maintenance handbook.

Passenger Seating Cabin seats are designed with integral headrests, single lever backrest recline control, single lever track and swivel control, seat base stowage and seat belt. Single seats are designed with integral adjustable headrests, single lever track and swivel control and seat belt and shoulder harness restraint systems. Single seats are equipped with seat pan lifters and are fully berth able. Double seats are built to the same specification as a single seat but with limited lateral tracking and swivel capability. Double passenger seats are equipped with drop-down inboard and center armrests. Life vests are stored in each seat base.

The lavatory is completed by addition of the enclosures and accessories as dictated by individual aircraft specifications. As a minimum the completed lavatory must comply with the provisions of AD 74-08-09 R2.

Forward Lavatory The forward lavatory, if installed, is located on the left side and aft of the main entry door.

Aft Lavatory The aft lavatory is located on right side of aircraft at the rear of the passenger compartment.

The divans provide seating for four passengers and are certified for passenger use during taxi, takeoff and landing. They can be converted to sleepers and are joined when each is fully berthed. Additionally, each divan incorporates end cabinets for miscellaneous storage. The passenger compartment may also be equipped with tables, credenzas and other amenities which will vary according to individual aircraft configuration.


25-15



STEPWELL STORAGE

TOW BAR

TAIL COMPARTMENT STORAGE

SPARE LIGHT BULBS

TAIL COMPARTMENT LADDER

MAIN ENTRANCE DOOR

BAGGAGE COMPARTMENT

Figure 25-9. Aircraft Storage Areas

25-16




STORAGE

Tail Compartment

Flight Station Storage Areas

The Gulfstream G500/G550 main accessory compartment is located aft of the baggage compartment pressure bulkhead and is commonly referred to as the tail compartment. Access is via the tail compartment door located on the bottom of the aircraft. The compartment houses the aircraft auxiliary power unit, pneumatic components and other production installed equipment. Several items of aircraft ground handling equipment are also stored in this compartment. The aircraft tail compartment ladder provides a means to enter the tail compartment once the tail compartment door has been opened. The ladder can be removed from the tail compartment and used for engine servicing and to gain access to the baggage door.

Gulfstream G500/G550 has provisions for storage of miscellaneous items that are built into the aircraft cockpit. A stepwell storage box is located aft of the center pedestal. A spare lamp storage panel is incorporated into the left console. Additional storage racks are dependent on aircraft configuration (Figure 25-9).

Stepwell Storage Box The stepwell storage box is located aft of the center pedestal between Fuselage Station (FS) 133 - FS 139. The box is rectangular in shape and constructed of metal alloy sheet metal. The lid, hinged to the forward edge of the box, has a sliding latch assembly and support to keep the lip in the open position when required.

Spare Lamp Storage Panel The spare lamp storage panel is located in the aft end of the pilot side console. The box is rectangular in shape and constructed of metal alloy sheet metal. The lid, hinged to the forward edge of the box, is secured with a wing head stud. The box is mounted to the side console with spring-loaded studs. The interior of the box has cutouts for inserting spare lamps.

Main Entrance Door The Gulfstream G500/G550 main entrance door is equipped with two small accessory compartments normally used for stowing of landing gear safety pins, door control valve pins and pitot probe covers. Additional accessory compartments are dependent on individual aircraft configuration

Cargo "Baggage" Compartment The cargo compartment, also referred to as baggage compartment, is located aft of the aircraft passenger compartment. Access is from the passenger compartment through a sliding door or the aircraft exterior via a slide up cargo door. The cargo compartment houses an aft electronic equipment rack and water storage tank and is also equipped with cargo nets and lifting devices.


25-17



FLASHLIGHTS

LIFE VEST

HALON EXTINGUISHER

INT'L FIRST AID KIT

LIFE RAFTS (BENEATH DIVAN)

PORTABLE OXYGEN STANDARD FIRST AID KIT (AT DIVAN ENDS)

WATER AND HALON FIRE EXTINGUISHERS

FLASHLIGHTS

Figure 25-10. Aircraft Emergency Equipment

25-18




EMERGENCY EQUIPMENT

Emergency First Aid Equipment

The Gulfstream G500/G550 emergency equipment consists of a variety of approved safety devices. The aircraft is equipped with flotation devices, fire and smoke protection devices, emergency locators and first aid kits. Emergency equipment is located throughout the aircraft and is dependent on individual aircraft configuration (Figure 25-10).

First aid kits are installed in the aircraft. Location is dependent on aircraft configuration.

Flashlights Rechargeable MAGLITE flashlights are installed on pilot and copilot side consoles and to the baggage compartment electronic equipment rack. The flashlights are clipped into combination mounting / recharging units which keep them fully charged.

Emergency Flotation Emergency flotation equipment consisting of inflatable life rafts and individual life vests are provided. Life vests are stored underneath passenger seats and in pouches on back of crew seats. Life rafts are normally stored beneath divans in the passenger compartment. Precise location is dependent on individual aircraft configuration.

Fire and Smoke Aircrew smoke goggles and an approved smoke hood in a sealed container are installed in the aircraft. Component location is dependent on aircraft configuration.

Emergency Locator The aircraft is equipped with an emergency locator system. Automatically activated in event of an aircraft crash the system transmits aircraft identification and position data to a geostationary satellite system that provides total worldwide coverage. A remote switch on the copilot instrument panel allows for manual activation of the emergency locator if required.

Evacuation Equipment Aircraft evacuation equipment is comprised of ditching lines and a fire / crash axe. Ditching lines are provided at each emergency window exit. They are accessed by removing the emergency exit window, pulling on the line to extend it and then exiting the aircraft. The fire /crash axe location is dependent on aircraft configuration.

Flight Compartment Pilot-Use Accessories Flight compartment pilot accessories are stored in storage compartments located in the cockpit. These accessories include smoke goggles and items of emergency equipment such as life vests.

Insulation Aircraft insulation is used extensively for sound dampening and environmental control purposes. Insulation is installed in walls, ceiling and under floor in both the flight crew and passenger compartments. Several types of insulating materials are utilized depending on aircraft location and specified requirements. Flame retardant foam, water resistant fiberglass batting and felt blankets are the most common materials used. The insulating material is cut into specified shapes for each location requiring insulation. Contact adhesive is applied to aircraft surface and the insulating material is affixed.


25-19



CHAPTER 26 FIRE PROTECTION CONTENTS Page INTRODUCTION ................................................................................................................. 26-1 GENERAL ............................................................................................................................ 26-1 FIRE DETECTION AND OVERHEAT SYSTEM............................................................... 26-3 Engine Fire and Heat-Sensing System........................................................................... 26-3 Pylon Heat-Sensing System......................................................................................... 26-17 APU Fire Detection System ........................................................................................ 26-21 Equipment Area Overheat Indication System ............................................................. 26-25 Smoke Detector ........................................................................................................... 26-27 Emergency Smoke Evacuation Panel .......................................................................... 26-29 Fire Detection Indication and Test System.................................................................. 26-31 FIRE-EXTINGUISHING SYSTEM................................................................................... 26-41 Engine Fire-Extinguishing System.............................................................................. 26-41 APU Fire-Extinguishing System ................................................................................. 26-47 Portable Fire-Extinguishing System............................................................................ 26-49 SUMMARY ........................................................................................................................ 26-49


26-i




Title

Page

26-1

Sensing Element..................................................................................................... 26-2

26-2

Engine Zones 1 and 2............................................................................................. 26-4

26-3

Fire Detection Sensor Locations ............................................................................ 26-6

26-4

Core Fire Detection Sensor Location..................................................................... 26-7

26-5

Fire Detection Control Unit Location .................................................................... 26-8

26-6

Firewire Control Unit........................................................................................... 26-10

26-7

Fire Panel Location .............................................................................................. 26-11

26-8

Fire Warning Block Diagram............................................................................... 26-12

26-9

Engine Fire Loop Short........................................................................................ 26-14

26-10

Pylon Thermal Switches ...................................................................................... 26-16

26-11

Pylon Heat-Sensing Diagram............................................................................... 26-18

26-12

APU Compartment Temperature Sensor ............................................................. 26-20

26-13

APU Fire Test Switch .......................................................................................... 26-20

26-14

APU Fire Warning Light...................................................................................... 26-21

26-15

APU Fire Detection Schematic............................................................................ 26-22

26-16

Thermal Switch Schematic .................................................................................. 26-24

26-17

Aft Baggage Compartment Smoke Detector ....................................................... 26-26

26-18

Smoke Evacuation Panel...................................................................................... 26-28

26-19

Engine Fire Test Diagram .................................................................................... 26-30

26-20

APU Fire Test Diagram ....................................................................................... 26-32

26-21

Fault Detector Test Diagram ................................................................................ 26-34

26-22

Aft Baggage Compartment Smoke Detector Schematic...................................... 26-36

26-23

Compartment Overheat Test Diagram ................................................................. 26-38


26-iii



26-24

Cockpit Fire Handle Locations ............................................................................ 26-40

26-25

Fire-Extinguishing Bottles ................................................................................... 26-42

26-26

Engine Fire-Extinguishing System—Shot 1........................................................ 26-44

26-27

Engine Fire-Extinguishing System—Shot 2........................................................ 26-45

26-28

APU Fire-Extinguishing System.......................................................................... 26-46

26-29

Halon and Water Portable Extinguisher Locations .............................................. 26-48

26-iv




CHAPTER 26 FIRE PROTECTION FIRE WARN

FIRE PULL

INTRODUCTION The Gulfstream G500/G550 fire protection system provides a means of detecting and alerting the crew of a fire in the engine/nacelle and APU area, an overheat condition in the equipment area, and smoke in the baggage area. The system also provides for the elimination of fire in the engine core, engine nacelle, and APU compartment. In addition, there are portable Halon and water fire extinguishers for use by the crew in the event of a fire in other areas.

GENERAL The fire protection system comprises two subsystems: fire, smoke, and overheat detection systems and a fire-extinguishing system. The fire, smoke, and overheat systems are applicable to the engines, pylons, APU,

aircraft equipment area, and baggage area. The fire-extinguishing system includes the on-board systems for the engine and APU and the portable fire bottles to be used for internal, crew-accessible areas.


26-1



RAIL ASSEMBLY

RESISTANCE

CAPACITANCE

DEC INC

CENTER WIRE LOOP A

GLASS/OXIDE

STAINLESS SHEATH

LOOP B TEMPERATURE INC

Figure 26-1. Sensing Element

26-2




FIRE DETECTION AND OVERHEAT SYSTEM

NOTES

ENGINE FIRE AND HEAT-SENSING SYSTEM Sensing Element The engine fire and heat-sensing system consists of two loops. The sensing element is constructed of an outer stainless-steel sheath containing a temperature-sensitive glass/oxide material and a coaxial cable center wire (Figure 26-1). The two loops, loop A and loop B, which are also called “fire rails,” are attached to a tube and routed in parallel one inch apart. As temperatures increase, the resistance between the center wire and the sheath decreases, while the capacitance between the center wire and the sheath increases, providing a basis for fault/fire discrimination.

CAUTION To prevent damage to the fire rails, do not use as hand-holds. Loop A and loop B for each engine are monitored by their respective fire detector control units.


26-3

26-4 ENGINE CORE FAIRINGS ZONE 1 AIR INLET


ZONE 2

ZONE 2

INTERSERVICES FAIRING

ZONE 2 AIR EXHAUST INTO THE BYPASS DUCT

international

FlightSafety

ZONE 1 AIR EXHAUST


EEC

ZONE 1 AIR INLET

Figure 26-2. Engine Zones 1 and 2



Nacelle Ventilation

NOTES

Zone 1 Fire and overheat detection is provided in designated engine zones. Engine zone No. 1 (Figure 26-2) is the annular space between the engine bypass duct and the cowl doors. It extends from the air intake cowl rear bulkhead to the exhaust unit assembly front bulkhead. Zone 1 is ventilated by ram air through two inlets in the upper cowl door which exhausts through a grill in the lower cowl door.

Zone 2 Engine zone No. 2 (Figure 26-2) is the annular section around the HP compressor, combustion section, and turbine casing and is covered by the core fairings. The bypass services fairing is open to and is, therefore, part of engine zone No. 2.


26-5



FWD

5

4

3

2

1

Figure 26-3. Fire Detection Sensor Locations

26-6




Fire Detection Rails Five fire detection rails support the sensing elements (Figure 26-3).

Zone 1 Rails Rail 1 covers the front gearbox-mounted accessories. Rail 2 is located aft of the gearbox and covers the zone 1 ventilation exhaust grill. Fire detection rail No. 3 is located aft of the integrated drive generator (IDG) at the 5th- and 8th-stage air takeoff connections. The rail provides coverage for the 5th- and 8th-stage takeoff ducts and high-stage air valve and provides partial coverage for the IDG.

Fire detection rail No. 4 is located aft of the interservice connections and provides coverage for the 8th-stage air pressure regulating shutoff valve and starter air valve. Fire detection rail No. 5 is attached to the fixed cowl structure and provides coverage for the front engine mount, thrust strut, and pylon in case of combustion burn-through.

Zone 2 Rails Fire detection rail No. 6 is separated into two sections and attached to the engine HP compressor casing on each side of the bypass services fairing. These sections are located within engine ventilation zone 2. They provide coverage for the VSV actuator, overspeed and splitter unit, and any hot-air leaks within zone 2 (Figure 26-4).

Figure 26-4. Core Fire Detection Sensor Location


26-7



FIRE DETECTION CONTROL UNIT

HF RECEIVER/ TRANSMITTERS

D

FW

Figure 26-5. Fire Detection Control Unit Location

26-8




Fire Detection Control Unit

NOTES

The fire detector control units are located on the aft side of the HF-R/T support shelf in the aft equipment compartment. The fire detector control unit (Figure 26-5) contains two separate, but identical, detection control circuits. Both ends of loop A sensing elements are connected to loop A control circuits, likewise for loop B. If either loop fails or is shorted for any reason, a fault warning illuminates in CAS. The crew may then turn off the faulty loop. The remaining loop will then carry the detection capabilities and act as a single loop detection system. The control unit monitors resistance and capacitance changes in the sensing elements, and if the resistance and capacitance of both loops fall within set boundaries for an element heated by fire, a fire alarm is generated. The affected engine’s fire-extinguisher switch and fuel shutoff switch are illuminated, the solenoid in the fire-extinguisher switch is activated to allow bottle release, and the CAS displays the red annunciation “Engine Fire Loop Alert” and “Left Engine Fire” or “Right Engine Fire”. Also, both upper loop A and lower loop B segments of the affected engine’s FIRE TEST–L ENG or FIRE TEST–R ENG switch illuminate. If both loops are energized and the resistance and capacitance of one of the loops falls outside the range, the control unit generates a fault alarm and sends signals to CAS, which generates the red annunciation “Engine Fire Loop Alert”. Also, the LOOP A or LOOP B segment of the FIRE TEST switch will illuminate, corresponding to the faulty loop. Additional protection against false alarms is achieved by requiring fire signals from both loops before the fire-extinguisher switch and fuel shutoff switch are illuminated and the lock solenoid is energized. A fire signal from either loop results in only CAS messages and one lighted segment of the FIRE TEST switch, unless one of the loops has been tuned off. In which case, the fire-extinguisher and fuel shutoff switches are illuminated and the lock solenoids are energized.


26-9

FIRE TEST LOOP A

LOOP A

LOOP B

LOOP B

FIRE DETECTION

LEFT

LOOP B

FAULT TEST

FAULT

FAULT

OFF

OFF

TEST

RIGHT LOOP A LOOP B FAULT

FAULT

FAULT

OFF

OFF

OFF

TEST

ST

E TO T

E

SS

LOOP B

FAULT

FAULT

FAULT

OFF

OFF

OFF

OFF

SS

E TO T

LOOP B

FAULT TEST

TEST

RIGHT LOOP A LOOP B

FAULT

FAULT

OFF

OFF

LO O P 'B'

LO O P 'A' 1

E

LOOP A

TEST

LOOP A FAULT OBSERVED BEFORE OR AFTER LEFT ENGINE FIRE TEST

3

E TO T

R ENG

FIRE DETECTION

LEFT LOOP A

FAULT

L OOP ' A ' 2

APU

E

SS

2

3

PR

SS

RIGHT LOOP A LOOP B

PR

ST

FAULT TEST

1

E

LOOP B

L OOP ' B '

3

E TO T

LOOP A

LOOP B

PR

SS

LOOP A

LOOP A FAULT OBSERVED DURING LEFT ENGINE FIRE TEST

PR

E

2

PR

PR


LOOP 'A' 1

LOOP B

OFF

LOOP 'B'

L ENG

FIRE DETECTION

LEFT LOOP A

FAULT

FAULT TEST NORMAL INDICATION

TEST

R ENG

E

E TO T

SS

E TO T

PO S. 2

N OR MA L OP E R AT I ON

N O R M A L O P E R AT I O N

N O R M A L O P E R AT I O N

FI R E W I R E C ON TR OL U NI T B I TE I N S TR U C TI ON S

F I R EWI R E C O N T R O L U N I T B I T E I N ST R U C T I O N S

F I RE W I RE CO NT RO L UNI T BI T E I NS T RUCT I O NS

After failure to test from Flight Deck, p r e s s l a m p s t o t e s t b u l b s . Un l e s s b o t h l a m p s l i g h t , r e p l a c e Co n t r o l U n it .

A f t e r f a i l u r e t o t e s t f r o m Fl i g h t De c k , p r e s s l a mp s t o t e s t b u l b s . Un l e s s b o t h l a mp s l i g h t , r e p l a c e Co n t r o l Un i t .

After failure to test from F light D eck, press lam ps to test bulbs. U nless both lam ps light, replace C ontrol U nit.

O b s e r v e l a m p o n FA U LT Y l o o p o n l y.

Ob s e r v e l a mp o n FA ULTY l o o p o n l y.

O bserve lam p on FAU LT Y loop only.

I L L UM I NAT E D – S w i t c h t o P o s i t i o n 1 . Loop lamp lit–Replace Control Unit. loop lamp not lit–Loop short circuit.

I L L UMI NATE D– S wi t c h t o P o s i t i o n 1 . Loop lamp lit–Replace Control Unit. l o o p l a mp n o t l i t – L o o p s h o r t c i r c u i t .

ILLU M IN AT ED –Sw itch to Position 1. Loop lam p lit–R eplace C ontrol U nit. loop lam p not lit–Loop short circuit.

N O T I L L UM I NAT E D – S w i t c h t o P o s i t i o n 3 Loop lamp lit–Loop open circuit. L o o p l a m p n o t l i t – Re p l a c e C o n t r o l U n i t .

NOT I L L UMI NATE D– S wi t c h t o P o s i t i o n 3 Loop lamp lit–Loop open circuit. Loop lamp not lit–Replace Control Unit.

N O T ILLU M IN AT ED –Sw itch to Position 3 Loop lam p lit–Loop open circuit. Loop lam p not lit–R eplace C ontrol U nit.

FAULT TEST NORMAL INDICATION

LOOP OPEN CIRCUIT

FAULTY CONTROL UNIT

Figure 26-6. Firewire Control Unit

international

P OS . 2

FlightSafety

POS. 2


LOOP A

FIRE TEST

APU

ST

TEST

LOOP B

L ENG

ST

LOOP A

R ENG

ST

APU

ST

26-10

FIRE TEST L ENG



When the ENG FIRE TEST switch is selected, both loop A and B should illuminate. If either loop light does not illuminate, then there is either an open loop, no continuity, or a defective fire detection circuit. Performing the FAULT TEST checks the condition of only the control unit fault detection circuit, not the loops. In order to determine what has to be replaced, select the toggle switch on the control unit (Figure 26-6) to position 3 and observe the loop A and B lights on the control unit. If the loop light illuminates, then the loop has an open circuit. A loop light not illuminating indicates the control unit needs to be replaced.

When a fault light on the cockpit overhead panel illuminates before or after the fault test is accomplished, then there is either a shorted loop or a defective control unit. To determine either case, move the toggle switch to position 1. Should a loop light illuminate, replace the control unit. If the loop light does not illuminate, then the loop has a short circuit.

System Controls The fire detection system control panel is located on the upper left side of the overhead panel (Figure 26-7).

Figure 26-7. Fire Panel Location


26-11

26-12

LEFT FIRE DETECT LEFT ESS DC BUS

LOOP A RIGHT FIRE DETECT

RIGHT FIRE DETECT CONTROL UNIT

LEFT FIRE DETECT LOOP B RIGHT FIRE DETECT

RIGHT ENGINE FIRE LOOPS MASTER WARN


MAU 2 SINGLE GENERIC I/0 MODULE 4 SLOT 12 R ESS

AURAL WARN

MAU 1 SINGLE GENERIC I/0 MODULE 1 SLOT 3 L ESS

R Engine Fire R Engine Fire Loop Alert

ENGINE CONTROL PANEL

FIRE TEST L ENG LOOP A LOOP B

APU

R ENG

TEST

LOOP A LOOP B

LOOP B

FAULT TEST

FAULT

FAULT

OFF

OFF

TEST


FAULT

OFF

OFF

Figure 26-8. Fire Warning Block Diagram

international

FIRE TEST PANEL

FlightSafety

FIRE DETECTION

LEFT LOOP A


RIGHT ESS DC BUS



The fire detection system is tested from this control panel. Each ENG FIRE TEST switch has a split-lens, with LOOP A displayed in the upper half and LOOP B in the lower half. The crew may check for normal operation of the system by depressing and holding either ENG FIRE TEST switch (Figure 26-8).

NOTES

The fault detection switches are also located on this control panel and are powered by the left and right essential buses. A fault detection test is performed by depressing and holding the FIRE DETECTION FAULT TEST switch.

System Displays The cockpit fire warning indications are the red illumination of the left or right fire handle, the left or right fuel control switch, and the loop A/B segments of the FIRE TEST switch (Figure 26-8). Other fire warning indications are as follows: • Audible warning tones (three chimes) • Master warning lights illuminate. • Engine fire, overheat, and fault indication messages are displayed on the CAS. • Solenoid in fire handle is activated to allow handle to be pulled.


26-13

26-14 LEFT FIRE DETECT LEFT ESS DC BUS



LEFT FIRE DETECT

RIGHT ENGINE FIRE LOOPS

LOOP B RIGHT FIRE DETECT

MASTER WARN

GROUNDED


MAU 2 SINGLE GENERIC I/0 MODULE 4 SLOT 12 R ESS

AURAL WARN

MAU 1 SINGLE GENERIC I/0 MODULE 1 SLOT 3 L ESS

MASTER CAUTION

Engine Fire Loop Alert Fire Detect Loop Fault


FIRE TEST L ENG LOOP A

R ENG

TEST

LOOP A LOOP B

FIRE DETECTION

LEFT LOOP A

LOOP B

FAULT TEST

FAULT

FAULT

OFF

OFF

TEST


FAULT

OFF

OFF

Figure 26-9. Engine Fire Loop Short

international

FIRE TEST PANEL

FlightSafety

LOOP B

APU


RIGHT ESS DC BUS



System Operation

NOTES

The fire detection control unit receives its power from the left and right essential DC buses (Figure 26-8). It detects changes in the resistance and capacitance of the fire loop due to a fire or overheat condition. The CAS receives a signal from the fire detection control unit and displays the following: • Upper LOOP A and lower LOOP B lights on the overhead test panel • Master warning light (red “W”) • “Left or Right Engine Fire” message on the CAS • Red “Engine Fire Loop Alert” message on the CAS • Red lights illuminate in the fire handle • Red lights illuminate in the fuel control switch • Master warning chime (three chimes) • Solenoid in fire handle is activated to allow handle to be pulled. If the fire detection control unit detects a loop (short) failure, a signal is sent to the CAS, and the following are displayed (Figure 26-9): • Master warning light (red “W”) • Red “Engine Fire Loop Alert” message Either the LOOP A or the LOOP B test light may illuminate on the overhead test panel, depending on which loop is faulty, and the nature of the failure, e.g., a short or open circuit.


26-15



FS 860

FS 782

UP

IN

BD

D FW

LEFT PYLON LOOKING AFT (THIRD SWITCH NOT SHOWN IN PHOTOGRAPH)

SWITCH DETAIL

Figure 26-10. Pylon Thermal Switches

26-16




PYLON HEAT-SENSING SYSTEM

NOTES

Thermal Switches The pylon fire and heat detection system alerts the crew to an overheat condition within the pylon. Three thermal switches are in each pylon. Two are located aft of the precooler assembly, while a third is located forward of the precooler (Figure 26-10).


26-17

26-18 28VDC

250°F

250°F

250°F

MAU 1 MAU–GEN2 DUAL GENERIC I/0 SLOT 9 L ESS

250°F

MAU 2 MAU–GEN2 DUAL GENERIC I/0 SLOT 7 R ESS

PYLON THERMAL SW’S

L Pylon Hot


28VDC

RIGHT WARN LTS ESS PWR NO. 1 BUS

250°F

250°F

PYLON THERMAL SW’S

international

FlightSafety

AUDIBLE WARNING 3 CHIMES


LEFT WARN LTS ESS PWR NO. 2 BUS

Figure 26-11. Pylon Heat-Sensing Diagram



System Operation

NOTES

The three, normally open pylon thermal switches on each engine receive power from the left and right 28-VDC essential buses. When any one of the switches reaches the 250° F trip point, the switch closes, and the fire detection system is activated, which results in the following (Figure 26-11): • A red “Left or Right Pylon Hot” message appears on the CAS. • A master warning red “W” is displayed on the instrument panel. • A three-chime master warning tone is audible in the cockpit.


26-19



Figure 26-12. APU Compartment Temperature Sensor

FIRE TEST L ENG

APU

R ENG

LOOP A

TEST

LOOP A

LOOP B

LOOP B

FIRE DETECTION

LEFT LOOP A

LOOP B

FAULT TEST

FAULT

FAULT

OFF

OFF

TEST


FAULT

OFF

OFF

Figure 26-13. APU Fire Test Switch

26-20




APU FIRE DETECTION SYSTEM APU Compartment Sensing A seven-foot, temperature-sensitive pneumatic detector secured to the top of the APU compartment provides fire detection for the APU (Figure 26-12). The detector consists of a helium-filled tube and detector switch. The switch has two trip points: a 450°F area overheat trip point and a 1,000°F spot detection trip point. The tube installation instructions specify a minimum bend radius of .375 inch (99.5 mm). The sensor tubes can be bent without causing fault alarms. Dents, nicks, or pits are not cause for replacement unless a sensor tube wall has been open, causing a gas leak. In the event of a gas leak, an amber “APU Fire Detect Fail” message appears on the CAS.

APU Fire Detection Test Relay

APU FIRE light on the APU control panel to illuminate. A 28-VDC signal is also sent to both MAU 1 SGIO 2 and MAU 3 SGIO 6 to display the red APU FIRE message and removes the 28-VDC input to MAU 1 SGIO 1 and MAU 2 SGIO 4 to display the amber “APU Fire Detect Fail” message. If the aircraft is in the weight-on-wheels mode, the MAUs will drive the fire bell in the nose wheel well.

APU Fire Test Switch The APU fire test switch is located on the FIRE TEST panel on the cockpit overhead panel. It initiates a test of the fire warning system (Figure 26-13).

APU Fire Warning Light The APU FIRE warning light is located on the APU control panel. The light provides an indication of an APU fire or APU compartment overheat condition (Figure 26-14).

The APU fire detection test relay provides a 28-VDC signal to the annunciator light/dim/test box, which causes the

Figure 26-14. APU Fire Warning Light


26-21

26-22 APU CONTROL 28 VDC TO THE APU A/C FUEL SHUTOFF RLY AND THE ECU WHEN A FIRE IS SENSED

TEST OFF

*

APU FIRE DET TEST RELAY


APU FIRE TEST SW TO ENGINE FIRE DETECTION SYS

MAU 2 SINGLE GENERIC I/O MODULE 4 SLOT 12 R ESS APU Fire Detector Fail MAU 3 SINGLE GENERIC I/O MODULE 6 SLOT 12 L ESS APU Fire

APU CONTROL 28 VDC FROM THE APU CONT NO. 1 OR APU CONT NO. 2 CIRCUIT BREAKERS

APU FIRE WARNING LTS PWR

MAU’S WILL ALSO DRIVE AN APU FIRE BELL TONE TO THE NOSE WHEEL WELL SPEAKER.

*

ANNUN LTS DIM/TEST CTRL OPN/GND OPN/GND

FIRE

THE AIRPLANE IS * WHEN ON THE GROUND,

OPN/GND

OPN/GND

OPN/GND

OPN/28 VDC

SENSOR SW

INTEGRITY SW

ANNUN LTS PWR

SENSOR TUBE

APU FIRE DET SENSOR APU ENCLOSURE FIRE EXT DISCHD

APU FIRE EXT

TO ENGINE FIRE EXT SYS

“BLUE” OUTLET R ENG (+) “BLUE” OUTLET R ENG (–)

TO ENGINE FIRE EXT SYS

“RED” OUTLET L ENG (+) “RED” OUTLET L ENG (–)

APU FIRE EXT DISCHARGE SW

“YELLOW” OUTLET APU (+) “YELLOW” OUTLET APU (–) TO ENGINE FIRE EXT SYS

PRESSURE SWITCH GROUND

Figure 26-15. APU Fire Detection Schematic

international

L FIRE EXT BOTTLE

FlightSafety

L ESS 28 VDC


WARNING LTS PWR

MAU 1 SINGLE GENERIC I/O MODULE 1 SLOT 3 L ESS APU Fire Detector Fail MAU 1 SINGLE GENERIC I/O MODULE 2 SLOT 12 R ESS APU Fire



APU Fire Warning Speaker

NOTES

An APU fire warning speaker is externally mounted in the nose landing gear wheel well in order to alert ground personnel to an APU fire during ground operations. This speaker is used only when the aircraft is on the ground.

APU Compartment Sensing System Operation The APU fire warning system is powered by the 28-VDC essential bus through the APU CONTROL No. 1 and No. 2 circuit breakers located on the left and right EER. When a fire or overheat condition is detected, a signal is sent to the APU control unit which shuts down the APU. A signal is sent in parallel to the MAU 1 SGIO 2 and MAU 3 SGIO 6 of the CAS, which then displays the red annunciation “APU Fire”. The signal is routed to the annunciator lights dim/test box, which in turn causes the “APU Fire” light in the cockpit overhead to illuminate. The MAUs will drive the fire bell if the aircraft is in the weight-on-wheels mode, which results in an audible tone from the APU fire warning speaker in the nose wheel well. If the aircraft is in the air, the MAUs will not drive the nose wheel well fire bell (Figure 26-15). If the helium charge leaks, an integrity switch in the detector opens, signaling a fault to MAU 1 SGIO 1 and MAU 2 SGIO 4. An amber “APU Fire Detect Fail” message is displayed on the CAS. In the event of an actual fire or an APU fire test, a red “APU Fire” message is displayed. An amber “APU Fire Detect Fail” message appears for a failure of the detector or when the system is tested.


26-23

26-24 >250°F

>250°F

>250°F

L PYLON THERMAL SWITCHES

>150°F

>150°F

>150°F


L EER THERMAL SWITCHES

>150°F

>150°F

BAG EER THERMAL SWITCHES

ANNUN LTS PWR

TEST OFF

ANNUN LTS DIM/TEST CTRL DIM TEST OPN/GND OPN/28 VDC

>250°F

>250°F

AFT EQUIPMENT THERMAL SWITCHES

>250°F

>250°F

>250°F

L AFT FLOOR THERMAL SWITCHES

L PYLON HOT (R)

>250°F

>250°F

>250°F

R PYLON THERMAL SWITCHES

R PYLON HOT (R)

MAU 1 MAU–GEN2 DUAL GENERIC I/O SLOT 10 R ESS L EER HOT (A) BAGGAGE EER HOT (A) AFT EQUIPMENT HOT (R) L AFT FLOOR HOT (R) C AFT FLOOR HOT (R) R AFT FLOOR HOT (R)

>150°F

>150°F

MAU 2 MAU–GEN2 DUAL GENERIC I/O SLOT 7 L ESS L PYLON HOT (R) R PYLON HOT (R)

MAU 2 MAU–GEN2 DUAL GENERIC I/O SLOT 8 L ESS AFT EQUIPMENT HOT (R) L AFT FLOOR HOT (R) C AFT FLOOR HOT (R) R EER HOT (A) FWD FLOOR AREA HOT (A) R AFT FLOOR HOT (R)

>250°F

CENTER AFT FLOOR THERMAL SWITCHES

>150°F

R EER THERMAL SWITCHES

>150°F

>150°F

>150°F

>150°F

FWD FLOOR THERMAL SWITCHES

>250°F

>250°F

>250°F

L AFT FLOOR THERMAL SWITCHES

international

FlightSafety

>250°F

MAU 1 MAU–GEN2 DUAL GENERIC I/O SLOT 9 L ESS

28 VDC WARNING LTS PWR


28 VDC WARNING LTS PWR

Figure 26-16. Thermal Switch Schematic



EQUIPMENT AREA OVERHEAT INDICATION SYSTEM Thermal Switches The system contains 28 normally open thermal switches that detect high ambient temperatures,with the 16 250°F switches connected to both the MAU 1 DGIO 1 and MAU 2 DGIO 2, and the 12 150°F switches connected to either MAU 1 DGIO 1 or MAU 2 DGIO 2 (Figure 26-16). The pylon overheat switches (three left and three right) are also part of this system. Each switch is preselected for location and is not interchangeable with another. As indicated in Figure 26-16, the 22 thermal switches are located as follows (the six pylon switches were previously discussed): • Five thermal switches are in the electronic equipment racks, three in the left EER and two in the right EER. The trip point is set at 150°F. • Two of the thermal switches are in the baggage compartment EER. The trip point is set at 150°F. • Two of the thermal switches are in the aft equipment compartment in the vicinity of the hot-air ducts to alert the crew to any leaks in the hot-air manifold. The trip point is set at 250°F. • Three more switches are below the aft cabin floor, on the left side in the vicinity of the hot-air ducts. They alert the crew to any leaks or breaks in the hotair ducts. The trip point is set at 250°F.

System Operation The normally open switches receive power from the left and right 28-VDC essential buses. These switches supply input to the MAU 1 DGIO 1 and MAU 2 DGIO 2 (Figure 26-16). When the switch trip point is reached, they route signals to the CAS. The following messages are displayed to the crew: • An amber “Left or Right EER Hot” message is displayed as a result of an overheat condition in the left or right electronic rack. The trip point is set at 150°F. • An amber “Baggage EER Hot” message is displayed as a result of an overheat condition in the baggage compartment equipment rack. The trip point is set at 150°F. • An amber “Forward Floor Area Hot” message is displayed as a result of an overheat condition below the cabin floor, in the area between the electronic racks. The trip point is set at 150°F. • A red “Left, Right or Center Aft Floor Hot” message is displayed as a result of an overheat condition below the floor areas. The trip point is set at 250°F. • A red “Aft Equipment Hot” message is displayed as a result of an overheat condition in the tail compartment. The trip point is set at 250°F.

• Two thermal switches are below the aft cabin floor in the vicinity of the hot/cold air mixing manifold in the environmental control system. They alert the crew to hot-air leaks and are set at 250°F. • Three switches are below the aft cabin floor, on the right side in the vicinity of the hot-air ducts. The trip point is set at 250°F. • Five switches are below the forward cabin floor, between the left and right EER. The trip point is set at 150°F. FOR TRAINING PURPOSES ONLY

26-25



Figure 26-17. Aft Baggage Compartment Smoke Detector

26-26




SMOKE DETECTOR

NOTES

The smoke detection system is made up of a photoelectric smoke detector located in the baggage compartment (Figure 26-17). The galley and lavatories may also contain smoke detectors. The exact location of the smoke detectors is determined by the agency that installs them, because they are not part of the production aircraft.

Smoke Detector Operation The smoke detector receives its power from the left essential DC bus. The smoke detector is a photoelectric cell that emits a steady beam of light across a white surface. Smoke entering the detector causes the light beam to be broken, thus alerting the crew by a red “Aft Baggage Smoke” message on the CAS and the master warning light panel.


26-27



EMERGENCY SMOKE EVACUATION VALVE

RESET

EMERGENCY SMOKE EVACUATION VALVE BAG COMPT VENT VALVE RESET NORM OPS

VENT/SMOKE BAG COMPT VENT VLV SENSING PORT

WARNING: DO NOT BLOCK PORT

HOLD TOGGLE UP FOR 10 SEC TO PRESSURIZE BAG COMPT

Figure 26-18. Smoke Evacuation Panel

26-28




EMERGENCY SMOKE EVACUATION PANEL

NOTES

Located on the forward (cabin) side of the secondary pressure bulkhead, this panel contains the emergency smoke evacuation valve (Figure 26-18). Rotation of this valve to the VENT/SMOKE position will deflate the external baggage door air seal, allowing the baggage compartment to depressurize, venting any smoke overboard. Rotation of the valve back to the NORM OPS position will allow the baggage door seal to reinflate. To repressurize the baggage compartment, the baggage compartment vent valve reset switch must be held in the up position for at least 10 seconds. Prior to flight, the flight crew must verify the emergency smoke evacuation valve is in the NORM OPS position.The valve remains in this position through all phases of flight, unless needed otherwise.


26-29

26-30




LEFT FIRE DETECT



MASTER WARN


MAU 1&2 SINGLE GENERIC I/O

AURAL WARN

Right Engine Fire Engine Fire Loop Alert


FIRE TEST L ENG

APU

LOOP A LOOP B

TEST

R ENG LOOP A LOOP B

FIRE DETECTION

LEFT

LOOP B

FAULT

FAULT

OFF

OFF

FAULT TEST

TEST


FAULT

OFF

OFF

FIRE TEST PANEL

international

FlightSafety

LOOP A


RIGHT ESS DC BUS

Figure 26-19. Engine Fire Test Diagram



FIRE DETECTION INDICATION AND TEST SYSTEM

NOTES

The detection indication and test system control panel provides a central point for the testing of the overheat and fire detection systems. The engine fire tests are for loop continuity, and the fire loop fault tests are for control box circuitry. The APU fire test is for system integrity, and the equipment overheat test checks the MAU message output.

System Operation The fire detection indication and test system provides the crew with a way to test the fire detection system. Pressing the left engine fire test switch starts the system test. When the test is activated, the following occurs (Figure 26-19): • Loop A and B lights on the test panel for the respective engine illuminate. • Master warning lights on the glareshield illuminate. • CAS displays the engine fire messages in red for the respective engine. • Respective fire handle and fuel shutoff switch illuminate red. • Master warning tone sounds. • Fire handle locking release solenoid energizes (audible click).

NOTE T h e o t h e r e n g i n e a n d A P U fi r e detection system tests are deactivated during the test.


26-31

26-32



WARN INHIBIT

MASTER WARN

GPWS O'RIDE

INHIBIT

W

RAD ALT

BELOW G/S G/S INHIBIT

C

APU Fire APU Fire Detector Fail

APU CONTROL PANEL

Figure 26-20. APU Fire Test Diagram

international

APU FIRE WARNING LIGHT AND EXTINGUISHER SWITCH

FlightSafety

FIRE TEST PANEL


APU FIRE WARNING SPEAKER (NOSE WHEEL WELL)

VOICE O'RIDE



APU Fire Test

NOTES

Pressing the APU fire test switch starts an APU fire detection system test, the TEST switchlight illuminates red. The glareshield master warning light illuminates, and the aural warnings in the cockpit and the nose landing gear wheel well bell sound (ground mode only). Red “APU Fire” and amber “APU Fire Detector Fail” messages are displayed on the CAS, and a red FIRE message is displayed on the APU control panel (Figure 26-20). Pressing the APU fire test switch energizes the APU fire test relay. The relay energizes the APU fire detection system, excluding the ECU and fuel shutoff. The APU fire test relay also removes power from the MAUs, causing the amber “APU Fire Detect Fault” message to appear. When the test switch is activated, the following occurs: • APU TEST switch illuminates red • APU FIRE indicator in the APU control panel illuminates red • Master warning and master caution lights on the glareshield illuminate • CAS displays a red “APU Fire” and an amber “APU Fire Detect Fault” • APU fire bell in the nose wheel well sounds


26-33

26-34




LEFT FIRE DETECT



MASTER WARN


MAU 1&2 SINGLE GENERIC I/O

AURAL WARN

Fire Detect Loop Fault


FIRE TEST L ENG

APU

LOOP A LOOP B

TEST

R ENG LOOP A LOOP B

FIRE DETECTION

LEFT

LOOP B

FAULT TEST

FAULT

FAULT

OFF

OFF

TEST


FAULT

OFF

OFF

FIRE TEST PANEL

international

FlightSafety

LOOP A


RIGHT ESS DC BUS

Figure 26-21. Fault Detector Test Diagram



Fault Detection Test

NOTES

The fault detection switches are located in the cockpit overhead panel and are powered by the left and right essential buses. The fault detection test is performed by depressing and holding the FIRE DETECTION FAULT test switch. When the test is activated, the four loop fault lights and the test light on the panel illuminate. An amber “Fire Detect Loop Fault” message is displayed on the CAS, an aural warning tone sounds, and the master caution lights on the glareshield illuminate (Figure 26-21).

NOTE This procedure checks the system fire detection control unit for faulty circuitry, not for fire loop continuity.


26-35



AFT BAGG SMOKE DET LEER-CS LEFT ESS 28 VDC DC PWR DIST

1

F

28 VDC INPUT

B

TEST

H

GND

G

CHASSIS GND

D

ALARM (CONSTANT 28 VDC OUT)

ANN LTS PWR #7 REER-B24

65 TEST

CH TEST DIM 203 36 0/GND 0/GND ANNUN LTS DIM & TEST

SMOKE DET TEST 057DS1 LOC:COP

MAU 2 SINGLE GENERIC I/O MODULE 12

AFT BAG COMPT SMK DET MAU 3 SINGLE GENERIC I/O MODULE 12

Figure 26-22. Aft Baggage Compartment Smoke Detector Schematic

26-36




Smoke Detector Test

NOTES

The smoke detector test switch is located in the cockpit overhead panel and is labeled “SMOKE DET.” The system may be tested by depressing this switch. This action connects the test pin to a ground, which causes the smoke detector to output an alarm signal consisting of an audible warning tone and a red “Aft Baggage Smoke” CAS message (Figure 26-22) illuminates.


26-37



L–R Pylon Hot L–C–R Aft Floor Hot Aft Equipment Hot

Figure 26-23. Compartment Overheat Test Diagram

26-38




Compartment Overheat Test

NOTES

The compartment overheat warning system can be tested from the system test panel located in the cockpit overhead panel. Pressing the OVHT switch causes the following to occur (Figure 26-23): • A blue light appears in the test switch. • The red compartment overheat lights illuminate on the CAS: • “Left–Right Pylon Hot” • “Left–Center–Right Aft Floor Hot” • “Aft Equipment Hot” • Both master warning lights illuminate on the glareshield. • A three-chime warning tone sounds.

NOTE Only the red messages are displayed on the CAS.


26-39



DISCH

1

2

SOLENOID RELEASE

Figure 26-24. Cockpit Fire Handle Locations

26-40




FIRE-EXTINGUISHING SYSTEM

NOTES

The subsystems of the fire-extinguishing system are the engine fire, APU fire, and portable fire-extinguishing systems.

ENGINE FIRE-EXTINGUISHING SYSTEM The engine fire-extinguishing system consists of two identical, single-shot fire-extinguishing bottles containing a fire-extinguishing agent and propellant, mounted in the tail compartment. The bottles are interchangeable and are discharged using the fire handle rotary switches located on the center console.

Engine Fire Pull Handles The engine fire pull handles are located on the forward left and right corners of the center console (Figure 26-24). They are used to shut off electrical power, hydraulic fluid, and fuel to the engine in the event of a fire. The pull handles are also used to release the extinguishing agent when the handles are rotated. Each fire handle incorporates a locking solenoid that prevents inadvertent operation. A manual override release button is located below each handle in case of a solenoid malfunction.


26-41



RIGHT ENGINE PRESSURE RELIEF VALVES SHOT NO. 1

SHOT NO. 2

APU

LEFT ENGINE

SHOT NO. 1

SHOT NO. 2 PRESSURE SWITCH

APU LEFT ENGINE

APU (NOT USED)

RIGHT ENGINE

RIGHT ENGINE

FWD

Figure 26-25. Fire-Extinguishing Bottles

26-42


LEFT ENGINE



Engine Fire-Extinguishing Bottles

NOTES

Two fire-extinguishing bottles are located in the tail compartment. Each bottle stores the fire-extinguishing agent Halon 1301 (monobromotrifluoromethane). The pressure in each bottle is 600 psi with a capacity of 224 cubic inches and a service life of five years. On the left bottle, a third cartridge is connected to the feed-line which is routed to the APU. Thermal relief occurs when the bottle pressure exceeds 1,400 psi and the bottle dumps fireextinguishing agent into the aft compartment. This action requires the fire bottle to be recharged (Figure 26-25). Each fire bottle has the capability of housing three electrically activated explosive cartridges which, when fired, release the bottle’s contents into a bonnet which routes the agent to the engine or APU. The cartridges represent SHOT 1 on the right bottle and SHOT 2 on the left bottle for both engines. On the right bottle the third cartridge is unused. The firing squibs for each engine and the APU are different part numbers and are keyed differently to prevent improper connections. A pressure switch on each fire bottle sends a signal to the CAS if the nitrogen propellant pressure falls to the trip point of 200 ±25 psi.


26-43



SHUTTLE VALVE

LEFT ENG

CAS

CAS

L FIRE BTL DSCHG

R FIRE BTL DSCHG

LEFT FIRE EXT AGENT BOTTLE

SHOT #2 LEFT ENG

SHUTTLE VALVE

RIGHT FIRE EXT AGENT BOTTLE SHOT #1 LEFT ENG

SHOT #2 RIGHT ENG

RIGHT ENG

SHOT #1 RIGHT ENG

APU APU CONTROL PANEL

FIRE (HANDLE OUT)

LEFT ESS

HYDRAULICS SHUT-OFF

FUEL SHUT-OFF CONTROL

} }

POWER SOURCE

CCW

CW

SHOT #1

SHOT #2

FIRE EXT SHOT #1

FIRE (HANDLE OUT)

RIGHT ESS

FIRE EXT SHOT #2

HYDRAULICS SHUT-OFF

CW SHOT #2

}POWER SOURCE }FUEL SHUT-OFF

CCW SHOT #1

CONTROL NORMAL (HANDLE IN)

NORMAL (HANDLE IN)

LEFT FIRE EXT SWITCH

RIGHT FIRE EXT SWITCH

Figure 26-26. Engine Fire-Extinguishing System—Shot 1

26-44




System Operation

Rotating the left handle clockwise connects the cartridge to the right 28-VDC essential bus and discharges the left bottle into the left engine (Figure 26-26).

When an engine fire is detected, the lock release solenoid is energized, allowing the fire handle to be pulled. Pulling the left handle and rotating it counterclockwise connects the cartridge to the left 28-VDC essential bus. This action results in the right fire bottle being discharged into the left engine (Figure 26-24).

SHUTTLE VALVE

LEFT ENG

APU CONTROL PANEL

CAS

L FIRE BTL DSCHG

R FIRE BTL DSCHG


SHOT No. 2 LEFT ENG

APU

CAS

SHUTTLE VALVE

RIGHT FIRE EXT AGENT BOTTLE SHOT No. 1 LEFT ENG

SHOT No. 2 RIGHT ENG

FIRE (HANDLE OUT)

LEFT ESS

HYDRAULICS SHUTOFF } }

POWER SOURCE FUEL SHUTOFF CONTROL

LOWPRESSURE SWITCH

CCW

CW

SHOT No. 1

SHOT No. 2

SHOT No.1 RIGHT ENG

FIRE (HANDLE OUT)

RIGHT ESS

FIRE EXT SHOT No. 1

FIRE EXT SHOT No. 2

RIGHT ENG

HYDRAULICS SHUTOFF CCW SHOT No. 2

CW

} POWER SOURCE } FUEL SHUTOFF CONTROL

SHOT No. 1

NORMAL (HANDLE IN)

NORMAL (HANDLE IN)


RIGHT FIRE EXT SWITCH

Figure 26-27. Engine Fire-Extinguishing System—Shot 2


26-45

26-46 FROM RIGHT BOTTLE

LOW PRESSURE SWITCH

TO RIGHT ENGINE

CAS

SHUTTLE VALVE

LEFT ENG

APU FIRE EXT

APU

FROM RIGHT FIRE HANDLE

TO ANN LTS PWR TO RIGHT BOTTLE

ANN LTS DIM/TEST BOX

ESS FIRE EXT SHOT #2

FIRE (HANDLE OUT) HYDRAULICS SHUT-OFF POWER SOURCE FUEL SHUT-OFF CONTROL

}

LEFT ESS 28 VDC BUS

SHOT No. 2 RIGHT ENG

}

A

SHOT No. 2 LEFT ENG

ESS

}


FIRE EXT DSCHG


CCW

CW

SHOT No. 1

SHOT No. 2

FIRE EXT SHOT No. 1


international

Figure 26-28. APU Fire-Extinguishing System

TO RIGHT SYSTEM

FlightSafety

NORMAL (HANDLE IN)


L FIRE BTL DSCHG



APU FIRE-EXTINGUISHING SYSTEM

NOTES

The APU fire-extinguishing system consists of a single shot, utilizing the left fire bottle mounted in the tail compartment. A guarded momentary pushbutton switch labeled “FIRE EXT DISCHD” is mounted in the APU control panel located in the cockpit overhead panel.

System Operation When the crew is alerted to an APU fire by the APU fire detection system, momentarily depressing the FIRE EXT DISCHD switch sends a 28-VDC signal to an electroexplosive cartridge on the left fire bottle. This action allows the extinguishing agent to flow into the APU enclosure. A pressure switch sends a signal to MAU 1 SGIO 1 and MAU 2 SGIO 4, and it displays an amber annunciator light (Figure 26-28).

NOTE Only the left fire bottle can be used for extinguishing an APU fire.


26-47



HALON FIRE BOTTLE • A,B, AND C CLASS • 8.2 LB

WATER FIRE EXTINGUISHER • CLASS A FIRES • 7 LB

Figure 26-29. Halon and Water Portable Extinguisher Locations

26-48




PORTABLE FIREEXTINGUISHING SYSTEM

NOTES

There are two types of portable fire extinguishers on the G500/G550 aircraft: halon and water (Figure 26-29).

Portable Halon Fire Extinguisher The aircraft is equipped with a portable Halon 1211 fire extinguisher. The extinguisher is a swing-horn type weighing 8.2 pounds when fully charged. It has an operating range of –40°F to 130°F (–40°C to 54.4°C) and is used on class A, B, and C fires. It is mounted on a quick-release bracket on the right forward side of the fuselage in the cockpit.

Portable Water Fire Extinguisher The aircraft is equipped with a portable water fire extinguisher that uses a carbon dioxide cartridge to pressurize the water when the carrying handle is twisted. When fully charged with a solution of antifreeze and water, the extinguisher weighs about seven pounds and is mounted in the vertical position, aft of the cockpit. This extinguisher is designed to combat class A fires.

SUMMARY The fire detection system provides a means of detecting and alerting the crew to a fire in the engine, engine nacelle, and APU area. The system is capable of detecting an overheat condition in the equipment area and can detect smoke in the baggage area. The fireextinguishing system provides for fire elimination in the engine nacelle and APU compartment. The aircraft is also equipped with portable Halon and water fire extinguishers.


26-49



CHAPTER 27 FLIGHT CONTROLS CONTENTS Page INTRODUCTION ................................................................................................................. 27-1 GENERAL ............................................................................................................................ 27-1 PRIMARY FLIGHT CONTROLS ........................................................................................ 27-3 Lateral Control System .................................................................................................. 27-3 Aileron Power Disconnect System .............................................................................. 27-19 Aileron Trim Control System ...................................................................................... 27-21 Aileron Hardover Prevention System .......................................................................... 27-23 Longitudinal Control System....................................................................................... 27-31 Elevator Power Disconnect System ............................................................................. 27-37 Stall Barrier System..................................................................................................... 27-39 Elevator Trim System .................................................................................................. 27-45 Elevator Hardover Prevention System ......................................................................... 27-51 Directional Control System ......................................................................................... 27-59 Yaw Damper System ................................................................................................... 27-65 Standby Rudder System............................................................................................... 27-69 Rudder Trim Control System....................................................................................... 27-71 Rudder Hardover Prevention System .......................................................................... 27-73 SECONDARY FLIGHT CONTROLS..................................................................................27-77 Ground Spoilers and Speedbrakes ............................................................................... 27-77 Flap/Horizontal Stabilizer System............................................................................... 27-97 Gust Lock System ..................................................................................................... 27-111


27-i




Title

Page

27-1

Flight Controls System........................................................................................... 27-2

27-2

Flight Controls ....................................................................................................... 27-3

27-3

Lateral Control Fuselage Components................................................................... 27-4

27-4

Fuselage-to-Wing Interface.................................................................................... 27-6

27-5

Wing Rear Beam .................................................................................................... 27-8

27-6

Lateral Control Linkages ..................................................................................... 27-10

27-7

Flight Spoiler Actuator ........................................................................................ 27-11

27-8

Aileron Actuator .................................................................................................. 27-12

27-9

Aileron and Flight Spoiler RVDTs ...................................................................... 27-14

27-10

Lateral Control System Block Diagram............................................................... 27-16

27-11

Aileron Power Disconnect ................................................................................... 27-18

27-12

Aileron Trim Control ........................................................................................... 27-20

27-13

Aileron Force Link............................................................................................... 27-22

27-14

Deactivation Valves.............................................................................................. 27-24

27-15

Aileron HOPS Schematic .................................................................................... 27-26

27-16

Aileron and Flight Spoiler Indications................................................................. 27-28

27-17

Longitudinal Control System............................................................................... 27-30

27-18

Elevator Actuator Assembly ................................................................................ 27-32

27-19

Longitudinal Control System Components.......................................................... 27-34

27-20

Elevator Disconnect Handle................................................................................. 27-36

27-21

Stick Shaker Motors............................................................................................. 27-38

27-22

Stall Barrier Actuator........................................................................................... 27-40

27-23

Autopilot/Stall Barrier Disconnect Switch .......................................................... 27-41


27-iii



27-24

Stall Barrier System ............................................................................................. 27-42

27-25

Stall Barrier Control System................................................................................ 27-43

27-26

Elevator Trim Control Wheels ............................................................................. 27-44

27-27

Elevator Trim Actuator ........................................................................................ 27-46

27-28

Elevator Trim System........................................................................................... 27-48

27-29

Elevator HOPS Switches...................................................................................... 27-50

27-30

Elevator Deactivation Valves ............................................................................... 27-52

27-31

Elevator HOPS Schematic ................................................................................... 27-54

27-32

Longitudinal Control Indications......................................................................... 27-56

27-33

Rudder Pedal Linkage.......................................................................................... 27-58

27-34

Rudder Actuator Assembly.................................................................................. 27-60

27-35

Rudder Horn, Tube and Stops.............................................................................. 27-62

27-36

Rudder Actuator Components.............................................................................. 27-64

27-37

Yaw Damper Block Diagram ............................................................................... 27-66

27-38

Standby Rudder System....................................................................................... 27-68

27-39

Rudder Trim Control............................................................................................ 27-70

27-40

Rudder HOPS Schematic..................................................................................... 27-72

27-41

Rudder Control Indications.................................................................................. 27-74

27-42

Ground Spoiler/Speedbrake Panels...................................................................... 27-76

27-43

Ground Spoiler/Speedbrake Actuator .................................................................. 27-78

27-44

Secondary Flight Control Components................................................................ 27-80

27-45

Spoiler Control Shutoff Valves ............................................................................ 27-82

27-46

Ground Spoiler Schematic ................................................................................... 27-84

27-47

Primary and Secondary Control Valves ............................................................... 27-86

27-48

Ground Spoiler Switches ..................................................................................... 27-88

27-iv




27-49

Ground Spoiler RVDT and Stow Switch ............................................................. 27-90

27-50

Automatic Spoiler Control System Operation ..................................................... 27-92

27-51

Ground Spoiler Control System Indications ........................................................ 27-94

27-52

Flap/Stabilizer Component Locations.................................................................. 27-96

27-53

Flap/Stabilizer Control Unit................................................................................. 27-98

27-54

Flap Power Drive Unit ....................................................................................... 27-100

27-55

Flap System Torque Tube and Flap Actuators................................................... 27-102

27-56

Horizontal Stabilizer Actuator........................................................................... 27-104

27-57

Emergency Stabilizer Switch............................................................................. 27-106

27-58

Flap/Stabilizer Indications ................................................................................. 27-108

27-59

Gust Lock Handle .............................................................................................. 27-110

TABLE Table 27-1

Title

Page

Flap/Horizontal Stabilizer Synchronization Schedule....................................... 27-105


27-v



CHAPTER 27 FLIGHT CONTROLS

20

20 10

10

G S

5

5

5

5 10 20

L

O

C

INTRODUCTION This chapter covers the primary and secondary flight controls designed for the Gulfstream G500/G550 aircraft. All values, such as pressures, temperatures, rpm, and power, are used for their illustrative meanings only. The current manufacturer’s Maintenance Manual must be consulted for all maintenance specifications and tolerances, and the actual values must be determined from approved Gulfstream reference material.

GENERAL The G500/G550 flight controls are hydraulically boosted systems, which allow the pilot, through mechanical linkages, pushrods, and cables, to operate hydraulic actuators to move the control surfaces. Hydraulic power is provided by the left (No. 1) and right (No. 2) hydraulic systems (Figure 27-1). The hydraulic actuators used in the flight control systems are

tandem; therefore, a loss of one hydraulic system will have no effect on the flight controls. If either system fails, the remaining system is capable of maintaining actuator load capacity. Should both hydraulic systems fail, the flight controls will automatically revert to manual operations.


27-1



ELEVATORS STALL BARRIER RUDDER YAW DAMP 1

ELEVATORS STALL BARRIER RUDDER YAW DAMP 2

AILERONS FLT SPOILERS/ SPEED BRAKES GROUND SPOILERS

AILERONS FLT SPOILERS/ SPEED BRAKES GROUND SPOILERS L SYS OR PTU OR AUX PRESS SIGNAL REQ'D FOR R SYS USE

GND SPLR SERVO PRESS WING FLAPS

E L E V

LIFT

LIFT

D I S C

AIL DISC

STBY RUD

PWR XFR UNIT

ON

OFF/ARM

ON

NOT ARM

ON

AUX PUMP ACCUM

LEFT ENG PUMP

AUX SOV AUX BOOST PUMP

ACCUM

RIGHT ENG PUMP

DISCH

DISCH

1

2

2

L

AUX PUMP OFF/ARM

ON

NOT ARM

ON

1

R R SYS

L SYS

AUX L SYS R SYS PTU AUX

L SYS/PTU L SYS/PTU/AUX NITROGEN MECH CONT

ELECT CONT CHECK VALVE SHUTOFF VALVE

Figure 27-1. Flight Controls System

27-2


(

FLOW

)



PRIMARY FLIGHT CONTROLS

LATERAL CONTROL SYSTEM

The primary flight controls direct movement of the aircraft control surfaces responsible for lateral, longitudinal, and directional control. They consist of the ailerons, flight spoilers, elevators and rudder. The secondary flight controls consist of the ground spoilers and speedbrakes, flaps and horizontal stabilizer system, and gust lock. These components assist the primary flight controls in lift augmentation, aerodynamic and rollout braking, and control surface damage protection on the ground (Figure 27-2).

The lateral control system provides a means for conjunctive movements of the ailerons and flight spoilers, causing the aircraft to rotate about its longitudinal (roll) axis.

Component Locations and Functions Control Surfaces The lateral control surfaces are the ailerons and the flight spoilers (Figure 27-2). The ailerons are located on the outboard trailing edge of each wing. The two flight spoiler panels are located on the aft upper surface of each wing, forward of the flaps.

ELEVATOR TRIM TAB

VERTICAL AXIS

ELEVATOR ELEVATOR TRIM TAB AILERON RUDDER

FLAP LATERAL AXIS

SPOILERS AILERON TRIM TAB

AILERON

LONGITUDINAL AXIS

Figure 27-2. Flight Controls


27-3



AUTOPILOT SERVOS

CONTROL COLUMN

LATERAL CONTROL CRANK

DISCONNECT ASSEMBLY

FORWARD SECTOR

CROSSOVER PUSHROD

D

FW

Figure 27-3. Lateral Control Fuselage Components

27-4




Fuselage Components

NOTES

The pilot and copilot control wheels, which are mounted on their respective control columns, allow lateral control inputs to the control surfaces. They impart motion to the aileron and flight spoiler control linkages through the lateral control cranks, located at the base of each control column (Figure 27-3). The control wheels are connected between the control columns by a lateral crossover p u s h r o d . I n c o r p o r a t e d i n t h e c r o s s ove r pushrod is a disconnect mechanism to be used in the event of a jammed or inoperable aileron. This mechanism is the lateral control-disconnect assembly. The control wheels are limited to 90° left and right rotation by nonadjustable internal stops. The aileron autopilot Smart Servos are located aft of each control column in the cockpit step area. The servos are the interface between the autopilot and flight control systems. The cable sectors are attached by a pushrod to the control column lateral control crank. Pulleys located downstream of the cable sectors maintain cable tension and aid in directional control. The cables are 7 x 19, 3/16 inch (7 strands, with 19 wires per strand, 3/16 inch outside diameter) and are routed from the forward sector to the left and right main wheel wells, where pulley assemblies route the cables to the inboard wing sector crank assemblies.


27-5

27-6 LATERAL CONTROL SYSTEM PULLEY FOD COVER

LATERAL CONTROL SYSTEM PULLEYS

international

FlightSafety



Figure 27-4. Fuselage-to-Wing Interface



Fuselage-to-Wing Interface

NOTES

The lateral control system cables are routed through the fuselage into each main wheel well. Pulleys mounted on each main wheel well aft bulkhead direct the cables from the wheel wells to the lateral control system inboard input sector crank on each wing rear beam.


27-7



TO COCKPIT

AILERON CONTROL CABLE

LOAD AND MOTION RELIEF BUNGEE INBOARD INPUT SECTOR/CRANK

RIG PIN HOLE

Figure 27-5. Wing Rear Beam

27-8




Wing Rear Beam

NOTES

Load and Motion Relief Bungee Inputs from the control wheels are transmitted to the inboard input sector crank assembly where the input is split into two commands (Figure 27-5). One command repositions the ailerons via cable loop, and the other repositions the flight spoilers via mechanical linkage. The load and motion relief bungee is located in the inboard portion of the wing rear beam and imparts motion through mechanical linkages to the flight spoiler actuator control valve. The load and motion relief bungee provides protection to the mechanical linkages from hard control inputs and wing flex.


27-9



LOAD AND MOTION RELIEF BUNGEE

GROUND SPOILER SPEEDBRAKE ACTUATOR

BUNGEE

FLIGHT SPOILER ACTUATOR PUSHROD MIXING SUMMING LINK FLIGHT SPOILER ACTUATOR PUSHROD

FLIGHT SPOILER ACTUATOR

Figure 27-6. Lateral Control Linkages

27-10




Mixing/Summing Linkage The mixing/summing linkage is located in the wing rear beam area (Figure 27-6). The link receives lateral control inputs via the inboard input sector crank and mechanical linkages, and enables the correct combining of aileron and flight spoiler position to achieve the desired input roll rate. Through the mixing/summing linkage the flight spoiler actuator pushrod also receives input from the ground spoiler/speedbrake actuator. A bungee assembly is mounted on the lateral control mechanical linkage in the rear beam area of each wing. The spring design protects linkage components from shock damage during lateral control system operation.

Flight Spoiler Hydraulic ServoActuators The flight spoiler hydraulic servo-actuators are mounted on the wing rear beam and are tandem actuators that incorporate mechanical feedback (Figure 27-7). The two actuators deploy the four flight spoiler surfaces in response to the aircrew and autopilot lateral inputs. The lateral control inboard sector crank controls input to the flight spoiler actuator servo control valve input lever/feedback link via mechanical linkages.

Input Lever/Feedback Link The input lever/feedback link is mounted on the flight spoiler actuator. The link provides input to the flight spoiler actuator servo control valve and stops actuator retraction when the flight spoilers are deployed to a position corresponding to aileron travel, thus providing mechanical feedback.

FLIGHT SPOILER ACTUATOR PUSHROD

FLIGHT SPOILER RVDT

INPUT LEVER FEEDBACK LINK FLIGHT SPOILER ACTUATOR

Figure 27-7. Flight Spoiler Actuator


27-11



AILERON INPUT SECTOR CRANK

AILERON DRIVE LINK

AILERON ACTUATOR

LOAD RELIEF BUNGEE

AILERON OUTPUT BELLCRANK (SLOPPY LINK)

Figure 27-8. Aileron Actuator

27-12




Lateral Control Outboard Sector

NOTES

The inboard sector/crank also imparts movement to the aileron input sector/crank, commonly called the lateral control outboard sector, via a cable loop system (Figure 27-8). The aileron input sector transmits motion/force through sectors and a pushrod to displace the aileron actuator servo control valve.

Aileron Hydraulic Servoactuators The aileron hydraulic servoactuators are mounted on the fixed trailing edge of each wing (Figure 27-8). Each aileron actuator is a movingbody-type actuator that receives hydraulic pressure from two sources (tandem actuator) and incorporates mechanical feedback. When the actuator extends or retracts, motion is transmitted through the aileron output bellcrank (sloppy link).

Aileron Actuator Load Relief Bungee The aileron actuators incorporate a load relief bungee that provides protection and acts as a fixed-link input for the actuator control valve. The load relief bungee also provides “artificialfeel” inputs to the control wheels. In addition, the load relief bungee must be properly rigged to the aileron actuator input crank to ensure that the control wheels return to center and that the crew receives artificial-feel inputs for the surface aerodynamics.

Aileron Drive Link When the aileron actuator is extended or retracted, motion is transmitted through the output bellcrank (sloppy link) and a pushrod to the aileron drive link, which positions the aileron control surface.


27-13



AILERON RVDT

FLIGHT SPOILER RVDT

Figure 27-9. Aileron and Flight Spoiler RVDTs

27-14




Surface Position Sensors (RVDTs)-

data recorder (FDR).

All of the flight control surface position sensors have been revised for 3 VAC, 2,048 Hz excitation and are factory calibrated for a ratiometric (i.e., three wire versus two wire) output. The MAU AFCS (AIOP) modules output a 3 VAC, 2,048 Hz reference signal that is used for position sensing of the flight control surfaces, spoilers, and elevator trim tabs. The signals representing the control positions are then directed back to the AFCS (AIOP) modules and used as position feedback indications for the AFCS and monitor and warning system. These signals are also used to generate the pictorial control surface positions on the FLIGHT CONTROLS synoptic page.

Aileron RVDTs An aileron rotary variable differential transducer (RVDT) is mounted near the inboard hinge point of each aileron control surface (Figure 27-9). The aileron RVDT is mechanically linked to the aileron drive crank and provides aileron position indications. The left aileron position transducer (RVDT) is connected to modular avionics unit No. 1 (A) AFCS1-B. slots 7/8, located in the left electronics equipment rack (LEER). The right RVDT is connected to modular avionics unit No. 2 (B) AFCS2-B, slots 13/14, located in the right electronics equipment rack (REER). The MAUs transmit the position data to EICAS for display on the FLIGHT CONTROLS synoptic page and to the flight data recorder (FDR).

Flight Spoiler RVDTs The flight spoiler RVDTs are aircraft-mounted and linked to the flight spoiler control surface (Figure 27-9). The left flight spoiler RVDT is connected to the MAU No. 1 (A), AFCS1-B, slots 7/8, and the right flight spoiler RVDT is connected to the MAU No. 2 (B), AFCS2-B, slots 13/14. The MAUs transmit the position data to EICAS for display on the FLIGHT CONTROLS synoptic page and to the flight


27-15



PILOT INPUT

COPILOT INPUT

AILERON INPUT

CONTROL VALVE +11° AILERON ACTUATOR

AUTOPILOT INPUT

0°

3 VAC 2048 Hz

RVDT

FLIGHT SPOILER ACTUATOR

CONTROL VALVE

LEGEND LEFT HYDRAULIC PRESSURE RIGHT HYDRAULIC PRESSURE MECHANICAL INPUT

Figure 27-10. Lateral Control System Block Diagram

27-16


+47°

0°



NOTES

Lateral Control System Operation Manual or autopilot inputs are transmitted to the aileron actuators and flight spoiler actuators through cables and mechanical linkages (Figure 27-10). Movement of the aileron actuator body drives the output crank and mechanical linkage to deflect the aileron flight control surfaces. Maximum aileron deflection is 11°±1 up and 11°±1 down. As the aileron actuator is extended, mechanical linkage provides simultaneous input to retract the flight spoiler actuator. The flight spoiler actuator is normally extended, which holds the spoilers down and flush to the wing. Mechanical linkage moves the flight spoiler actuator control valve, raising the flight spoilers to a maximum of 47° with maximum aileron travel. Aileron Travel

Spoiler Deployment

Up 1/2°

Begins to deploy

Up 1°

Deploy 5.5° Up

Up 7°

Deploy 28° Up

Up 11°

Deploy 47° Up


27-17



AIL DISC PULLING THE HANDLE SEPARATES LEFT AND RIGHT AILERON SYSTEMS. STOWING HANDLE RECONNECTS AILERON SYSTEMS.

POWER DISCONNECT ASSIST TRIGGER PULLING TRIGGER CAUSES A GAS SPRING CARTRIDGE TO FULLY EXTEND AIL DISC HANDLE AND SEPARATE LEFT AND RIGHT AILERON SYSTEMS.

Figure 27-11. Aileron Power Disconnect

27-18




Manual Reversion With no hydraulic pressure available when the input arrives at the actuator, the aileron actuator bypass valves will open, allowing manual operation (manual reversion) of the aileron control surface via the input and output crank assembly (sloppy link). During manual reversion, all lateral control system motion is accomplished by the ailerons only. The flight spoilers possess no manual reversion capabilities with a loss of hydraulic pressure.

NOTE

AILERON POWER DISCONNECT SYSTEM The aileron power disconnect system is provided to allow the flight crew to mechanically disconnect the left and right aileron and flight spoilers in the event of a mechanical jam at the control surface or in the linkages to the control surface.

Component Locations and Functions Aileron Disconnect Handle

Refer to the Maintenance Schematic Manual for corresponding schematics.

The aileron disconnect handle is located on the right side of the center pedestal (Figure 27-11). It is connected to the aileron disconnect assembly on the lateral control crossover pushrod by a flex cable (see Figure 27-3). During normal operation, the aileron disconnect handle is in the stowed position, and the lateral control crossover pushrod disconnect mechanism is connected.

Operation In the event of a jammed aileron or flight spoiler control surface, the aileron disconnect cover is lifted to gain access to the disconnect handle. The aileron disconnect handle is then pulled to unlock the lateral control crossover pushrod mechanical disconnect. The handle may be operated manually, or the power disconnect assembly (PDA) can be employed. The PDA is a trigger-actuated gas spring that serves as a booster to the crew in the disconnect operation. It provides a 150 ±15-pound force to the disconnect cable. If the PDA is used, a special tool is required to reset the system on the ground.


27-19



TRIM ACTUATOR (INSIDE AILERON)

AILERON TRIM CONTROL WHEEL

UP

AILERON TRIM CABLE

ELEVATOR TRIM CABLE

Figure 27-12. Aileron Trim Control

27-20


RUDDERTRIM CABLE

D FW



NOTES

AILERON TRIM CONTROL SYSTEM The purpose of the aileron mechanical trim control system is to maintain optimum aircraft attitudes about its longitudinal (roll) axis.

Component Locations and Functions Aileron Trim Control Wheel/Cable System The aileron trim control wheel is located on the aft end of the center pedestal (Figure 27-12). An indicator on the trim wheel shows the setting of the trim tab in the left or right wing down units (8 units = 15°). The trim control wheel is fully mechanical. The cable drum and cables transmit the trim commands from the cockpit through the fuselage and then along the rear face of the left wing rear beam. The last set of aileron trim system pulleys turns the cables 90°, so they travel aft and connect to the aileron trim tab mechanical actuator.

Aileron Trim Mechanical Actuator The aileron trim mechanical actuator is located in the leading edge of the left aileron. The trim actuator is protected from freezing by a self-regulating, ceramic resistance heater. The aileron trim tab is located on the inboard section of the left aileron only and provides roll trim control (Figure 27-12).

Operation Rotation of the aileron trim control wheel transmits rotary motion to a torque rod that is connected to the forward cable drum. From the forward cable drum, cables lead to the aileron trim actuator. Rotation of the cable drum produces movement of the push-pull rods, bellcranks, and trim tab. Trim tab deflection is reflected on the aileron trim control wheel by left wing down (LWD) or right wing down (RWD) units. The maximum aileron trim deflection is 15° up or down from trim tab neutral (8 units = 15°).The trim actuator heater is


27-21



IDLE CRANK

FORCE LINK

OUTPUT CRANK (SLOPPY LINK)

Figure 27-13. Aileron Force Link

27-22




self-regulating and prevents ice buildup on the actuator. The heater is powered by 115 VAC from the right main AC bus via the AIL TRIM HT circuit breaker. The heater temperature is 175° ±20°F.

NOTES

AILERON HARDOVER PREVENTION SYSTEM The aileron hardover prevention system (HOPS) is an automatic system that prevents powered lateral control system actuator hardovers due to actuator malfunction.

Component Locations and Functions Force Link The aileron hardover prevention system incorporates a force link to detect a hardover condition (Figure 27-13). The force link forms part of the input linkage between the idler crank and the aileron input crank (sloppy link). It acts as a fixed-length pushrod except during hardover conditions. The force link is a two-section tubular assembly that may change lengths (becoming shorter or longer) during a hardover condition, due to pilot and copilot aileron control input. Each force link contains two electrical microswitches. One switch is closed when the retraction force exceeds a preset tolerance. The other switch is closed when the extension force exceeds a preset tolerance.


27-23



LEFT HYDRAULIC SYSTEM VALVE RIGHT HYDRAULIC SYSTEM VALVE

Figure 27-14. Deactivation Solenoid Valves

27-24




NOTES

Deactivation Solenoid Valves Located in each wing rear beam area are two deactivation solenoid valves (Figure 27-14). The valves control the hydraulic pressure to the aileron actuators from the left and right hydraulic systems. When the aileron hardover prevention system detects a hardover condition, the valve solenoids are energized, shutting off both left and right hydraulic system pressure to the aileron actuators. The actuators will automatically revert to manual reversion. The deactivation solenoid valves are powered by the left and right essential DC busses respectively. The LEFT and RIGHT AIL HYD S/O circuit breakers are located on the cockpit overhead circuit breaker panels.


27-25

27-26 FOR TRAINING PURPOSES ONLY

LEFT ACTUATOR DEACTIVATION SOLENOID VALVE

LEFT AIL HYD S/O FORCE LINK RETRACT SWITCH

L ESS 28VDC

1/2 SEC DELAY ON OPERATE

SET

ELECTRICALLY LATCHED RELAY

TO RETURN

LEFT HYD SYS PRESS

TO RETURN

ANNUNCIATION "L Aileron Hydraulics Off" LEFT ACTUATOR (AMBER) DEACTIVATION SOLENOID VALVE RIGHT HYD SYS PRESS

MAU 2

TO ACT

R ESS 28VDC

FORCE LINK RETRACT SWITCH FORCE LINK EXTEND SWITCH


SET


SLOT 9 L ESS 28 VDC SLOT 10 R ESS 28 VDC

TO ACT

FORCE LINK EXTEND SWITCH

RIGHT AIL HYD S/O

DG I/O MODULE (1)

DG I/O MODULE (2)

RIGHT ACTUATOR DEACTIVATION SOLENOID VALVE SLOT 7 R ESS 28 VDC TO RETURN LEFT HYD SYS PRESS SLOT 8 R ESS 28 VDC ANNUNCIATION "R Aileron Hydraulics Off" TO ACT RIGHT ACTUATOR (AMBER) DEACTIVATION SOLENOID VALVE TO RETURN RIGHT HYD SYS PRESS

international

FlightSafety

NOTE: ALL VALVES WILL DEACTIVATE THEIR RESPECTIVE ACTUATOR HYDRAULIC SYSTEM WHEN ENERGIZED.


MAU 1

TO ACT

Figure 27-15. Aileron HOPS Schematic



NOTES

Operation The aileron hardover prevention system (HOPS) compares pilot input to actuator output by examining the force transmitted through the force link (Figure 27-15). During normal operation, the force link is almost completely unloaded, and the four microswitches are open. When the aileron actuator experiences a hardover, and control input applied to the force link exceeds a predetermined force threshold, the force link assembly will overcome the internal spring resistance by becoming shorter or longer. The change in length will activate one of the microswitches. When either one of the two switches remains closed for more than a half second, an electrical latch is triggered for both aileron actuators. The electrical latch energizes the actuator deactivation valves and hydraulic power is shut off to both aileron actuators. When the deactivation solenoid valves are energized, hydraulic power is removed from the actuators, the left and right system pressure is shut off, and any internal pressure is rerouted to the system return. The ailerons will revert to mechanical operation. Removing hydraulic power from the aileron actuators will have no effect on the flight spoilers. The aileron HOPS system can be reset in flight by cycling both the LEFT and RIGHT AIL HYD S/O circuit breakers simultaneously. The circuit breakers are located on the pilot and copilot overhead circuit breaker panels.

NOTE Refer to the GV Maintenance Schematic Manual for corresponding schematics.


27-27



L-R Aileron Hydraulics Off

Figure 27-16. Aileron and Flight Spoiler Indications

27-28




NOTES

Indications The amber “L-R Aileron Hydraulics Off” CAS message appears when the automatic hardover prevention system has disconnected hydraulic pressure to the aileron actuators (Figure 27-16). Invalid data from an RVDT results in an amber “X” over the affected control surface on the FLIGHT CONTROLS synoptic page. Aileron and flight spoiler control surface deflection indications are displayed on the FLIGHT CONT RO L S s y n o p t i c p a g e . T h e a i l e r o n t r i m position is not indicated.


27-29



COPILOT CONTROL COLUMN

DISCONNECT ASSEMBLY FRONT SECTOR

TRANSVERSE TORQUE TUBE PUSHROD EDDY CURRENT DAMPER

Figure 27-17. Longitudinal Control System

27-30




NOTES

LONGITUDINAL CONTROL SYSTEM The longitudinal control system provides a means for movement of the elevator flight control surfaces, causing the aircraft to move about its lateral (pitch) axis.

Component Locations and Functions Stabilizer-Elevator The longitudinal control of the Gulfstream G500/G550 aircraft is furnished by a conventional stabilizer-elevator combination. Displacement of the elevator from the neutral position will cause the aircraft to rotate about its lateral axis (see Figure 27-2).

Control Columns The conventional dual-control columns are connected to a common transverse torque tube located beneath the cockpit floor (Figure 27-17). Fore and aft movement of either control column provides longitudinal control of the elevators through the elevator actuators. Movement of the control columns is limited to 5 inches forward and 8 inches aft of neutral.

Mechanical Linkage The independent left and right mechanical linkages begin at the base of each control column and are connected to the front sectors with a pushrod. The front sectors are connected via cable runs to the actuator input sectors, and the actuators are connected to the elevator control surfaces via output cranks, pushrods, idlers, and cranks.

Eddy Current Dampers The eddy current dampers sense fore and aft motion of the control column and generate a resisting torque on the control column, proportional to how fast the control system is moving. This prevents flight control surface aerodynamic shock (flutter) in the control column output cranks. The eddy current dampers are located behind the left and right cheek panels (Figure 27-17). FOR TRAINING PURPOSES ONLY

27-31



ELEVATOR ACTUATOR

OUTPUT CRANK

AUTOPILOT SERVO

INPUT SECTOR

Figure 27-18. Elevator Actuator Assembly

27-32


LOAD RELIEF BUNGEE



NOTE

Autopilot Smart Servos

Refer to the Maintenance Schematic Manual for corresponding schematics.

NOTES

Located in the tail compartment forward and above the elevator actuators are the longitudinal control system autopilot Smart Servos (Figure 27-18). The autopilot inputs are transferred to the actuator through a cable and sector crank. The actuator is then displaced to obtain the desired attitude about the lateral axis called for by the autopilot system.

Servoactuators Located in the tail compartment and cradled in the power boost linkages between the input sectors and the output cranks are two movingbody-type hydraulic servoactuators that provide mechanical feedback (Figure 27-18). Each elevator has its own hydraulic servoactuator.

Load Relief Bungee A load relief bungee is located on each elevator actuator (Figure 27-18). It provides input and protection for the actuator control valve and provides artificial feel to the control columns.


27-33



AUTOPILOT SERVO

STABILITY (DOWN) SPRINGS

CABLE TENSION REGULATOR

Figure 27-19. Longitudinal Control System Components

27-34




NOTES

Stability Springs Two stability springs, commonly called “down springs,” are mounted on the forward side of the input cable sector (Figure 27-19). They introduce an approximate 13-pound pull force on the elevator input sector to drive the control columns forward. They also assist in providing artificial feel to the longitudinal control system.

Cable Tension Regulators Cable tension regulators are installed on the elevator input sectors. They maintain a constant cable tension regardless of dimensional changes caused by expansion and contraction of the aircraft due to temperature changes.

Elevator Position (RVDTs) There are two rotary variable differential transducers (RVDTs) mounted on the top of the vertical fin. One RVDT is connected to each of the elevator drive crank linkages. They provide elevator position signals to the MAUs, which transmit the data to the FDR, and for display on the FLIGHT CONTROLS synoptic page. The left RVDT receives excitation voltage from MAU No. 1 (A), AFCS1-B (AIOP) slots 7/8. The right RVDT from MAU No. 2 (B), AFCS2-B (AIOP) slots 13/14.

Operation The pilot or autopilot inputs cause rotation of the input cable sector, which is pushrod-connected to the actuator input crank. Rotation of the actuator input crank provides mechanical inputs to the actuator through the load relief bungee. The actuator output crank transmits motion to the elevators through a series of connecting pushrods and idlers that are connected to the elevator drive cranks at the top of the vertical fin. Maximum elevator deflection is 24° trailing edge up and 13° trailing edge down.

NOTE Refer to the Maintenance Schematic Manual for corresponding schematics.


27-35



ELEV DISC PULLING THE HANDLE SEPARATES LEFT AND RIGHT ELEVATOR SYSTEMS AT THE COCKPIT.

POWER DISCONNECT ASSIST TRIGGER PULLING TRIGGER CAUSES A GAS SPRING CARTRIDGE TO FULLY EXTEND ELEV DISC HANDLE AND SEPARATE LEFT AND RIGHT ELEVATOR SYSTEMS AT THE COCKPIT.

Figure 27-20. Elevator Disconnect Handle

27-36




NOTES

ELEVATOR POWER DISCONNECT SYSTEM The elevator power disconnect system provides the flight crew with a means to mechanically disconnect the left and right elevator control system in the event of a jammed elevator control surface or a mechanical jam of the elevator control linkages.

Component Locations and Functions Elevator Disconnect Control Handle The elevator disconnect control handle is located on the left side of the center pedestal (Figure 27-20). During normal operation the elevator-disconnect control handle is in the stowed position, and the longitudinal control transverse torque tube disconnect mechanism is connected (see Figure 27-17).

Operation In the event of a jammed elevator control surface or linkage, the elevator disconnect hand l e c ove r i s l i f t e d t o g a i n a c c e s s t o t h e disconnect handle. The handle is then pulled to unlock the longitudinal control transverse torque tube mechanical disconnect. The handle may be operated manually, or the power disconnect assembly (PDA) can be employed. The PDA is a trigger-actuated gas spring that serves as a booster to the crew in the disconnect operation. It provides a 150±15-pound force to the disconnect cable. If the PDA is used, a special tool is required to reset the system on the ground.


27-37



PILOT CONTROL COLUMN

SHAKER MOTORS

COPILOT CONTROL COLUMN

Figure 27-21. Stick Shaker Motors

27-38




NOTES

STALL BARRIER SYSTEM The stall barrier system provides a two-stage stall protection. The first level of protection warns the pilot that the aircraft is approaching a stall condition by shaking the control column. The second level of protection is provided prior to the aerodynamic stall by the operation of the stall barrier actuator.

Component Locations and Functions Stick Shaker Motors The stick shaker motors are attached to the pilot and copilot longitudinal control system front sector/cranks and provide a warning by shaking the control columns (Figure 27-21). They are located behind the left and right cheek panels. The motors are energized by the MAUs with 28 VDC through the left and right shaker relays. The left shaker motor is powered from the 28 VDC left essential bus via the SHAKER No. 1 circuit breaker. The right shaker motor is powered from the 28 VDC right main bus via the SHAKER No. 2 circuit breaker. Both circuit breakers are located on the cockpit overhead circuit breaker panel.


27-39



CRANK

PUSHROD

BUNGEE CRANK STALL BARRIER ACTUATOR

D FW CAM

Figure 27-22. Stall Barrier Actuator

27-40




NOTES

Stall Barrier Actuator Dual electrohydraulic servo valves control the stall barrier actuator (Figure 27-22). Valve No. 1 is energized by MAU No. 1 (A), AFCS1-A slots 5/6 and enabled by AFCS1-B slots 7/8. Valve No. 1 and allows left hydraulic system pressure to flow to the actuator. Valve No. 2 is energized by MAU No. 2 (B), AFCS2-A slots 9/10 and enabled by AFCS2-B slots 13/14. Valve No. 2 controls right hydraulic system pressure. The two valves are isolated to prevent single point failures. The stall barrier actuator is connected to cam assemblies which transmit actuator output to mechanical linkages.

Stall Barrier Bungee A bungee assembly is connected to each elevator control input sector (Figure 27-22). The bungees transmit the stall barrier actuator inputs to the elevator control linkages.

PITCH TRIM SWITCH

AUTOPILOT DISCONNECT SWITCH

Figure 27-23. Autopilot/Stall Barrier Disconnect Switch


27-41

27-42 MAU 1 AFCS 1-A SLOT 5/6 L ESS 28VDC

STALL BARR VALUE #1 L ESS 28VDC

STALL BARR 1 ENABLE PUSH VALVE OUT 28 VDC PWR PUSH VALVE 1

ANNUN LTS DIM & TEST PWR

(A)

OFF

OPN/GND

#1 PUSH RELAY

OPN/GND

ON

MAU 1 AFCS 1-B SLOT 7/8 L ESS 28VDC STALL BARR 1 ENABLE PUSH VALVE 1 RELAY MON PUSH VALVE 1 RELAY OUT 28 VDC PWR PUSH VALVE 1

(28V) (GND) (28V)


OFF

L SHAKER MOTOR ACTIVE MON (28V) L SHAKER CB MON

STALL BARRIER SW SHAKER #1

L ESS 28VDC

M L SHAKE RELAY

MAU 1 SINGLE GENERIC I/O MODULE (1) SLOT 3 L ESS L STICK SHAKE OUT (GND)

VALVE #1 PWR

PILOT SHAKER MOTOR

VALVE #1 PWR RTN VALVE #2 PWR

SHAKER #2

STALL BARRIER TANDEM ACTUATOR

VALVE #2 PWR RTN

R MAIN 28VDC

M R SHAKE RELAY

MAU 2 SINGLE GENERIC I/O MODULE (4) SLOT 12 R ESS R STICK SHAKE OUT (GND)

CO-PILOT SHAKER MOTOR

STALL BARR VALUE #2

MAU 2 AFCS 2-A SLOT 9/10 R ESS 28VDC

R ESS 28VDC

FLAP POSITION

AOA PROBE #2 FLAP STAB CHANNEL B

MAU 2 SINGLE GENERIC I/O MODULE (2) SLOT 7 R MAIN 28VDC SLOT 8 R ESS 28VDC AOA ARINC 429 IN FLAP POSITION

WOW

WOW

WOW

WOW

ASCB

NORMAL ACCEL/ STICK PUSH INHUBIT/ SELF TEST

ASCB

NORMAL ACCEL/ STICK PUSH INHUBIT/ SELF TEST

#2 PUSH RELAY

(28V) (28V)

MAU 2 AFCS 2-B SLOT 13/14 R ESS 28VDC STALL BARR 2 ENABLE PUSH VALVE 2 RELAY MON PUSH VALVE 2 RELAY OUT 28 VDC PWR PUSH VALVE 2

(28V) (GND) (28V)

R SHAKER MOTOR ACTIVE MON (28V) R SHAKER CB MON

international

FlightSafety

AOA PROBE #1 FLAP STAB CHANNEL A

MAU 1 DUAL GENERIC I/O MODULE (1) SLOT 9 L ESS 28VDC SLOT 10 R ESS 28VDC AOA ARINC 429 IN

STALL BARR 2 ENABLE PUSH VALVE OUT 28 VDC PWR PUSH VALVE 2


ANNUNCIATOR LTS DIM / TEST CONTROLLER

(28V) (28V)

Figure 27-24. Stall Barrier System




Autopilot/Stall Barrier Disconnect Switches On the outboard handgrip of each yoke are the autopilot/stall barrier disconnect switches (Figure 27-23). They can be used to override the stick pusher function while they are held in the depressed position. The stall barrier disconnect switches will not override the stick shaker motors.

NOTE Joint Aviation Authority (JAA) certified aircraft will have the stall barr i e r d i s c o n n e c t s w i t c h ove r r i d e function disabled.

STALL BARR OFF: ON:

SYSTEM OFF SWITCH LEGEND ILLUMINATION AMBER SYSTEM ON SWITCH LEGEND EXTINGUISHED

Figure 27-25. Stall Barrier Control System


27-43



Figure 27-26. Elevator Trim Control Wheels

27-44




Stall Barrier Control Switch

ELEVATOR TRIM SYSTEM

The stall barrier system is armed by the stall barrier control switch, which is located on the center pedestal ( Figure 27-25). The stall barrier control switch disables the stall barrier actuator in the event of a malfunction. The switch will not disable the stick shaker motors.

The elevator trim control system is designed to maintain optimum aircraft attitudes about its lateral (pitch) axis. Elevator trim can be employed three ways; manually, electrically, or automatically when the autopilot is engaged.

Operation

Component Locations and Functions

The stick shaker function will be armed after the aircraft transitions to weight off wheels (Figure 27-24). The MAUs compare the aircraft angle of attack (AOA), altitude, flap position and airspeed to compute a stall threshold for that aircraft configuration. This stall threshold will be used as a reference AOA for the stick push function of the stall barrier system. When a true wing (normalized) AOA exceeds 85% of this stall threshold, the No. 1 MAU energizes the pilot shaker motor, and the No. 2 MAU energizes the copilot shaker motor. The MDAUs automatically disengage the stick shaker function after the flight crew has reduced the angle of attack by approximately 2° below the stall threshold. If the normalized AOA exceeds the stall threshold, the MAUs energize the No. 1 and No. 2 stall barrier actuator solenoid valves.

Elevator Trim Control Wheels The elevator trim control wheels are interconnected in the center pedestal (Figure 2726). Trim wheel travel is indicated in degrees on each trim wheel (22° noseup, 8° nosedown).

Mechanical Linkage The trim control system is fully mechanical and consists of U-joints, torque tubes, cable d r u m s , c a b l e s , p u l l ey s , b e l l c r a n k s , a n d pushrods. Rotation of the elevator trim control wheels transmits motion to a cable loop system that is routed to a cable drum in the tail compartment. From the tail compartment a second cable loop transmits the input to the elevator trim actuators (see Figure 27-28).

When the actuator solenoid valves are energized, hydraulic pressure from the left and right hydraulic systems extend the actuator one inch. This causes the cam assemblies to make contact with mechanical linkages, which transmit the input through the stall barrier bungees to the elevator input sectors. The result is a pitch down force or “push” applied to the control columns. The stall barrier actuator will be de-energized when either the aircraft “g” loading reaches 0.5g or the normalized AOA has been reduced 3.6 degrees below the stall threshold.



27-45



ELEVATOR TRIM ACTUATOR RVDT

Figure 27-27. Elevator Trim Actuator

27-46




NOTES

Trim Actuator A mechanical trim actuator drum is mounted in each elevator (Figure 27-27). Rotation of the trim cables around the trim actuator drums positions the trim surface by the automatic mach trim from the autopilot system, the electrical pitch trim switches, and the mechanical trim control wheels. Each trim actuator incorporates a self-regulating heater that prevents ice buildup on the actuator. The heater is powered by 115 VAC from the right main bus. Its temperature is 175 ±20°F.

Elevator Trim (RVDTs) An elevator trim RVDT is mounted in each elevator control surface to sense the trim tab position (Figure 27-27). The elevator trim RVDTs are dual channel. Channel 1 of the left elevator trim RVDT receives excitation voltage from MAU No. 1 (A), AFCS1-1 slots 5/6 and channel 2 from MAU No. 1(A), AFCS1-B slots 7/8. Channel 1 of the right elevator trim RVDT receives excitation voltage from MAU No. 2 (B), AFCS2-A slots 9/10 and channel 2 from MAU No. 2 (B), AFCS2B slots 13/14. The MAUs will then transmit the trim tab position data to the FDR and the FLIGHT CONTROLS synoptic page.

Pitch Trim Engage Switch The pitch trim engage/disengage switch is located on the pilot’s lower instrument panel. It provides a means to engage the electrical elevator trim system without engaging the autopilot and disengaging the elevator trim system, if necessary.

Electrical Pitch Trim “Beep” Switches The electrical pitch trim switches are located on the pilot’s and copilot’s control wheels. They are used to send a signal to the MAUs to activate the elevator pitch trim servo motors and to automatically inhibit the autopilot mach trim system (see Figure 27-23).


27-47



ELEVATOR TRIM SERVO MOTORS

TRIM CONTROL WHEELS

Figure 27-28. Elevator Trim System

27-48




NOTES

Pitch Trim Servo Motors The pitch trim servo motors are located in the tail compartment (Figure 27-28). The motors are activated by the MAUs through use of the autopilot or when the pitch trim “beep switches” are actuated, provided the pitch trim engage switch is engaged. The pitch trim servo motors are connected to a dual cable drum via sprocket and chain. When the servo motors are energized, the dual cable drum is rotated, which transmits trim input to the trim actuators and back to the trim control wheels via cable loop.

Operation Elevator trim is initiated with two trim control wheels, two trim switches, or the autopilot system for Mach trim. The trim requirement is transmitted manually to the trim actuators, or electrically to the pitch trim servo motors, then to the trim actuators via mechanical linkage and cable loop (Figure 27-28). Rotation of the actuator cable drums produces movement of the trim tab bellcranks and pushrods. Since there are no stops in the trim actuators, shaft travel is determined by the integral stops in the elevator trim control wheels. The maximum trim tab deflection is 8° tab up and 22° tab down. When the pitch trim engage switch is in the engaged position, input to the pitch trim servo motors can come from the autopilot or from the pitch trim “beep” switches on the control wheel. When using the pitch trim “beep” switches, autopilot mach trim will be inhibited.


27-49



PILOT INPUT SWITCHES (4)

ELEVATOR ACTUATOR

DIFFERENTIAL PRESSURE SWITCHES

INPUT SECTOR

Figure 27-29. Elevator HOPS Switches

27-50




NOTES

ELEVATOR HARDOVER PREVENTION SYSTEM The elevator hardover prevention system (HOPS) is an automatic system that prevents powered longitudinal control system actuator hardovers due to actuator malfunctions.

Component Locations and Functions Pilot Input Switches The elevator hardover prevention system monitors pilot inputs via four double-pole switches mounted on the elevator output crank (Figure 27-29). The elevator actuator input and output cranks are mounted on a common shaft. During normal operation the two cranks stay closely aligned. When an actuator hardover occurs the input crank makes contact with the input switches.

Actuator Output Switches Elevator actuator output is monitored via differential pressure sensors, mounted internal to the body of each actuator (Figure 27-29). The sensors contain four separate switches that monitor actuator hydraulic pressure. They provide a signal to the elevator hydraulic shutoff relays when differential pressure exceeds 650 to 800 psid.


27-51



RIGHT DEACTIVATION VALVE LEFT DEACTIVATION VALVE

PR EC S

RTN

LEFT ELEV DEACT VALVE RIGHT SYS CYL

PR EC S

LEFT ELEV DEACT VALVE LEFT SYS HYD

ELEVATOR ACTUATOR FWD

Figure 27-30. Elevator Deactivation Valves

27-52




NOTES

Deactivation Solenoid Valves There are two deactivation solenoid valves for each elevator actuator. The valves are located in the tail compartment above each actuator assembly (Figure 27-30). When the valves are energized, both left and right hydraulic system pressure is shut off to the actuator.

Hydraulic Shutoff Relays The elevator hydraulic shutoff relays are located on the tail compartment junction boxes. The shutoff relays incorporate an electrical latch that will energize the deactivation shutoff valves when a hardover has occurred for more than 0.2 seconds.


27-53

27-54 L ESS 28VDC

RET

>650-800 PSI S3-RET SYS 2 <600 PSI

EXT

>650-800 PSI S2-EXT SYS 1 <600 PSI

RET

>650-800 PSI S1-RET SYS 1 <600 PSI

MAU 1

TO ACT 1/2 SEC DELAY ON OPERATE


SET

SET

EXT


LEFT ACTUATOR DEACTIVATION SOLENOID VALVE TO RETURN

TO ACT

ELECTRICALLY LATCHED RELAY TO RETURN

>650-800 PSI

LEFT HYD SYS PRESS

LEFT ACTUATOR DEACTIVATION SOLENOID VALVE RIGHT HYD SYS PRESS

DUAL GENERIC I/O MODULE (1) SLOT 9 L ESS 28 VDC SLOT 10 R ESS 28 VDC ANNUNCIATION "L Elevator Hydraulics Off" (AMBER)

MAU 2 S4-EXT SYS 2 <600 PSI

RIGHT ELEV HYD S/O R ESS 28VDC

TO ACT RET

>650-800 PSI S3-RET SYS 2 <600 PSI

EXT

>650-800 PSI S2-EXT SYS 1 <600 PSI

RET

>650-800 PSI S1-RET SYS 1 <600 PSI

EXT



SET

SET


TO RETURN

RIGHT ACTUATOR DEACTIVATION SOLENOID VALVE LEFT HYD SYS PRESS TO ACT

ELECTRICALLY LATCHED RELAY TO RETURN

>650-800 PSI

RIGHT ACTUATOR DEACTIVATION SOLENOID VALVE RIGHT HYD SYS PRESS

DUAL GENERIC I/O MODULE (2) SLOT 7 R MAIN 28 VDC SLOT 8 R ESS 28 VDC ANNUNCIATION "R Elevator Hydraulics Off" (AMBER)

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FlightSafety



LEFT ELEV HYD S/O

S4-EXT SYS 2 <600 PSI

Figure 27-31. Elevator HOPS Schematic



NOTES

Operation During normal operation, the elevator input and output cranks stay closely aligned and will not close either input switch pair. During a hardover condition, the mechanical pilot input switch pair make contact, and the two “cross-directional” hydraulic actuator output switches are energized (Figure 27-31). Contact between an input switch pair in concert with a “cross-directional” differential pressure switch activates the shutoff relays. If the relays stay energized for more than 0.2 second, an electrical latch automatically triggers. Activation of the latch will energize the deactivation solenoid valves, shutting off left and right hydraulic system pressure to the malfunctioning elevator actuator. Once hydraulic pressure is shut off to an actuator, a relay holding circuit maintains the shutoff condition until electrical power is recycled. To re-engage hydraulic power to the actuator, cycle the appropriate circuit breaker (LEFT ELEV HYD S/O on pilot’s overhead, RIGHT ELEV HYD S/O on copilot’s overhead) If the hardover situation is gone, the system resets, and the actuator operates normally. If the failure still exists, the system activates again.



27-55



L-R Elevator Hydraulics Off Elevator Trim 1-2 Fail Stall Barrier 1-2 Stall Barrier Off Stick Push 1-2 Fail Stick Push 1-2 Fault Stick Push Unavailable Elevator Trim Down Limit Elevator Trim Up Limit Pitch Trim 1-2 Power Fail Stick Shake 1-2 Fail

Figure 27-32. Longitudinal Control Indications

27-56




NOTES

Indications Elevator deflection and elevator trim position are displayed on the FLIGHT CONTROLS synoptic page. Invalid data from an RVDT results in an amber “X” over the affected control surface (Figure 27-32). The following list contains some of the possible CAS messages associated with the longitudinal control system:

Amber Messages L-R Elevator Hydraulics Off Elevator Trim 1-2 Fail Stall Barrier 1-2 Stall Barrier Off Stick Push 1-2 Fail Stick Push 1-2 Fault Stick Push Unavailable

Blue Messages Elevator Trim Down Limit Elevator Trim Up Limit Pitch Trim 1-2 Power Fail Stick Shake 1-2 Fail A configuration alarm sounds if the aircraft is weight on wheels, the throttles are greater than 19°, and the elevator trim is not in the g r e e n r a n g e b a n d . T h e r e d “A i r c r a f t Configuration” CAS message also appears.


27-57



TORQUE TUBE

CROSSOVER PUSHROD

PILOT RUDDER PEDALS

TORQUE TUBE LOWER CRANK ARM

UP D FW

RUDDER STEERING RVDT

FOWARD SECTOR/ CRANK

Figure 27-33. Rudder Pedal Linkage

27-58




NOTES

DIRECTIONAL CONTROL SYSTEM The directional control system provides a means of movement of the single rudder flight control surface, causing the aircraft to move about its vertical (yaw) axis.

Component Locations and Functions Rudder Control Surface The single rudder is hinged to the trailing edge of the vertical fin. The rudder is constructed from epoxy graphite.

Rudder Pedals Dual rudder pedals are mounted and supported by pedal hanger arms that pivot on horizontal support tubes (Figure 27-33). The pedals are spring-loaded aft and have nine fore and aft detent position adjustments. Crossover pushrods interconnect the pilot and copilot pedals. Movement of the pedals is limited by pedal stops.

Linkage Movement of the rudder pedals transmits motion through the lower crank arm and torque tube mechanical linkage that is connected to the forward cable sector/crank (Figure 2733). Rotation of the sector/crank is transmitted via cable loop and mechanical linkage to the rudder actuator servo control valve.


27-59



OUTPUT CRANK

FAILSAFE PUSHROD RUDDER DRIVE CRANK

INPUT SECTOR

RUDDER ACTUATOR

RUDDER FEEL BUNGEE

RUDDER TRIM ACTUATOR LINK RUDDER CABLE RUDDER TRIM CRANK ASSEMBLY

RUDDER TRIM CABLE RUDDER TRIM ACTUATOR

Figure 27-34. Rudder Actuator Assembly

27-60




NOTES

Rudder Servo-Actuator The rudder hydraulic servo-actuator is located in the aft, right corner of the tail compartment (Figure 27-34). The actuator is a dual tandem type, pressurized by the left and right hydraulic systems. Pressure-regulating valves reduce actuator output pressure to 1,500 psi. The actuator incorporates bypass valves to prevent a hydraulic lock when hydraulic pressure is lost. The bypass valves are fully open when hydraulic system pressure drops below 90 psi to allow manual control of the rudder surface. Above 125 psi the valves are fully closed.

Bungee The double-acting spring bungee is connected to the input sector at its upper end and to the trim actuator crank at its lower end. The bungee provides input to the rudder actuator from the rudder trim system, and acts as an artificial feel system for the rudder.

Rudder Position RVDT The rudder position RVDT is installed on the rudder actuator input sector. It receives excitation voltage from both MAU No. 1(A), AFCS1-A slots 5/6 and MAU No. 2 (B), AFCS2-A slots 9/10. The MAUs will transmit rudder position data to the FDR and the FLIGHT CONTROLS synoptic page..

Operation Induced motion from the rudder pedals is transferred via the lower crank arm, pushrods, bellcranks, forward input sectors, and cable loops to the actuator input sector. The input sector and the output crank are joined by a pin-in-slot arrangement. When the actuator is stroked, it drives the output crank, which is connected to the rudder drive crank (horn) assembly by a failsafe pushrod (Figure 27-34). The drive horn, in turn, drives the rudder surface. To protect the aircraft tail structure from aerodynamic sideloads, the rudder actuator incorporates hydraulic pressure regulating valves. Rudder surface movement is limited by these valves when airspeeds increase airloads


27-61



RUDDER DRIVE CRANK

RUDDER STOPS

Figure 27-35. Rudder Horn, Tube and Stops

27-62




against the rudder. When the rudder hinge movement limit is reached, the pressure regulating valves shift, allowing hydraulic pressure to bypass, limiting actuator load output. This action causes a blue RUDDER LIMIT message to be displayed on the CAS. Any increased pedal input cannot further displace the rudder.

Rudder Stops Maximum left and right rudder deflection is established by stops on the aircraft structure and rotation of the rudder horn (Figure 27-35). The adjustable rudder stops are set to maintain 22° maximum left/right rudder deflection.

NOTES


27-63



AFT MANIFOLD DIFFERENTIAL PRESSURE SENSORS PRESSURE REGULATING VALVES

FORWARD MANIFOLD

BYPASS VALVE

BYPASS VALVE MAIN CONTROL VALVE

SOLENOID VALVE LVDT

SOLENOID VALVE

EHSV

LVDT EHSV

MAIN RAM ACTUATOR ASSEMBLY

Figure 27-36. Rudder Actuator Components

27-64




NOTES

YAW DAMPER SYSTEM The yaw damping system provides automatic dutch roll damping.

Component Locations and Functions Modular Avionics Units The No. 1 and No. 2 modular avionics units (MAUs) provide primary automatic yaw damping control by computing the yaw damper command and providing signals to the yaw damper assemblies located on the rudder actuator.

Yaw Damper Assembly There are two yaw damper assemblies integral to the rudder actuator (YD No. 1 and YD No. 2). They convert electrical MAU inputs into rudder control surface movements. The yaw damper assemblies consist of solenoid control valves, electrohydraulic servo valves, and modulating valves with linear variable differential transducers (Figure 27-36).

Solenoid Valves The yaw damper solenoid valves are energized by 28 VDC from the MAUs to supply hydraulic pressure to the electrohydraulic servo valves (EHSVs).

Electrohydraulic Servo Valves The electrohydraulic servo valves (EHSVs) use electrical inputs from the MAU to hydraulically position the rudder actuator main control valve via the modulating valves and mechanical linkage.

Modulating Valves The modulating valves are positioned by hydraulic pressure coming from the EHSVs. Mechanical linkage connects the modulating valves to the rudder actuator servo control valve. The modulating valves will also reposition the LVDTs. The LVDT position is converted to an electrical signal that is sent back to the MAUs as rudder position.


27-65

27-66

FLIGHT GUIDANCE YD ENGAGE DISENGAGE YD ENGAGE DISENGAGE

YAW DAMP SERVO NO.1 PILOT D-6

MAU 1

RUDDER ACTUATOR

YAW DAMP COND

RIGHT ESS 28 VDC

YAW DAMP EHSV

YD SOL ENGAGE OUT

YD SCL 1 ENGAGE

YD SOL ENGAGE

LVDT NO.1 EXCITE

LVDT EXCITE

LVDT NO.1 OUTPUT

(NORMALLY CLOSED)

LEFT HYD SYS

MAU 1 YD SOLONOID VALVE NO.1


YD SOLONOID VALVE NO.2

MAU 2 SINGLE RUDDER

MAU 2

STBY RUDD HYD ON

OFF PRESSURE SWITCH

YD SOLONOID VALVE NO.1 YD SOLONOID VALVE NO.2

MAU 2 LVDT INPUT YAW DAMP SERVO NO.2 COPILOT D-6

LVDT NO. 2 OUTPUT

LVDT EXCITE

LVDT NO. 2 EXCITE

YD SOL ENGAGE OUT

YD SCL 2 ENGAGE

YD SOL ENGAGE RIGHT ESS 28 VDC

YAW DAMP COND

SOL ENGAGE RELAY TO ANNUNCIATOR LIGHT POWER

international

FlightSafety

YAW DAMP ENGDISENG STBY RUDD

YAW DAMP EHBV

RIGHT HYD SYS


LVDT INPUT CROSS YD PWR SENSE

AUX HYD SHUTOFF VALVE

28 VDC FROM AUX HYD PUMP YAW DAMP PITCH TRIM CONTROL PANEL

Figure 27-37. Yaw Damper Block Diagram



Yaw Damper Engage/Disengage Switch The yaw damper engage/disengage switch is located on the pilot’s instrument panel (just below the pilot’s NAV display) and allows the crew to engage the yaw damper without engaging the autopilot. The switch illuminates “DISENG” when the system is not engaged.

NOTES

Operation When engaged by the autopilot system or the YAW DAMP switch, the MAUs supply electrical power to the yaw damper assemblies (YD No. 1 and YD No. 2). YD No. 1 is powered by MAU No. 1(A), AFCS1-A slots 5/6 and AFCS1-B slots 7/8. YD No. 1 uses left hydraulic system pressure to position the actuator main control valve. YD No. 2 is controlled by MAU No. 2 (B), AFCS2-A slots 9/10 and AFCS2-B slots 13/14. YD No. 1 uses right hydraulic system pressure. When either system is activated, the MAU will energize the solenoid valve open, allowing hydraulic pressure to be sent to the electrohydraulic servo valve (EHSV) (Figure 27-37). The EHSV is electrically controlled by the MAU to hydraulically position the modulating valve. (The inactive EHSV will remain in a null position). Movement of the modulating valve mechanically repositions the actuator servo control valve which ports left and right hydraulic system pressure to extend or retract the actuator piston, thus repositioning the rudder. Modulating valve movement also repositions the linear variable differential transducer (LVDT), which sends an electrical signal back to the MAU, completing the electrical loop. The yaw damper assemblies will provide 5° of left or right rudder travel. The yaw damper assemblies operate as active/standby systems. One assembly is actively in control while the other is in passive/standby mode. Which assembly is active depends on which flight guidance computer is selected on the pilot’s display controller. If the active MAU/yaw damper assembly fails, control will automatically revert to the passive MAU/yaw damper assembly.


27-67

27-68 MOTOR INPUT VOLTAGE MAU 1

MOTOR ENABLE

AUX HYD PUMP

AUX BOOST INLET PRESS LOW

AUX BOOST FAIL L ESS DC

OVHD ANN LIGHT PWR

MOTOR PWR

NORMAL

MAU 2


OVERLOAD R ESS DC

AUX BOOST PUMP OVERLOAD SENSOR

ELECTRONIC MOTOR CONTROLLER

AUX BOOST PUMP OVERLOAD RELAY L ESS DC

STBY RUDDER HYD ON

STBY RUDDER SW ON (A)

ANN LTS PWR #10

TEST DIM NOSE WOW RELAY FROM LEFT SYSTEM PRESS

AIR GND

TO RUDDER ACTUATOR

M OVERLOAD NORMAL STBY RUDDER VALVE OVERL;OAD SENSOR

STBY RUDDER VALVE OVERL;OAD RELAY

INLET PRESS SW

359¡F

AUX PRESS XDUCER

341¡F

T

THERMAL SW

M

TO AUX SYSTEM (MED, BRAKES, FLAPS, GEAR NWS)

LETTERS AUX AUX SYSTEM BOOST PUMP

M

AUX SYSTEM MOTOR PUMP STANDBY RUDDER VALVE

international

FlightSafety

AUX SYSTEM SHUTOFF VALVE


LO PRESS < 20 PSI

Figure 27-38. Standby Rudder System



NOTES

STANDBY RUDDER SYSTEM The standby rudder system allows the auxiliary hydraulic pump to pressurize the rudder actuator and the No. 1 yaw damper assembly during flight when both the left and right hydraulic systems have failed.

Component Locations and Functions Standby Rudder Valve The standby rudder valve is located on the aft bulkhead of the left wheel well. Its purpose is to route hydraulic pressure from the auxiliary system to the rudder actuator.

Standby Rudder Switch The standby rudder valve is energized through the standby rudder switch, provided the nose weight on wheels relay is in the AIR mode.

Operation Depressing the STBY RUD switch to the ON position turns the auxiliary hydraulic system on. If the nose weight-on-wheels relay is in the AIR mode, the standby rudder valve is also energized, and the auxiliary hydraulic system pressure is routed exclusively to the left system pressure port of the rudder actuator. The auxiliary hydraulic system pressure is then routed from the actuator to the auxiliary hydraulic system return (Figure 27-38). The standby rudder system is used after a dual left and right hydraulic system failure.


27-69



RUDDER TRIM CONTROL • KNOB IS ROTATED IN DIRECTION OF DESIRED NOSE MOVEMENT. • TRIM AUTHORITY TO RUDDER ACTUATOR IS 7.5° LEFT OR RIGHT OF RUDDER NEUTRAL.

PILOT INPUT SWITCHES

RUDDER FEEL BUNGEE

RUDDER TRIM ACTUATOR LINK RUDDER CABLE RUDDER TRIM CRANK ASSEMBLY

RUDDER TRIM CABLE RUDDER TRIM ACTUATOR

Figure 27-39. Rudder Trim Control

27-70




NOTES

RUDDER TRIM CONTROL SYSTEM The mechanical rudder trim control system is designed to maintain optimum aircraft direction about its vertical axis.

Component Locations and Functions Rudder Trim Control Wheel The rudder trim control wheel is located on the aft end of the center pedestal (Figure 27-39). Rotation of the rudder trim control wheel transmits rotary motion to a torque rod connected to the forward cable drum.

Mechanical Linkage The trim control system linkage is fully mechanical, and trim cables transmit the trim commands from the cockpit through the fuselage to the rudder trim actuator.

Rudder Trim Actuator The rudder trim actuator cable drum is mounted on the lower right side of the tail compartment rear bulkhead (Figure 27-39). Its linear shaft is attached to the input link of the crank at the lower end of the artificialfeel bungee. Rotation of the rudder trim control wheel imparts rotational motion to the actuator cable drum, which produces rudder trim shaft linear travel and input crank movement. The vertical movement of the artificial-feel bungee adjusts the hydraulic rudder actuator servo control valve’s null point. There is no trim tab on the rudder. Directional trim is accomplished by varying the position of the rudder.


27-71

27-72 RIGHT AIL HYD S/O R ESS 28VDC

MAU 2

>275-375 PSI S3 RET SYS 2 <200 PSI

EXT

>275-375 PSI S4 EXT SYS 2 <200 PSI

EXT

>275-375 PSI S3 RET SYS 2 <200 PSI

TO ACT

EXT 1/2 SEC DELAY ON OPERATE

SET


TO RETURN

RUDDER ACTUATOR DEACTIVATION SOLENOID VALVE LEFT HYD SYS PRESS TO ACT


EXT

>275-375 PSI

SET


TO RETURN

PILOT ACTUATOR DIFFERENTIAL INPUT PRESSURE SWITCHES SWITCHES

RUDDER ACTUATOR DEACTIVATION SOLENOID VALVE RIGHT HYD SYS PRESS

DUAL GENERIS I/O MODULE (2) SLOT 7 R MAIN 28 VDC SLOT 8 R ESS 28 VDC ANNUNCIATION "Rudder Hydraulics Off" (AMBER)

international

FlightSafety



S4 EXT SYS 2 <200 PSI

Figure 27-40. Rudder HOPS Schematic



Operation

RUDDER HARDOVER PREVENTION SYSTEM

During normal operation, rotation of the rudder trim control wheel transmits motion through the torque rod to turn the forward cable drum (10 units = 7.5° left/right). From the cable drum, cables lead to the rudder trim actuator.

This automatic system prevents powered directional control system actuator hardovers due to actuator malfunction.

Rotation of the actuator produces linear travel, which is transmitted through input linkage to the rudder actuator servo control valve. The maximum rudder trim deflection is 7.5° left or right of neutral. The trim actuator for the rudder is not heated.

Input Switches

NOTES

Component Locations and Functions The rudder hardover prevention system monitors pilot inputs via four double-pole switches mounted on each side of the actuator input crank (see Figure 27-39).

Output Switches Actuator output is monitored via differential pressure sensors mounted on the body of the actuator. The sensors contain four separate switches that monitor actuator output. They provide a signal to the rudder hydraulic shutoff relays when the differential pressure exceeds 275 to 375 psid.

Hydraulic Shutoff Relays The rudder hydraulic shutoff relays are located in the left tail junction box in the tail compartment of the aircraft. The shutoff relays incorporate an electrical latch to energize the deactivation solenoid valves.

Deactivation Solenoid Valves There are two deactivation solenoid valves for the rudder actuator. They are located in the aft, left side of the tail compartment. The valves control the hydraulic power from the left and right hydraulic systems to the rudder actuator.


27-73



Rudder Hydraulics Off Yaw Damper Off Yaw Damper 1-2 Fail Rudder Limit Single Rudder Standby Rudder Hyd On

Figure 27-41. Rudder Control Indications

27-74




Operation

Indications

During normal operation, the rudder actuator input and output cranks stay closely aligned and will not close either input switch pair (Figure 27-40). During a hardover condition, the mechanical pilot input switch pair make contact, and two cross-directional hydraulic actuator output switches are activated. Contact between an input switch pair in concert with a cross-directional differential pressure switch activates the shutoff relays. If the relays stay energized for more than 0.5 second, an electrical latch automatically triggers.Activation of the latch will energize the deactivation solenoid valves, shutting off left and/or right system hydraulic pressure to the rudder actuator.

The rudder control surface position is indicated on the FLIGHT CONTROLS synoptic page (Figure 27-41); however, the rudder trim position is not indicated. Invalid data from the RVDT results in an amber “X” over the rudder on the FLIGHT CONTROLS synoptic page.

Once hydraulic pressure is shut off to the actuator, a relay holding circuit maintains the shutoff condition until electrical power is cycled. To re-engage hydraulic power to the rudder actuator, cycle the RUDDER HYD S/O circuit breaker on the copilot overhead panel. If the hardover situation has been corrected, the system resets, and the actuator operates normally. If the failure still exists, the system deactivates again.

The amber “Rudder Hydraulics Off” CAS message is displayed when the automatic hardover prevention system has disconnected hydraulic pressure to the rudder actuator. The blue “Rudder Limit” CAS message indicates maximum rudder displacement. When only one source of hydraulic pressure is supplied to the rudder actuator, a blue “Single Rudder” message appears. If the yaw damper ENG/DISENG switch is not powered, an amber “Yaw Damper Off” message is displayed on the CAS. The amber “Yaw Damper 1-2 Fail” message indicates a failure of the indicated yaw damper. The blue “Standby Rudder Hyd On” message is displayed when the standby rudder system is activated.

T h e r u d d e r h a r d ove r p r eve n t i o n s y s t e m (HOPS) has the capability to shut off one or both hydraulic systems to the rudder actuator. The amber “Rudder Hydraulics Off” message will be displayed on CAS when the HOPS system is activated. If only one hydraulic system is shut off, the “Rudder Hydraulics Off” message will be accompanied by a blue “Single Rudder” message.


27-75



OUTBOARD FLIGHT SPOILER INBOARD FLIGHT SPOILER GROUND SPOILER

GROUND SPOILER INBOARD FLIGHT SPOILER OUTBOARD FLIGHT SPOILER

Figure 27-42. Ground Spoiler/Speedbrake Panels

27-76




NOTES

SECONDARY FLIGHT CONTROLS The secondary flight controls consist of the ground spoilers and speedbrakes, automatic ground spoilers, flap/horizontal stabilizer system, and gust lock. These components assist the primary flight controls in lift augmentation, aerodynamic and rollout braking, and provide control surface damage protection on the ground. This section provides an introduction to the major components that make up the secondary flight control system, their functions, and their visual indications.

GROUND SPOILERS AND SPEEDBRAKES Component Locations and Functions Spoiler Panels Six spoiler segments are used to dump wing lift, provide in-flight deceleration, and control airspeed during descent (Figure 27-42). Three spoiler panels are on each wing. The two outboard flight spoilers assist in controlling the aircraft laterally. All three spoiler panels are used during ground spoiler and speedbrake operations.


27-77



SPEEDBRAKE CONTROL ROD

GROUND SPOILER SERVO PRESSURE PORT

Figure 27-43. Ground Spoiler/Speedbrake Actuator

27-78




NOTES

Ground Spoiler/Speedbrake Actuator The hydraulic actuator for the ground spoilers and speedbrakes is located at the rear beam area of each wing (Figure 27-43). These actuators position the inboard spoiler surfaces during speedbrake operation or when deployed by the automatic ground spoiler control system. The ground spoiler/speedbrake actuators receive a mechanical control input for speedbrake operations, and a hydraulic control input for automatic ground spoiler deployment. The ground spoiler/speedbrake actuators also provide input to the flight spoiler actuators during ground spoiler/speedbrake operation. The flight spoiler actuator input is provided through the spoiler bungee, the mixing/summing linkage, and a pushrod.



27-79



GPWS/GND SPLR FLAP ORIDE ON:

SPOILER CONTROL ON:

OFF:

AMBER OFF LEGEND EXTINGUISHED. HYDRAULIC POWER IS PROVIDED TO FLIGHT SPOILERS, GROUND SPOILERS AND SPEEDBRAKES AMBER OFF LEGEND IS ILLUMINATED. HYDRAULIC POWER IS REMOVED FROM FLIGHT SPOILERS, GROUND SPOILERS, AND SPEEDBRAKES.

OFF:

AMBER ON LEGEND ILLUMINATED. AUTOMATIC GROUND SPOILER DEPLOYMENT WILL OCCUR IF FLAPS ARE LESS THAN 22° AND ALL OTHER PARAMETERS ARE SATISFIED. GPWS VOICE ALARM, "TOO LOW, FLAPS" IS INHIBITED. AMBER ON LEGEN IS EXTINGUISHED. AUTOMATIC GROUND SPOILER DEPLOYMENT WILL NOT OCCUR FROM WHEEL SPIN-UP IF FLAPS ARE LESS THAN 22° EVEN IF ALL OTHER PARAMETERS ARE SATISFIED. GPWS VOICE ALARM, "TOO LOW, FLAPS" IS NOT INHIBITED.

SPEEDBRAKE HANDLE EXTEND:

SPEED BRAKES WILL EXTEND ALL SIX SPOILERS 30°. LIGHT IN THE HANDLE WILL ILLUMINATE AS A REMINDER TO FLIGHT CREW. RETRACT: MOVES ALL SIX SPOILERS INTO A STOW POSITION.

FLAP HANDLE PROVIDES INPUT TO FCU. FOUR POSITIONS: UP (0°) 10° T/O APP (20°) DOWN (39°)

Figure 27-44. Secondary Flight Control Components

27-80




NOTES

Ground Spoiler Flap Override Switch The ground proximity warning system (GPWS) GND SPLR FLAP ORIDE switch is located on the center pedestal, forward of the flap handle (Figure 27-44). The switch provides the means to override the requirement for the flaps to be fully down before the ground spoilers can deploy from wheel spin-up. The switch also inhibits the GPWS aural warning “Too low, flaps”.

Spoiler Control Switch The spoiler control switch is located on the speedbrake control handle panel and is used to energize and deenergize the spoiler control shutoff valves (Figure 27-44).


27-81



LEFT SPOILER SHUTOFF VALVE

SPOILER PRESSURE CONTROL MODULE

Figure 27-45. Spoiler Control Shutoff Valves

27-82


RIGHT SPOILER SHUTOFF VALVE



NOTES

Spoiler Control Shutoff Valves The spoiler control shutoff valves are mounted to the aft side of the main wheel well rear bulkhead (Figure 27-45). They are normally open and are energized closed when the spoiler control switch is positioned OFF. When the valves are closed, hydraulic pressure is removed from the ground spoiler/speedbrake actuators and the flight spoiler actuators. The spoilers do not deploy in manual reversion. An amber “Spoilers Hydraulics Off” message is displayed on CAS when the valves are closed.

Ground Spoiler/Speedbrake Pressure Control Module A ground spoiler/speedbrake pressure control module is also mounted to the aft side of the main landing gear wheel well rear bulkhead (Figure 27-45). The module is plumbed into the ground spoiler/speedbrake actuator hydraulic pressure lines. To prevent structural damage when the speedbrakes are deployed at high airspeeds, the pressure control module reduces hydraulic system pressure to 1500 psi when both the left and right hydraulic systems are operating. If one hydraulic system fails, the remaining system’s pressure is returned to 3000 psi.

Speedbrake Handle The speedbrake control handle is located on the center pedestal (see Figure 27-44). It establishes the spoiler control surface maximum travel and has an infinite selection between stowed and full deployment. The speedbrake handle movement is transferred via a 7x7, 3/32-inch cable loop system, which extends from the speedbrake handle aft. The cables are then routed from the left and right wheel wells outboard through a series of pulleys on each wing and are connected to the speedbrake sector crank assemblies. The speedbrake sector crank repositions a control pushrod connected to the ground spoiler/speedbrake actuator servo valve input lever (see Figure 27-43).


27-83



LEFT PTU PRESSURE AUX PRESSURE L AND R THROTTLES

AIR

L WOW

LEFT WHEEL SPINUP

GND

IDLE

PRIMARY CONTROL VALVE ARMED

> 53 kts

GRD SPOILER FLAPS > 22°

PRESSURE SWITCH

RIGHT WHEEL SPINUP 28 VDC RIGHT ESS BUS

EGPWS GRD, SPRL FLAP O'RIDE SWITCH AIR R WOW

> 53 kts

GND

GND SPLR

L AND R THROTTLES

IDLE

SECONDARY CONTROL VALVE ARMED

TO LEFT ACTUATOR

Figure 27-46. Ground Spoiler Schematic

27-84


TO RIGHT ACTUATOR



Speedbrake Operation Retraction of the ground spoiler/speedbrake actuator rotates a bellcrank that moves a pushrod, which deploys the ground spoiler panel. Also attached to the bellcrank is a spoiler bungee, which, acting as a fixed rod, repositions the mixing/summing linkage and a pushrod so that ground spoiler/speedbrake actuator retraction simultaneously provides control inputs to the servo valves on the flight spoiler actuators (see Figure 27-6). The mixing/summing link ensures that the flight spoiler panels are deployed to the same angle as the ground spoiler panel. To slow the aircraft in flight using speedbrakes, the speedbrake control handle is moved more than 1° aft from the stowed detent position, which causes a light in the handle to illuminate. Power for the light is provided by the right essential 28 VDC bus. With the handle in the fully extended position, all six spoiler surfaces deploy to a maximum 30°. If lateral control inputs are executed with the speedbrake deployed, the flight spoilers on the side of the turn will surpass the 30° limit to a maximum of 55 ±4°.


Automatic Ground Spoiler Control The automatic ground spoiler control system is used to shorten aircraft ground braking distances by spoiling wing lift and quickly placing the full weight of the aircraft onto the landing gear. This decreases the chance of “ballooning.” The system actuates all spoiler panels to the full 55° position. The ground spoiler primary and secondary control valves will be energized if one of the following conditions exist:(Figure 27-46): CONDITION 1 • GND SPLR switch is in the ARM position • Both throttles retarded to IDLE position • WOW is positive (aircraft on ground), or flaps extended more than 22° with wheel spinup greater than 53 knots CONDITION 2 • GND SPLR switch in the ARMED position • Both throttles retarded to IDLE position • WOW is positive (aircraft on ground), or flaps extended less than 22° with wheel spin-up greater than 53 knots and GPWS/GND SPLR FLAP ORIDE switch ON


27-85



PRESSURE SWITCH

PRIMARY CONTROL VALVE

Figure 27-47.

27-86

SECONDARY CONTROL VALVE

Primary and Secondary Control Valves (Left Main Wheel Well)




NOTES

Primary and Secondary Control Valves The ground spoiler primary and secondary control valves are located on the left main wheel well aft bulkhead (Figure 27-47). They control the left/PTU or auxiliary hydraulic system control (pilot) pressure to the ground spoiler/speedbrake actuator servo control valves. The valves are hydraulic in series, and each has its own electrical control circuit. Both control valves must be electrically energized for ground spoiler deployment.

Ground Spoiler Pressure Switch A ground spoiler control system pressure switch is located between the primary and secondary control valves in the left main wheel well (Figure 27-47). The pressure switch will trigger a red “Ground Spoiler” message on CAS on the ground when the system is unarmed and the following conditions exist: • One or both control valves are energized. • Ground spoiler pressure switch indicates pressure. • One or both ground spoiler throttle monitor points are energized in the control circuit. • One or both ground spoiler panels are unstowed with the speedbrake handle in the stowed detent. A red “Ground Spoiler” warning will also trigger in flight when the same conditions exist and one or both ground spoiler touchdown points are energized in the control circuit. A “Ground Spoiler” warning will also occur with the system armed and either throttle lever out of idle position if one or both control valves are energized or one or both ground spoiler throttle monitor points are energized in the control circuit. One or both ground spoilers unstowed with the throttle levers out of idle on the ground will trigger a “Ground Spoiler” warning any time.


27-87



MASTER WARN (Glareshield) PUSH BUTTON, RESET (RED) W. PILOT AND COPILOT ALERT OF WARNING MESSAGE.

NO GND SPLRS ILLUMINATES RED IF: • OPERATIONAL LOGIC PARAMETERS ARE SATISFIED. • GROUND SPOILER SWITCH SELECTED TO "ARMED". • LEFT WING, RIGHT WING OR BOTH WING GROUND SPOILERS DID NOT DEPLOY.

GND SPLR TEST ILLUMINATES "IN TEST" (BLUE) WHEN SELECTED. THE FOLLOWING ACTIONS OCCUR: • GROUND SPOILERS REMAIN STOWED. • "NO GND SPLR" (RED) LIGHTS (2) ARE ILLUMINATED. • "GROUND SPOILER" (RED) MESSAGE DISPLAYED ON CAS. • "MASTER WARN" LIGHTS ILLUMINATE (2). • 3 CHIME AURAL WARNING TONE SOUNDS. GND SPLR OFF: • AMBER "OFF" LEGEND IS ILLUMINATED. • INDICATES GROUND SPOILER SYSTEM IS NOT ARMED. ARMED: • BLUE "ARMED" LEGEND IS ILLUMINATED. • GROUND SPOILER SYSTEM IS ARMED.

Figure 27-48. Ground Spoiler Switches

27-88




NOTES

Ground Spoiler Switch The ground spoiler switch is located on the center pedestal and has two modes: OFF and ARMED (Figure 27-48). In the ARMED position, the primary and secondary solenoid control valves will energize and provide hydraulic control (pilot) pressure to the ground spoiler/speedbrake actuators when all conditions are satisfied. In the OFF position, the primary and secondary solenoid control valves remain deenergized and a blue “Ground Spoiler Unarm” message is generated with the left main landing gear down and locked.

Ground Spoiler Test Switch The ground spoiler test switch is located on the center pedestal, immediately above the ground spoiler OFF–ARMED switch (Figure 27-48). With the aircraft properly configured and the test switch selected, a preflight functional checkout of the automatic ground spoiler control and warning system is accomplished. When the test is in progress, an IN TEST light illuminates on the switch, along with a NO GND SPLRS center post warning light and a red “Ground Spoiler” message on CAS. The master warning light also illuminates.

NOTE Refer to the Aircraft Maintenance Manual for proper ground spoiler testing procedures.


27-89



RVDT

STOW SWITCH

Figure 27-49. Ground Spoiler RVDT and Stow Switch

27-90




NOTES

Ground Spoiler Stow Switches The ground spoiler stow switches are mounted on the wing rear beams and illuminate the red NO GND SPLRS warning on the windshield center post when the ground spoilers are not automatically deployed after landing (Figure 2749). The switch also triggers a red “Ground Spoiler” CAS message in flight when the speedbrake handle is in the retract detent and the spoiler panels are not stowed.

Ground Spoiler RVDTs Ground spoiler rotary variable differential transducers (RVDTs) are located outboard of the stow switches and are connected to the ground spoiler panels. The left ground spoiler RVDT receives excitation voltage from MAU No. 1(A), AFCS1-A slots 5/6. The right ground spoiler RVDT receives excitation voltage from MAU No. 2 (B), AFCS2-A slots 9/10. The MAUs transmit ground spoiler position data to the FDR and the FLIGHT CONTROLS synoptic page.


27-91



SPOILER CONTROL SWITCH

LEFT HYDRAULIC SYSTEM (P1)

RIGHT HYDRAULIC SYSTEM (P2) SPOILER CONTROL SHUTOFF VALVE

SPOILER CONTROL SHUTOFF VALVE

TO FLIGHT SPOILER ACTUATORS (3000 PSI)

TO FLIGHT SPOILER ACTUATORS (3000 PSI) GROUND SPOILER PRESSURE CONTROL MODULE (1500 PSI)

55°

SERVO VALVE

STOW SWITCH

RVDT

GROUND SPOILER/SPEEDBRAKE ACTUATORS SECONDARY CONTROL VALVE

PRESSURE SWITCH

GROUND SPOILER OFF/ARMED SWITCH (ARMED)

PRIMARY CONTROL VALVE LEFT/PTU/AUX PRESSURE

Figure 27-50. Automatic Spoiler Control System Operation

27-92




NOTES

Automatic Spoiler Control System Operation With the spoiler control shutoff valves open, right and left hydraulic systems continuously provide actuating pressure to the flight spoiler actuators. Hydraulic pressure is also directed from the shutoff valves through the spoiler pressure control module (Figure 27-50). The pressure control module limits and directs 1,500-psi pressure to the ground spoiler/speedbrake actuators. Control (pilot) pressure is provided continuously to the primary control valve from the left/PTU or auxiliary hydraulic system. Upon touchdown, if all requirements for ground spoiler deployment are met, the primary and secondary control valves are electrically energized, and pilot pressure is ported to the ground spoiler/speedbrake actuator servo control valves. The left and right system pressure will then retract the actuators, and all six spoiler surfaces deploy to 55°.



27-93



Ground Spoiler Speed Brake Extended Spoilers Hydraulics Off Ground Spoiler Unarm Single Speed Brake Speed Brake Extended

Figure 27-51. Ground Spoiler Control System Indications

27-94




NOTES

Ground Spoiler Control System Indications The ground spoiler/speedbrake surface position is indicated on the FLIGHT CONTROLS synoptic page. Invalid data from an RVDT results in an amber “X” over the spoiler indication (Figure 27-51). CAS messages associated with the ground spoiler/speedbrake system are listed below:

Red Messages Ground Spoiler Amber Messages Speed Brake Extended Spoilers Hydraulics Off Blue Messages Ground Spoiler Unarm Single Speed Brake Speed Brake Extended A red NO GND SPLRS message will illuminate on the windshield center post when the ground spoilers do not deploy after all conditions are met. The red “Aircraft Configuration” CAS message will illuminate: • On the ground if the speedbrakes are deployed and either throttle is advanced toward takeoff • In flight if the flaps are fully extended or the landing gear is extended with the speedbrakes deployed.


27-95



FLAP

FLAP CONTROL HANDLE

POWER DRIVE UNIT

PDU

FCU

FLAP CONTROL UNIT

HSA HORIZONTAL STABILIZER ACTUATOR

Figure 27-52. Flap/Stabilizer Component Locations

27-96




NOTES

FLAP/HORIZONTAL STABILIZER SYSTEM The flaps provide lift augmentation for takeoff and landing. The flap system converts aircraft-supplied hydraulic power into a mechanical actuation force to position the flaps. As the flaps are extended or retracted, the horizontal stabilizer is positioned leading edge up (LEU) or leading edge down (LED) to compensate for the aerodynamic changes created by flap movement.

Component Locations and Functions Flap Control Handle The flap control handle is located on the right side of the center pedestal and commands the flap control unit (FCU) to place the flaps in one of the four following positions (see Figure 27-44): 0° (fully up) 10° 20° (takeoff/approach) 39° (fully down).


27-97



AUTORIG SWITCHES

FLAP/HORIZONTAL STABILIZER CONTROL UNIT AUTORIG DATA MODULES

Figure 27-53. Flap/Stabilizer Control Unit

27-98




NOTES

Flap/Stabilizer Control Unit The flap control unit (FCU) is a dual channel controller located in the baggage compartment electronic equipment rack (BEER) (Figure 27-53). Each channel of the FCU uses both AC and DC power. The FCU provides independent dual-channel operational control, protection, and calibration of the flap and stabilizer operations. This completely eliminates the need for an emergency flap system. Each channel consists of a microprocessor controller that is isolated from the other electrically and mechanically. The flap actuator system (FAS) and horizontal stabilizer actuator (HSA) are controlled by both channels of the FCU, simultaneously. Each channel controls one channel of the FAS and one channel of the HSA. The unit also performs operational status testing with built-in test and funct i o n a l fa u l t m o n i t o r i n g f o r t h e f l a p a n d stabilizer systems. Channel 1 of the FCU is powered by the left standby 115 VAC bus and the left essential 28 VDC bus. Channel 2 is powered by the right standby 115 VAC bus and the right essential 28 VDC bus. Two removable autorig data modules are installed on the front of the FCU and are used during the autorig function to record the flap and stabilizer rig positions and provide the offset data necessary for the FCU to accurately define the positions (Figure 27-53). Upon receipt of the autorig command, the FCU calibrates the flap and stabilizer resolver positions through the use of the autorig mode select, enable, and command switches. The FCU also provides flap/stabilizer control surface position indications to the MAUs via the ARINC 429 bus. The MAUs transmit the data to the FDR and the FLIGHT CONTROLS synoptic page.

NOTE Refer to the Aircraft Maintenance Manual for proper flap and stabilizer rigging procedures.


27-99



ENABLE SOLENOID

RETRACT SOLENOID VALVE MODULE POWER DRIVE UNIT

HYDRAULIC MOTOR

GEAR BOX

Figure 27-54. Flap Power Drive Unit

27-100




NOTES

Power Drive Unit The power drive unit (PDU) is mounted at the centerline of the aircraft, at the upper forward end of the torque box between the main wheel wells (Figure 27-54). The PDU consists of a control valve module and an integrated hydraulic motor/gearbox and provides hydraulic control and power conversion for the flap system. A fixed-displacement, axial hydraulic motor is splined to the gearbox and converts aircraft hydraulic power into bidirectional rotary shaft power to drive the mechanical flap actuators via the torque tubes. Independent hydraulic power sources (left/PTU or auxiliary) exist for failure redundancy. Two output shafts, driven from a single reduction-type gearbox, rotate to drive the torque tubes that operate the left and right flap ballscrew actuators.


27-101



RESOLVER

OUTBOARD ACTUATOR

TORQUE TUBES (7)

INBOARD ACTUATOR

PILLOW BLOCK (5)

PDU

Figure 27-55. Flap System Torque Tubes and Flap Actuators

27-102




NOTES

Flap System Torque Tubes The aluminum torque tubes transmit the power drive unit torque to the inboard and outboard actuators (Figure 27-55). The seven tubes are attached to the rear beam structure of each wing and are supported at five points by fixed pillow block radial bearing assemblies. Torque tube assemblies are bolted at one end and “float” in the splined connection at the other end. Universal joints are incorporated in each torque tube.

Flap Actuators Each flap is positioned by two ballscrew actuators, which contain a force limiter that protects the aircraft from structural damage if a flap roller or track jam occurs (Figure 27-55). The force limiter is bidirectional and is selfresetting by reversing the direction of rotation. The actuators also have a no-back device that prevents airloads on the flap panels from attempting to retract the flaps. A resolver is mounted on the outboard side of each outboard actuator and is directly coupled to the actuator spline drive. The resolver is dual-channel and supplies flap position and offset data to the FCU. The FCU will then transmit the position information to the MAUs.


27-103



UPPER MOUNT

BALLSCREW

ACTUATOR

RESOLVER MANUAL DRIVE NUT ACCESS CAP

UPPER GEARBOX HOUSING

MANUAL DRIVE UNIT LOWER GEARBOX HOUSING

AC MOTOR ASSEMBLY LOWER MOUNT

Figure 27-56. Horizontal Stabilizer Actuator

27-104




NOTES

Horizontal Stabilizer The horizontal stabilizer is located at the top of the vertical fin and pivots at the aft mounting attachment point (Figure 27-56). Extension of the Fowler-type flaps increases wing chord, thus moving the center of lift aft. This change in lift is compensated by adjusting the stabilizer leading edge down (LED) according to a schedule (Table 27-1).

Horizontal Stabilizer Actuator The horizontal stabilizer actuator (HSA) is an electrically-driven ballscrew actuator that is located in the vertical fin leading edge below the horizontal stabilizer (Figure 27-56). The actuator uses a dual-channel AC motor to position the stabilizer in accordance with FCU commands and cockpit input. A resolver is incorporated on the actuator to provide position feedback information to the FCU. A force limiter prevents structural damage in the event of a jammed actuator. The actuator also has a dual “no-back” device, which locks the actuator in the desired position, preventing air loads from moving the horizontal stabilizer.

Table 27-1. FLAP/HORIZONTAL STABILIZER SYNCHRONIZATION SCHEDULE FLAP POSITION vs HORIZONTAL STABILIZER ANGLE 0° 10° 20° 39°

–1.5° LED –2.6° LED –3.6° LED –4.6° LED


27-105



EMER STAB ARM: • AMBER "ARM" LEGEND IS ILLUMINATED. • "EMER STAB" ,MODE IS ENABLED. • STABILIZER POSITIONED BY USING ELECTRIC PITCH TRIM SWITCH ON CONTROL WHEELS. OFF • AMBER "ARM" LEGEND IS EXTINGUISHED. • "EMER STAB" MODE IS DISABLED. STABILIZER MOVES CORRESPONDING TO FLAP POSITION.

Figure 27-57. Emergency Stabilizer Switch

27-106




NOTE

Flap/Horizontal Stabilizer Operation

If the stabilizer actuator no-back mechanism trips, the stabilizer will lock in the full leading edge down (LED) position, preventing further stabilizer movement. The procedure to reset the actuator no-back can be found in the Aircraft Maintenance Manual.

The flaps are normally powered by the left hydraulic system, but can be powered by the PTU or auxiliary systems. Flap commands are initiated by repositioning the flap handle. The flap handle is connected to a dual-channel RVDT, which is powered by 28 VDC excitation voltage from the FCU. The command signal then passes from the RVDT back to the FCU. The FCU compares the commanded signal to the flap and stabilizer position transmitted by the resolvers. If the flap position does not match the handle position, the FCU energizes the solenoid control valves within the hydraulic control valve module of the PDU to port system pressure to the hydraulic motor, which drives the gearbox. The gearbox rotates the torque tubes connected to the flap actuators. The FCU will simultaneously energize the dual channel AC-powered stabilizer motor that will position the stabilizer actuator to a position corresponding to flap position. The basic principle of operation of the inboard and outboard actuators is identical. The length and pitch of the ballscrews are different between the inboard and outboard actuators to accommodate the different stroke requirements. When the flaps and stabilizer reach the commanded position, the FCU deenergizes the solenoid control valves and the stabilizer motor.

Emergency Stabilizer Switch The emergency stabilizer switch is a guarded switch located on the center pedestal ( Figure 27-57). The switch decouples the automatic synchronization between the horizontal stabilizer and the flaps. When the EMER STAB switch is selected to ARM, it enables the electric pitch trim (beep) switches on the control wheels to send commands to the FCU to move the stabilizer independently of the flaps. The electric pitch trim system is immediately disengaged when the EMER STAB switch is armed.

Emergency Stabilizer Operation The stabilizer emergency mode is selected to allow the flight crew to get more pitch trim au-


27-107

27-108 3.50

8.5 Down


HDG

TO

TO

129

+800

100 00

80 60

20

20

10

10

-900

500

6 2 1

1 40 9

5

5

0 60

0.03M

-15

10

10

1

20

20

2 6

29.92 in

-5

+13 No Bearing Targets RA 6.1nm -08 TA 2.5nm -15

AOA 0.17 MAG1

358 N

5 5

33 30

UP

0

T / 0

10

DN STB

APU

RPM 101.0

DBN

STBY TERMINAL

3

Open

6

RNP 1.00 EPU 0.03 VOR1 SAV/112.7 Ident 068/ 7nm

20 39 FLP

EGT 495

VOR2 ALD/116.7 Ident 357/ 54nm

Right 45 psi

Left 45 psi Bleed Air Pressure

Figure 27-58. Flap/Stabilizer Indications

international

FLAP/STABILIZER POSITION

FMS2

86.8 nm

FlightSafety

HDG 358

DTRK 293


Emergeny Stabilizer On Flap Asymmetry Flap Command Invalid Flaps Failed Flap/Stab Independent Op Flap/Stab Miscompare Flap/Stab Sync Fail Stabilizer Failed Uncommanded Flaps Uncommanded Stabilizer Flap/Stab Maint Reqd A-B Flap/Stab Rig Complete Flap/Stab System Fail A-B Stabilizer Syncing A-B



thority in the event of a jammed elevator control surface. With this mode selected, the flap/stabilizer synchronization is decoupled. The control wheel pitch trim switches are used to command stabilizer movement via the FCU. The FCU directs the stabilizer motors to drive the stabilizer at a constant rate. The range of stabilizer travel is increased in the emergency mode of operation. When moving the horizontal stabilizer with the pitch trim switches, the stabilizer can be moved from –4.6° LED to a +1.5° leading edge up (LEU). The EMER STAB switch can also be used when moving the flaps with the auxiliary hydraulic system for maintenance checks. Selecting EMER STAB to ARM will allow the flaps to move while the horizontal stabilizer remains stationary. When the EMER STAB switch is selected OFF, the FCU will synchronize the horizontal stabilizer to the selected flap position provided AC power is applied to the aircraft.

Flap/Stabilizer Indications The FCU provides position and status information for display on the CAS and PFD (Figure 27-58). The flaps display on the FLIGHT CONTROLS synoptic page has a green raster, and digital readout. The stabilizer position is also displayed in degrees Leading Edge Down. The stabilizer position indication will turn amber when EMER STAB switch is in the ARM position. The PFD displays a green line for flap handle position and arrows for STB and FLP position. Some of the flap/stab system amber and blue CAS messages are listed below:

CAS Amber Messages Emergency Stabilizer On Flap Asymmetry Flap Command Invalid Flaps Failed

Flap/Stab Troubleshooting Troubleshooting of the flap/stab system requires the use of the CMC parameters pages. The FCU utilizes the inputs from several different LRUs. Some of these send inputs in degrees of position, and the accurate input of that position is critical to the function of the system. The FCU requires the proper response from the PDU, along with the proper position, to continue the movement of the flaps and the horizontal stabilizer. It should be noted that once an initial failure occurs and is indicated, the FCU may log several subsequent failures as a result. For example, if one of the flap resolvers were improperly set, the result would be a stoppage of movement of the flap and stab. Not only would the stoppage be indicated, but several other CMC faults may be displayed as well. Understanding the initial failure will keep the technician from troubleshooting the wrong message or failure. Consulting the Fault Isolation Manual is critical in the troubleshooting of the flap/stab system.

Flap/Stab Independent Op Flap/Stab Miscompare Flap/Stab Sync Fail Stabilizer Failed Uncommanded Flaps Uncommanded Stabilizer

CAS Blue Messages Flap Stab Maint Reqd A-B Flap/Stab Rig Complete Flap/Stab System Fail A-B Stabilizer Syncing A-B The FLIGHT CONTROLS synoptic page will also indicate invalid display data or loss of data (Figure 27-58).


27-109



GUST LOCK HANDLE • FORWARD AND DOWN ("OFF" POSITION) RELEASES GUST LOCKS. • AFT AND UP ("ON" POSITION) ENGAGES THE GUST LOCKS, LOCKING AILERONS AND RUDDER IN NEUTRAL POSITION AND ELEVATORS IN TRAILING EDGE DOWN POSITION. IN ADDITION, POWER LEVER MOVEMENT IS RESTRICTED TO NO MORE THAN 6° ABOVE GROUND IDLE.

Figure 27-59. Gust Lock Handle

RR0728B

27-110




NOTES

GUST LOCK SYSTEM The gust lock system protects the ailerons, elevators, and rudder from wind gusts up to 60 knots while the aircraft is on the ground.

Component Locations and Functions Aileron Gust Lock The aileron gust lock mechanism is installed below the cabin flooring and secures the ailerons at the wing faired position.

Elevator/Rudder Gust Locks and Bungees The elevator and rudder gust locks are located in the tail compartment. The locks secure the rudder at the faired position and the elevators with the trailing edge down. A bungee is installed prior to each gust lock latch. With the gust lock handle in the OFF position, the bungees act as fixed rods to unlock the latches. The bungees allow the gust lock handle to be engaged while the surfaces are out of the locked position.

Gust Lock Control Handle The gust lock control handle is located on the right side of the center pedestal and locks or unlocks the mechanical latches (Figure 2759). A spring latch, located underneath on the forward side of the handle, must be unlocked before the control handle can be moved in either direction. A mechanical interlock is incorporated between the gust lock handle and the throttles, which restricts throttle lever movement to a maximum of 6° above ground idle when the gust lock handle is in the locked position. The gust lock handle cannot be moved when the throttles are out of the 6° range.


27-111



CHAPTER 28 FUEL CONTENTS Page INTRODUCTION ................................................................................................................. 28-1 GENERAL ............................................................................................................................ 28-1 FUEL STORAGE SYSTEM ................................................................................................. 28-5 General........................................................................................................................... 28-5 Wing Tanks .................................................................................................................... 28-5 Fuel Hopper ................................................................................................................... 28-7 Gravity Fueling .............................................................................................................. 28-9 Water and Fuel Drainage ............................................................................................. 28-11 Fuel Ventilation............................................................................................................ 28-13 Heated Fuel Return System ......................................................................................... 28-17 Fuel Storage System Operation ................................................................................... 28-19 FUEL DISTRIBUTION SYSTEM ..................................................................................... 28-21 General......................................................................................................................... 28-21 Pressure Fueling and Defueling................................................................................... 28-21 Pressure Fueling Shutoff.............................................................................................. 28-23 Components ................................................................................................................. 28-25 Operation ..................................................................................................................... 28-29 Fuel Crossflow and Intertank Transfer System ........................................................... 28-31 Engine and APU Fuel Distribution System ................................................................. 28-35


28-i



FUEL INDICATION SYSTEM .......................................................................................... 28-47 General......................................................................................................................... 28-47 Components ................................................................................................................. 28-47 Operation ..................................................................................................................... 28-57

28-ii





Title

Page

28-1

Wing Fuel Load...................................................................................................... 28-2

28-2

Pressure Fueling Adapter ....................................................................................... 28-3

28-3

Wing Tank .............................................................................................................. 28-4

28-4

Fuel Hopper............................................................................................................ 28-6

28-5

Gravity Fueling Port............................................................................................... 28-8

28-6

Drain Valve Cross-Section ................................................................................... 28-10

28-7

Vent Ducts, Plenum, and Valves .......................................................................... 28-12

28-8

Inboard Vent Valves ............................................................................................. 28-14

28-9

Vent Duct Drain Valve ......................................................................................... 28-14

28-10

Engine Heated Fuel System Block Diagram ....................................................... 28-16

28-11

Backup Heated Fuel Return Valve—Pylon Location .......................................... 28-18

28-12

Pressure Fueling/Defueling Components ............................................................ 28-20

28-13

Ground Service Control Panel ............................................................................. 28-22

28-14

Fueling Shutoff Control—Block Diagram........................................................... 28-24

28-15

Fuel Shutoff Control—System Components ....................................................... 28-26

28-16

Manual Precheck Valves ...................................................................................... 28-28

28-17

Fuel Crossflow and Intertank Transfer System.................................................... 28-30

28-18

Intertank and Crossflow Switch........................................................................... 28-32

28-19

Fuel Boost Pumps ................................................................................................ 28-34

28-20

Boost Pump Control Schematic ........................................................................... 28-36

28-21

EMI Filter............................................................................................................ 28-37

28-22

Fuel Boost Pump Manifold.................................................................................. 28-38

28-23

Fuel Ejector Pumps .............................................................................................. 28-40


28-iii



28-24

Engine Fuel Shutoff Valve ................................................................................... 28-42

28-25

APU Fuel Shutoff Valve ...................................................................................... 28-42

28-26

Engine and APU Fuel Distribution ...................................................................... 28-44

28-27

Fuel Indication System Block Diagram............................................................... 28-46

28-28

Signal Conditioner Location ................................................................................ 28-47

28-29

Fuel Probes........................................................................................................... 28-48

28-30

Densitometer ........................................................................................................ 28-49

28-31

Dummy Probe and High-Level Sensor ................................................................ 28-50

28-32

System Monitor Test Panel .................................................................................. 28-52

28-33

Fuel Boost Pump Pressure Switch ....................................................................... 28-54

28-34

Fuel Synoptic Page .............................................................................................. 28-56

28-35

Fuel Quantity Block Diagram.............................................................................. 28-58

28-36

Refuel/Test Switch ............................................................................................... 28-59

28-37

High-Level Warning Light and Switch ................................................................ 28-60

28-iv




CHAPTER 28 FUEL

6

4 MAIN FUEL 2

8

LBS X 100 0

10

INTRODUCTION The Gulfstream G500/G550 fuel system provides fuel for the two Rolls-Royce Deutschland BR710 turbofan engines and the auxiliary power unit (APU), as well as continuous fuel quantity and system information to the crew. There are two integral (wet wing) fuel tanks, each formed by the respective wing’s structure.

GENERAL A single-point pressure-fueling adapter is provided for fueling the tanks (Figure 28-2). These tanks can also be fueled from two overwing (gravity) fuel ports. Total fuel capacity of the G550 is 41,300 pounds for a total of 5,118 U.S. gallons (23,127 liters).

A fuel hopper, which is an isolated compartment with a capacity of 190 U.S. gallons within each wing tank, provides fuel for each engine and the APU. Four fuel boost pumps remove the fuel in the hoppers through fuel feed lines to the APU and the left and right engines.


28-1



TOTAL FUEL LOAD = 41,300 LBS (6,118 U.S. GALLONS)

20,650 LBS

20,650 LBS

APU

Figure 28-1. Wing Fuel Load

28-2




The wing tank vent system provides sufficient venting while the aircraft is on the ground and during flight. It also lightly pressurizes each wing tank during flight. The fuel quantity is measured by 38 transistorized capacitancetype fuel level probes, with 19 per tank. The left and right fuel tank probes operate independently and supply separate signals to individual left and right fuel quantity processors within the fuel signal conditioner. Fuel quantity is displayed on the CAS, along with fuel flow and fuel temperature.The MCDU

also has a backup fuel quantity indication. The CAS display contains the fuel synoptic page, which provides the flight crew with specific information about the fuel system. The ground service control panel, located on the front of the LEER circuit-breaker panel, provides control for automatic pressure-fueling operations. It displays fuel quantity on the ground only, provides a diagnostic check of the fuel system, and allows for preselecting fuel quantities during refueling.

PRESSURE FUELING ADAPTER

Figure 28-2. Pressure-Fueling Adapter


28-3

28-4

BL 0.0

BL 62.9

PROBES RBS 135.0

WING TANK

7

RBS 214.0


RBS 294.0

6

5 8

RBS 373.5

RBS 453.0

4 DENSITOMETER (Left Hopper Only)

3

RBS 506.0 2 WING ACCESS DENSITOMETER (Right Wing Only) PANEL

1

GRAVITY REFUELING ADAPTER

international

FlightSafety

HIGH-LEVEL FUEL SENSOR


DENSITOMETER

WS 520.36 (VOUGHT) BL 531.00 (88.5FT)

Figure 28-3. Wing Tank



FUEL STORAGE SYSTEM

Twenty-one wing access panels are installed in the lower wing skin of each tank. They permit access to the full internal part of the wing tank for inspection, maintenance, and repair.

GENERAL The purpose of the fuel storage system is to provide the necessary fuel storage for the aircraft. The system also provides gravity-filling capability, drainage, venting, and fuel heating.

NOTES

The fuel storage system consists of the following: • Wing fuel tanks • Fuel hoppers • Gravity fueling • Fuel drainage • Fuel ventilation • Heated fuel return system

WING TANKS The wing fuel tanks carry the total fuel load for the two turbofan engines and the APU. Each wing fuel tank consists of an integral fuel tank and baffle ribs. The wing tank is composed of the wing’s front and rear spars, the upper and lower skins, and closure ribs at BL 0.0 and RBS 506.0 (Figure 28-3). Each wing tank is divided into seven compartments by six baffle ribs. The baffle ribs prevent a sudden shift in weight due to fuel movement. Three to five flapper-type check valves are installed near the bottom of each rib. These check valves prevent fuel flow outboard and permit fuel to gravity-drain inboard during gravity fueling and engine operation. Weep holes are provided at the bottom of the ribs to prevent fuel and water from being trapped there. They allow fuel and water to drain inboard to the drain valves. Small openings are provided at the top of each baffle rib to allow fuel flow between compartments during pressure fueling.


28-5



FUELING CHECK VALVE (TYPICAL 8 PLACES) HOPPER ACCESS PANEL (TYPICAL 4 PLACES)

BL 35.8 BL 62.9 RIB

BL 18.5 BL 6.0

UP

D FW

BL 0.0

EJECTOR PUMP (TYPICAL 2 PLACES

REAR SPAR

ENGINE FUEL SHUTOFF VALVE (TYPICAL 2 PLACES) APU SHUTOFF VALVE

CROSSFLOW SHUTOFF VALVE

INTERTANK SHUTOFF VALVE

FUEL BOOST PUMP (TYPICAL 4 PLACES)

BL 62.9 RIB HYDRAULIC HEAT EXCHANGER (TYPICAL 2 PLACES)

Figure 28-4. Fuel Hopper

28-6




FUEL HOPPER

NOTES

The fuel hopper (Figure 28-4) is an isolated compartment in each fuel tank, located on each side of the centerline rib (BL 0.0). The fuel hoppers are made by the rear spar, the centerline rib, the baffle rib at BL 62.9, and a front wall located 43.5 inches forward of the rear spar. Two access panels are installed in the forward wall of each hopper. These access panels are removed for inspection and repair procedures. The components of each fuel hopper are as follows: • Fuel boost pump inlet filter screen and manifold • Fuel ejector • Crossflow shutoff valve (left hopper) • Hydraulic fluid heat exchanger (radiator type) • Intertank valve (right hopper) The following components are located in the hoppers, but not shown: • Densitometer (left hopper) • Low level sensor • Fuel tank temperature probe • Compensators • Fuel quantity probes • Water/fuel drain valve


28-7



GRAVITY FUELING PORT

GRAVITY FUELING PORT

3-INCH FILLER CAP D

FW

OPEN D

OS E

CL

ADAPTER AND FILTER SCREEN

Figure 28-5. Gravity Fueling Port

28-8




GRAVITY FUELING

NOTES

Two overwing gravity fueling ports (Figure 285) are located on top of the wing, just outboard of RBS 479.5. Gravity fueling provides a means to refuel the tanks if pressure fueling is not available. Each fueling port contains a gravity-fueling adapter assembly, which consists of a 3-inchdiameter filler cap and an adapter with a filter screen. The assembly is installed in the wing tank upper skin and supplies a mount for the filler cap and filter screen. The assembly prevents the fueling nozzle from being inserted too far into the tank and possibly damaging the lower wing skin. It also prevents large objects from going into the wing tank during gravity fueling. A refuel safety grounding jack is located in the wing’s leading edge, close to the gravity fueling adapter. It is used to ground the aircraft during fueling operations.


28-9



NUT

LOCK WASHER

0-RING

VALVE BODY

POPPET VALVE RETAINER

NUT

SPRING

SCREEN

0-RING

Figure 28-6. Drain Valve Cross-Section

28-10




WATER AND FUEL DRAINAGE The water and fuel drain valves are flushmounted to the lower wing skins of each tank. They are used for draining water and fuel. On aircraft SNs 5001–5034, two valves are located forward of the main gear well; the forward valve drains the tank area in front of the hopper, and the rear valve drains the hopper. The third drain valve, located near the wingtip outboard of RBS 506.0, is used for draining fuel from the vent plenum. On aircraft SNs 5035 and subsequent, there are three drains forward of the main wheel well. Two valves drain the fuel tank forward of the hopper and one valve drains the hopper. Each drain valve contains the following (Figure 28-6): • Poppet valve • Retainer • Screen

A secondary valve seat on the valve stem permits replacement of the O-ring with the tanks fueled. Using a Phillips screwdriver, push the valve stem in slightly (about 1/32 inch); rotate it clockwise until it stops, and then release. The internal spring moves the valve stem below the skin surface, where its O-ring can be replaced. As the stem moves down, the secondary valve seat engages the valve body and prevents fuel flow from the tank. A small quantity of fuel may leak as the valve stem drops down. Install the new O-ring carefully, as it can be pinched and could cause leaking. After the O-ring has been installed, return the valve to its original condition. Using a Phillips screwdriver, push the poppet valve stem all the way in to a hard stop. Then rotate it counterclockwise until it stops. The new valve O-ring engages the valve body to stop fuel flow. Perform a final check for any fuel leaks.

• Spring

NOTES

• O-ring • Valve body • Nut The nut contains the screen to keep contamination out of the poppet valve. It also secures the drain valve to the lower wing skin. To drain water and residual fuel from the tank, open the poppet valve on 79C3T2 with a Phillips screwdriver. Push the valve stem all the way in with the screwdriver to a hard stop. Then rotate it 90° counterclockwise and release. The pin engages a slot in the retainer to hold the valve in the open position. Holes in the valve body allow fuel to flow out of the t a n k . U s e a 1 / 8 - i n c h h e x d r iv e o n 79C20E/79C20H valves. To close the drain valve, push the stem of the poppet valve all the way in with a screwdriver to a hard stop. Then rotate it 90° clockwise again and release. The spring closes the poppet valve, and the valve O-ring engages the valve body to stop fuel flow.


28-11



GASKET

VENT FLOAT VENT FLOAT (VENT DUCT) VALVE (2) VENT FLOAT (AIR PASSAGE) VALVE (2)

VENT PLENUM

OVERBOARD VENT/RAM AIR INLET

Figure 28-7. Vent Ducts, Plenum, and Valves

28-12




FUEL VENTILATION

Vent Float Non-Relieving Valves (Air Passage)

General The purpose of the fuel ventilation system is to allow air and fuel vapor to flow in and out of the fuel tanks. It also prevents excessive pressure or vacuum in the fuel tanks and allows fuel to flow overboard to relieve pressure if necessary.

Components The fuel ventilation components consist of the following: • Vent ducts • Outboard vent plenum

Two vent float non-relieving valves are attached to the upper air passage of the tank between RBS 479.5 and RBS 506.0. These valves are not connected to the forward and aft vent ducts; however, they are connected directly to the vent plenum. The valves open and close as necessary to allow air and fuel vapor to flow in and out of the fuel tanks. This prevents too much pressure or vacuum in the tanks. When the aircraft is fueled to a full load, these valves are the last to close immediately before fuel shutoff. When these valves close, fuel is prevented from entering the vent plenum. As fuel is used, these valves are the first to open and permit air to flow into the tank from the vent plenum.

• Overboard vent and ram-air inlet

Vent Plenum

• Outboard vent valves

The vent plenum (Figure 28-7) is an isolated compartment located at the upper end of each wing tank. The plenum is formed by the front and rear beams, the upper and lower skin planks, and closure ribs at RBS 506.36 and WS 520.36. The vent plenum supplies the 2% fuel expansion space required by the FAA. The plenum permits air to flow in and out of the wing tank and collects fuel from the vent ducts. The overboard vent allows fuel to be discarded overboard if the plenum is full.

• Inboard vent valves • Vent drain valves

Vent Ducts There are two forward and two aft vent ducts. They are attached to the inner top part of each tank and extend parallel from the inboard vent valves to the outboard vent plenum (Figure 287). Two pipes are attached to the outboard end of the vent ducts and extend down to within 1/2 inch of the vent plenum floor.

Vent Float Non-Relieving Valves (Vent Duct) The vent float non-relieving valves (Figure 28-7) are installed on the bottom of the forward and aft vent ducts between RBS 479.5 and RBS 506.0. The valves open and close as necessary to allow air and fuel vapor to flow in and out of the fuel tanks. This prevents excess pressure or vacuum in the fuel tanks. These valves vent the tanks during fueling, ground operations, level flight, and descent.

Overboard Vent and Ram-Air Inlet The overboard vent and ram-air inlet is installed on the bottom of each wing, immediately outboard of RBS 453.0. The inlet is connected directly into the vent plenum by a tube assembly. The tube enters the vent plenum and turns upward to near the top of the plenum. The inlet allows air to enter and exit the tank through the plenum.


28-13



3-INCH VENT FLOAT VALVE

AFT VENT DUCT

FORWARD VENT DUCT

1-INCH VENT FLOAT VALVE

Figure 28-8. Inboard Vent Valves

Figure 28-9. Vent Duct Drain Valve

28-14




Vent Float and Pressure-Relief Valves

NOTES

The 3-inch (main) and 1-inch (auxiliary) vent float and pressure-relief valves are installed at the inboard end of the fuel ventilation system (Figure 28-8). A 3-inch vent float and pressure-relief valve is installed in the inboard, upper forward part of each tank. The valves open and close as necessary to allow air and fuel vapor to flow in and out of the fuel tanks. The 3-inch valve allows the tank to vent during maximum fuel supply to the engines. A 1-inch vent float and pressure-relief valve is installed in the inboard, upper forward part of each tank. The 1-inch valve helps venting during maximum fuel supply. The fuel tank venting is supplied through these valves during takeoff and climb. These valves also prevent overpressurization in the fuel tank by releasing pressure (fuel) overboard through the vent ducts. If high pressure occurs in the tanks, the force pushing on the valve element for opening is greater than the float force trying to keep the valve element closed. This allows the vent element to open, and the unwanted fuel (pressure) is permitted to flow through the valve and into the vent ducts.

Vent Drain Valves There are three vent duct drain valves (Figure 28-9). Two are installed on the bottom of the forward and aft vent ducts between RBS 161.5 and RBS 188.0. The third vent drain valve is installed at the lowest point in the vent system (Y-pipe at approximately BL 12.0). The primary function of these ball-type check valves is to allow fuel in the vent ducts and tubing to drain back into the tank. They also supply some tank venting when fuel level is below these valves.


28-15



FUEL RETURN

REMOTE FUELING

OFF/AUTO

L SHUTOFF

R SHUTOFF

OFF

CLSD

CLSD

WING TANK OPERATING RANGE 0°C TO 10°C

ENGINE

FCOC

LP PUMP

F I L T E R

HP FMU PUMP

HFRS CONTROL VALVE

PER HOP

FUE

L

HFRS B/U VALVE

HEATED FUEL

Figure 28-10. Engine Heated Fuel System Block Diagram

28-16




HEATED FUEL RETURN SYSTEM General The heated fuel return system (HFRS) provides fuel from the engines to heat fuel tanks during high altitude flying and cold weather conditions to decrease fuel viscosity caused by cold-soaked wing tanks. A heated fuel control valve is located in the fuel line between the spill diverter valve in the fuel management unit (FMU) an the fuel cooled oil cooler (FCOC) inlet. The control valve is a solenoidcontrolled, two-position, three-port valve which switches the flow path from the FCOC inlet to the return-to-tank plumbing. This valve is controlled by the full authority digital engine control (FADEC) based on temperatures sensed in each fuel hopper. A second, backup valve is installed in the fuel line to serve as a cockpit-controlled shutoff valve. This backup valve is controlled by a twoposition switch located on the cockpit overhead panel (COP), labeled FUEL RETURN. The switch positions are: AUTO (light out) and OFF (amber light). Placing the switch to AUTO opens the shutoff valve. Once the system is armed, the return flow will be automatically controlled by the FADEC through the heated fuel control valve.

Heated Fuel Return to Tank Valve The heated fuel return to tank valve is located at the 7-o’clock position on the engine bypass duct, below the FCOC. It is a solenoidpowered shutoff valve which is controlled through the EEC from several inputs. These inputs include the HFRS OFF/ARM switch on the cockpit overhead panel, fuel tank temperature sensor, fire handle, engine fuel low-pressure sensor, fuel low-quantity sensor, HP shutoff valve, fuel filter ∆P, fuel flow exceeding the threshold of 2,250 ±30 pph, and fuel crossflow valve.

Backup Heated Return to Tank Valve The backup heated fuel return to tank valve is located in the forward section of each engine pylon. It ensures that the return of fuel to the tank will not occur during operating conditions under which it would be unsafe to spill fuel back into the tanks. Such conditions include takeoff and landing maneuvers, during which the possibility of negative-g forces causing the ingestion of air bubbles in the fuel line must be guarded against. The BHFRV also provides additional security against failure of the HFRCV or its drive. The return of fuel to the tank can still be prevented by the BHFR gate independent valve.

To provide adequate fuel heating without overheating, the temperature switch opens and closes based on sensed fuel temperatures; opening at approximately 0°C and closing at approximately 10°C. The returned fuel temperature is approximately 50°C.

Components The HFRS consists of a heated fuel return-totank valve, backup heated fuel return to tank valve, FUEL RETURN switch on the cockpit overhead panel, and wing pipework.


28-17



FS 720

FS 742

FS 767

FS 782

ELECTRICAL CONNECTION

OUTPUT (DRAIN) ELBOW COUPLING

HEATED FUEL RETURN VALVE

FUEL LINE

INPUT (DRAIN) ELBOW COUPLING

Figure 28-11. Backup Heated Fuel Return Valve—Pylon Location

28-18




FUEL RETURN Switch The FUEL RETURN switch is located on the cockpit overhead panel. It controls the backup heated fuel return valve and annunciates an amber OFF legend when deselected. When the ARM position is selected, the switchlight is extinguished.

Wing Pipework The wing pipework (Figure 28-11) is made up of piccolo-type tubing mounted throughout the integral wing fuel tank. The purpose of the tubes is to carry heated fuel from the engine to each wing tank.

FUEL STORAGE SYSTEM OPERATION As fuel is added through the gravity-fueling port, the wing tanks fill in the following sequence. First the fuel flows inboard to the center of the wing. Since the wing dihedral is 3°, the fuel flows to the lowest point, which is located at the center of the wing.

Overpressurization of the fuel tanks will result in fuel being forced past the 3-inch and 1-inch vent float pressure-relief valves. The fuel flows through the vent ducts, into the vent plenum, and then overboard if necessary. Most of the fuel in the vent plenum is then pushed back into the vent ducts by the ram-air pressure. Excess fuel in the vent plenum can be drained through the water and fuel drain valve. Be sure to use a suitable container to collect the drained fuel. During flight, baffle ribs in the fuel tank restrict the movement of fuel as the aircraft executes turns and banks. This helps prevent weight shifts from side to side. As the fuel moves, the vent float valve opens and closes as necessary to prevent fuel from entering the vent plenum and allow the fuel tank to vent. If the fuel level in the tank is low, fuel in the vent ducts drains back into the tank through the vent duct drain valves.

NOTES

Next the fuel flows from the center of the wing, passing through flapper-type check valves and then into the fuel hopper. The fuel tank will fill from the center of the wing outboard. As the fuel rises, air and fuel vapor flow into the vent ducts through all of the vent duct valves and vent drain valves. The fuel venting continues from the vent ducts to the vent plenum and then overboard through the overboard vent and ram-air inlet. Venting continues until the fuel level reaches a point that causes all the valves to close. In order to prevent fuel from entering the vent plenum, the two vent float valves located outboard in the wing’s upper air passage are the last to close.

NOTE Gravity fueling must be done on level ground. If one wing is lower than the other, the fuel load in the wing tanks may not be balanced.


28-19



MAIN PRESSURE FUELING/DEFUELING LINE BL 0.0 RIB PRESSURE FUELING LINE DEFUELING CHECK VALVES

SUCTION DEFUELING LINES PRESSURE FUELING LINE PRESSURE FUELING SHUTOFF CONTROL PRESSURE FUELING/DEFUELING ADAPTER

Figure 28-12. Pressure Fueling/Defueling Components

28-20




FUEL DISTRIBUTION SYSTEM GENERAL The fuel distribution system provides for pressure refueling/defueling and supplies fuel to the engines and the APU. The system functions are grouped as follows: • Pressure fueling and defueling • Fueling shutoff • Fuel crossflow and intertank

Defueling Check Valves The defueling check valves are flapper-type valves located downstream of the pressure fueling and defueling adapter in the defueling lines going to each fuel hopper. These check valves prevent fuel from entering the hoppers during pressure refueling. They also open to allow fuel to be removed from the main tanks through hoppers during suction defueling.

Fueling and Defueling Lines The fueling and defueling lines consist of tubing connecting the main fill line from the pressure fueling and defueling adapter to a cross-fitting.

• Engine and APU distribution

Fueling and Defueling Operation

• Fuel filtration

PRESSURE FUELING AND DEFUELING General The purpose of pressure refueling and defueling is to supply a single point in the fuel system where the wing fuel tanks can be fueled and defueled.

Components The pressure fueling and defueling system consists of the following three components (Figure 28-12): • Pressure-fueling adapter • Defueling check valves • Fueling and defueling lines

Pressure Fueling and Defueling Adapter The pressure fueling and defueling adapter supplies a connection point for the fueling and defueling nozzle.

First, fuel enters the pressure-fueling adapter at a pressure between 35 and 55 psi. Fuel then flows through the main fill line to a cross-fitting, where the fuel flows through the right and left tank fill lines and then through pressurefueling shutoff valves, which control the flow of fuel into the tank. There are three methods for defueling. The first method is accomplished through the pressurefueling adapter using suction from a fuel tanker. Fuel is drawn through the suction defuel line installed in each hopper. This allows both tanks to be defueled at the same time but does not result in a completely empty tank. The remaining fuel (approximately 11 gallons) must be removed through the wing water and fuel drain valves in the lower skin. The second method uses the engine fuel line drain valve for suction defueling. A one-inch drain hose is connected from a defueling tanker to the right or left engine fuel line drain valve, located in the respective wheel well. The drain valve is opened, and the tanker removes fuel from the tank. The same procedure is repeated for the other tank. The small amount of fuel remaining is then drained through the water and fuel drain valves in the lower skin.


28-21



30000

Figure 28-13. Ground Service Control Panel

28-22




The third method uses the boost pumps for defueling. A one-inch hose is connected from a storage tank to the right or left engine fuel line drain valve. The drain valve is opened and the respective side boost pumps energized. This method results in the least amount of fuel remaining in the tanks. The small amount of fuel remaining is then drained through the water and fuel drain valves in the lower skin.

NOTES

NOTE Open the engine line drain valve before turning on the boost pumps, and turn off the boost pumps before closing the drain valve. If the proper procedure is not followed, the O-ring on the valve may be ruptured.

PRESSURE FUELING SHUTOFF General The components are designed to automatically shut off pressure-fueling flow when the wing tanks are full, the wing tanks are filled to a predetermined level, or a wing tank overpressure condition occurs. The pressure-fueling shutoff components also allow fueling shutoff from the cockpit overhead panel or the ground service control panel (GSCP), located on the forward side of the left EER (Figure 28-13).


28-23



INBOARD HIGH-LEVEL VALVE PILOT PRECHECK FLOAT PRECHECK VALVE PRECHECK TANK PRESS

28 VDC GROUND SERVICE BUS

REMOTE SHUTOFF SWITCH (COCKPIT)

FUEL QTY SIGNAL CONDITIONER SOLENOID SHUTOFF VALVE

OUTBOARD HIGH-LEVEL VALVE PILOT

CHECK VALVE

TANK PRESSURE

PRESSURE SENSING VALVE

FUEL OUT (TO TANK) AMBIENT PRESSURE

FUEL IN (FROM FUELING ADAPTER)

NOTE: SCHEMATIC IS TYPICAL FOR BOTH WINGS.

Figure 28-14. Fueling Shutoff Control—Block Diagram

28-24


PRESSURE FUELING SHUTOFF VALVE



COMPONENTS

CAUTION

The pressure-fueling shutoff components consist of the following (Figure 28-14):

Ve r i f y t h a t t h i s o p e n i n g i s n o t blocked.

• Pressure-fueling shutoff valve • Pressure-sensing valves

NOTES

• High-level pilot valves • Pressure-fueling solenoid shutoff valves • Manual precheck valves • Fuel quantity signal conditioner (via the GSCP)

Pressure-Fueling Shutoff Valves One pressure-fueling shutoff valve (SOV) is located in each wing tank, mounted to the baffle rib at BL 62.9. The pressure-fueling SOVs control the fuel entering the tanks until closed by inputs from the high-level pilot valves, sensing valve, manual precheck valves, or solenoid shutoff valve.

Pressure-Fueling Sensing Valve The pressure-fueling sensing valve is located at the structural rib at BL 90.0. The valve senses the difference between the outside ambient pressure and the pressure in the tank during pressure fueling. The valve functions to stop the pressure-fueling process if the pressure in the tank increases to the maximum differential value of the valve, such as a vent system blockage. Air pressure within the tank is sensed on the diaphragm of the pressure-sensing valve. If the tank pressure does not balance with the outside ambient pressure, the valve closes. This, in turn, causes back pressure to build up in the pressure-sensing line to the pressure-fueling SOV, causing it to shut off. The ambient vent for the pressure-fueling sensing valve is a small opening exiting out of the bottom surface of the wing skin, which must remain clear. This opening can be easily overlooked. Regular inspection of this opening will prevent anomalies during pressure refueling.


28-25



PRESSURE SENSE CHECK VALVE PRESSURE SENSING VALVE TO OUTBOARD HIGH-LEVEL PILOT VALVE

AMBIENT PRESSURE LINE PRESSURE FUELING

BL 0.0 RIB INBOARD HIGH-LEVEL PILOT VALVE

SHUTOFF VALVE

INBOARD HIGH-LEVEL PILOT VALVE SOLENOID SHUTOFF VALVE PRESSURE FUELING/ DEFUELING CONTROL

PRECHECK VALVES

PRESSURE FUELING SHUTOFFVALVE

SOLENOID SHUTOFF VALVE

3 11 LEGEND 12 1. ELECTRICAL CONNECTOR 2. FUEL INLET LINE 6 5 2. FUEL OUTLET LINE 4. BOLT (2) 5. WASHER (4) 6. NUT (2) 7. REMOTE FUELING SOLENOID 10 8. NUT 8 9. ELBOW 10. O-RING 11. UNION 12.. O-RING 9 RH SHUTOFF VALVE

5

4

TO OUTBOARD HIGH-LEVEL PILOT VALVE

AMBIENT PRESSURE LINE

7 1 9 LH SHUTOFF VALVE 2

Figure 28-15. Fueling Shutoff Control—System Components

28-26


PRESSURE SENSE CHECK VALVE



High-Level Pilot Valves

NOTES

The inboard high-level pilot valves are located in each wing (Figure 28-15). The outboard highlevel pilot valves are located at the outboard end of the wing by the gravity-fueling adapter. The inboard and outboard high-level pilot valves shut off pressure fueling when the wing tank is full and must be closed to stop fueling. This allows the airplane to be fueled at any ramp attitude, resulting in a balanced fuel load.

Remote Fueling Solenoid Shutoff Valves Two remote fueling solenoid shutoff valves are located on the front spar of the right wing. (Figure 28-15). These shutoff valves stop the fueling by causing the pressure-fueling shutoff valves to sense the tank as full. The solenoid SOVs are energized by the REMOTE FUELING L SHUTOFF or R SHUTOFF switches on the cockpit overhead panel or an automatic preselected fuel-level signal from the fuel quantity signal conditioner.

NOTE The remote fueling shutoff and signal conditioner control of the pressure-fueling shutoff functions only when there is power on the aircraft.


28-27



TANK PRESSURE LINE

FUEL FLOAT

FUEL T FLOA

PRECHECK

TANKS PRES

K CHEC HECK PREC

PRESSURE FUELING/DEFUELING ADAPTER

TANK PRESS

CHECK

PRESSURE FUELING/DEFUELING PRECHECK SELECTOR VALVES

TANK FLOAT LINE

Figure 28-16. Manual Precheck Valves

28-28




Manual Precheck Valves Two manual precheck valves are mounted on the pressure-fueling adapter in the fueling compartment (Figure 28-19). These three-position valves are used to make sure the pressure-sensing valves and high-level pilot valves close correctly before pressure fueling is completed. They use the fuel pressure from the pressure-fueling adapter for precheck shutoff operations.

NOTE The valves remain in the last position selected.

Manual Precheck Valve Operational Check PRESS Position The PRESS position on the manual precheck valve allows a small quantity of fuel to flow through the precheck valve to the pressuresensing valve, causing an artificial buildup of tank pressure, which closes the valve. This causes the pressure-fueling SOV to close in a manner similar to the normal closing of the high-level pilot valves. FLOAT Position The FLOAT position on the manual precheck valve allows a small quantity of fuel to flow through the precheck valve to the float cage of each high-level pilot valve. Fuel goes into each float cage faster than it can drain off, causing the floats to rise and stopping the pressure-sensing flow. This causes the pressure-fueling SOV to close in a manner similar to the normal closing of the high-level pilot valves and provides a time delay for operation.

OPERATION Normal Pressure Shutoff Fuel enters the pressure-fueling adapter and then flows through the main fill line to the pressure-fueling SOV in each wing tank. Air

in the tank is vented to the vent ducts and overboard as fuel flows inboard through the flapper valves in the baffle ribs, to the wing’s centerline, and then on to the outboard section of the wing tanks. As fuel fills the inboard end of the wing, the inboard high-level pilot valve float rises and closes the valve. Pressure fueling continues because the outboard high-level pilot valve is still open. When fuel fills the outboard end of the wing, the outboard high-level pilot valve closes, creating a back pressure on the pressurefueling shutoff valve. This causes the pressurefueling SOV to close.

Remote Fueling Shutoff (Power On) Pressure fueling may be stopped from the cockpit overhead panel by pushing the REMOTE FUELING L or R SHUTOFF pushbutton. This action energizes the fueling solenoid SOV, causing the pressure-fueling SOV to close in a manner similar to the normal closing of the high-level pilot valves. When using this procedure, keep the power on until pressure fueling is completed.

Emergency Vent If the pressure-fueling shutoff valves fail, the tank fuel pressure continues to increase until the vent float pressure-relief valves are forced open by the fuel pressure buildup. Fuel flows into the vent system, filling the vent plenum. When the vent plenum is full, fuel will then exit through the wing vent.

Overpressure Shutoff The pressure-sensing valve senses tank pressure and external (ambient) pressure. If the vent system becomes blocked, air pressure in the tank increases, causing tank pressure on the diaphragm of the pressure-sensing valve to increase. Sensing an imbalance with outside ambient pressure, the valve closes. Fuel-sensing pressure on the pressure-fueling shutoff valve increases and causes it to close in a manner similar to closing the high-level pilot valves.


28-29



REAR SPAR RBS 36 CROSSFLOW SHUTOFF VALVE

RH MANIFOLD

CROSSFLOW LINE

BL 0.0

RSB 36


INTERTANK LINE

FW

D

VIEW LOOKING AFT

Figure 28-17. Fuel Crossflow and Intertank Transfer System

28-30




NOTES

Automatic Refueling Mode When the predetermined value is reached during automatic pressure fueling, the fuel quantity signal conditioner provides a signal to close the fuel solenoid shutoff valves. This causes the pressure-fueling shutoff valve to close in a manner similar to the normal closing of the high-level pilot valves.

FUEL CROSSFLOW AND INTERTANK TRANSFER SYSTEM General The fuel crossflow components allow either fuel tank to supply fuel to both engines simultaneously. Fuel intertank transfer components allow fuel to flow between each tank (Figure 28-17).

Components The fuel crossflow and intertank transfer components consist of the following: • Intertank valve • Crossflow valve • Tubing


28-31



Figure 28-18. Intertank and Crossflow Switch

28-32




Intertank Valve The intertank valve is a motor-driven butterflytype valve located on the rear wall of the right hopper. Its function is to allow fuel from the left and right hoppers to combine. The intertank control pushbutton is a guarded switch located in the fuel section of the cockpit overhead panel. It displays a solid white bar when the switch is engaged and the valve is open. A blue “Fuel Intertank Open” message appears on the CAS when the intertank is open to allow fuel transfer between hoppers. The CAS fuel synoptic page also shows the valve position (green valve).

In the event of boost pump failure, depressing the crossflow pushbutton causes the crossflow valve to open, enabling fuel to feed from one boost pump manifold to the other. A white bar illuminates on the crossflow pushbutton when the valve has opened, and a signal is sent to the CAS via the MAU. 1 DGIO 1 to display a “Fuel Crossflow Open” message.

NOTES

Crossflow Valve The crossflow valve is a motor-driven butterfly-type valve located on the rear wall of the left hopper. This valve enables the fuel to feed from the pump manifold pressure chamber on one side to the opposite side. The crossflow valve motor is located in the wheel well, while the valve is in the hopper. The crossflow pushbutton control and indicator switch, located in the fuel section of the cockpit overhead panel, energizes the crossflow valve (Figure 28-18). The indicator portion of the switch displays a solid white bar when the valve is open, independent of switch position. A blue “Fuel Crossflow Open” message appears on the CAS when the crossflow valve is open, allowing fuel transfer between manifolds. An arrow depicting the direction of flow is displayed on the fuel synoptic page, along with the pump manifold in the crossflow condition (green valve).

Operation Control power for the fuel intertank valve and fuel crossflow valve is supplied by the left essential 28-VDC bus. Depressing the intertank pushbutton causes the intertank valve to open, allowing fuel from the left and right hoppers to combine. This action evens the fuel load on both sides. The white bar illuminates on the intertank pushbutton when the valve has opened, and a signal is sent to the CAS via the MAU. 2 DGIO 2 to display a “Fuel Intertank Open” message.


28-33



HOPPER ACCESS PANEL (TYPICAL 4 PLACES)

FUELING CHECK VALVE (TYPICAL 8 PLACES)

BL 35.8 BL 62.9 RIB BL 18.5 BL 6.0

UP

D FW

BL 0.0

EJECTOR PUMP (TYPICAL 2 PLACES

REAR SPAR

ENGINE FUEL SHUTOFF VALVE (TYPICAL 2 PLACES)

APU SHUTOFF VALVE

CROSSFLOW SHUTOFF VALVE


FUEL BOOST PUMP (TYPICAL 4 PLACES) BL 62.9 RIB HYDRAULIC HEAT EXCHANGER (TYPICAL 2 PLACES)

Figure 28-19. Fuel Boost Pumps

28-34




ENGINE AND APU FUEL DISTRIBUTION SYSTEM

NOTES

General The engine and APU fuel distribution system supplies fuel from the wing tank hoppers to the engines and APU.

Components The engine and APU fuel distribution system components consist of the following: • Fuel boost pumps • EMI filters • Boost pump manifolds • Ejector pumps • Engine fuel shutoff valves • APU fuel shutoff valve • Boost pump fuel pressure switch

Boost Pumps Two pumps, main and alternate, are mounted in pairs on the rear spar (rear wall of each hopper) in the main wheel well area on the left and right sides of the centerline rib (Figure 2819). They are identified as the left and right main fuel boost pump (inboard) and the left and right alternate fuel boost pump (outboard). These four pumps are plug-in type pumps with the suction and discharge ports penetrating the rear beam and into the fuel pump manifold. Each pump is equipped with a pressure switch, a vapor separator, and a shaft seal drain. A portion of the fuel flows through the pump for cooling and is returned to the hopper via the vapor return line. The boost pumps utilize an AC motor for operation. DC power is supplied to the pump assembly. An inverter inside the pump converts the DC power to AC power for pump motor operation.


28-35

28-36 <10.5 PSI

ANNUN LTS PWR

OFF

L ALT BOOST PUMP SW

L ALT FUEL BOOST PUMP CONTACTOR

OFF

SINGLE GENERIC I/O MODULE (2) SLOT 12 L ESS

L MAIN PUMP ON

OFF

L MAIN BOOST PUMP SW

L MAIN FUEL PUMP

ANNUNCIATOR LTS DIM / TEST CONTROLLER L MAIN FUEL BOOST OPN/GND OPN/28 VDC PUMP CONTACTOR OPN/GND

OPN/28 VDC

OPN/GND

OPN/28 VDC

OPN/GND

OPN/28 VDC

MAU 3

L BOOST PUMP EMI FILTER

SINGLE GENERIC I/O MODULE (6) SLOT 12 L ESS

<10.5 PSI

L MAIN FUEL BOOST PUMP

>16 PSI

L MAIN PUMP ON

L MAIN FUEL PRESS SW

MAU 1 SINGLE GENERIC I/O MODULE (1) SLOT 3 L ESS

R MAIN PUMP ON R ALT PUMP ON L ALT PUMP ON MAU 2

ANNUN LTS PWR

SINGLE GENERIC I/O MODULE (4) SLOT 12 R ESS

OFF OFF

R ESS 28 VDC

R MAIN FUEL PUMP

R MAIN PUMP ON R ALT PUMP ON L ALT PUMP ON

>16 PSI

R MAIN FUEL R MAIN FUEL PRESS SW BOOST PUMP

MAU 2

ON

R MAIN BOOST PUMP SW

R MAIN FUEL BOOST PUMP CONTACTOR

OFF R MAIN 28 VDC

R ALT FUEL PUMP

DUAL GENERIC I/O MODULE (2) SLOT 7 R MAIN/8 R ESS

R MAIN BP SEL ON R ALT BP SEL ON CROSSFLOW VLV OPN

NOT OPEN

R BOOST PUMP EMI FILTER

<10.5 PSI

OPEN ON

FUEL CROSSFLOW VALVE POSITION SW

Figure 28-20. Boost Pump Control Schematic

>16 PSI

R ALT FUEL PRESS SW

international

R ALT BOOST PUMP SW

L ALT FUEL BOOST PUMP CONTACTOR

R ALT FUEL BOOST PUMP

FlightSafety

OFF

<10.5 PSI

INVERTER


L ESS 28 VDC ON

R ALT PUMP CONT R MAIN 28 VDC

INVERTER

MAU 1

ANNUN LTS PWR

ANNUN LTS PWR

L ALT BP SEL ON L MAINBP SEL ON CROSSFLOW VLV OPN


L ALT FUEL PUMP

L ALT FUEL BOOST PUMP INVERTER

L MAIN 28 VDC

R MAIN PUMP CONT R ESS 28 VDC

L ALT FUEL PRESS SW

DUAL GENERIC I/O MODULE (1) SLOT 9 L ESS/10 R ESS

ON

L ALT PUMP CONT L MAIN 28 VDC

>16 PSI

MAU 1

OFF

INVERTER

L ALT PUMP CONT L MAIN 28 VDC



NOTES

The main boost pumps are powered by the essential DC bus, while the alternate boost pumps are powered by the main DC bus (Figure 28-20).

EMI Filters Two EMI filters are mounted in each wheel well, one for the left-side boost pumps and one for the right-side boost pumps (Figure 28-21).

NOTE A faulty EMI filter is indicated by an inoperative boost pump. Do not replace a suspect boost pump until its EMI filter has been confirmed as not faulty.

+28 VDC

NEGATIVE

A

B

+28 VDC CONTROL

A

+28 VDC POWER

C

NEGATIVE

B

INPUT +28 VDC

NEGATIVE

ALT PUMP OUTBD

D

B

+28 VDC CONTROL

A

+28 VDC POWER

C

NEGATIVE

MAIN PUMP INBD

C EMI FILTER

Figure 28-21. EMI Filter


28-37



PRESSURE SWITCH

LINE TO EJECTOR

CHECK VALVE ANTIBLEED VAPOR RETURN LINES

PUMP PRESSURE SWITCH PUMP DISCHARGE PORT AND FLAPPER CHECK VALVE

CROSSFEED LINE

FUEL INLET

SUCTION BYPASS VALVE

PUMP PRESS CHAMBER FUEL INLET SUCTION CHAMBER

Figure 28-22. Fuel Boost Pump Manifold

28-38


PUMP INLET PORT AND FLAPPER CHECK VALVE



Fuel Boost Pump Manifolds

NOTES

Two fuel boost pump manifolds are mounted to the rear spar inside each hopper, with each located so that it will receive one pair of boost pumps (Figure 28-22). Spring-loaded check valves, mounted to gland fittings within the manifold, allow the pumps to be removed and installed without defueling the tank. Each manifold has three chambers: one pressure chamber and two suction chambers. The pressure chamber, common to the discharge ports of both pumps, is fitted with a fuel feed line, crossflow line, motive-flow line, and pump suction bypass line. Each of the suction chambers has a suction line and a screen inlet extending forward inboard and downward to approximately 3/8 inch from the bottom of the hopper. Each suction chamber mates with the suction port of the fuel boost pump.


28-39



ENGINE FUEL SHUTOFF VALVE APU FUEL SHUTOFF VALVE CROSSFLOW SHUTOFF VALVE

FUEL BOOST PUMP

FUEL BOOST PUMP VAPOR RETURN CHECK VALVE (TYPICAL 2 PLACES) FUEL BOOST PUMP PRESSURE SWITCH) (TYPICAL 2 PLACES)

EJECTOR PUMP MOTIVE FLOW CHECK VALVE BL 0.0 FUEL BOOST PUMP MANIFOLD

PUMP INLET BYPASS LINE

PUMP INLET LINES

FUEL EJECTOR PUMP

INLET SCREENS

MOTIVE FLOW LINE

VIEW LOOKING AFT

Figure 28-23. Fuel Ejector Pumps

28-40


NOTE: LH SHOWN; RH OPPOSITE, EXCEPT FOR CROSSFLOW VALVE



NOTES

Fuel Ejector Pumps The aircraft is equipped with two fuel ejector pumps; both are low-pressure, high-volume jet pumps with flapper check valves. They are mounted on the forward wall of each hopper and maintain a full level of fuel in the hoppers at all times (Figure 28-23). The pumps’ internal flapper check valves operate when they receive the high-pressure fuel from the boost pump. The ejector pump is protected from blockage by a screen over the inlet.


28-41



Figure 28-24. Engine Fuel Shutoff Valve

Figure 28-25. APU Fuel Shutoff Valve

28-42




Engine Fuel Shutoff Valves

Pump Discharge Check Valves

The engine fuel shutoff valves (Figure 28-24) prevent fuel flow from the tank into the manifold pressure chamber when not pressurized. They are located between each respective pair of boost pumps and are mounted to the rear spar in each main gear wheel well. Normally open, the valves control the fuel feed line to the respective engine. When closed, the valves supply a ground to the MAU 1 DGIO 1 for the left side and MAU 2 DFIO 2 for the right side to generate the “Left or Right Fuel Shutoff Valve Closed” message on the CAS. The fuel shutoff valves are also controlled by the associated FIRE handle in the cockpit.

The pump discharge check valves are flappertype check valves installed in the manifold pressure chamber. Opened by fuel pressure from the boost pumps, they prevent fuel from entering a pump while it is not operating, such as in a crossflow condition or with only one pump operating on the same side.

The engine fuel shutoff valves are essentially gate valves operated by a DC motor that contains an internal EMI noise filter. The engine fuel shutoff valves have a valve position lever that can be manually operated as necessary for maintenance. The shutoff valves are designed so that the DC motor can be replaced without defueling.

Fuel Vapor Return Check Valves The fuel vapor return check valves are installed on the rear spar for each fuel boost pump. The check valves prevent fuel flow from the hopper when the vapor return line is disconnected from a pump.

Pump Suction Bypass Check Valve The pump suction bypass check valve is part of a suction bypass line that extends down to the bottom of the hopper. If both pumps stop operating, the engine will be able to siphon fuel through this line.

Fuel Pressure Switch

APU Fuel Shutoff Valve The APU fuel shutoff valve (Figure 28-25) is located inboard of the left boost pumps and is mounted to the rear spar in the left main gear wheel well. Normally energized open, the APU fuel shutoff valve controls the fuel feed line to the APU. The valve is solenoid-operated and is controlled by the APU CONTROL MASTER switch on the cockpit overhead panel. The valve is also closed when an APU fire is detected.

A fuel pressure switch is installed on each boost pump and actuates when fuel pressure increases to 16 ±2 psig. If fuel pressure drops to 9 ±1.5 psig, a signal is supplied (to the MAU 1 SGIO 2 and MAU 3 SGIO 6 for the left side and MAU 2 SGIO 4 and MAU 1 SGIO 1 for the right side) that the pump has stopped or fuel pressure is too low.

Pump Suction Check Valves The pump suction check valves are springloaded, flapper check valves installed in the pump manifold (see Figure 28-24). They permit the boost pumps to be removed and installed with fuel in the tanks. The valves are held open by the pump inlet ports as long as the pumps are installed. The valve closes when the pump is removed, preventing fuel from flowing from the tank.


28-43



TO BOOST PUMP

FUEL INLET ASSEMBLY

FILTER SCREEN (TYPICAL 2 PLACES EACH HOPPER

Figure 28-26. Engine and APU Fuel Distribution

28-44




Operation

NOTES

During normal operation, fuel flows by gravity into each hopper through three flappertype check valves installed in the hopper’s forward wall (Figure 28-26). Also installed in the hopper’s forward wall is a fuel ejector pump that operates on fuel flow (motive flow) supplied from the fuel boost pump. The ejector pump transfers fuel into the hopper at a rate of 4,550 pph, which is in excess of engine demand at maximum cruise power. Excess fuel is spilled back into the tank through a slot at the top of the forward hopper wall. Because the hopper remains full, the APU or engines are ensured an adequate supply of fuel as long as either pump on a side is operating or if the opposite pumps are operating with the crossflow valve open.

Fuel Feed The main and alternate boost pumps draw fuel from the hoppers into their respective manifolds. The fuel boost pump manifolds deliver fuel under pressure to the left and right main engines. The ejector pumps replenish the hopper, utilizing the motive flow supplied from the boost pump manifold. Motive flow from the left manifold supplies fuel for APU operation. The main and alternate fuel boost pump switches are located on the overhead panel and are labeled as L ALT, L MAIN, R MAIN, and R ALT. The fuel synoptic page provides indications of the on and off status of the fuel distribution system components.

Fuel Filtration System The fuel filtration system consists of overwing fueling adapter screens that keep foreign object debris (FOD) from entering the fuel tanks during gravity fueling. The system also consists of a fuel boost pump inlet filter screen that filters the fuel drawn by the boost pumps through their intake lines.


28-45



COMPENSATORS: • 3 TOTAL • MEASURES FUEL DIELECTRIC CONSTANT

GSCP AVIONICS RACK

SIGNAL CONDITIONER

RIGHT 28 VDC

DATA BUS LEFT FUEL QUANTITY PROCESSOR

OIL QUANTITY INDICATOR

ARINC 429

LEFT 28 VDC

ARINC 429

EICAS

RIGHT FUEL QUANTITY PROCESSOR

DATA BUS

TEST

OIL QYT PROBES; LEFT, RIGHT, APU

CONFIG

CONFIG

HIGH LEVEL WARN REFUEL VALVE OIL PUMP ENABLE

HIGH LEVEL WARN REFUEL VALVE

TANK UNITS

TANK UNITS C

H

L

D

LEVEL SENSORS

L

C

CD

LEVEL SENSORS

Figure 28-27. Fuel Indication System Block Diagram

28-46


H



FUEL INDICATION SYSTEM

• Compensators

GENERAL

• Low-level sensor

The fuel indication system provides the crew with a continuous indication of the amount of fuel remaining in the left and right wing tanks, the total fuel remaining in the system, a readout of the fuel temperature in each tank, a readout of fuel pressure, and faults that may occur in the fuel system (Figure 28-27).

• Ground service control panel

COMPONENTS The fuel quantity and indication system components, controls, and indicators are as follows: • Signal conditioner • Fuel quantity probes

• Densitometers • High-level sensor

• Fuel temperature bulb • Fuel pump pressure switch

Signal Conditioner Located in the left electronics equipment rack, the signal conditioner (Figure 28-28) provides independent left and right fuel quantity processing. It also provides independent and isolated low- and high-level processing for the left and right fuel tanks and digital data bus outputs to the CAS and ground service control panel. Two processors are contained in the signal conditioner unit.

SIGNAL CONDITIONER BEHIND CB PANEL

Figure 28-28. Signal Conditioner Location


28-47



Figure 28-29. Fuel Probes

28-48




Fuel Level Probes

Densitometers

Distributed throughout each wing are 19 capacitance-type fuel level probes, which are numbered from inboard to outboard (Figure 2829). They are used to accurately measure tank contents at different attitudes and fuel levels.

The fuel system has two densitometers. One is located in the left hopper, and the other one in the right wing tank. The right wing densitometer measures the density of the fuel being loaded during fueling. It accounts for the effect of temperature stratification, particularly during refueling operations. The left hopper densitometer measures the density of existing fuel in the aircraft (Figure 28-30).

Compensators Three compensators are included in the fuel indication system. One is located in each of the two hoppers, and the third one is located in the right wing tank. The compensators measure the fuel’s dielectric constant during refueling.

NOTE Refer to MSM, Chapter 28 for corresponding schematics.

Figure 28-30. Densitometer


28-49

28-50

international



Figure 28-31. Dummy Probe and High-Level Sensor



High-Level Sensors

NOTES

There are two high-level sensors per wing. One is located inboard, just forward of the hopper, and the other is in the vent plenum (Figure 28-31). The inboard high-level sensor is not used by the signal conditioner for determining a high-level condition. The outboard high-level sensor’s input passes through a “dummy probe” located in cell No. 1. A high-level condition exists when the outboard sensor is wet with fuel.

Low-Level Sensors There are two low-level sensors per aircraft located in the hopper of each tank. The lowlevel sensors provide the flight crew with warning of low fuel level when the fuel level falls to 650 pounds (96 gallons).


28-51



30000

Figure 28-32. System Monitor Test Panel

28-52




Ground Service Control Panel (GSCP)

NOTES

The ground service control panel is located on the forward side of the left EER. The GSCP display consists of three amber, eight-character, dot-matrix LED display modules. The upper display provides the left tank fuel quantity and left high-level warning. The middle display provides the right tank fuel quantity and right high-level warning. The lower display shows the preselected quantity value for automatic refueling set by the INCR–DECR switch, along with the fuel imbalance indication. The refuel/test switch is used to display fault codes and messages and initiate auto refueling.


28-53



PRESSURE SWITCHES

Figure 28-33. Fuel Boost Pump Pressure Switch

28-54




Fuel Temperature Bulb

NOTES

One fuel temperature bulb is located in each fuel tank hopper. The bulb contains a resistive element which varies with temperature. The measured temperature is displayed on the CAS fuel and summary synoptic pages.

Fuel Pump Pressure Switches Four fuel pump pressure switches are located on each fuel boost pump (Figure 28-33). These switches provide for monitoring of the fuel pressure to the engines by providing a 28VDC signal to the MAU SGIOs when the fuel pressure is greater than or equal to 16 ±2 psig. When pressure drops below 9 ±1.5 psig, the switch provides an open signal (fault state) to the MAUs. The fuel pump pressure switches also provide a maintenance advisory feedback indication when stuck in a pressurized state. This is accomplished by comparing the boost pump operating condition to the pressure switch condition and sets the appropriate bits on the ARINC 429 bus for monitoring by the CMC.


28-55



FUEL FLOW TANK FUEL QUANTITY

TOTAL FUEL QUANTITY INTERTANK VALVE

FUEL TANK TEMPERATURE

BOOST PUMPS

CROSSFLOW VALVE

HEATED FUEL RETURN

ENGINE SHUTOFF VALVE

ENGINE FUEL TEMPERATURE

Figure 28-34. Fuel Synoptic Page

28-56




OPERATION

NOTES

The fuel indication system provides indication on the following fuel system displays: the fuel synoptic page, summary page, ground service page, engine instruments displays, and MCDU. The fuel synoptic page provides indications for fuel flow, fuel quantity, fuel tank temperature, fuel boost pumps, fuel intertank valve position, fuel crossflow valve position, engine fuel shutoff valve position, engine fuel temperature, fuel imbalance (mismatch), and the heated fuel return system (HFRS) (Figure 28-34). The CAS summary page displays fuel quantity and fuel tank temperature. The ground service page provides fuel quantity information. The engine instruments display provides fuel quantity, fuel flow, and fuel tank temperature. The system test panel, located on the cockpit overhead panel, incorporates a fuel TEST switch which checks the fuel information on the engine instruments display for proper operation. When the fuel TEST switch is pressed, the fuel total should read amber “14000,” L and R should read amber “7000,” and the CAS message should read amber “Left or Right Fuel Level Low.” The MCDU displays fuel flow and quantity information.


28-57

28-58 FUEL QUANTITY SIGNAL CONDITIONER GND

DATA BUS A

TO LEFT FUELING SHUTOFF VALVE (CLOSES VALVE WHEN GND OUTPUT

DATA BUS B

TO RIGHT FUELING SHUTOFF VALVE (CLOSES VALVE WHEN GND OUTPUT

POWER IN PDM/FSK DATA BUS B

L FUEL QTY (GND) GRND SERV BUS 28 VDC


R EMERG DC BUS 28 VDC

PDM/FSK DATA BUS A

AIR GND SERVICE WOW RELAY

GROUND SERVICE PANEL (PILOT AFT BULKHEAD)

R FUEL QTY (AIR)

LEFT POWER IN RIGHT POWER IN

R FUEL QTY (GND)

DOOR OPEN

GRND SERV BUS 28 VDC FUEL HIGH LEVEL WARNING LIGHT FUEL HIGH LEVEL WARNING LIGHT CONTROL RELAY

LOW LEVEL SENSOR

REFUELING DOOR SWITCH

LEFT WING FUEL TANK FUEL HIGH LEVEL WARNING (GND = HIGH LEVEL) FUEL QUANTITY TRANSMITTER NO. 1

FUEL HIGH LEVEL WARNING TEST SW

FUEL QUANTITY TRANSMITTER NO. 2

MAU 1 DUAL GENERIC L/O MODULE (1) SLOT 9 L ESS/10 R ESS

GND WEIGHT ON WHEELS (GND – AIR) SYSTEM TEST

FUEL QTY NO. 1

MCDU NO.1 FUEL QTY NO.1 MAU 2 DUAL GENERIC L/O MODULE (2) SLOT 7 R MAIN/8 R ESS

REFUELING PANEL DOOR POSITION (GND = OPEN DOOR)

FUEL QTY NO.2

AIR GND SERVICE ANNUN WOW RELAY LTS PWR (B)

TEST

19 TANK QTY UNITS DENSITOMETER (HOPPER) COMPENSATOR HIGH LEVEL SENSOR HIGH LEVEL SENSOR

ANNUNCIATIR LTS DIM / TEST CONTROLLER OPN/GND

OPN/GND

FUEL QTY NO. 2

LOW LEVEL SENSOR

RIGHT WING FUEL TANK

international

FlightSafety

FUEL SYSTEM TEST SWITCH

19 TANK QTY UNITS DENSITOMETER (WING) COMPENSATOR COMPENSATOR HIGH LEVEL SENSOR HIGH LEVEL SENSOR


L EMERG DC BUS 28 VDC

L FUEL QTY (AIR)

Figure 28-35. Fuel Quantity Block Diagram



Refueling During refueling, two modes can be used: manual and automatic. The fuel quantity signal conditioner controls the automatic refueling mode (Figure 28-35). The ground service control panel (GSCP) inputs the amount of fuel desired. When the desired level of fuel is reached, the signal conditioner closes the remote fueling solenoid SOV, which receives power from the ground service bus. During refueling, the auto refuel/test switch on the ground service control panel initiates the automatic refueling mode (Figure 28-36).

The TEST/RESET position of this switch is used to test the indicator, scroll fault messages, and reset faults. The OFF position stops automatic refueling. The switch should remain in the OFF position when automatic refueling is complete. The INCR–DECR switch is spring-loaded and momentary in both the INCR and DECR positions. When released, it returns to the center-off position. When held up or down, this switch updates the preselect display in three increments: 100-pound, 200-pound, and 500pound increments.

NOTE

NOTE Upon power-up the default refueling mode is manual refueling. The refuel switch must transition to the OFF position and back to auto REFUEL to initiate the automatic refuel mode.

The maximum fuel quantity that can be selected is 50,000 pounds.

30000

Figure 28-36. Refuel/Test Switch


28-59



Figure 28-37. High-Level Warning Light and Switch

28-60




When the system is powered up, the PRESEL line is flashing, indicating the last selected value used for automatic refueling. The left and right lines indicate the present fuel quantity that is in each tank. To select a new PRESEL quantity, move the INCR–DECR switch in the direction needed. As soon as the INCR–DECR switch is moved, the PRESEL value quits flashing, displaying the new PRESEL fuel quantity.

NOTE If the value that is displayed is the value wanted, then quickly toggle the INCR–DECR switch in either direction and back to the center position to get the display to stop flashing. If the PRESEL value changed during the toggling of the switch, move the switch in the direction of the desired value until the value wanted is indicated in the PRESEL line. The PRESEL value is equally split between the left and right systems when refueling begins.

To initiate the automatic mode of refueling, move the refuel/test switch to the auto REFUEL position.

NOTE An asterisk (*) appears on the PRESEL line, indicating that the system is in the automatic refueling mode. The fueling pressure can be started at this time.

During refueling, the left and right fuel quantities are continuously updated. When the desired quantity is reached, the fuel quantity signal conditioner stops the refueling by energizing the fueling solenoid shutoff valve to the closed position. Stop the fueling pressure before moving the refuel/test switch to the OFF position.

NOTE If the fueling pressure is on when the switch is moved to the OFF position, pressure refueling starts again. When the refuel/test switch is moved to OFF or is already in the OFF position, the fueling valves are open.

If a 500-pound or greater difference exists between the left and right fuel quantities, the signal conditioner automatically closes the fueling valve of the high side, letting the low side continue refueling until it equals the high side. The high-side refueling valve then opens, and both sides continue refueling. This allows for refueling when the wings are not level .If the imbalance reaches 1,000 pounds, the PRESEL line flashes WARN and then IMBAL. When a high-level condition is encountered, the appropriate fuel quantity display flashes HI, and the PRESEL line flashes WARN, the high-level warning light on the pressure fueling access door will illuminate, but fueling will not terminate. Automatic refueling cannot be initiated if a high-level condition exists. The high-level warning system includes a high-level test switch and a red high-level indication light mounted on the pressure-fueling port access door (Figure 28-37). If at any time the ground service control panel does not receive data from the signal conditioner for more than 15 seconds, all the displays show dashes. Holding the refuel/test switch in the TEST/RESET position for two seconds causes the display to indicate its test pattern. If the refuel/test switch is held in the TEST/RESET position for more than five seconds, the display indicates a fault code summary, telling the number of stored or active faults. If there are no stored or active faults, the message reads NO FAULTS. Faults are sequenced by toggling the refuel/test


28-61



switch from the center position to TEST. When faults are displayed, the upper display contains the line replaceable unit (LRU) nomenclature, the middle display contains the error nomenclature, and the lower display shows the failure status (active or not) and the number of occurrences. After the last fault code is displayed, the message CLEAR appears for five seconds. If the refuel/test switch is held in the TEST/RESET position for five seconds, all the fault codes are cleared. During this time, CLEARING is displayed to indicate that the codes are being cleared. After the codes have cleared, CLEARED is displayed for two seconds.

NOTE While displaying fault codes, a twominute time-out causes the GSCP to exit the fault code display mode. If the INCR–DECR switch is used, the fault code display terminates, and the display returns to normal mode. Ten faults per side is the maximum number of faults that can be stored. More than ten faults result in the oldest fault being removed from memory to make room for storing a new fault. Once cleared, all messages are lost.

28-62


NOTES



CHAPTER 29 HYDRAULIC POWER CONTENTS Page INTRODUCTION ................................................................................................................. 29-1 GENERAL ............................................................................................................................ 29-1 PLUMBING FEATURES ..................................................................................................... 29-2 Hydraulic Line Identification ........................................................................................ 29-2 Approved Hydraulic Fluids ........................................................................................... 29-2 LEFT AND RIGHT HYDRAULIC SYSTEM COMPONENTS ......................................... 29-3 General........................................................................................................................... 29-3 Reservoirs ........................................................................................................................29-5 Reservoir Components................................................................................................... 29-5 Hydraulic Shutoff Valve ................................................................................................ 29-9 Engine-Driven Pumps.................................................................................................... 29-9 Engine-Driven Pump Isolation Check Valves ............................................................... 29-9 Acoustical Filters (Hydraulic Mufflers) ........................................................................ 29-9 Accumulators.................................................................................................................29-11 System Pressure Transducer........................................................................................ 29-11 Pump Pressure Switches.............................................................................................. 29-11 Filter Manifold Assemblies ......................................................................................... 29-13 Left System Pressure Relief Valve .............................................................................. 29-15 Heat Exchangers ............................................................................................................29-17 Hydraulic Ground Service Panel ...................................................................................29-17 Hydraulic Replenisher................................................................................................. 29-17


29-i



HYDRAULIC SYSTEM OPERATION ............................................................................. 29-19 Electrically Powered Reservoir Supply Shutoff Valve................................................ 29-19 Operation of the Engine Hydraulic System................................................................. 29-19 Engine-Driven Pumps ................................................................................................. 29-21 Engine-Driven Pump Isolation Check Valves ............................................................. 29-21 Acoustical Filter .......................................................................................................... 29-21 Distribution.................................................................................................................. 29-21 CONTROLS AND INDICATORS ..................................................................................... 29-21 Synoptic Pages ............................................................................................................ 29-21 Left and Right System Pressure Transducers .............................................................. 29-25 Pump Pressure Switches.............................................................................................. 29-25 Left and Right Temperature Transducers .................................................................... 29-25 Quantity Transducers................................................................................................... 29-27 Hydraulic Quantity Compensation.............................................................................. 29-27 Left and Right System Quantity Indicator .................................................................. 29-27 POWER TRANSFER UNIT (PTU) SYSTEM................................................................... 29-29 General ........................................................................................................................ 29-29 PTU System Components ........................................................................................... 29-29 PTU Hydraulic System Operation............................................................................... 29-31 AUXILIARY HYDRAULIC SYSTEM.............................................................................. 29-33 General ........................................................................................................................ 29-33 Auxiliary Hydraulic System Components................................................................... 29-33 Auxiliary Hydraulic System Operation....................................................................... 29-34 FAULT INDICATIONS ...................................................................................................... 29-37 Left and Right System Fault Indications ..................................................................... 29-37 Auxiliary System Fault Indications ............................................................................. 29-39

29-ii




HYDRAULIC SUBSYSTEM ............................................................................................. 29-41 Main Entry Door.......................................................................................................... 29-41 Locking Handles.......................................................................................................... 29-41 Main Entry Door Hydraulic Components ................................................................... 29-43 Main Entry Door Control Components ....................................................................... 29-45 Warning System Components...................................................................................... 29-47 Main Entry Door Control and Warning System Operation ......................................... 29-49 Main Entry Door Operation......................................................................................... 29-49


29-iii




Title

Page

29-1

Simplified Hydraulic System ................................................................................ 29-4

29-2

Hydraulic System Reservoirs................................................................................. 29-6

29-3

Engine-Mounted Hydraulic Components .............................................................. 29-8

29-4

Hydraulic Accumulator Components—Left System Shown............................... 29-10

29-5

Left System Filter Manifold ................................................................................ 29-12

29-6

Right System Filter Manifold .............................................................................. 29-14

29-7

Wing Fuel Hopper ............................................................................................... 29-16

29-8

Heat Exchanger.................................................................................................... 29-16

29-9

Hydraulic Shutoff Control ................................................................................... 29-18

29-10

Hydraulics Synoptic Page.................................................................................... 29-20

29-11

Engine Start Synoptic Page ................................................................................. 29-20

29-12

Summary Synoptic Page...................................................................................... 29-22

29-13

Ground Service Synoptic Page ............................................................................ 29-22

29-14

Hydraulic Pressure Indication.............................................................................. 29-24

29-15

Hydraulic Quantity/Temperature Indication........................................................ 29-26

29-16

Power Transfer Unit (PTU) ................................................................................. 29-28

29-17

Hydraulic Control Panel ...................................................................................... 29-30

29-18

Hydraulics Synoptic Page—PTU Operation ....................................................... 29-30

29-19

Auxiliary Hydraulic System ................................................................................ 29-32

29-20

Hydraulics Synoptic Page—Left and Right System Fault Messages .................. 29-36

29-21

Hydraulics Synoptic Page—PTU Fault Message ................................................ 29-36

29-22

Hydraulics Synoptic Page—Auxiliary System “Fail” Message .......................... 29-38

29-23

Main Door Components ...................................................................................... 29-40


29-v



29-24

Main Door Locking Handles .............................................................................. 29-41

29-25

Main Door Hydraulic Components ..................................................................... 29-42

29-26

Main Door Control Panel..................................................................................... 29-44

29-27

External Door Switch Location ........................................................................... 29-46

29-28

Secondary Handle................................................................................................ 29-46

29-29

Doors Synoptic Page............................................................................................ 29-48

29-30

Door Hydraulic Schematic .................................................................................. 29-48

TABLES Table 29-1

29-vi

Title

Page

Approved Type IV Fluids....................................................................................... 29-2




CHAPTER 29 HYDRAULIC POWER

INTRODUCTION The hydraulic system of the Gulfstream G500/G550 aircraft comprises four subsystems: the left and right engine-driven systems, the power transfer unit, and the auxiliary hydraulic system. The hydraulic system provides power to operate the landing gear, close the cabin door, operate flight controls, and deploy the thrust reversers, as well as providing pressure to operate the hydraulic motor generator (HMG). If a fault occurs in the left or right engine-driven system, PTU, or auxiliary system, the fault indications are displayed on the engine indicating and crew alerting system (EICAS).

GENERAL The Gulfstream G500/G550 aircraft hydraulic system provides Type IV phosphate esterbased fluid under pressure to a common point for further distribution to its subsystems. The subsystems that make up the hydraulic system

include the left and right engine-driven systems, the power transfer unit (PTU), and the auxiliary hydraulic system. The left and right enginedriven systems are the primary means of providing hydraulic power.


29-1



Each aircraft engine has a hydraulic pump mounted on the front of its accessory gearbox. The hydraulic systems operate the primary flight controls and thrust reversers. In addition, system pressure is used for operation of the landing gear, brakes, nosewheel steering, flaps and hydraulic motor generator (HMG). The auxiliary system is used to close the main entrance door and charge the parking brake accumulator.

PLUMBING FEATURES The Gulfstream G500/G550 hydraulic plumbing system consists of steel, titanium, and aluminum lines. The steel and titanium lines are used as pressure lines, whereas, depending on the size and location of the line, aluminum and steel are used as return lines. Steel is also used where there is a risk of foreign object debris (FOD) damage, such as the landing gear wheel well areas and the aft wing beam. Fittings are either Cryofit® or Perma-swage®. Cryofit® are standard AN fittings with a cryogenic sleeve, which is made of a shape-memory alloy called “Tinel.” The fittings are manufactured 3% smaller than the desired line-connection size. They are frozen in liquid nitrogen, expanded larger than the line size by 5%, and then shipped frozen for installation. Cryofit® fittings are installed on all hydraulic lines within the wings. Cryolive® B-nut fittings are installed on the rest of the airplane in place of Perma-swage® B-nut fittings. All other fittings (elbows, splices, T-fittings, etc.) are Perma-swage® fittings.

HYDRAULIC LINE IDENTIFICATION To facilitate maintenance of the hydraulic system, each hydraulic line is identified with a decal-type band that indicates the drawing number, line number, line code, and tubing size. The bands are placed near both ends of each line to provide ready identification. Some hydraulic lines installed on the wings do not have a decal but have been laser-etched with the same identification.

APPROVED HYDRAULIC FLUIDS The hydraulic systems are designated for use with Type IV phosphate ester (or equivalent) hydraulic fluid and for operational temperatures from –40°F to 491°F (–40°C to 255°C). Type IV fluids have all the beneficial characteristics of Type II fluids, plus they provide erosion resistance to hydraulic components and improved thermal stability (Table 29-1). HyJet IV and Skydrol LD-4 are lower density, resulting in significant weight savings. All approved phosphate ester fluids are compatible and can be safely mixed. However, up to 15% HyJet IV in Skydrol 500 may cause an increase in valve erosion. There is no restriction on HyJet IV-A mix. All other concentrated mixtures of approved phosphate ester fluids have complete compatibility.

Table 29-1. APPROVED TYPE IV FLUIDS IDENTIFICATION

29-2

COLOR

VENDOR

HyJet IV

Purple

Exxon Chemical Company

HyJet IV-A+

Purple

Exxon Chemical Company

Skydrol LD-4

Purple

Monsanto Chemical Company

Skydrol 500B-4

Purple

Monsanto Chemical Company




NOTE Mineral oils MIL-H-5606 and MILH-83282 are used only in landing gear struts of the G500/G550 and should never be mixed with phosphate ester hydraulic fluids. All recommended safety and handling precautions should be adhered to when servicing aircraft or handling hydraulic fluids.

WARNING Skydrol Type IV phosphate ester hydraulic fluid is combustible and can be a health hazard. Inhalation of vapor and contact with skin and eyes should be avoided. The fluid should not be exposed to extreme heat or open flames. All material safety data sheet (MSDS) recommendations for health and safety precautions should be followed. To prevent injury to personnel and damage to equipment, protective caps should be installed on all open electrical disconnects, open hoses and ports.

LEFT AND RIGHT HYDRAULIC SYSTEM COMPONENTS GENERAL The left and right main hydraulic systems are the primary sources of hydraulic power for the Gulfstream G500/G550 . They store, pressurize, and deliver hydraulic fluid to the using system components. Both left and right hydraulic systems supply pressure to operate the tandem actuators of the primary flight controls and stall barrier. The primary flight controls consist of the ailerons, ground and flight spoilers, rudder, yaw damper, and left and right elevators. The left system also supplies pressure to operate the hydraulic motor generator (HMG), flaps, landing gear, nosewheel steering, and toe brakes. The left thrust reverser is pressurized solely by the left system. The right system, however, supplies pressure to the PTU motor, as well as the right thrust reverser (Figure 29-1).

CAUTION Skydrol Type IV phosphate ester hydraulic fluid can damage paints, rubber and plastic materials. Care must be taken to prevent spillage from remaining on surfaces or damage may result.


29-3



RIGHT REVERSER

LEFT REVERSER ELEVATORS STALL BARRIER AILERONS FLIGHT SPOILERS/ SPEED BRAKES RUDDER

YD 1

YD 2

GROUND SPOILERS L SYS OR PTU OR AUX PRESS REQ'D FOR R SYS USE

GND SPLR SERVO

MAIN ENTRANCE DOOR

NOSEWHEEL STEERING

N2

LANDING GEAR

ACCUM

NORMAL BRAKES

PARKING BRAKES

WING FLAP

STANDBY ELECTRICAL POWER MASTER

ON PWR XFR UNIT

HMG MOTOR

OFF/ARM

ON

NOT ARM

ON

ACCUM STBY RUD PTU PUMP

ON

PTU MOTOR

ACCUM

AUX PUMP

L

AUX BOOST PUMP

LEFT ENG PUMP

DISCH 2

R

AUX PUMP ON

NOT ARM

ON

RIGHT ENG PUMP

FIRE HANDLE

FIRE HANDLE

OFF/ARM

1

L SYS

R SYS

AUX L SYS R SYS AUX L SYS/PTU

L SYS/PTU/AUX NITROGEN MECH CONN ELECT CONT

CHECK VALVE

(

FLOW

)

GND SERVICE VALVE

Figure 29-1 Simplified Hydraulic System

29-4


SHUTOFF VALVE

BOOT STRAP

2

BOOT STRAP

DISCH 1



The components of the engine-driven hydraulic system include the following: • Reservoirs • Hydraulic shutoff control valve • Engine-driven pumps • Check valves • Acoustical filter (muffler)

The left system reservoir contains a separate subchamber that supplies the auxiliary system. The left system chamber holds 5.7 gallons, and the auxiliary system subchamber holds 2.0 gallons, for a total reservoir capacity of 7.7 gallons. The servicing level for the left reservoir is 4.8 gallons. The total system capacity is 20.6 gallons. The right system reservoir has a single system chamber with a capacity of 1.8 gallons and a total system capacity of 7.0 gallons. The right system reservoir is serviced to 1.5 gallons.

• Accumulators

RESERVOIR COMPONENTS

• Nitrogen servicing panel

Manual Bleed and Relief Valve

• Pressure transducers • Pump pressure switches • Filter manifolds • Left system pressure relief valve • Heat exchangers • Ground service panel

Located on the top of each reservoir is the bleed and pressure relief port (Figure 29-2). Manually bleeding the reservoir allows air in the system to be bled when the system is pressurized. This action transfers fluid to the reservoir overboard drain. On the bottom forward end of the reservoir is the scupper, a drain port from which reservoir internal leakage drains into an overboard drain.

• Hydraulic replenisher

WARNING Hydraulic fluid can exit the reservoir overboard drain at high pressures. Ensure personnel are clear of the overboard drain before depressing bleed valve.

RESERVOIRS General NOTE Refer to the Maintenance Schematic Manual for corresponding schematics.

There are two hydraulic reservoirs located in the tail compartment. Their purpose is to store hydraulic fluid under pressure for use by the pumps. Each reservoir consists of a low-pressure system chamber and a high-pressure bootstrap chamber (Figure 29-2). An individual cylinder forms each chamber. A rod connects pistons in each chamber so that hydraulic pressure applied to the high-pressure bootstrap piston causes a reaction by the larger lowpressure piston to maintain a constant pressure inside the system chamber.

Both hydraulic reservoirs contain a pressure relief valve, which is built into the reservoir housing. The pressure relief valve prevents damage to the reservoir from over-servicing or overpressurizing. The valve vents into the overboard drain through the relief and bleed port if internal reservoir pressure exceeds 110 psi.


29-5



MANUAL BLEED RELIEF VALVE

SYSTEM CHAMBER

BOOT STRAP CHAMBER

LVDT

UP

FWD

QUANTITY INDICATOR AUX SUB-CHAMBER

LEFT SYSTEM RESERVOIR

MANUAL BLEED RELIEF VALVE BOOT STRAP CHAMBER

SYSTEM CHAMBER LVDT FWD

UP

QUANTITY INDICATOR

RIGHT SYSTEM RESERVOIR

Figure 29-2 Hydraulic System Reservoirs

29-6




Linear Variable Differential Transducer (LVDT)

NOTES

An electronic quantity transducer is located on the aft end of the bootstrap cylinder. In the center of the piston rod is a linear variable differential transducer (LVDT) (Figure 29-2). Electronic quantity readouts are displayed on the HYDRAULICS synoptic page and also on the digital quantity indicator on the replenisher panel. In order to provide an accurate reading, the reservoir must be pressurized.

Quantity Indicator Reservoir quantity is displayed on the directreading quantity indicator (Figure 29-2). The indicator has a cable measuring assembly connected to the forward side of the low-pressure piston. Piston position is translated into a quantity readout. For accurate measurement readouts, the reservoir must be pressurized.

Reservoir Temperature Transducers Reservoir temperature transducers, mounted on the reservoirs, provide discrete temperature data to the modular avionics units (MAUs). The left reservoir temperature transducer provides temperature data to MAU No 1 (B), dual generic I/O No. 1 (DGIO-1) in slots 9/10. The right temperature transducer provides data to MAU No. 2 (A), DGIO-2 in slots 7/8. The MAUs will also transmit the data to the flight data recorder (FDR) and to the synoptic pages (See Figure 29-15). When the reservoir fluid temperature reaches 104°C (220°F), an amber “L-R Hydraulic Reservoir Hot” message is displayed on the CAS to alert the crew of an overheat condition in the reservoir. The message goes off when the reservoir fluid temperature decreases to 80°C (175°F). If the temperature transducer fails, a blue “L-R Hydraulic Temp Sen Fail” message is displayed on the CAS.


29-7



HYDRAULIC ACOUSTICAL FILTER

PRESSURE LINE

EDP SUCTION LINE

CHECK VALVE

CASE DRAIN LINE

ENGINE-DRIVEN PUMP

VIEW FROM FRONT RIGHT ENGINE CROSS SECTION

Figure 29-3. Engine-Mounted Hydraulic Components

29-8




HYDRAULIC SHUTOFF VALVE Each system shutoff valve is located below and outboard of its respective reservoir. Two 28VDC motor-driven shutoff valves control fluid flow to the suction ports of the engine-driven pumps. The shutoff valves are powered from the left and right 28-VDC essential busses respectively. The valves incorporate a manual operating lever.

ACOUSTICAL FILTERS (HYDRAULIC MUFFLERS) An acoustical filter is located on the top of each engine, in the pump pressure line (Figure 29-3). The filter dampens the harmonic imbalances of pump output pressure to help reduce cabin noise levels.

Emergency shutoff of the left and/or right hydraulic system is controlled by the engine fire handles in the cockpit. Hydraulic fluid is shut off downstream of the hydraulic reservoirs in the tail compartment. This prevents any hydraulic fluid from going out to the engine area, where the main engine-driven hydraulic pumps are located.

NOTES

ENGINE-DRIVEN PUMPS The engine-driven hydraulic pumps are located on the forward side of each engine’s accessory gearbox (Figure 29-3). Each pump is rotary-piston, pressure compensated, variable volume and displacement, and rotates any time the engine is turning. Four ports are incorporated into each hydraulic pump: a suction (inlet) port, a pressure (outlet) port, a case drain (bypass) port, and a shaft seal drain port. Each pump will produce 18 gpm flow with the engine at idle rpm and up to 28 gpm at rated thrust, at a pressure of 3000 + 200-100 psi.

ENGINE-DRIVEN PUMP ISOLATION CHECK VALVES Isolation check valves are located on the outboard side of each engine on the intermediate case (Figure 29-3). Each pump has its own check valve in its pressure line. The check valves prevent “backflow” to the pump when the hydraulic system is pressurized using the ground servicing cart.


29-9



LEFT SYSTEM ACCUMULATOR

LEFT SYSTEM FILTER MANIFOLD

SERVICING VALVES

Figure 29-4. Hydraulic Accumulator Components—Left System Shown

29-10




ACCUMULATORS There are two 50 cubic-inch accumulators, one for each hydraulic system, located on either side of the tail compartment (Figure 294). Their purpose is to dampen hydraulic pressure surges within the system. The accumulators are serviced with 1200 psi of nitrogen (at 70°F) via the nitrogen servicing panel. This nitrogen precharge ensures proper operation of the accumulator. If the accumulator is not properly serviced, hydraulic line damage may occur.

SYSTEM PRESSURE TRANSDUCERS The pressure transducers transmit system pressure and are located outboard of their respective filter manifolds (Figure 29-5). They are powered by the ESS DC buses through the HYD PRESS circuit breakers. With pressure applied, the transducers transmit pressure data to the MAUs. The left pressure transducer provides data to MAU No. 1 (B), DGIO-1 in slots 9/10. The right pressure transducer provides data to MAU No. 2 (A), DGIO-2 in slots 7/8. A snubber is installed in line with the transducer to smooth out oscillations.

NOTE For proper servicing procedures, cons u l t t h e A i rc r a f t M a i n t e n a n c e Manual.

Nitrogen Servicing Panel Located on the left side of the tail compartment is the nitrogen servicing panel that is used to service the two hydraulic accumulators. The panel consists of two pressure gages that incorporate filler valves (Figure 29-4).

PUMP PRESSURE SWITCHES Each engine-driven pump has its own pressure switch located in the tail compartment on the outboard side of that system’s (left or right) filter manifold. The switch provides an indication of pump operation on the HYDRAULICS synoptic page (Figure 29-5).

Care must be taken to ensure the hydraulic systems are not pressurized whenever checking or servicing the accumulator nitrogen precharge pressure.

NOTE For proper servicing procedures, cons u l t t h e A i rc r a f t M a i n t e n a n c e Manual.


29-11



L SYS CASE DRAIN (HIDDEN)

PTU CASE DRAIN

L PUMP PRESSURE SWITCH

FLUID SAMPLE PORTS (3)

D.P.I. (5)

L SYS PRESSURE SWITCH

L SYS RETURN

AUX RETURN

Figure 29-5. Left System Filter Manifold

29-12


L SYS PRESSURE TRANSDUCER

L SYS PRESSURE



FILTER MANIFOLD ASSEMBLIES

drain, and PTU case drain filters. The sampling and bleed valve is a manually operated valve that permits fluid samples and/or air to be removed from the circuit.

General The purpose of the main hydraulic system filter manifold subsystem is to provide the using system components with filtered fluid and control the direction of fluid travel. The left and right hydraulic system filters are incorporated into filter manifold assemblies and are located in the tail compartment. All the filters, which are disposable, provide 3-micron filtration and incorporate differential pressure indicators. If the filter becomes clogged, the red button on the differential pressure indicator (DPI) extends approximately 3/16 of an inch. The system return and case drain filters are bypass-type filters. The filter bowls provide for filter removal without loss of fluid and incorporate a ratchet lever-lock device.

The left system pressure switch, also located on the left filter manifold (Figure 29-5), provides for automatic PTU operation. The pressure switch senses left system pressure down stream of the system pressure filter. When the PTU is ARMED and left system pressure drops below 1500 psi, the pressure switch opens the PTU shutoff valve, initiating PTU operation.

NOTES

Both filter manifolds incorporate low-pressure relief valves. These valves protect the pump cases by permitting fluid to be released to the return circuit as thermal expansion occurs in the pump case drain line. The relief valves open at 150 psi, and close at 115 psi.

Left System Filter Manifold The left system manifold (Figure 29-5) is made up of four sections: the main pressure filter, the main return filter, the auxiliary return filter, and the case drain filter sections. Four bypass-type filters and one nonbypass-type filter are in the manifold. The nonbypass filter is used in the left system pressure, whereas the bypass filters are used in the left system return line, left pump case drain, PTU pump case drain, and auxiliary return circuits. Check valves are incorporated into the system to control the direction of fluid flow. Sampling and bleed valves (fluid sample ports) are installed on the hydraulic system filter manifold. The sampling ports are located upstream of the left system return filter, left system pump case


29-13



R SYS PRESSURE TRANSDUCER

R SYS PRESSURE

D.P.I. (3)

R SYS HIGH PRESSURE RELIEF VALVE

R SYS/PTU MOTOR CASE DRAIN (HIDDEN)

R SYS RETURN

Figure 29-6. Right System Filter Manifold

29-14


R PUMP PRESSURE SWITCH

FLUID SAMPLE PORTS (2)



Right System Filter Manifold

NOTES

The right system manifold (Figure 29-6) is made up of three sections: the main pressure filter, the main return filter, and the enginedriven pump/PTU motor case drain and bypass filter section. There are three filters in the manifold: one nonbypass-type, used for right system pressure, and two bypass-type, used in the right system return line and the right engine pump/PTU motor case drain. Two sampling and bleed valves (fluid sample ports) are located in the right system return filter and in the right system pump case drain/PTU case drain. The sampling and bleed valves are manually operated valves that permit fluid samples and/or air to be removed from the circuit. Check valves are incorporated into the hydraulic system to control the direction of fluid flow. The right system has a high-pressure relief valve installed internally. The relief valve protects the system from overpressure by permitting high pressure to be released to the return circuit. The valve opens at 3,850 psi, allowing up to 27 gpm fluid flow to return, and closes at 3,200 psi.

LEFT SYSTEM PRESSURE RELIEF VALVE The left system pressure relief valve protects the system from overpressure. It is located in the right main landing gear wheel well and is mounted on the aft inboard bulkhead. The valve opens at 3,850 psi, routing fluid back to the reservoir through the system return filter, and closes at 3,200 psi. Maximum flow rate at 3,850 psi is 27 gpm. Even when the relief valve is open to relieve excessive pressure, it does not prevent fluid pressure from operating the system components.


29-15



Figure 29-7. Wing Fuel Hopper

Figure 29-8. Heat Exchanger

29-16




HEAT EXCHANGERS The hydraulic pumps rotate and generate pressure whenever the engines are running. If there are no demands on the system, system pressure stabilizes at 3,000 +200/–100 psi. A quick overheat would occur if there were no transfer of fluid out of the pump when it turns with no demand on the system. To prevent an overheat and to lubricate the pump, some hydraulic fluid from the suction port is allowed to bypass into the pump housing. This bypassed fluid is then routed to the heat exchanger through the pump’s case drain port. Each system contains its own heat exchanger that cools the hydraulic case drain fluid while heating the fuel. The heat exchanger is a radiator-type cooler, immersed in fuel, and is located in each wing fuel hopper (Figures 29-7 and 29-8). The left hydraulic system heat exchanger is located in the right fuel hopper and the right system heat exchanger is in the left fuel hopper. The heat exchanger cools the bypass fluids from the engine-driven pumps and the PTU motor and pump. The cooled fluid is returned from the heat exchangers to the reservoirs through the respective system’s return filters.

HYDRAULIC GROUND SERVICE PANEL

pressure lines. A 3,000-psi hydraulic test rig, filled with Type IV phosphate-ester hydraulic fluid, should be used when operating the hydraulic systems.

HYDRAULIC REPLENISHER Located on the right side of the tail compartment are the hydraulic replenishing pump and the reservoir servicing system panel. The system includes a manual selector valve, an electric-driven pump, an ON–OFF momentary-hold replenisher pump control switch, a 1.5-gallon hydraulic reservoir, and a digital quantity indicator. The replenishing system is used to service small amounts of hydraulic fluid into the reservoirs. The selector valve is used to select either the left or right system reservoir, and the ON–OFF switch controls pump operation. The hydraulic quantity indicator panel gives digital readouts of the left and right hydraulic reservoir fluid quantity. The quantity is displayed in a bar graph format as well as numerically in gallons. For numerical readout, the display will show the quantity of the selected reservoir in 0.1-gallon increments. The replenisher system is powered by the ground service bus (See Figure 29-15).

The ground service panel is located on the underside of the airplane, forward of the tail compartment access. It consists of five quickdisconnects; left and right system suction, left and right system pressure, and a reservoir filler and bypass fitting. The quick-disconnects permit attachment of an external hydraulic test rig to operate the hydraulic system during aircraft ground maintenance, and are also used to service the system reservoirs. (A reservoir selector valve is installed in the tail compartment for reservoir servicing using the ground service panel.) During such operations, the engine-driven pumps are isolated from the system by a check valve installed in their respective main system


29-17



DISCH 1 2

OPEN

OPEN

M L HYD S/O

NORM

CLOSE

L ESS 28 VDC PULLED

L FIRE EXT SW NO. 5 (FIRE HANDLE)

DISCH 2 1

L HYD SHUTOFF VALVE (SHOWN IN OPEN POSITION)

OPEN

M NORM

CLOSE

R ESS 28 VDC PULLED

R FIRE EXT SW NO. 5 (FIRE HANDLE)

CLOSED R HYD SHUTOFF VALVE (SHOWN IN CLOSED POSITION)

Figure 29-9. Hydraulic Shutoff Control

29-18

L HYD VALVE OPEN L HYD VALVE CLOSED R HYD VALVE OPEN R HYD VALVE CLOSED

CLOSED

OPEN

R HYD S/O

MAU 1 SINGLE GENERIC 1 1/O MODULE SLOT 3


MAU 2 SINGLE GENERIC 4 1/O MODULE SLOT 12 L HYD VALVE OPEN L HYD VALVE CLOSED R HYD VALVE OPEN R HYD VALVE CLOSED



HYDRAULIC SYSTEM OPERATION

NOTES

ELECTRICALLY POWERED RESERVOIR SUPPLY SHUTOFF VALVE Hydraulic fluid is supplied to the enginedriven pumps by the reservoirs via the supply shutoff valves. The supply shutoff valve closes when the engine fire handle is pulled. Left and right electrical control comes from the left and right essential 28-VDC busses respectively. Circuit protection is provided by the appropriate L HYD S/O and R HYD S/O circuit breakers located on the cockpit overhead panel. The shutoff valve internal switch monitors valve position and transmits a discrete signal to the MAUs for use on the HYDRAULICS synoptic page and the primary engine instrument display (Figure 29-9).

OPERATION OF THE ENGINE HYDRAULIC SYSTEM The engine-driven pumps start generating flow as soon as the engines are started. The pumps create a suction and pull fluid from the reservoirs through the fire handle shutoff valve. Bypass fluid exits the pump via the case drain line to the heat exchanger for cooling, then back to the reservoir via the return filter for cooling and lubrication of the pumps.

NOTE Refer to the Maintenance Schematic Manual for corresponding hydraulic schematics.


29-19



Figure 29-10. Hydraulics Synoptic Page

Figure 29-11. Engine Start Synoptic Page

29-20




ENGINE-DRIVEN PUMPS There is one rotary piston, engine-driven hydraulic pump on each engine. The hydraulic pump is manufactured by the Abex Aerohydraul Division of Parker Aerospace. It is mounted on the forward left side of the accessory gearbox. The engine-driven pump provides the associated hydraulic system with pressurized hyd r a u l i c f l u i d . T h e va r i a b l e vo l u m e a n d displacement pressure compensated pumps rotate any time the engines are turning. Four ports are incorporated into the pump: a suction (inlet) port, a pressure (exit) port, a case drain (bypass) port, and a shaft seal drain port. The pump will produce 18 gpm of flow at engine idle rpm and 28 gpm at rated thrust at 3,000 +200/–100 psi. There is no method to shut down the pumps during engine operation.

ENGINE-DRIVEN PUMP ISOLATION CHECK VALVES Fluid from each pump flows through a pump isolation check valve before flowing through the acoustical filter. The valve isolates the engine-driven pumps from the system when the system is being powered via a ground service test stand.

DISTRIBUTION NOTE Refer to the Maintenance Schematic Manual for corresponding hydraulic schematics.

After entering the tail compartment, pump output is directed to the system accumulator. The accumulator absorbs pressure surges in the system. Fluid is then routed to the pump pressure switch and system pressure transducer and is then filtered by the system pressure filter. A left system pressure switch is incorporated into the left filter manifold. This switch provides the logic for automatic PTU operation when the cockpit PTU switch is in the ARM position and left system pressure is below 1500 psi. From the filter manifold assembly, the pressure is sensed by the system pressure relief valves. The right system relief valve is located on the right filter manifold, and the left system relief valve is located in the right main wheel well. System pressure is then routed to the using components and returns to the reservoirs via the return filters.

CONTROLS AND INDICATORS

ACOUSTICAL FILTER The acoustic filter smoothes and quiets hydraulic fluid flow by dampening any remaining pulses induced by the pumps. From the acoustical filter, the fluid pressurizes the engine thrust reverser system. It then enters the fuselage through a quick-disconnect fitting in the pylon.

SYNOPTIC PAGES The HYDRAULICS synoptic page displays data for the left and right hydraulic systems (Figure 29-10). This data includes reservoir fluid temperature, reservoir fluid quantity, hydraulic shutoff valve position, engine-driven pump operation, and system pressure (digital indication). In addition, the Engine Start page shows system pressure digitally (Figure 29-11).


29-21



Figure 29-12. Summary Synoptic Page

Figure 29-13. Ground Service Synoptic Page

29-22




SYNOPTIC PAGES—CONT.

NOTES

Along with the EICAS and HYDRAULICS synoptic page, system pressure can also be found on the SUMMARY and GROUND SERVICE synoptic pages (Figures 29-12 and 291 3 ) . T h e S U M M A RY p a g e a l s o s h o w s reservoir quantity. The GROUND SERVICE synoptic page displays left and right system reservoir quantity both in digital form and with a dim green line on a bar graph.


29-23

29-24

L HYD PRESS L ESS 28 VDC

28 VDC IN 28 VDC RETURN CHASSIS (GND)

0-5 VDC OUT+ 0-5 VDC OUT–

MAU 1 DUAL GENERIC 1/O MODULE (1) SLOT 9 L ESS/10 R ESS L HYD PRESSURE DC (H) L HYD PRESSURE DC (L)



AUX HYD PRESS L ESS 28 VDC R HYD PRESS R ESS 28 VDC

AUX HYD PRESSURE DC (H) AUX HYD PRESSURE DC (L)

CHASSIS (GND) 28 VDC RETURN 28 VDC IN AUX HYD SYS PRESSURE TRANSDUCER 28 VDC IN 28 VDC RETURN CHASSIS (GND)


MAU 2 DUAL GENERIC 1/O MODULE (2) SLOT 7 R MAIN/8 R ESS R HYD PRESSURE DC (H) R HYD PRESSURE DC (L)

R HYD SYS PRESSURE TRANSDUCER

R ESS 28 VDC

CHASSIS (GND) 28 VDC RETURN 28 VDC IN PTU HYD SYS PRESSURE TRANSDUCER

Figure 29-14. Hydraulic Pressure Indication

international

PTU HYD PRESS

PTU HYD PRESSURE DC (H) PTU HYD PRESSURE DC (L)

FlightSafety



L HYD SYS PRESSURE TRANSDUCER



LEFT AND RIGHT SYSTEM PRESSURE TRANSDUCERS

NOTES

The left and right system pressure transducers are powered by the on-side essential 28VDC busses through the L and R HYD PRESS circuit breakers. They provide discrete pressure inputs to the MAUs. The MAUs transmit pressure data to the FDR and to the synoptic pages (Figure 29-14).

PUMP PRESSURE SWITCHES The pump pressure switches are located on the same pressure line as the pressure transducers. The switches close when pressure exceeds 2600 psi and open when pressure drops below 2000 psi. The HYDRAULIC synoptic page displays an amber impeller symbol when pump pressure is below 2000 psi. The impeller is shown green when pump pressure is above 2600 psi (Figure 29-15).

LEFT AND RIGHT TEMPERATURE TRANSDUCERS The left and right system temperature transducers are powered by the left essential 28VDC bus through the HYD CONT circuit breaker. They provide discrete temperature inputs to the MAUs. The MAUs transmit temperature data to the FDR and to the synoptic pages (Figure 29-15).


29-25

29-26

ANNUNCIATOR LTS DIM/TEST CONTROLLER

ANNUN LTS DIM/TEST PWR

OPEN/GND

ON

OPEN

3.0 A MAX 28 VDC

MAU 1 DUAL GENERIC I/O MODULE SLOT 9/10

OPEN/GND

NORM

OPEN

L HYD TEMP DC (H) L HYD TEMP DC (L) L HYD QUANTITY DC (H) L HYD QUANTITY DC (L)

M

HYD CONT

CLOSE

CLOSED

PTU HYD PUMP NORM/ON SW

NOT OPEN

OPEN/GND

R HYD TEMP DC (H) R HYD TEMP DC (L) R HYD QUANTITY DC (H) R HYD QUANTITY DC (L) PTU HYD SOV CLOSED

PTU SHUTOFF VALVE (SHOWN IN CLOSED POSITION)

ANNUNCIATOR LTS DIM/TEST CONTROLLER NOT ARM


OPEN

ON

OPEN/GND


> 2600 PSI

NOT ARM < 2000 PSI

> 1900 PSI

HYD PUMP 1 PRESS SW < 1500 PSI

L SYS PRESS SW ARM

PTU ARMING CONT RLY

PTU HYD PUMP OFF/ARM SW

L ESS 28 VDC R ESS 28 VDC

> 2600 PSI

L SYS LOW PRESS RLY

28 VDC IN 28 VDC RETURN CASE GRND


L SYS TEMP TRANSDUCER 28 VDC IN 28 VDC RETURN CASE GRND

R HYD QTY

LEFT HYD PUMP FAIL RIGHT HYD PUMP FAIL PTU PUMP CONTROL GRND IF L RSVR LOW OR R RSVR HOT CLOSES PTU SOV

< 2000 PSI

HYD PUMP 2 PRESS SW

L HYD QTY

MAU 1 SINGLE GENERIC 1 I/O MODULE SLOT 3


MAU 2 SINGLE GENERIC 4 I/O MODULE SLOT 12 LEFT HYD PUMP FAIL RIGHT HYD PUMP FAIL PTU PUMP CONTROL GRND IF L RSVR LOW OR R RSVR HOT CLOSES PTU SOV

R SYS TEMP TRANSDUCER

GRND SRVC BUS TRANSFER RLY ENABLED

L RESERVOIR QTY TRANSDUCER OFF ON

OILER PUMP ENABLE RELAY


REPLENISHER PUMP ON/OFF SW

M


R RESERVOIR QTY TRANSDUCER

REPLENISHER MOTOR PUMP

Figure 29-15. Hydraulic Quantity/Temperature Indication

SERVICE PANEL HYDRAULIC RESERVOIR QUANTITY L HYD QUANTITY DC (L) L HYD QUANTITY DC (H) CASE GRND R HYD QUANTITY DC (L) R HYD QUANTITY DC (H) 28 VDC IN 28 VDC RETURN

international

TO ENGINE OILER GRND SRVC BUS 28 VDC


FlightSafety



L ESS 28 VDC



QUANTITY TRANSDUCERS The left and right system quantity indication LVDTs are powered by the on-side essential 28-VDC bus or by the 28-VDC ground service bus. The quantity LVDTs provide discrete quantity inputs to the MAUs and to the replenisher panel quantity indicator. The MAUs transmit quantity data to the FDR and the synoptic pages (Figure 29-15).

LEFT AND RIGHT SYSTEM QUANTITY INDICATOR The hydraulic replenisher panel also incorporates the left and right systems quantity indicator and provides left and right hydraulic reservoir fluid quantity indication. Quantity is displayed in bar graph form and digitally in gallons (Figure 29-15). Reservoir status is displayed as follows:

HYDRAULIC QUANTITY COMPENSATION Temperature Compensation The EICAS synoptic page displays are compensated to account for the expansion and contraction of the fluid with changes in temperature. As the reservoir temperature varies from the baseline temperature of 21°C, the calculated fluid expansion or contraction is subtracted or added to the measured reservoir quantity. The result is a quantity indication adjusted to always display what it would be if the reservoir temperature were 21°C. This eliminates the perceived decrease in hydraulic quantity that occurs as the aircraft is cold soaked at high altitude on long flights.

• LO indication: • Left quantity—Below 3.9 gallons • Right quantity—Below 1.2 gallons • OK indication: • Left quantity—Between 3.9 and 4.8 gallons • Right quantity—Between 1.2 and 1.5 gallons • HI indication: • Left quantity—Above 4.8 gallons • Right quantity—Above 1.5 gallons In addition, each reservoir contains a directreading quantity gage.

Landing Gear Actuator Compensation The EICAS synoptic page left system quantity display is compensated to account for the gallon of hydraulic fluid that leaves the reservoir and stays out in the landing gear actuators during gear retraction. This compensation eliminates a misleading low quantity indication by accounting for the fluid that is still in the system, but temporarily stored in the landing gear actuators. The compensation is accomplished by adding 1.0 gallon back to the reservoir quantity whenever the landing gear is retracted.


29-27



CHECK VALVE (ASC-1)

TO PTU CASE DRAIN FILTER

TO R SYS EDP CASE DRAIN FILTER

BEARING WEAR INDICATOR

CASE DRAIN

PUMP OUTLET

PUMP INLET

22.5 GPM

BOOST PUMP

SHAFT SEAL

CHECK VALVE HYDRAULIC PUMP

HYDRAULIC MOTOR

FLOW REG

MOTOR INLET

HYD FUSE

MOTOR OUTLET

TO LEFT HYDRAULIC SYSTEM

TO RIGHT HYDRAULIC SYSTEM

OVERBOARD DRAIN

LEGEND LEFT SYSTEM FLUID RIGHT SYSTEM FLUID

Figure 29-16. Power Transfer Unit (PTU)

29-28




POWER TRANSFER UNIT (PTU) SYSTEM GENERAL The PTU provides backup pressure for the following systems if the left system enginedriven pump should fail:

Fluid flow greater than 34 gpm activates the fuse, and the PTU motor is shut down. The flow regulator limits maximum rpm to 3,900 by regulating right system flow to 28 gpm. The PTU pump incorporates an integral boost pump and a redundant integral check valve. It uses fluid from the left system reservoir and produces 22.5-gpm maximum flow (Figure 29-16).

NOTE

• Flaps • Landing gear • Nosewheel steering • Normal toe brakes • HMG • Ground spoiler control pressure

PTU SYSTEM COMPONENTS The components of the PTU system, located in the tail compartment, include the PTU shutoff valve, motor/pump assembly, restrictor, pressure transducer, pressure filter, and pump case drain filter located on the left filter manifold.


PTU Motor and Pump The primary component of the PTU system is the motor and pump assembly. It is located in the tail compartment on the right forward side. The PTU consists of two fixed-displacement components: a hydraulic motor rated at 28 gpm and a hydraulic pump that produces 22.5 gpm at 3,900 rpm. The bearing wear indicator is mounted on the top of the motor and pump assembly and gives a visual indication of excessive motor bearing wear (Figure 29-16). The PTU motor is connected to a common drive shaft that turns the pump. Right hydraulic system pressure from the PTU shutoff valve drives the motor via a hydraulic fuse and flow regulator. The fuse ensures that adequate flow and pressure are available.

Aircraft Service Change 1 (ASC-1) installs a check valve on the PTU motor case drain outlet port. This check valve prevents hydraulic pressure back flow to the PTU motor when the right hydraulic system is pressurized by the right EDP and the PTU is not operating. This ASC is applicable for aircraft SNs 5001–5014.

PTU Shutoff Valve The PTU motor-driven shutoff valve is located in the tail compartment below the PTU motor/pump assembly. It provides right system pressure to the PTU motor, controls the operation of the PTU, and provides for shutoff in case of left system fluid loss or right system overtemperature. The shutoff valve is powered by the L 28-VDC essential bus through the HYD CONT circuit breaker (See Figure 29-15).

PTU Flow Restrictor The PTU flow restrictor creates a constant pressure demand on the PTU system. It bleeds PTU pressure to the left system return and prevents the PTU motor/pump from stalling for the purpose of reducing starting torque on the shaft when both right system pressure and PTU pressure are at 3,000 psi.

Pressure Transducer and Snubber The PTU system contains a pressure transducer and snubber. It is located near the motor and pump assembly and provides indication of PTU pressure on the synoptic pages.


29-29



Figure 29-17. Hydraulic Control Panel HYDRAULICS

Left 75 °C

Right 75 °C

Full

Full

Low

Low 1.5g

4.6g

Aux 3000 psi

Left 0 psi L T/R

Right 3000 psi R T/R

Aileron Elev Flt Spl Gnd Spl Stl Bar YD1

Rudder

YD2

NWS PTU 3000 psi

Main Door Flaps Gnd Spl Ctrl

HMG

Brakes ACCUM 3000 psi

Ldg Gear BOTTLE

3100 psi

Figure 29-18. Hydraulics Synoptic Page—PTU Operation

29-30




PTU Filters A PTU pressure filter is located in the tail compartment, downstream of the pump, and filters the PTU output pressure. The filter is a full-flow type, nonbypassing, and incorporates a 3-micron disposable filter element. The filter bowl contains a ratchet lock mechanism for ease of maintenance. A PTU pump case drain filter, which is a bypass-type filter, is incorporated in the left system filter manifold assembly.

PTU HYDRAULIC SYSTEM OPERATION

The PTU pressure transducer provides a PTU signal to the EICAS. It is powered by the right essential 28-VDC bus through the PTU HYD PRESS circuit breaker, which provides a discrete pressure signal to MAU-2 (See figure 2914). The MAU transmits pressure data to the FDR and the synoptic pages. The Hydraulics synoptic page displays PTU operation showing the following: the PTU shutoff valve open, the PTU motor and pump operating, PTU pressure, and the systems powered by the PTU. It also displays a blue “PTU Hydraulic On” message on the CAS (Figure 29-18).

Automatic Operation

NOTES

PTU and auxiliary hydraulic control is provided by the hydraulic control panel located on the cockpit overhead panel (Figure 29-17). The PTU has two modes of operation: automatic and manual. To select the automatic mode of operation on the hydraulic control panel, place the PTU ON switch in the NORM (dark) position and place the OFF–ARM switch in the ARM position. The PTU is now armed and ready for automatic activation. Automatic operation of the PTU system occurs when the left system pressure drops below 1,500 psi. The PTU motor-driven shutoff valve is energized to the open position. When the PTU is in automatic operation and the fluid temperature in the right reservoir exceeds104°C (220°F), or the left system fluid quantity is less than 1.5 gallons, the PTU system shuts off automatically (See Figure 29-15).

Manual Operation If the PTU system has been shut off due to hot fluid in the right system reservoir, the automatic shutoff can be overridden by selecting the PTU ON–NORM switch to the ON position. The PTU system will again operate with hot fluid in the right system reservoir. Selecting the PTU NORM–ON switch to the ON position will open the shutoff valve any time the right system is powered.


29-31

29-32

OFF

TO APU CONT

AUX PWR CNTCR

OFF ON

OPEN

ANNUN LTS PWR

ANNUNCIATOR LTS DIM/TEST CONTROLLER OPEN/GND

ON

HYD CONT

NOT ARM

OPEN/28VDC

NORM

L MN BAT SW


OPEN M CLOSE

OPEN/GND OPEN/GND

ON

CLOSED

TO BRAKE CTRL AND INDICATION

ARM

AUX HYD PUMP OFF/ARM SW

TO DC PWR

CLOSED NOT CLOSED


OPENS >356°F

P/O AUX HYD PUMP ARMING RLY

CLOSES <341°F AUX PUMP OVRLD SNSR RLY

M

THERMAL SWITCH

AUX SYS MOTOR/PUMP

DR CLSD COMMAND

WARN LTS PWR TO DOOR WARN AND CONT

AUX PUMP CTRL RLY NO.1

AUX HYD PUMP OVERLOAD SENSOR

BAT BUS 28 VDC REPLENISHER PUMP ON/OFF SW

TO DC PWR

NOT ARM

AUX PUMP CTRL RLY NO.2

OUTSIDE AUX HYD PUMP SW GRND SRVC VLV

R MN BAT SW

AUX HYD PUMP CONTACTOR

MAU 1 DUAL GENERIC I/O MODULE SLOT 9/10 AUX HYD SOV CLOSED AUX HYD PUMP OVLD AUX HYD PUMP HOT AUX HYD BOOST INLET PRESS LOW AUX HYD BOOST OUTLET PRESS LOW

MOTOR INPUT VOLTAGE MOTOR ENABLE AUX BOOST PUMP OVRLD SNSR

ANNUN LTS PWR


AUX BOOST FAIL AUX BOOST PUMP OVRLD SNSR RLY

MOTOR PWR AUX BOOST M PUMP

ON OPEN/GND

AIR ON

AUX SYS NORM/STHBY RUD VLV

Figure 29-19. Auxiliary Hydraulic System

STBY RUD HYD ON

international

STANDBY RUDDER NOSE WOW RLY SWITCH L SYS STBY RUD VLV STBY RUD VLV OVRLD RLY OVRLD SNSR


GRND

FlightSafety

OPEN/GND

ON


L ESS 28 VDC

ANNUN LTS PWR

ON



AUXILIARY HYDRAULIC SYSTEM

Located in the tail compartment are the following: • Left system reservoir

GENERAL

• Suction boost pump

The auxiliary hydraulic system provides 3000 psi pressure for normal operation of the entrance door close system and charges the park and emergency brake accumulator any time the auxiliary pump is operating. It also provides a backup source of pressure to the flaps, nosewheel steering, toe brakes, and ground spoiler control in the event that the left system and PTU lose pressure.

• Auxiliary hydraulic shutoff valve

On the ground, the auxiliary hydraulic system is used to pressurize the landing gear system for maintenance via the ground service valve. In the event of a total loss of hydraulic pressure when the aircraft is in flight, it can be used to pressurize the rudder and yaw damper actuator. Hydraulic pressure will come from the auxiliary system via use of the standby rudder switch.


AUXILIARY HYDRAULIC SYSTEM COMPONENTS The auxiliary hydraulic system consists of the following components, which are located in the main landing gear wheel well: • Electric motor-driven pump • Acoustical filter • Pressure filter

• Auxiliary return filter (left system filter manifold) • Left system/PTU pressure switch

Left System Reservoir The auxiliary hydraulic system reservoir consists of a chamber which is located inside the left system reservoir. It is a 2.0 gallon compartment that stores fluid for auxiliary hydraulic system operation. A bulkhead inside the reservoir partially isolates hydraulic system fluid from left system fluid. When the left hydraulic system or PTU is pressurized, bootstrap pressure forces left system fluid through a spill-over at the top of the bulkhead, insuring that the auxiliary chamber remains full. The components that utilize the auxiliary hydraulic system return the fluid to the auxiliary chamber.

Suction Boost Pump The suction boost pump, located in the tail compartment, pressurizes the auxiliary pump suction line to 100 psi to prevent the auxiliary pump from cavitating. The suction boost pump (SBP), contains an internal pressure switch. This switch, located in the boost pump inlet, controls operation of the pump. When the auxiliary hydraulic system is running and suction line pressure drops below 20 psi, the SBP starts operating. The SBP shuts off when suction pressure is greater than 25 psi (left system reservoir bootstrapped).

• Pressure relief valve • Pressure transducer • Standby rudder valve


29-33



Auxiliary Hydraulic Shutoff Valve The electrical motor-driven supply shutoff valve controls the fluid flow from the reservoir to the auxiliary pump. It is energized to the open position when the auxiliary hydraulic system is turned on. The valve is located in the tail compartment, on the left side floor level and aft of the left battery.

Motor and Pump Assembly The auxiliary hydraulic system motor and pump assembly is mounted to the aft bulkhead in the right main wheel well. A 28-VDC motor powers the variable-displacement, pressure-compensated, piston pump. Its maximum rated flow is 2 gpm. The current draw for continuous running is 168 amps at maximum flow rate. The auxiliary hydraulic system motor and pump assembly incorporates an internal cooling fan and a thermal overheat switch that will cause an amber message to appear on the EICAS if the motor overheats.

Standby Rudder Valve The standby rudder valve is a solenoid-operated fluid directional control valve. It normally directs fluid flow from the auxiliary pump outlet to the auxiliary hydraulic system components. When energized via the standby rudder switch, it sends fluid flow to the left system hydraulic pressure port on the rudder/yaw damper actuator.

Auxiliary System Acoustical Filter

ter incorporated in the left system filter manifold (See Figure 29-5). Both filters are bypass type and provide 3-micron filtration. They have a differential pressure indicator and have disposable filter elements. The filter bowl incorporates a ratchet lock for ease of maintenance.

Left System/PTU Pressure Switch The left system/PTU pressure switch provides input to the auxiliary hydraulic system automatic “latch-on” feature when the left system or PTU pressure is not available (below 1,500 psi) and one of the brake pedals is depressed. It is located on the left side of the tail compartment in the left system reservoir bootstrap pressure line.

Pressure Relief Valve The pressure relief valve, located in the left main wheel well’s aft bulkhead, returns any excess pressure (3-gpm full flow rate) directly back to the left reservoir’s auxiliary chamber via the auxiliary return filter. It opens at 3,850 psi and resets at 3,500 psi.

Pressure Transducer The auxiliary system pressure transducer is located in the left wheel well’s aft bulkhead. The pressure transducer is powered by the L 28-VDC essential bus through the AUX HYD PRESS circuit breaker. It provides auxiliary system pressure data to MAU No. 1 (See Figure 29-14).

The auxiliary hydraulic system also contains its own acoustical filter, located on the left wheel well’s aft bulkhead. The acoustical filter absorbs output pressure harmonics from the hydraulic pump to reduce cabin noise level.

AUXILIARY HYDRAULIC SYSTEM OPERATION

Filters

The auxiliary hydraulic system operates in two modes: automatic and manual. In the automatic mode the operation is controlled by two switches on the cockpit overhead panel: NORM–ON and OFF–ARM (see Figure 29-17).

The auxiliary hydraulic system incorporates two filters: a pressure filter, mounted on the aft bulkhead in the left wheel well, and a return fil-

29-34

Auxiliary System Automatic Mode




Auxiliary System Armed (Automatic Mode) When the auxiliary hydraulic system control panel is not armed (manual mode), the automatic operation is disabled. To enable the automatic system, select the NORM position on the NORM–ON switch, and select the ARM position on the OFF–ARM switch. The auxiliary hydraulic system is now armed and ready for automatic operation.

loaded to the OFF position and, with the lever manually held OPEN, engages the auxiliary system switch. (The ground service valve is located on the aft right corner of the nose wheel well.) The battery tie bus must have power for the ground service valve to operate.

NOTE If using the batteries to power the auxiliary hydraulic system, the battery switches must be ON.

The auxiliary hydraulic system activates automatically if the left system and PTU pressure is less than 1,500 psi, as sensed by the left system/PTU pressure switch, and a brake pedal is depressed greater than 10°. The auxiliary hydraulic system is also activated when the standby rudder switch is selected to the ON position (See Figure 29-19).

Due to current draw from batteries, auxiliary hydraulic system use should be limited to prevent draining the batteries.

Standby Rudder Solenoid Valve

NOTE

The standby rudder solenoid valve is energized when the standby rudder switch in the cockpit is selected ON and there is no nose weight-on-wheels signal. This action allows the auxiliary hydraulic system to power the rudder actuator in the event of a dual hydraulic system failure.

Electrical power can be applied from the batteries, externally, or by onboard power.

The valve also sends a discrete input to MAU No. 2 when the standby rudder switch is ON. MAU No. 2 will also provide indications on the HYDRAULICS synoptic page and CAS when the standby rudder system is selected ON. The standby rudder valve is protected by an overload sensor. If a current draw of 7 amps is detected in the solenoid, the solenoid will de-energize (See Figure 29-19).

Auxiliary System Manual Mode Auxiliary hydraulic system manual control (manual mode) is provided by two switches: the NORM–ON switch, located on the cockpit overhead panel, and the ground service valve. The ground service valve is spring-

CAUTION

Activation of the auxiliary hydraulic system is also accomplished when the entrance door control system is energized. An external switch panel incorporates the external battery switch and the outside door switch. The outside door switch automatically energizes the battery tie bus and the auxiliary hydraulic system until the primary door lock is engaged. If the inside door switches are used, the battery switch must be hard-selected to power the door closed using the auxiliary system. The auxiliary hydraulic system remains energized until the primary door lock is engaged. The two inside door switches are located on the cockpit overhead panel and on the left electronic equipment rack forward of the main entry door. Pressure in the system is read in t h e AU X P R E S S w i n d ow o f t h e E I C A S . Auxiliary pressure is displayed in the same places as for the left and right hydraulic sys-


29-35



104

0.0

Figure 29-20. Hydraulics Synoptic Page—Left and Right System Fault Messages

24

4.8

Figure 29-21. Hydraulics Synoptic Page—PTU Fault Message

29-36




NOTES

tems.

FAULT INDICATIONS LEFT AND RIGHT SYSTEM FAULT INDICATIONS The CAS display and HYDRAULICS synoptic page display fault indication messages. The following is a list of the CAS messages and a description of the faults associated with them (Figures 29-20 and 29-21): • L-R Hydraulic System Fail—Left or right system pressure is less than 1,500 psi, as indicated by the left/right pressure transducer. • L-R Hydraulic Reservoir Hot—Left or right reservoir fluid temperature has exceeded 104°C (220°F) ±5°, as sensed by the reservoir temperature sensor. • L-R Hydraulic Quantity Low—Fluid level in the left and right system reservoirs is low, as sensed by the quantity LVDT (below 2.8 gallons for left system and below 1 gallon for right system). • PTU Hydraulic Fail—Amber message is displayed when one of the following conditions occurs (Figure 29-20): • PTU shutoff valve is open, PTU pressure is less than 1,500 psi, and right system pressure is greater than 1,500 psi. (Pressure inputs come from the PTU and right system pressure transducers.) • PTU shutoff valve is closed, and PTU pressure is greater than 1,500 psi.


29-37



Figure 29-22. Hydraulics Synoptic Page—Auxiliary System “Fail” Message

29-38




AUXILIARY SYSTEM FAULT INDICATIONS

NOTES

An amber “Aux Hydraulic Hot” message is displayed when the auxiliary hydraulic system motor thermal switch exceeds 356°F (Figure 29-22). The message turns off when the temperature is less than 341° F. An amber “Aux Hydraulic Fail” message appears if the auxiliary shutoff valve is not closed and the pressure is less than 1,500 psi (Figure 29-22). The message is also displayed if the auxiliary shutoff valve is closed and the pressure is greater than 1,500 psi. An amber “Aux Hydraulic Pump Overload” message appears if a pump overload greater than 200 amps has occurred, causing the system and the pump to shut down (Figure 29-22). An amber “Aux Hydraulic Boost Fail” message appears when a boost pump overload greater than 15 amps is detected (Figure 29-22).


29-39



SLIDING BAYONET SLIDING BAYONET

MAIN CABIN DOOR INTERIOR HANDLE LIFT HANDLE TO OPEN MAIN CABIN DOOR. CONTROLS SIX SLIDING BAYONETS WHICH LOCK MAIN CABIN DOOR.

SLIDING BAYONET

SECONDARY HANDLE BOTTOM PORTION OF HANDLE MUST BE DEPRESSED TO OPERATE LARGER PRIMARY HANDLE.

MAIN CABIN DOOR PRIMARY HANDLE FOUND ON OUTSIDE LOWER EDGE OF DOOR. IS THE EXTERIOR PRIMARY LOCKING AND UNLOCKING MECHANISM

Figure 29-23. Main Door Components

29-40


KEY-OPERATED LOCK



HYDRAULIC SUBSYSTEM MAIN ENTRY DOOR The main entry door allows access to the passenger cabin and cockpit. The door free-falls open and is hydromechanically closed. The door components consist of the door steps, lower airstair, and mechanical linkage.

is open, the airstair is mechanically locked into position. As the door is closed, the airstair is hydraulically unlocked.

Mechanical Linkage The sequence and control of the door operation are controlled by mechanical linkage, which consists of bellcranks, pushrods, and sector crank assemblies (Figure 29-23).

Door The door is located on the forward left side of the aircraft (Figure 29-23). The door assembly is 36 inches in width and 62 inches in length. The construction is rib and beam, with a stressed skin covering. The assembly consists of a door structure and upper steps and is attached to the fuselage by a hinge. A noninflatable door seal provides sealing for cabin pressurization.

Lower Airstair The lower airstair is attached to the upper stairs and provides support to the airstair and accessibility into the aircraft. When the door

LOCKING HANDLES The locking handles, both internal and external, consist of primary and secondary handles (Figures 29-23 and 29-24). When the primary locking handle is rotated to the closed position, it mechanically actuates six bayonets that lock the door in place. The six locking bayonets are located within the door structure to lock the door closed (Figure 29-23). The purpose of the secondary locking handle is to lock the primary door handle in the open and closed position. When the primary handle is latched, the auxiliary pump is deenergized.

SECONDARY HANDLE SWITCH

SECONDARY LOCKING HANDLE

PRIMARY LOCKING HANDLE

Figure 29-24. Main Door Locking Handles


29-41



RESTRICTOR

DOOR CONTROL VALVE

PRESSURE REDUCER

DOOR ACTUATOR

Figure 29-25. Main Door Hydraulic Components

29-42




MAIN ENTRY DOOR HYDRAULIC COMPONENTS

NOTES

Pressure Reducer The pressure reducer is located on the fuselage bulkhead, aft of the main entry door. The purpose of the pressure reducer is to control the door rate of closure by reducing the hydraulic pressure to approximately 2,300 psi (Figure 29-25).

Door Control Solenoid Valve The door control solenoid valve is also located on the fuselage bulkhead, aft of the main entry door. It is electrically energized and ports hydraulic fluid for door closure. If the control solenoid valve fails, there is a manual override (Figure 29-25).

Restrictor The restrictor is located on the main door actuator and restricts hydraulic fluid flow during opening and closing. It regulates how quickly the door free-falls to the open position (Figure 29-25).

Airstair Locking Actuator The airstair locking actuator is located inside the main door, on the right side, where the two sections meet. The actuator is springloaded to the extend position and is hydraulically retracted to unlock the lower airstair during the door closure cycle.

Main Door Actuator The main door actuator is located on the fuselage bulkhead, aft of the main entry door. To close the door, the actuator extends. An internal snubbing device slows the door during closing (Figure 29-25).


29-43



Figure 29-26. Main Door Control Panel

29-44




MAIN ENTRY DOOR CONTROL COMPONENTS

NOTES

The internal door switch panel is located on the cockpit overhead panel and consists of two push button switches (Figure 29-26). One switch serves as the door safety switch while the other is the door control switch. There is also a red guarded door control switch on the aft bulkhead of the left electronic equipment rack. When energized, the DOOR SAFETY switch will prevent any DOOR CLOSE switch from closing the door.

Cabin Door Close Switch The cabin door close switch is used to close the main cabin door from the cabin area. The switch is normally in the guarded (off) position. The switch is mounted on the aft bulkhead of the left electronic equipment rack.


29-45



Figure 29-27. External Door Switch Location

Figure 29-28. Secondary Handle

29-46




External Door Control Switch

NOTES

An external door control switch is located in the forward external service panel, just forward of the main entrance door. The switch connects the right battery to the battery tie bus for auxiliary pump operation. Turning the battery on for this operation is not necessary. The external service panel door opens outward to gain access to the switch (Figure 29-27).

WARNING SYSTEM COMPONENTS The warning system provides an indication if the door is not closed or locked.

Main Doorlock Switches There are six main doorlock switches mounted on the door frame. The switches are mounted in series and are actuated by the six locking bayonets located within the door structure.

Secondary Door Handle Switch The secondary door handle switch is mounted in the door at the internal secondary handle (Figure 29-28). It is connected in series with the other six switches and is actuated by the primary door handle. The only function of the secondary door handle switch is to provide a warning if the door is not properly locked (See Figure 29-24).


29-47



Figure 29-29. Doors Synoptic Page 28 VDC ESS BUS LOWER AIRSTAIR RELEASE ACTUATOR

DOOR SAFETY SWITCH

DOOR CONTROL SWITCH

RETURN DOOR ACTUATOR

RESTRICTOR

MANUAL OVERRIDE DOOR CONTROL VALVE

PRESSURE REDUCING VALVE

Figure 29-30. Door Hydraulic Schematic

29-48


OFF

OUTSIDE DOOR SWITCH

CLOSE



MAIN ENTRY DOOR CONTROL AND WARNING SYSTEM OPERATION With the DOOR SAFETY switch in the off position, actuating any one of the door control switches activates the auxiliary hydraulic pump to close the door. When the cockpit door switch is pushed, the switchlight illuminates amber until the door movement is complete. When the door is closed and the handle is locked, the auxiliary hydraulic pump shuts down. The movement of the door can be stopped by pressing the DOOR SAFETY switch. If any doorlock control switch or the secondary door handle switch is placed in the unlocked position, a red MAIN DOOR message appears on CAS (Figure 29-29).

WARNING Prior to opening the main entry door, check to ensure the main engine and APU bleed-air switches are in the OFF/extended position and the cabin differential pressure indicator located on the cockpit overhead panel reads 0.00. Failure to do so may result in injury to personnel and/or damage to equipment.

To open the main entry door, lift the secondary handle and push upward on the primary handle with an open hand. When the primary handle is in the open position, the secondary handle is lowered to lock the primary handle in the open position. The door will free-fall to the extended position.

MAIN ENTRY DOOR OPERATION

WARNING

The auxiliary hydraulic system provides power to unlock the lower airstair latches and retract the door. A reducing valve drops the pressure from 3,000 to 2,300 psi. This reduced pressure controls the rate of door closure. The fluid is then routed to the door control valve (Figure 29-30). The control valve is energized when any DOOR CLOSE switch is activated. With the valve energized, the fluid return port is closed, the pressure port is opened, the lower airstair latches are unlocked, and pressure is routed through the restrictor to the main door actuator. The actuator will extend, closing the door. When the door is flush with the fuselage, the secondary handle is lifted, releasing the primary handle. The primary handle can then be rotated downward to the closed position, extending the locking bayonets, locking the door. (The bayonet fittings should be checked to ensure proper door seating, which is indicated by a visible orange dot on the bayonet fittings). The secondary handle is then lowered to lock the primary handle in the closed position.

Always ensure that the door area is clear of personnel and equipment prior to operating the door.

WARNING If the door is closed manually, without hydraulic pressure, placard the handle and use extreme caution when opening the door. Use three or four people to lower the door manually.


29-49



CHAPTER 30 ICE AND RAIN PROTECTION CONTENTS Page INTRODUCTION ................................................................................................................. 30-1 GENERAL ............................................................................................................................ 30-1 ICE DETECTION SYSTEM ................................................................................................ 30-3 General........................................................................................................................... 30-3 Components and Controls.............................................................................................. 30-3 Operation and Indications.............................................................................................. 30-9 WING ANTI-ICE SYSTEM............................................................................................... 30-13 General......................................................................................................................... 30-13 Components and Controls ........................................................................................... 30-13 Operation and Indications............................................................................................ 30-19 ENGINE COWL ANTI-ICE SYSTEM .............................................................................. 30-19 General......................................................................................................................... 30-19 Components and Controls ........................................................................................... 30-21 Operation and Indications............................................................................................ 30-25 PROBE ANTI-ICE SYSTEM ............................................................................................. 30-27 General......................................................................................................................... 30-27 Components and Controls ........................................................................................... 30-27 Operation and Indications............................................................................................ 30-33 WINDSHIELD/WINDOW ICE AND RAIN SYSTEM..................................................... 30-35 General......................................................................................................................... 30-35


30-i



Windshield Blower System ......................................................................................... 30-35 Windshield Heat System.............................................................................................. 30-37 Cabin Window Heat System........................................................................................ 30-41

30-ii





Title

Page

30-1

Ice Detection System............................................................................................. 30-2

30-2

Anti-Ice Control Panel ........................................................................................... 30-3

30-3

Ice Detection Schematic......................................................................................... 30-4

30-4

Ice Detectors........................................................................................................... 30-5

30-5

Ice Detector Test Switch ........................................................................................ 30-6

30-6

Cowl Anti-Ice Operation........................................................................................ 30-8

30-7

Wing Anti-Ice Operation ..................................................................................... 30-10

30-8

Wing Anti-Ice Temperature Sensors.................................................................... 30-12

30-9

ECS/PRESS Synoptic Page ................................................................................. 30-14

30-10

Wing Anti-Ice Ducting......................................................................................... 30-16

30-11

Wing Anti-Ice Ducting Components ................................................................... 30-17

30-12

Wing Anti-Ice Control Schematic........................................................................ 30-18

30-13

Engine Cowl Anti-Ice System.............................................................................. 30-19

30-14

Cowl Anti-Ice Control—Valve Open................................................................... 30-20

30-15

Cowl Anti-Ice Control—Valve Closed ................................................................ 30-20

30-16

Cowl Anti-Ice Control—Valve Regulating.......................................................... 30-21

30-17

Pressure Transducer ............................................................................................. 30-22

30-18

Cowl Anti-Ice Components ................................................................................. 30-24

30-19

Probe Anti-Ice System ......................................................................................... 30-26

30-20

Upper Pitot Static and No. 1 TAT Probe Schematic ............................................ 30-28

30-21

AOA and Pitot Probe............................................................................................ 30-30

30-22

Probe Heater Control Panel ................................................................................. 30-30

30-23

AOA Probe Heat Schematic................................................................................. 30-32


30-iii



30-24

Windshield Rain Removal Schematic.................................................................. 30-34

30-25

Windshield Blower System.................................................................................. 30-35

30-26

Windshield Heat Control Units............................................................................ 30-36

30-27

Windshield Heat Sensor Schematic ..................................................................... 30-38

30-28

Cabin Window Heat Switch................................................................................. 30-40

30-29

Window Heat Ground Bypass Switch.................................................................. 30-41

30-30

Emergency Exit Window ..................................................................................... 30-42

30-31

Cabin Window Diagram ...................................................................................... 30-43

30-32

Cabin Window Heat System Schematic .............................................................. 30-44

30-iv




CHAPTER 30 ICE AND RAIN PROTECTION

INTRODUCTION The Gulfstream G500/G550 ice and rain protection systems are designed for the detection, prevention, and removal of ice, rain, and fog. The components and operation of the following systems are discussed: ice detection, wing anti-ice, engine cowl anti-ice, probe anti-ice, windshield blower, windshield heat, and cabin window heat.

GENERAL There are four anti-ice areas on the Gulfstream G500/G550 aircraft: the wing leading edges, engine cowls, windshield, and air data sensors. The term “anti-ice” refers to the prevention of ice formation on these areas b y

maintaining the areas’ temperature at a high enough level to prevent icing. If icing should occur, it could adversely affect flight handling characteristics and engine performance.


30-1



Figure 30-1. Ice Detection System

30-2




ICE DETECTION SYSTEM GENERAL The primary purpose of the ice detection system is to sense icing conditions on the aircraft and to provide visual alerts to the flight crew.

head panel. When the switches are in the ON or OFF position, they provide manual control to the normally automatic anti-ice system.

The secondary purpose of the ice detection system is to provide automatic control of the wing and cowl anti-ice systems when icing conditions are detected (Figure 30-1).

In the AUTO select mode, anti-ice is not enabled until the aircraft ascends above 1,500 feet above ground level (AGL) after takeoff and below 35,000 feet pressure altitude to allow for maximum performance. In AUTO mode, the wing anti-ice system receives a discrete input from the ice detector system and provides control of the related anti-ice valve to maintain 130 ±10° F wing anti-ice temperature during icing conditions. The cowl antiice system, in the AUTO select mode, is also enabled based on a signal from the ice detector system.

COMPONENTS AND CONTROLS Wing and Cowl Anti-Ice Control Panel The wing and cowl anti-ice switches (Figure 30-2) are three-pole rotary switches located on the anti-ice control panel in the cockpit over

Figure 30-2. Anti-Ice Control Panel


30-3

30-4 ESS ØB 115-VAC BUS

MOUNTED ON LEFT SIDE OF AIRCRAFT LEFT ICE DETECTOR 115-VAC INPUT AC RETURN TEST INPUT SIGNAL RETURN


CASE GROUND STATUS SIGNAL (GND–OK) ICE SIGNAL NO. 1 (GND–ICE) ICE SIGNAL NO. 2 (COWL) (GND–ICE)

ICE DET ANNUNCIATOR LTS 28-VDC POWER

TEST

ANNUNCIATOR LTS DIM AND TEST CONTROLLER ANNUN LT GND GROUND TRIP (GND ACTIVATED)

MAU 1 DUAL GENERIC I/O SLOT 9 L MAIN

MAU2 DUAL GENERIC I/O SLOT 7 R MAIN

L ICE DETECTOR FAIL (A) L ICE DETECTED (A)


R ICE DETECTOR FAIL (A) R ICE DETECTED (A)

ICE SIGNAL NO. 3 (WING) (GND–ICE) L ICE DET CONT LEFT ESS 28-VDC BUS COWL ICE DETECTED L ICE DET AUTO COWL AI RELAY WING ICE DETECTED

CAI SYS CNTRL

MAU 1 CUSTOM I/O SLOT 4 L ESS RA BELOW 1,500 FT OUT

MAU 2 CUSTOM I/O SLOT 11 R ESS RA BELOW 1,500 FT OUT TO R BLEED AIR OFF RELAY

<35,000 FT WAI SYS CNTRL

L PACK CONT RLY SW’D 28 VDC >35,000 FT L BYPASS VLV CONT RELAY

L ICE DET AUTO COWL AI RELAY WAI/CAI SYS CNTRL

MOUNTED ON RIGHT SIDE OF AIRCRAFT

ESS ØC 115-VAC BUS

RIGHT ICE DETECTOR 115-VAC INPUT AC RETURN TEST INPUT SIGNAL RETURN CASE GROUND STATUS SIGNAL (GND–OK) ICE SIGNAL NO. 1 (GND–ICE) ICE SIGNAL NO. 2 (COWL) (GND–ICE)

PROBE PROTRUDES OUT INTO AIRSTREAM

ICE SIGNAL NO. 3 (WING) (GND–ICE) R ICE DET CONT RIGHT ESS 28-VDC BUS COWL ICE DETECTED

<35,000 FT R PACK CONT RLY SW’D 28 VDC >35,000 FT R BYPASS VLV CONT RELAY WOW 28-VDC POWER ON THE GROUND FROM THE WOW SYSTEM

R ICE DET AUTO WING AI RELAY

international

AUTO AI ALT INHIBIT RELAY

WING ICE DETECTED

FlightSafety

AUTO AI INHIBITED

R ICE DET AUTO COWL AI RELAY


PROBE PROTRUDES OUT INTO AIRSTREAM

R ICE DET

L ICE DET

Figure 30-3. Ice Detection Schematic



Ice Detectors Two ice detectors, located on the forward fuselage, receive power from the 115-VAC essential AC buses (Figures 30-3 and 30-4). The ice detectors are microprocessors whose probes operate on a resonant frequency of 40,000 hertz (Hz). Ice accumulation on the probe will result in a decrease of the resonant frequency. The ice detectors provide a visual warning on the CAS display when icing has been detected after one icing cycle is complete and the aircraft is in flight. An icing cycle occurs when ice develops on the ice detector probe, causing the resonant frequency to drop by approximately 133 Hz. The cycle begins with a 0.020-inch accumulation of ice on the probe and ends when the probe heats and melts the accumulated ice. When ice is detected, the ice detectors provide the ground for three outputs. Output No.

1 provides the MAU1 DGIO1 for the left side and MAU2 DGIO2 for the right side with a ground for the amber L-R ICE DETECTED message on the engine indicating and crew alert system (CAS), output No. 2 provides a ground for the left and right ice detection automatic cowl anti-ice relay, and output No. 3 provides a ground for the left and right ice detection automatic wing anti-ice relay. The 28-VDC power source for the ice detection automatic anti-ice relays is provided by the left and right essential DC buses. Ice detectors activate the respective anti-ice systems when the following conditions exist: the wing and cowl anti-ice switches are in AUTO mode, icing has been detected, and aircraft altitude is greater than 1,500 feet AGL after takeoff but less than 35,000 feet pressure altitude. ICE DETECTOR

Figure 30-4. Ice Detectors


30-5



Figure 30-5. Ice Detector Test Switch

30-6




Automatic Anti-Icing Altitude Inhibit Relay

NOTES

The automatic anti-icing altitude inhibit relay is located on the 347 relay panel forward of the left electronic equipment rack (LEER). It inhibits the automatic operation of the wing and cowl anti-ice systems when the aircraft altitude is less than 1,500 feet AGL after takeoff or greater than 35,000 feet pressure altitude.

Ice Detector Test Switch The ICE DET TEST switch is located on the SYSTEM TEST panel (Figure 30-5). When pressed, the switch illuminates TEST (blue) and initiates a test of the ice detection circuitry. The amber “L-R Ice Detect Fail” and “L-R Ice Detected” messages are displayed on the CAS as a result of this test. S e l e c t i n g t h e w i n g a n d / o r c ow l a n t i - i c e switches to AUTO energizes the wing and/or cowl anti-ice system for three seconds. As a result, the CAS displays the blue”L-R Cowl Anti-Ice On” and “L-R Wing Anti-Ice On” messages.


30-7

30-8 MAU 2 DUAL GENERIC I/O MODULE SLOT 7/8 R COWL A/I ON (B) OPEN

ANNUN LTS PWR

R COWL VLV FAIL CL (A)

R ESS 28 VDC

VLV OPEN

R COWL VLV FAIL OP (A)

POSITION SWITCH

WARN LTS PWR AUTO

R ESS 28 VDC

VLV CLSD

ON

ENER. TO CLOSE

CLSD


R COWL ANTI-ICE

OFF AUTO

R MAIN 28 VDC ON

NORMAL ICE DETECTED R ICE DETECT COWL A/I RLY

R COWL A/I SEL SWITCH

NORMAL

ON

ICE DETECTED

OFF

L ICE DETECT COWL A/I RLY

SOLENOID R COWL A/I VALVE MAU 1

R COWL A/I CONT RELAY

DUAL GENERIC I/O MODULE SLOT 9/10

>1500 FT <35000 FT

L COWL A/I ON (B) L COWL VLV FAIL CL (A)

<1500 FT >35000 FT AUTO A/I ALT INHIBIT RLY

L COWL VLV FAIL OP (A) OPEN

ANNUN LTS PWR

VLV OPEN

L ESS 28 VDC

POSITION SWITCH

OFF WARN LTS PWR AUTO L ESS 28 VDC

VLV CLSD ON ENER. TO CLOSE CLSD

L COWL ANTI-ICE L MAIN 28 VDC

OFF AUTO

ON

NORMAL

ICE DETECTED

ICE DETECTED L ICE DETECT COWL A/I RLY

R ICE DETECT COWL A/I RLY >1500 FT <35000 FT

SOLENOID L COWL A/I VALVE

ON

OFF L COWL A/I CONT RELAY

COWL ANTI-ICE CONTROL AND RIGHT COWL ANTI-ICE SELECTED AUTO • LEFT RELAYS AND SOLENOIDS EMERGIZED • CONTROL INHIBIT RELAY NOT EMERGIZED, NO ICE DETECTED •

Figure 30-6. Cowl Anti-Ice Operation

international

<1500 FT >35000 FT AUTO A/I ALT INHIBIT RLY

FlightSafety

L COWL A/I SEL SWITCH

NORMAL


OFF



OPERATION AND INDICATIONS

NOTES

Normal Operation During normal in-flight operation, the ice detectors provide an alert to the CAS display when ice buildup is detected after one complete cycle. When AUTO mode is selected, the ice detection circuitry provides control over the wing and cowl anti-ice system. This occurs when icing conditions are detected above 1,500 feet AGL after initial takeoff and below 35,000 feet pressure altitude.

Cowl Anti-Ice Automatic Operation The left and right cowl anti-ice valve system (Figure 30-6) is selected via a three-pole, three-position (ON, OFF, AUTO) rotary switch located on the anti-ice control panel in the cockpit overhead console. When the switch is in the ON position, the system is enabled, which deenergizes the valve solenoid and allows the valve to open. In the AUTO select mode, cowl anti-ice is not enabled until the aircraft ascends above 1,500 feet AGL to allow for maximum takeoff performance. When the switch is set to the AUTO mode, the valve is opened and closed based on the input of the left and right ice detectors. If either the left or right cowl anti-ice is selected in the AUTO or ON mode, a blue L-R COWL A/I ON message appears on the CAS.


30-9

30-10 R BLD AIR/WING A/I CONTROLLER

R BLEED AIR CONTROL R ESS 28 VAC

+28 VDC TORQUE MOTOR CONTROL 0-100 mA

R WING A/I TEMP SENSOR TO BLEED AIR CONTROL GAP BAND NO.2

DISPLAY/FAULT DATA OUT

(PROVIDES ACTUAL DUCT TEMP)

ARINC 429

R WING A/I ON/OFF COMMAND OFF

CL


AUTO

WARN LTS POWER

ON

R WG/AI SYS FL (A)

OFF

R WG/AI ON (B)

AUTO

180°F

WARN LTS POWER

R ESS 28 VAC

MAU 2

OP

ON

R WING ANTI-ICE

OFF

GAP BAND NO.1, 2, 4

AUTO

ON

WAI CONTROL RELAY

OVERTEMP SWITCHES

>1500 FT <35000 FT

MAU 1 <1500 FT >35000 FT

100°F

AUTO A/I ALT INHIBIT RELAY

GAP BAND NO.2 NORMAL

A/C IN FLIGHT ABOVE 1500' AGL, BELOW 35,000 FT.

R WING A/I VALVE

ICE DETECTED

L ICE DETECTOR AUTO WAI RELAY

UNDERTEMP SWITCH

SINGLE GENERIC I/O MODULE SLOT 3 R WING TEMP LO (A) MAU 2 SINGLE GENERIC I/O MODULE SLOT 12 R WING TEMP LO (A)

NORMAL

ICE DETECTED

R ICE DETECTOR AUTO WAI RELAY

TORQUE MOTOR CONTROL 0-100 mA

international

FlightSafety

WAI SWITCHES SELECTED TO OFF

DUAL GENERIC I/O MODULE SLOT 7/8

SOLENOID

R WING A/I SELECT SW

CONTROLLER POWERED, SENSOR POWERED

R WG/AI HOT (A)


WING A/I TEMP SENSOR INPUT (CONTROLS TO 130 ± 10°F

Figure 30-7. Wing Anti-Ice Operation



Wing Anti-Ice Automatic Operation

NOTES

The wing anti-ice system (Figure 30-8) is activated by the left and right three-pole, threeposition rotary switch on the wing anti-ice section of the cockpit overhead panel. The AUTO position is considered the normal mode of operation. The left wing anti-ice system is powered by the L WING ANTI-ICE circuit breaker and bused to the left essential 28-VDC bus. The right wing anti-ice is powered by the R WING ANTI-ICE circuit breaker and bused to the right essential 28-VDC bus. The wing anti-ice control relay is energized to provide 28 VDC to the wing anti-ice control valve solenoid and 28 VDC to the MAU1 DGIO1 for the left system and MAU2 DGIO2 for the right system for the “L-R Wing Anti-Ice On” message to the CAS. If no icing is detected after the anti-ice systems have been activated due to ice accumulation, the “Ice Detected” CAS message deactivates after one minute, the cowl anti-ice system deactivates after three minutes, and the wing anti-ice system deactivates after five minutes. In all modes, the “Ice Detected” message is inhibited with weight on wheels by the on-side MAUs. Failure of an ice detector results in a message that appears on the CAS display. Double-redundant operation allows one ice detector to control both anti-ice systems. It also will provide messages to the CAS should the other detector fail. When the ICE DET TEST switch is pressed, the ice detection circuitry is tested, and all associated CAS messages are displayed. The wing and cowl anti-ice systems are activated for three seconds. If the “Ice Detector Fail” message remains on after the test is complete, power on the system should be cycled and the test repeated.


30-11



180°F

No. 3 R O'HEAT SWITCH R BAC/WAI CONTROLLER

No. 2 R O'HEAT SWITCH

180°F

100°F

LEADING EDGE TEMPERATURE SENSOR

No. 1 R O'HEAT SWITCH

ANTI ICE

OFF

180°F

L WING

L COWL

AUTO

AUTO

ON

OFF

R COWL ON

OFF

AUTO

R WING ON

OFF

AUTO

ON

WING A/I CHECK VALVE

WING A/I CONTROL VALVE

WING A/I CHECK VALVE

WING A/I CONTROL VALVE

WING A/I CROSSOVER DUCT

180°F

R UNDERTEMP SWITCH

No. 1 L O'HEAT SWITCH

100°F

LEADING EDGE TEMPERATURE SENSOR

L UNDERTEMP SWITCH

180°F

No. 2 L O'HEAT SWITCH L BAC/WAI CONTROLLER 180°F

No. 3 L O'HEAT SWITCH

Figure 30-8. Wing Anti-Ice Temperature Sensors

30-12




WING ANTI-ICE SYSTEM GENERAL The purpose of the wing anti-ice system is to continually monitor and control the wing temperature to 130 ±10° F. Should the pressure or temperature rise above or fall below set parameters, the system notifies the crew via the CAS.

COMPONENTS AND CONTROLS Wing Anti-Ice Control Switches The wing anti-ice control switches are located on the anti-ice section of the cockpit overhead panel (see Figure 30-2). Each switch has three positions: ON, OFF, and AUTO. In the ON position the controller provides control of the related valve to maintain the appropriate wing anti-ice temperature during all icing conditions. The temperature as sensed at each wing return duct is controlled to 130 ±10° F. In the AUTO mode the controller and valve function in the same manner as in the ON position, except the controller receives the ON command only when ice is detected by the ice detector system. When the switch is selected to the OFF position, the control relay is deenergized, removing DC power from the controller command channel and causing the wing antiice valve to move to the fully closed position.

Bleed-Air Controllers Two bleed-air controllers are located in the aft baggage compartment. The bleed-air controllers maintain the wing anti-ice air temperature by adjusting the valve position through the torque motor, based on inputs from the temperature sensors located in the wing leading edge. Each controller provides anti-ice system status to the MAU1 DGIO1 on the ARINC 429 bus for the left system and MAU2 DGIO2 on the Arinc 429 bus for the right system for display on the CAS.

Wing Anti-Ice Temperature Sensors The wing anti-ice temperature sensors (Figure 30-8) are negative temperature coefficient thermistor sensing elements. They are located inside the inboard leading-edge access point (gap band No. 2) of each wing and are mounted in the wing return duct to monitor the return duct temperature. The sensors also provide for a signal to the bleed-air controller to maintain the anti-ice temperature at 130 ±10° F and to provide temperature indications for display on the CAS at all times.

Wing Anti-Ice Control Valves The wing anti-ice control valves are 3-inch in diameter, spring-loaded-closed, modulating pressure regulator and shutoff valves mounted in the left and right bleed-air manifolds. Each valve incorporates a torque motor and a solenoid, which controls airflow to the leading edge of the wings for anti-icing. The torque motor is controlled by the bleed-air controller. The solenoid is controlled by the wing antiice control switch. The valve acts as a pressure regulator with a maximum downstream pressure of 35 ±5 psi.


30-13



Figure 30-9. ECS/PRESS Synoptic Page

30-14




Wing Anti-Ice Temperature Switches

NOTES

The wing anti-ice temperature switches are located in the left and right wing leading edge at gap bands No. 1, No. 2, and No. 4. Each wing leading edge contains four temperature switches which send information to the MAUs. Three of the switches are normally open and will close if the wing leading edge temperature reaches 180°F. One of the switches is normally closed during wing anti-ice operation, and will open if the wing leading edge temperature falls below 100°F (Figure 30-9). The left and right wing 100-degree switch provides their inputs to MAU2 SGIO4 and MAU1 SGIO1. The left wing 180-degree switches provide their inputs to MAU1 DGIO1. The right wing 180-degree switches provide their inputs to MAU2 DGIO2. The MAUs use these inputs for fault indications, and for indication on the ECS/Press synoptic page in the form of varying colored lines to represent the following conditions: • Blue lines—Wing anti-ice system less than 100°F during warm-up period • Amber lines—Undertemperature conditions (less than 100°F after two minutes) • Green lines—Normal temperature (100 to 180°F) • Amber lines—Overtemperature conditions (greater than 180°F)


30-15



WING ANTI-ICE CHECK VALVES

CROSSOVER DUCT LEFT WING ANTI-ICE VALVE

RIGHT WING ANTI-ICE VALVE

LEFT BLEED AIR MANIFOLD

RIGHT BLEED AIR MANIFOLD

Figure 30-10. Wing Anti-Ice Ducting

30-16




Wing Anti-Ice Ducting The wing anti-ice ducting (Figure 30-10) provides a path for bleed air from the bleed-air manifold to be used for wing anti-icing. The supply ducting runs from the wing anti-ice control valves to the leading edges of each wing. The crossover duct permits bleed air to be distributed to both wings, should one system be inoperative or under pressure. The crossover duct is located under the baggage compartment floor, just aft of the secondary pressure bulkhead and forward of the wing anti-ice check valves.

Wing Anti-Ice Check Valves There are two wing anti-ice check valves located under the baggage compartment floor

(FS 694). The check valves permit the flow of air from the bleed-air manifold to the wing anti-ice ducting during normal operation. When the left and right wing anti-ice bleedair pressures are not equal, the valves prevent reverse airflow down the respective ducting. Bleed air is conveyed by the main duct to the leading edge of the wings and the landing lights. It is distributed to the wing leading edges and landing lights through holes in the piccolo assemblies (Figure 30-11). The return air is routed through the return ducts and vented overboard through exhaust vents located on the wing body fairing, just aft of the main landing gear.

ANTI-ICING SUPPLY DUCT (REF)

LANDING LIGHT LENS ASSEMBLY

PICCOLO TUBES

Figure 30-11. Wing Anti-Ice Ducting Components


30-17

30-18

R BLD AIR WING A/I CONTROLLER

R BLEED AIR

R ESS 28 VDC

+28 VDC

R WING A/I TEMP TO BLEED AIR

(CONTROLS TO ±130


OFF AUTO

R WING A/I ON/OFF

WARN LTS

R WG/AI SYS FL (A) R WING A/I ON (B) OP

ON OFF AUTO ON

ARINC 429

CL

ON OFF AUTO

R ESS 28 VDC

DISPLAY/FAULT DATA OUT

WARN LTS

OVERTEMP SWITCHES

>1500 FT

R WING A/I SELECT 100 F

NORMAL ICE

L ICE DETECTOR

GAP BAND No.2

UNDERTE MP

MAU 1 SINGLE GENERIC I/O MODULE SLOT 3 R WING TEMP LO MAU 2 SINGLE GENERIC I/O MODULE SLOT R WING TEMP LO

NORMAL

Figure 30-12. Wing Anti-Ice Control Schematic

SOLENOI TORQUE MOTOR CONTROL 0-100

international

ICE

R ICE DETECTOR

R WING A/I VALVE

FlightSafety

CONTROLLER POWERED, SENSOR POWERED WAI SWITCHES SELECTED TO OFF A/C IN FLIGHT ABOVE 1500' AGL, BELOW 35,000 FT.

R WING HOT (A) MAU 2 DUAL GENERIC I/O

GAP BAND No.1, No.2, No.4

WAI CONTROL

<1500 FT AUTO A/I ALT INHIBIT

180 F


WING A/I TEMP SENSOR INPUT GAP BAND No.2 (PROVIDES ACTUAL DUCT

R WING ANTI-ICE

TORQUE MOTOR



OPERATION AND INDICATIONS The wing anti-ice system receives 28-VDC power from the left and right essential DC buses through the circuit breakers on the LEER and REER circuit-breaker panels (Figure 3012). When the wing anti-ice system is in the AUTO position and ice is detected, the ice detector completes the ground to the control relay, which activates the system. Below 1,500 feet AGL after takeoff and above 35,000 feet pressure altitude, the wing anti-ice system automatic operation is inhibited. If the ice detector test is initiated while the wing anti-ice switches are selected to the AUTO position, the wing anti-ice system will be engaged for three seconds. The ice detector test overrides the altitude inhibit function. When the wing anti-ice system switches are in the ON position, the control relay is energized, and the anti-ice control valve opens, pro-

viding bleed air to the wing leading edges regardless of icing conditions. Wing anti-ice system status appears on the EI display. When a failure is detected, a message appears on the CAS display. Failures in the wing anti-ice system are recorded and can be displayed by the MDAU.

ENGINE COWL ANTI-ICE SYSTEM GENERAL The purpose of the engine cowl anti-ice system is to provide ice protection for each engine intake by distributing engine bleed air to maintain cowl inlet temperature. This can be accomplished either automatically or by manual selection (Figure 30-13

Figure 30-13. Engine Cowl Anti-Ice System


30-19



SOLENOID CONTROL VALVE DE-ENERGIZED CLOSE PRESSURE CONTROL PRESSURE

PILOT PRESSURE REGULATOR

SPRING LOAD AMBIENT PRESSURE ACTUATING PRESSURE DOWNSTREAM PRESSURE SENSE PORT HP5 AIRFLOW FROM ENGINE

TO ENGINE COWL VALVE OPEN

Figure 30-14. Cowl Anti-Ice Control—Valve Open SOLENOID CONTROL VALVE ENERGIZED CLOSE PRESSURE



VALVE CLOSED

Figure 30-15. Cowl Anti-Ice Control—Valve Closed

30-20


TO ENGINE COWL



COMPONENTS AND CONTROLS

NOTES

Cowl Anti-Ice Valves There is one cowl anti-ice valve located on each engine (Figure 30-13). Each electrically operated solenoid valve regulates the airflow to the anti-ice ducting. With power or actuating pressure from the 8th stage duct removed, the valve is spring-loaded to the fail-safe open position (Figure 30-14). The valve is closed when the solenoid is energized and air pressure is available (Figure 30-15). The regulating pressure for each valve is 25 psi (Figure 30-16). Operational status is displayed on the engine indicating (EI) display.

SOLENOID CONTROL VALVE DE-ENERGIZED CLOSE PRESSURE CONTROL PRESSURE



TO ENGINE COWL 25 PSI PRESSURE REGULATING

Figure 30-16. Cowl Anti-Ice Control—Valve Regulating


30-21



DOWNSTREAM PRESSURE SENSING LINE (TO TAI CONTROL VALVE)

FLEXIBLE TUBES

TAI DUCT TO INLET COWL CONNECTOR

SYSTEM PRESSURE SENSING OFFTAKE LINE "T" CONNECTOR

ELECTRICAL CONNECTOR

CASE VENT

PRESSURE TRANSDUCER

Figure 30-17. Pressure Transducer

30-22


TAI DUCT PRESSURE CONNECTION



Cowl Anti-Ice Control Switches

NOTES

The cowl anti-ice control switches are located on the anti-ice control panel (see Figure 302). They function in a similar manner as the wing anti-ice system control switches. The cowl anti-ice switches are three-position rotary switches which provide the crew with manual and automatic control of the cowl antiice system. In the AUTO mode, cowl anti-ice will not enable until the aircraft ascends above 1,500 feet AGL to allow for maximum takeoff performance. Also, the valve is opened and closed based on the input from the left and right ice detectors. While in the ON mode, both the control relay and the solenoid on the cowl anti-ice valve are deenergized, allowing the valve to remain open. When the switch is in the OFF position, the solenoid is energized, closing the valve.

Cowl Anti-Ice Pressure Transducers One cowl anti-ice pressure transducer (Figure 30-17) is located on each engine cowl bulkhead. The purpose of the pressure transducer is to monitor bleed-air pressure through the cowl anti-ice system by sensing the regulated pressure downstream of the cowl anti-ice control valve. The pressure transducer transmits pressure data to the MAU1 DGIO1 for the left side and MAU2 DGIO2 for the right side for fault detection and pressure display on the ECS/PRESS synoptic page. It also sends an output voltage signal to the MAUs for comparison and display on the CAS.


30-23



LIPSKIN

SPRAY RING

VEE CLAMP

INNER DUCT

EXHAUST DUCT

EXHAUST GRILL

NOSE COWL OUTER SKIN DUCT

COUPLING

NOSE COWL THERMAL ANTI-ICE VALVE

Figure 30-18. Cowl Anti-Ice Components

30-24




Cowl Anti-Ice Ducting The cowl anti-ice ducting provides a path for engine fifth-stage bleed air to be used for inlet cowl anti-icing. The engine-mounted ducting includes a venturi which limits the total airflow should any part of the ducting downstream rupture during operation. Also, piccolo assemblies, located under the lipskin of the engine inlet cowl, distribute bleed air throughout the inlet cowl leading edge (Figure 30-18).

OPERATION AND INDICATIONS The left or right (on-side) cowl anti-ice system receives 28-VDC power from the on-side essential DC bus through the cowl anti-ice circuit breaker. The cowl anti-ice pressure transducers are powered from the on-side main 28-VDC bus through the cowl anti-ice pressure transducer circuit breakers. System operation and fault data display come from hard-wired discretes to the MAUs. When the cockpit cowl anti-ice control switches are selected to the manual OFF position, the cowl anti-ice control relay and the valve solenoid are energized. With the valve position switch in the closed position, muscle air is provided to the valve. The blue “L-R Cowl AntiIce On” message on the CAS is extinguished. If the valve does not close, the switch logic sends a discrete 28-VDC signal to the MAUs, providing a fail message to the CAS. When the cowl anti-ice control switches are selected to the manual ON position, the cowl anti-ice control relay is deenergized. The valve solenoid is deenergized, and the valve position switch is in the open position. The “L-R Cowl Anti-Ice On” message appears, and a 0- to 5VDC signal is sent to the MAU1 DGIO1 for the left side and MAU2 DGIO2 for the right side. When the cockpit control switches are selected to the AUTO position, the valve solenoid is energized and deenergized dependent on an input from the ice detectors. The cowl antiice system automatic operation is inhibited below 1,500 feet AGL after initial takeoff and above 35,000 feet pressure altitude.

When icing conditions are detected and the aircraft altitude is greater than 1,500 feet AGL after initial takeoff but less than 35,000 feet pressure altitude, the following events occur: • Left and right ice detectors’ automatic cowl anti-ice detection relays are energized. • Cowl anti-ice control relays are deenergized. • Valve solenoids are deenergized, and the valve opens. • Blue “L-R Cowl Anti-Ice On” message appears on the CAS. • Fifth-stage engine bleed air is supplied to the engine cowls, and the cowl antiice pressure is indicated on the ECS/PRESS synoptic page. The cowl anti-ice AUTO operation is inhibited below 1,500 feet AGL after takeoff and above 35,000 feet pressure altitude. When the cockpit anti-ice control switches are in the AUTO mode and icing conditions are detected, the automatic anti-ice altitude inhibit relay is energized. The cowl anti-ice control relay and valve solenoid are also energized, closing the anti-ice valve. Like the wing anti-ice system, when the cowl anti-ice switches are set in the AUTO position during the ice detector test, the cowl anti-ice system is engaged for three seconds. The cowl anti-ice system status appears on both the engine indicating display and the crew alerting system display. The MAUs compare cowl anti-ice pressure inputs from the pressure transducers. If the pressure difference between the two systems is greater than 7 psi, a blue “Cowl Anti-Ice Miscompare” message is displayed on the CAS. When the air pressure is greater than 33 psi, an amber” L-R Cowl Anti-Ice High” message is displayed.


30-25



TOTAL AIR TEMPERATURE PROBE

ANGLE OF ATTACK PROBE

PITOT PROBE

Figure 30-19. Probe Anti-Ice System

30-26




When the cowl anti-ice system is in the ON or AUTO mode and the valve fails to open within 15 seconds (as timed by the on-side DAU), an amber “L-R Cowl Valve Failed Closed” message is displayed on the CAS.

NOTES

When the cowl anti-ice system is turned OFF and the anti-ice valve fails to close within 15 seconds, an amber “L-R Cowl Valve Failed Open” message is displayed on the CAS. Also, a pressure indication is displayed on the ECS/PRESS synoptic page.

PROBE ANTI-ICE SYSTEM GENERAL The probe anti-icing system prevents the icing of pitot-static probes, total-air-temperature (TAT) probes, and angle-of-attack (AOA) probes (Figure 30-19).

COMPONENTS AND CONTROLS Pitot-Static Probe Heaters There are four pitot-static probe heaters: two upper probes and two lower probes. To prevent icing, heater power is applied whenever the bus is powered and the controls are selected to ON. The upper left and lower right pitot-static probes are dedicated as the standby pitot-static probes for the standby airspeed/altimeter indicator unit. The four pitot-static probe heaters receive 115-VAC power from the following sources (Figures 30-20 and 30-21): • Upper left—Left standby AC bus • Upper right—Essential AC bus • Lower left—Essential AC bus • Lower right—Right standby AC bus


30-27

30-28 R UPPER PITOT HEATER ANNUNCIATOR LTS DIM AND TEST CONTROLLER

MAU 2 DGI/0 ADM 2–SBY PITOT FAIL X1 X2

DIM TEST

TRIPLE CHANNEL CURRENT SENSOR NO. 1

ON

R UPPER PITOT HEATER RTN R UPPER PITOT HEATER RTN (IN) R UPPER PITOT HEAT FAIL (OUT) +28 VDC +28 VDC RTN

R UPPER OFF PITOT HEAT PWR RLY


TAT NO. 1 HEAT CTRL LEFT ESS 28-VDC BUS

AIR

MAU 1 DGI/0 TAT PROB HT FAIL X1 X2

GND WOW RELAY

TAT PROBE NO. 1 HEATER POWER L STBY 115 VAC ØC

TAT NO. 1 PROBE HEATER

L UPPER PITOT HEAT POWER

CHASSIS GND

TAT NO. 1 OFF PROBE HEAT PWR RLY

GROUNDED BYPASS SWITCH (SWITCH GUARD TURNS SWITCH OFF) ON

OFF

TAT NO. 1 HEATER RTN TAT NO. 1 HEATER RTN (IN)

L UPPER PITOT HEATER MAU 1 DGI/0 ADM 1–3 PITOT FAIL X1 X2

L UPPER PITOT HEATER RTN L UPPER PITOT HEATER RTN (IN) L UPPER PITOT HEAT FAIL (OUT)

ON

L UPPER OFF PITOT HEAT PWR RLY

international

FlightSafety

L STBY 115 VAC ØA

ON

TAT NO. 1 HEAT FAIL (OUT)


R UPPER PITOT HEAT PWR ESS 115 VAC ØA UPPER PROBE ANTI-ICE HTR SWITCH TO ANNUNCIATOR LTS OFF DIM AND TEST

Figure 30-20. Upper Pitot-Static and No. 1 TAT Probe Schematic



Total Air Temperature (TAT) Probe Heaters

NOTES

The total air temperature (TAT) probe heaters are located on the lower left and right sides of the forward fuselage. Heater elements are an integral part of the TAT probes and receive 115VAC power from the left and right standby AC buses. They are normally powered when the upper and lower probe switches are selected to ON and the aircraft has weight off wheels (Figure 30-20). A TAT probe ground bypass switch provides the capability to power the TAT probe heating elements on the ground. The ground bypass switch is guarded by a red cover to the OFF position and provides for troubleshooting or deicing of the TAT probes while the aircraft is on the ground. The anti-ice upper and lower probe switches must be selected to ON in order to activate the TAT probes.

Current Monitors Two triple-channel current monitors are located in the forward nose compartment. The No. 1 current monitor (left side) monitors the left and right upper pitot-static probe heaters and the No. 1 TAT probe heater. The No. 2 current monitor (right side) monitors the left and right lower pitot-static probe heaters and the No. 2 TAT probe heater. Each triple-channel current monitor senses the current on the AC power ground return of each probe heater AC power circuit. When the heaters are operating correctly, the current monitors will sense greater than 0.5 amp. They provide discrete outputs to the MAU1 DGIO1 for the left upper and lower pitot-static probes, along with the left TAT probe, and MAU2 DGIO2 for the right upper and lower pitotstatic probes, along with the right TAT probe for fault indication when the heater return AC power is less than 0.5 amp and the controls are turned on.


30-29



ANGLE OF ATTACK PROBE

PITOT PROBE

Figure 30-21. AOA and Pitot Probes

WSHLD BLWR ON

LEFT

OFF

AOA

ANTI ICE HTR RIGHT

UPPER

OFF

OFF

PROBE

LOWER

OFF

Figure 30-22. Probe Heater Control Panel

30-30




Angle-of-Attack (AOA) Probe Heaters An angle-of-attack (AOA) probe heater is located on each side of the forward fuselage (Figure 30-21). Each probe incorporates a case heater and a vane heater as an integral part of the probe. The heaters prevent icing of the AOA vanes and cases. The left (No. 1) and right (No. 2) AOA probe heaters are powered from the on-side 28-VDC essential bus.

The left and right AOA heater switches control power to the angle-of-attack heater elements. The AOA probes differ by having their own internal overtemperature sensing devices and do not require external switching through relays. Both sets of pushbutton switches are illuminated/extended when they are in the amber OFF position and extinguished/depressed when in the ON position.

NOTES

Each AOA probe contains internal current monitors that provide heating element status to the on-side DAU over an ARINC 429 data bus. The AOA probes differ from the pitot and TAT probes by having their own internal overtemperature sensing devices and do not require external switching through relays. Loss of heating power or the ARINC 429 data bus causes an amber “AOA Probe 1-2 Fail” message to appear on the CAS.

Probe Heater Control Panel The probe heater control panel (Figure 3022) is located on the cockpit overhead panel. The probe heater pushbutton switches provide power to the pitot-static, TAT, and AOA probe heater elements. The upper and lower probe heat switches control power to the pitot-static and TAT probe heaters. The upper probe switch controls power to the upper left and right pitot-static and No. 1 (left) TAT probe heaters. The lower probe switch controls power to the lower left and right pitot-static and No. 2 (right) TAT probe heaters. Control power for each probe is provided by the on-side 28-VDC bus. When the probe heat pushbutton switch is deactivated, 28-VDC bused power is switched from the dedicated current sensor to the dedicated power relay coil, energizing the relay, pulling the contacts open, and removing the 115-VAC power source from the heating elements.


30-31

30-32

ï Left AOA Probe Heat Control Switch Selected ON ï Control Relay De-Energized ï Heat Power Applied

ï Right AOA Probe Heat Control Switch Selected ON ï Control Relay Energized ï Heat Power Removed

#1 AOA HTR

#2 AOA HTR R ESS 28VDC

ANNUN LTS DIM & TEST PWR FOR TRAINING PURPOSES ONLY

L AOA ANTI-ICE HTR SW

R AOA ANTI-ICE HTR SW

OFF

OFF

OFF

OFF

ON

ON

L AOA ANTI-ICE CNTRL RLY ON

ANNUNCIATOR LTS DIM/TEST CONTROLLER GND TRIGGER GND TRIGGER

R AOA ANTI-ICE CNTRL RLY ON

OFF

OFF

ARINC 429

R AOA HTR/SENSOR

CASE HEATER

CASE HEATER

VANE HEATER

VANE HEATER

MAU 2 ARINC 429

DUAL GENERIC I/O MODULE SLOT 7/8 AOA PROBE 2 FL (AMBER)

international

FlightSafety

L AOA HTR/SENSOR

MAU 1 DUAL GENERIC I/O MODULE SLOT 9/10 AOA PROBE 1 FL (AMBER)

ANNUN LTS DIM & TEST PWR


L ESS 28VDC

Figure 30-23. AOA Probe Heat Schematic



OPERATION AND INDICATIONS

NOTES

The probe anti-ice switches control power to the heat power relays (Figure 30-23). When selected OFF, the control relays are energized. Power to the heating elements and the triplechannel current monitor is removed, and any fault messages are inhibited. If 28-VDC control power is lost or a relay fails in the deenergized position, the system defaults to the ON position and cannot be turned off during normal operation unless the circuit breaker for the affected probes is opened. The TAT probe heaters are inhibited on the ground when the probe heat switches are selected ON. Selecting the TAT probe ground bypass switch and the system monitor test panel switch to the ON position deenergizes the TAT probe’s heat power relays and provides power to the TAT probe’s heating element. By selecting the AOA probe heater switches to ON, the heater power control relay is deenergized, and the probe heating elements are powered. When the switches are selected OFF, the heater power control is energized. During operation, the current monitor senses induced loads from the probe heating element returns, and if an element or other part of the system malfunctions, amber caution messages are displayed on the CAS. CAS messages are not displayed when the associated control switch is selected to OFF and the heater power relay is energized.


30-33

30-34 W/S BLWR FAN RELAY

VANE-AXIAL FAN

WSHLD BLWR LEER—F5

L MAIN AC BUS ØB 115/200 VAC 400 HZ ØC


10 (RADOME)

W/S BLWR ACT. RELAY BLOWER NOZZLE ACTUATOR

W/S BLWR CONT SW ON OFF

CWOW RELAY NO. 9 AIR

CLSD

E M OPEN I

ON

(COP) (RADOME)

GRND

(345 PANEL)

(345 PANEL)

M

international

FlightSafety

GROUND SWITCH SELECTED ON

WSHLD BLWR CONT LEER—E6 L MAIN +28 VDC 5


ØA

Figure 30-24. Windshield Rain Removal Schematic



WINDSHIELD/WINDOW ICE AND RAIN SYSTEM GENERAL The windshield and window ice and rain protection systems consist of three major subsystems: the windshield blower system, the windshield heat system, and the cabin window heat system. The purpose of the systems is to provide windshield/window heat and to remove moisture from the pilot’s and copilot’s windshields during inclement weather (Figure 30-24).

WINDSHIELD BLOWER SYSTEM Components and Controls The windshield blower switch is located on the cockpit overhead panel. It is a two-position ON–OFF switchlight which provides operational control and illuminates a blue “ON” when selected on the ground. In addition to the control switch, the system includes a vane-axial fan, ducting, two hinged nozzles, and an actuator that retracts or extends the nozzles through teleflex cables (Figure 30-25). Relay logic ensures that the fan can be operated only with weight on wheels in the GROUND mode. If the system is left on and the aircraft becomes airborne, the system automatically shuts down, and the nozzles retract.

WINDSHIELD BLOWER ASSEMBLY

WINDSHIELD BLOWER NOZZLES

Figure 30-25. Windshield Blower System


30-35



RIGHT FRONT LEFT SIDE WINDSHIELD HEAT CONTROL UNIT

LEFT FRONT RIDE SIDE WINDSHIELD HEAT CONTROL UNIT

Figure 30-26. Windshield Heat Control Units

30-36





Windshield Heat Control Units

The windshield blower fan is powered by the left main AC bus, and the nozzle actuator is powered by the left main DC bus (see Figure 30-24). When the control switch is pressed, the blue “ON” illuminates, the actuator extends the left and right blower nozzles, and the fan starts.

There are windshield heat control units for the left front/right side window (LF/RS) and for the right front/left side window (RF/LS). Both windshield heat control units are located forward and inboard of the copilot’s rudder pedals (Figure 30-26).

The nozzles direct high-velocity air to the bottom of each windshield. The high-velocity airstream removes beaded water droplets from the windshield and provides unobstructed visibility to the crewmembers during ground operation.

Each windshield heat control unit contains two circuit card assemblies: one circuit card for the windshield and one for the side window. Each circuit card controls the current to the heater element and monitors the resistance through one of two sensors mounted on the window. At initial power-up, the windshield heat control unit automatically performs a self-test of each circuit.

WINDSHIELD HEAT SYSTEM General The purpose of the windshield heat system is to provide heater power to the windshields and side windows for deicing and defogging.

Components and Controls The windshield heat control panel is located in the lower right corner of the overhead control panel. The control switches apply power to the heater control units and the heater power relays.


Each pushbutton indicator controls power to one of the two independent windshield heat systems. They provide control for the left front/right side (LF/RS) and the right front/left side (RF/LS) windows. Indicator lamps provide windshield heater power and fault indications for their respective windows to the flight crew.

The windshield heat control unit incorporates red LEDs for each of the four control circuits, which illuminate when a fault is detected, as follows: • Film range—Indicates that the heating element is above the acceptable tolerance: 22 ±1 ohms for the windshield and 100 ±5 ohms for the side window. • Command error—Indicates the absence of current detected when heat is commanded, current detected when no heat is commanded, or an overcurrent is detected. • Sensor error—Indicates a fault with the selected sensor on the window/windshield application, causing the windshield heat control unit to send a ground fault discrete to the MAU1 DGIO1 (LF/RS) or MAU2 DGIO2 (RF/LS), the CAS to display a blue L-R WSHLD FAULT message, and the windshield heat control unit to automatically switch to the other sensor. If both sensors fail, an amber L-R WSHLD FL message is displayed on the CAS. • Overtemperature—Indicates that a heating element overtemperature condition has been detected by the windshield heat control unit (143 ±4°F).


30-37

30-38 ANNUN LTS DIM AND TEST PWR

L FRONT WSHILD

OFF OFF LEFT FRONT, RIGHT SIDE WSHILD HEAT CONTROL UNIT

LF W/S HT CONT RLY ON

R SIDE WSHILD


L MAIN 115 VAC PH A RS W/S HT CONT RLY ANNUN LTS DIM AND TEST PWR

OFF

R FRONT WSHILD

OFF

R MAIN 115 VAC PH B

ON

R MAIN 115 VAC PH C

R FRONT HEATING INDICATOR L SIDE HEATING INDICATOR

R F WSHILD SENSOR FL T(B) R F WSHILD SENSOR FL (A) L S WSHILD SENSOR FL T(B) L S WSHILD SENSOR FL (A)

MOV

RIGNT FRONT WINDSHIELD

ANNUN LIGHTS DIM AND TEST

RIGHT FRONT, LEFT SIDE WSHILD HEAT CONTROL UNIT

MOV

SENSOR FAULT HEATER POWER

SENSOR SELECT

HEATER POWER SENSORS IN

WINDSHIELD HEAT CONTROL SYSTEM

Figure 30-27. Windshield Heat Sensor Schematic

HEAT SENSOR

RIGHT SIDE WINDOW

international

HEAT SELECTED ON • WINDSHIELD • HEATER POWER BEING APPLIED

MOV

POWER

FlightSafety

SENSOR POWER FAULT ON IND BOTH SENSORS FAIL

LS W/S HT CONT RLY

MOV

HEAT SENSOR POWER

115 VAC R MAIN 115 VAC PH C

POWER P STATIC GRND

P STATIC GRND

SENSORS IN BOTH SENSORS FAIL POWER ON IND RF W/S HT CONT RLY

LEFT SIDE WINDOW

HEAT SENSOR

208 VAC

L SIDE PH A WSHILD

POWER HEAT SENSOR

LEFT FRONT WINDSHIELD

L FRONT HEATING INDICATOR R SIDE HEATING INDICATOR

L F WSHILD SENSOR FL T(B) L F WSHILD SENSOR FL (A) R S WSHILD SENSOR FL T(B) R S WSHILD SENSOR FL (A)

SENSOR FAULT BOTH SENSORS FAIL SENSORS IN 208 VAC HEATER POWER POWER ON IND SENSOR SENSOR FAULT SELECT 115 VAC BOTH SENSORS FAIL HEATER POWER POWER SENSORS ON IND IN

L FRONT WSHILD



L MAIN 115 VAC PH C




Windshield Heater Sensor Elements The windshield heater/sensor elements are an integral part of the left and right windshields and side windows (Figure 30-27). The heater elements provide the deicing heat to the windshields and side windows. Power to the heater elements is controlled by the heater control unit through the power relays and the metal oxide varistors (MOVs). MOVs are used to protect the window heating control units from power spikes due to static electrical buildup by providing an electrical short to ground. The sensor elements provide temperature inputs to the heater control unit in the form of a variable line resistance.

The following noncritical conditions will cause the power-on light to flash at a 3-Hz rate for 90 seconds: • Wi n d o w f i l m normal range

resistance

above

• P-static protection circuit damage detected • Single sensor failure of the nonselected sensor During all other faults, the controller removes power from the affected window, and the indicator flashes at a 1-Hz rate for 90 seconds. These fault indications are also displayed on the LEDs of the windshield heat control units. Also, windshield heat faults appear on the CAS display and are transmitted to the CMC.


NOTES

The windshield heat system is powered by the left and right main AC buses (LF/RS by left main AC bus and RF/LS by right main AC bus). The windshield heat system is controlled by the switches on the cockpit overhead panel. The LF/RS and RF/LS switchlights illuminate amber “OFF” when selected off and are extinguished when selected on. A blue annunciation appears on the appropriate power indicator light when power is being applied to the window or windshield. When the windshield heat system is turned on, the switchlight extinguishes, the control relays are energized, and the controller is powered. At initial power-up, the windshield heat control unit performs a self-test. If the selftest is passed, the front windshield is then powered by ramping from 0 to 100% in four minutes. However, full power is applied to the side window as soon as the controller completes its BIT check. The windshield heater control unit applies power to the heater when the sensor element resistance falls below 326 ohms or approximately 104°F. The annunciator lights illuminate while power is being applied. The windshield heater control unit removes power from the heater when the sensor element resistance reaches 334 ohms or approximately 114°F.


30-39




Figure 30-28. Cabin Window Heat Switch

30-40




CABIN WINDOW HEAT SYSTEM General The cabin window heat system electrically defogs each of the cabin windows when the aircraft is airborne. There are five fixed cabin windows and two emergency windows on each side of the aircraft.

Components and Controls Cabin Window Heat Switch The cabin window heat switch, labeled “CABIN WDO HT,” is located on the cockpit overhead panel (Figure 30-28). It provides the crew the ability to activate the cabin window heat system. The control switch is a two-position ON–OFF switch. When in the OFF position, a blue OFF legend is illuminated in the switch. With the switch selected to ON, the blue OFF legend extinguishes.

Window Heat Ground Bypass Switch The window heat ground bypass switch is located on the systems monitor test panel (see Figure 30-29). The ground bypass switch, in conjunction with the cabin window heat switch, provides the ability to perform ground checks of the cabin window heat system.

Figure 30-29. Window Heat Ground Bypass Switch


30-41



EMERGENCY EXIT HANDLE SWITCH

Figure 30-30. Emergency Exit Window

30-42




Window Heat Control Relays There is one window heat control relay for each side of the aircraft. They are located in the left and right junction and relay panels forward of the left and right EERs. The relays supply 115VAC power to the window heating elements when the cabin window heat switch is selected to ON. The relays are enabled only when weight on wheels is removed (aircraft is airborne). The relays can also be enabled when the cabin window heat bypass switch is selected to ON to perform ground checks.

Cabin Emergency Exit Window Handle Switches The cabin emergency exit window handle switches (Figure 30-30) are located on the sixth and seventh windows on both the left and right sides of the aircraft. When the handles are pulled to unlock the exit window, the

ACOUSTICAL LAYER (INSIDE PANEL)

LAMINATED HEATER FILM

switch removes the 115-VAC heater power from the windows. The switch also provides a ground to the MAUs to indicate that the window is unlocked. An amber “Cabin Window Unlock” message then appears on the CAS display. Each cabin window has an electrically heated acoustical window pane (Figure 30-31). The panes incorporate a heater film sandwiched between the inner and outer plies. The inner panes can be easily removed from the inside of the aircraft without removing the outer window. The eight fixed panes are connected to the 115-VAC heater power via standard barrier-type terminal blocks. The four emergency exit windows incorporate disconnect electrical window contacts that facilitate easy rem ova l o f t h e w i n d ow i n t h e eve n t o f a n emergency.

STRUCTURAL LAYERS (OUTER PANE)

Figure 30-31. Cabin Window Diagram


30-43

30-44 AIRCRAFT IN FLIGHT SYSTEM SELECTED ON CABIN WDO HT (B) R FWD WDO HEAT CONT

EMERG EXIT

ANN. LIGHTS DIM & TEST CONTROLLER

OFF

RIGHT NO. 1

DIM TEST

OFF

RIGHT NO. 2

RIGHT NO. 3

RIGHT NO. 4

RIGHT NO. 5

RIGHT NO. 6 WINDOW HANDLE SWITCH

0/GND 0/28V

LINK L WDO HEAT CONT

1

LK

LINK

2

ON

L MAIN 28 VAC

RIGHT NO. 7 WINDOW HANDLE SWITCH

R WDO HT CONT RELAY

LOC. COP

2

LINK

LINK LK

OFF

1

LK

LK

R WDO HEAT 2&5


R MAIN 115 VAC PH A R WDO HEAT 3&6 R MAIN 115 VAC PH B R MAIN 115 VAC PH C

R WDO HEAT 4&7

FWD WDO HEAT R&L

R MAIN 115 VAC PH C

ON AIR

GND P/O CWOW RELAY NO. 10

OFF

ON ON

WINDOW HANDLE SWITCH

R/L FWD WDO HT CONT RELAY

L WDO HEAT 4&7

OFF CABIN WDO HTRS WOW BYPASS SW

LINK

WINDOW HANDLE SWITCH

LK

LINK

LK

OFF

2

2

LINK

L MAIN 115 VAC PH C

LINK LK

LK

L WDO HEAT 3&6 L MAIN 115 VAC PH B

1 LEFT NO. 1

L WDO HEAT 2&5

L MAIN 115 VAC PH A

LEFT NO. 2

LEFT NO. 3

LEFT NO. 4

ON AIR

P/O CWOW RELAY NO. 23

L WDO HT CONT RLY

MAU 1

MAU 2

SINGLE GENERIC I/O MODULE (1) SLOT 3 CABIN WDO UNLOCKED (CAS MSG)

SINGLE GENERIC I/O MODULE (4) SLOT 12 CABIN WDO UNLOCKED (CAS MSG)

GND/OPEN

GND/OPEN

LEFT NO. 6

1 LEFT NO. 7

EMERG EMERG EXIT EXIT CONNECTORS ARE BREAKAWAY TYPE 1 FOR EMERGENCY EXIT WINDOWS 2 WINDOW HANDLE SWITCH PLUNGER IS ACTUATED WHEN THE WINDOW IS INSTALLED AND LOCKED IN PLACE

international

FlightSafety

GND

LEFT NO. 5


R MAIN 28 VAC

EMERG EXIT

Figure 30-32. Cabin Window Heat System Schematic




NOTES

The cabin window heater film is powered by the left and right main AC buses (Figure 3032). The cabin window heater control is powered by the left and right main DC buses. The heater control relays are enabled when weight on wheels is removed or when the ground bypass switch is in the ON position for ground checks. Power is applied when the cabin window heat control switch is selected to ON, extinguishing the amber OFF switch legend. The control relays apply the heater power to the electrically heated panes through the power fuses. If an emergency exit window handle is pulled to unlock the exit window, the window handle switch removes 115-VAC power from the breakaway electrical contacts to the heater film. The switch also supplies a ground to the MAUs, which generate a “Cabin Window Unlock” message on the CAS display.


30-45



CHAPTER 31 INDICATING AND RECORDING SYSTEMS CONTENTS Page INTRODUCTION ................................................................................................................. 31-1 GENERAL ............................................................................................................................ 31-3 CMC COMPONENT DESCRIPTIONS ............................................................................... 31-5 Central Maintenance Computer Module........................................................................ 31-5 Left Emergency Power Battery Pack ............................................................................. 31-7 Cockpit Displays............................................................................................................ 31-9 Cursor Control Device (CCD) ..................................................................................... 31-10 Ethernet Ports .............................................................................................................. 31-11 FDR/CMC Event Switch ............................................................................................. 31-12 Remote Terminal ......................................................................................................... 31-13 Data Management Unit................................................................................................ 31-14 Cockpit Printer............................................................................................................. 31-15 Database Module ......................................................................................................... 31-16 Member Systems ......................................................................................................... 31-17 Ground Maintenance Test Switch ....................................................................................... 31-19 CMC SYSTEM OPERATION ............................................................................................ 31-21 CMC Control and Display........................................................................................... 31-21 CMC Communication.................................................................................................. 31-48 Remote Terminal Requirements .................................................................................. 31-59


31-i




Title

Page

31-1

Central Maintenance Computer Interface .............................................................. 31-2

31-2

Central Maintenance Computer Module................................................................ 31-4

31-3

Left Emergency Power Battery Pack ..................................................................... 31-6

31-4

Cockpit Displays .................................................................................................... 31-8

31-5

Cursor Control Device ......................................................................................... 31-10

31-6

Ethernet Ports....................................................................................................... 31-11

31-7

FDR/CMC Event Switch ..................................................................................... 31-12

31-8

Remote Terminal.................................................................................................. 31-13

31-9

Data Management Unit ........................................................................................ 31-14

31-10

Cockpit Printer ..................................................................................................... 31-15

31-11

Database Module ................................................................................................. 31-16

31-12

Loadable Diagnostics Information Database ....................................................... 31-17

31-13

Ground Maintenance Test Switch ........................................................................ 31-18

31-14

CMC Control and Display ................................................................................... 31-20

31-15

CMC Control Using the CCD.............................................................................. 31-22

31-16

CMC Main Menu Display ................................................................................... 31-24

31-17

CMC Maintenance Messages Display ................................................................. 31-26

31-18

CMC Active Maintenance Messages Display ..................................................... 31-28

31-19

CMC Present Log Maintenance Messages Display............................................. 31-30

31-20

CMC Historical by Date Maintenance Messages Display................................... 31-32

31-21

CMC System Diagnostics Menu Display ............................................................ 31-34

31-22

CMC Select a System Menu Display................................................................... 31-35

31-23

CMC Rigging Pages ............................................................................................ 31-36


31-iii



31-24

System Test Pages................................................................................................ 31-37

31-25

CMC Extended Maintenance Menu Display ....................................................... 31-38

31-26

Member System Status ........................................................................................ 31-39

31-27

CMC File Transfer Menu Display ....................................................................... 31-40

31-28

CMC Configuration Menu Display ..................................................................... 31-42

31-29

CMC System Configuration Pages ...................................................................... 31-43

31-30

CMC Reports Menu Display ............................................................................... 31-44

31-31

CMC Reports Pages ............................................................................................. 31-45

31-32

CMC Menu Displays (Air Mode) ....................................................................... 31-46

31-33

CMC Communication Diagram........................................................................... 31-48

31-34

CMC Flight Deck Effect and Maintenance Messages Correlation...................... 31-49

31-35

CMC CAS Messages ........................................................................................... 31-50

31-36

CMC Parameter Page Viewing ............................................................................ 31-52

31-37

Aircraft Condition Monitoring Function ............................................................. 31-53

31-38

ACMF Reports Viewing ...................................................................................... 31-54

31-39

Communication Management Displays............................................................... 31-56

31-40

Remote Terminal Setup Displays......................................................................... 31-58

31-41

Remote Terminal Connection Settings ................................................................ 31-61

31-42

Edit ACMF Data Display..................................................................................... 31-63

31-iv




CHAPTER 31 INDICATING AND RECORDING SYSTEMS

INTRODUCTION The central maintenance computer (CMC) is an aircraft diagnostic and maintenance system (ADMS). It is a centralized interface that monitors the data reported by aircraft member systems. The member systems continuously transmit and report fault information to the CMC via the MAU virtual backplane. The central maintenance computer is a data recorder only, that is utilized by technicians to perform most maintenance activities on the aircraft. The CMC performs the following functions: • Maintenance Message Display • System Diagnostic Display • System Configuration Display

• Flight Deck Effect to Maintenance Message Correlation • Fault Processing • Aircraft Condition Monitoring Function

• File Transfer Display • Member System Status Display (Parameters)

• Initiating Member System Tests • Performing System Rigging Function (VGS)


31-1



MODULAR AVIONICS UNITS (MAU 1) DATABASE MODULE (DBM)

TLD Data

FHDB

ARINC Member Systems

ARINC 429 I/O MODULE (running Digital Engine Operating System)

Cockpit Event Button

I/O Software

Software Services (Core S/W, PDD)

I/O Process

Remote Terminal (Laptop)

RS-422 Data Bus RIB

ADVANCED GRAPHICS MODULE (AGM)

CCD Data Management Unit

VIDEO MODULE

Remote Image Bus

VIRTUAL BACKPLANE BUS (PCI)

Main Test Switch

CMC MODULE (running COTS Operating System) TLD Data

Backplane Interface Controller Central Maintenance Software Fault Processing I/O Processing CMCF Utilities

Aircraft FHDB

Aircraft Condition Monitoring

Cockpit Printer

ACMF Data LDI Databases

Left E-Batt Other MAUs and MRCs Frame Buffer

Network Interface Controller (NIC)

Ethernet LAN ASCB-D

Figure 31-1. Central Maintenance Computer Interface

31-2




GENERAL

NOTES

Software contained on the CMC module can be categorized into two categories. These are the central maintenance function and the aircraft condition monitor function. The central ,maintenance function (CMF) collects fault information and stores it in the fault history database. This data can be viewed and retrieved by aircraft technicians when needed. The aircraft condition monitor function (ACMF) operates in parallel with the CMF and it monitors defined parameters on ASCB. It then collects that information and evaluates it based on defined trigger expressions. Once a trigger expression becomes true, the ACMF will create a report of the condition. The CMC Function is comprised of the following system components (hardware and software): • Central Maintenance Computer (CMC) Module • Left Emergency Power Battery Pack • Cockpit Displays • Cursor Control Device (CCD) • Ethernet Ports • FDR/CMC Event Switch • Remote Terminal Platform (Laptop with RT Software) • Data Management Unit • Cockpit Printer • Database Module (DBM) • Loadable Diagnostic Information (LDI) Database • Member Systems • Ground Maintenance Test Switch


31-3



Figure 31-2. Central Maintenance Computer Module

31-4




CMC COMPONENT DESCRIPTIONS

NOTES

CENTRAL MAINTENANCE COMPUTER MODULE The central maintenance computer module is the “home” of the maintenance function. The CMC module is located in the LEER in MAU No. 1 Slot 11. It hosts the central maintenance function (CMF) and the aircraft condition monitoring function (ACMF) software applications. The CMC Module provides fault processing capabilities and generates ACMF reports. It also stores the loadable diagnostic information (LDI) database, generated ACMF reports, and the fault history database. Storage capabilities are provided by a memory disk contained on the module itself.


31-5



Figure 31-3. Left Emergency Power Battery Pack

31-6




LEFT EMERGENCY POWER BATTERY PACK

NOTES

The Left E-Batt is located in the LEER. It provides the CMC Module 24 VDC power through the CMC module front connector. The CMC will utilize the E-Batt to power down the CMC module operating system. It will need to have this backup E-Batt power for 2 minutes during the shutdown cycle.


31-7



DU1

1155A1A11

CMC NO.1 OUTPUT

155A3A9

CMC NO 1 INPUT DU NO 2 SELECTED VIDEO OUTPUT

DU2

PRIMARY FLIGHT DISPLAY

I-NAV

MAU 3 AGM 1

MAU 3 AGM 2

DU3 PRI FLT INST SEC ENG INST

CREW ALRT SYS

DU4

I-NAV

PRIMARY FLIGHT DISPLAY

MAU 2 AGM 3

MAU 3 AGM 1

DU NO 1 SELECTED VIDEO OUTPUT VGM (1) CONTROL I/O (3)

155A2A5

155A3A14

AGM 3

AGM 1 155A3A5

AGM 2 MAU NO. 3

155A2A15 331A1P31

DU NO 3 SELECTED VIDEO OUTPUT

332A1P31

CMC NO.2 OUTPUT

CMC NO 2 INPUT

CMC

331A1P30

332A1P30

155A1A14

DU NO 4 SELECTED VIDEO OUTPUT VGM (2) CONTROL I/O (2)

AGM 4

MAU NO. 2

MAU NO. 1

LEFT ELECTRONIC EQUIPMENT RACK (LEER)

RIGHT ELECTRONIC EQUIPMENT RACK (REER) REMOTE IMAGE BUS FIBER CHANNEL

Figure 31-4. Cockpit Displays

31-8




COCKPIT DISPLAYS

NOTES

CMC information can viewed by maintenance crews on display units 2 and 3 as a 2/3 synoptic page or display units 1 and 4 on aircraft battery power. All four units are located on the flight panel in the cockpit. Once the CMC has been selected for display the CCD is utilized to manipulate the displayed information. Video data is supplied to the display system via the two RIB (remote image bus) outputs from the CMC module. The remote image bus runs from the CMC module to two control I/Os with video and then it is distributed to the four advanced graphic modules


31-9



Cursor Control Device (CCD) The pilot’s and copilot’s CCDs are required to manipulate the menus generated by the CMC. The CCDs are located on the left and right side panels in the cockpit. This interface is provided by the RS-422 data connection from the CCD to the AGM.

The CCD is comprised of: •The cursor positioning device, which controls the menu choice selections •The rotary dial, which controls the main menu selections (moving the green bar) •The “enter” button, which selects the highlighted function

Figure 31-5. Cursor Control Device

31-10




ETHERNET PORTS Two ethernet ports are located on the system monitor test panel which is mounted on the forward side of the REER above the observer seat. These two ports are used to connect the remote terminal (laptop) to the central maintenance computer via the local area network

(LAN). The local area network is used to transmit electronic terminal charts to the advanced graphics module, and is used for all aircraft software data loading. It is very important to maintain the integrity of the LAN throughout the modular avionics units (MAUs), data management unit, cockpit printer, central maintenance computer, and modular radio cabinets.

Figure 31-6. Ethernet Ports


31-11



FDR/CMC EVENT SWITCH The FDR/CMC event switch is located on the cockpit overhead panel. When selected, the CMC will create an ACMF report based on predefined parameters. The triggers and report parameters are defined by Gulfstream Engineering.

The ACMF report will consist of a 200-second “time-series” recording; 100 seconds before the event and 100 seconds after the event. When the recording is taking place, a blue advisory CAS message will be displayed. This message is “Event Record”.

Figure 31-7. FDR/CMC Event Switch

31-12




REMOTE TERMINAL The remote terminal (RT) platform is a laptop with a local area network (LAN) ethernet 10base2 connection capability and a CD-ROM drive hosting the Honeywell-provided remote terminal software. The RT platform allows the technician access to the CMC without being in the cockpit. To accomplish this the technician will need to have the remote terminal platform, a Linksys combo ethernet PC card, coax cable, and a

BNC T connector with a 50-ohm male terminator. These are supplied with aircraft delivery and operationally will be discussed later in the lesson. The CMF can support up to 3 simultaneous RT connections (including the cockpit function running). With the exception of initiated ground tests, the RTs shall run independently. Due to the communication protocol structure, only one RT can command a ground test or rigging at one time.

Figure 31-8. Remote Terminal


31-13



DATA MANAGEMENT UNIT The data management unit (DMU) is located in the cockpit on the pilot’s side console. It is mounted on the wall of the panel and not on top as in other aircraft typical installations. Care should be taken when opening the DMU because it could be damaged if the pilot’s seat is moved backward during a DMU loading procedure. The data management unit (DMU) provides the CMC with a LAN connection to a DVD-ROM drive and a PCMCIA data storage card. The data management unit’s PCMCIA slot 2 (bot-

tom slot) can be used to transfer the fault history database reports and the aircraft condition monitoring function reports from the CMC memory to a PCMCIA card for viewing on a personal computer. There are lights next to the PCMCIA slots on the DMU. When the DMU door is opened, the lights will turn amber indicating that the PCMCIA card is shutting down. The card should not be removed until this process is complete and the light turns green. The PCMCIA card will not power back up until the DMU door is closed and secure.

(MOUNTED ON WALL, NOT ON TOP) SLOT 2

Figure 31-9. Data Management Unit

31-14




COCKPIT PRINTER

These reports can be sent to the cockpit printer for printing:

The cockpit printer is located on the copilot’s side console. It is connected to the LAN and can be utilized to print various reports as commanded through the CMC menu interface.

•Active Maintenance Messages •Stored Maintenance Messages by ATA •Stored Maintenance Messages by date

NOTE

•System Configuration Data

The cockpit printer is a thermal printer. The paper width is 8.5”. The paper is on a roll stored within the printer.

Figure 31-10. Cockpit Printer


31-15



DATABASE MODULE Database module No. 2 in MAU No. 1 Slot 16 is utilized by the CMC function. The database module provides the CMC with 120 MB of storage, accessible via file transfer protocol (FTP) services. The CMC uses this storage space primarily for redundancy. It stores a backup of the fault history database in this area. This occurs in two phases: ground to air and air to ground transitions. If the CMC FHDB is lost or corrupt, it will download the most recent copy from the database module.

Loadable Diagnostic Information Database (LDI) The loadable diagnostic information (LDI) database contains a model of the aircraft's onboard systems. Gulfstream controls the LDI database and provides it to Honeywell as a separate database file. Honeywell adds their portion of the diagnostic information (Planeview Maintenance). Once this is accomplished, all of the separate database files shall then be combined into one LDI database. The LDI database is not part of the CMC executable code and is maintained as a separate

Figure 31-11. Database Module

31-16




file. This design approach provides a means for the aircraft manufacturer to add or remove member systems from the Planeview CMC without incurring the cost of software modifications. This also allows Gulfstream to update the maintenance LDI without re-certifying the flight software. For the CMC to operate properly, the LDI database must be loaded into the system. As upgrades to the file occurs, the LDI can be upgraded to maintain the CMCs maintenance functionality.

MEMBER SYSTEMS Member systems are not physically connected to the CMC, but the systems communicate CMC data via ARINC 429 through I/O modules or via ASCB through the NIC modules. As a valid member system, the CMC will be able to provide fault information,real time data viewing, and initiate built-in tests for the system.


TLD Data

FHDB




I/O Software


I/O Process


RS-422 Data Bus ADVANCED GRAPHICS MODULE (AGM)

RIB


VIDEO MODULE

Remote Image Bus


Main Test Switch



Aircraft FHDB


Cockpit Printer




Ethernet LAN ASCB-D

Figure 31-12. Loadable Diagnostic Information Database


31-17



Figure 31-13. Ground Maintenance Test Switch

31-18




GROUND MAINTENANCE TEST SWITCH

NOTES

The CMC has the ability to provide the user an interface to perform various actions on the member systems (i.e IBIT). Before any of these actions can be performed on the associated member system, the ground maintenance test switch must be initiated to enable these modes in the CMC. The switch sends a ground discrete to the SGIO No. 1 MAU No. 1 Slot 3 and SGIO No. 4 MAU No. 2 Slot 12. If both of these discretes are not received, the associated action that is being requested will not be performed. To validate these discretes, the CMC will display the status of the SGIO input on the bottom of the Systems Diagnostic main display page (ACTIVE or NOT ACTIVE).


31-19



PFD

MAP

1/6

SYS

2/3

BARO

ON SENSOR

FLT REF

TRS

TEST

CHKLST

HUD

PFD

MAP

1/6

< DU2 MAIN:

I NAV

< DU3 MAIN:

I NAV

SET

BRT

SYS

2/3

BARO

ON SENSOR

FLT REF

TRS

TEST

CHKLST

HUD

PUSH STD

NAV

SUMMARY

PUSH STD

CMC

NAV

ECS/PRESS FLT CTRLS

BRT

RETURN

CMC

SET

NEXT

CMC

CMC MAIN MENU

MAINTENANCE MESSAGES

SYSTEM DIAGNOSTICS

EXTENDED MAINTANANCE

END OF FLIGHT PAGE

FAULT HISTORY DATABASE CURRENTLY

CMC

DB CONFIG

CMS

EVS

75% FULL

Figure 31-14. CMC Control and Display

31-20




CMC SYSTEM OPERATION

NOTES

CMC CONTROL AND DISPLAY With only DU No. 1 and No. 4 powered, CMC information can be accessed on either DU No. 1 or DU No.4, but not both. The information can be selected for display by selecting the MAP button on the respective display controller. The INAV synoptic page will then appear on the respective DU No. 1 or DU No. 4. Access is then identical to having all four displays powered. With all four display units powered, CMC information can be accessed on DU No. 2 or DU No. 3. The information can be selected for display using the display controller or the cursor control device. To access the information utilizing the display controller: 1. Select the 2/3 menu select key 2. Select DU2 or DU3 3. Select line select key 2L, which corresponds to the CMC display


31-21



CMC

CMC

CMC MAIN MENU

MAINTENANCE MESSAGES

SYSTEM DIAGNOSTICS

EXTENDED MAINTANANCE

END OF FLIGHT PAGE

FAULT HISTORY DATABASE CURRENTLY

CMC

DB CONFIG

CMS

EVS

75% FULL

Figure 31-15. CMC Control Using the CCD

31-22




To access the CMC using the CCD:

NOTES

1. Select the DU2 or DU3 using the display selection keys. The CCD cursor will appear on that display. 2. Select the enter button on the CCD while the cursor is positioned on the INAV screen to access synoptic page menu 3. Then place the cursor on CMC using the cursor positioning device and select the enter button on the CCD.

CMC Menu Displays (Ground Mode) The CMC is powered and functional inflight, but full maintenance functionality is only realized when the aircraft is on the ground. While in ground mode, the CMC provides the following functions: • Maintenance Message Display

Active maintenance messages Historical maintenance messages Flight deck effect to maintenance message correlation • System Diagnostics Display Test pages Rigging Pages • System Configuration Display PlaneView module serial numbers Planeview module numbers • File Transfer Display

hardware

Transferring NVM from Planeview modules • Member System Status Display These displays and functions are built into the CMC and are displayed through several menu selection options. These menus and selections will be discussed on the following pages.


31-23



1.

2.

3.

4.

5.

6.

Figure 31-16. CMC Main Menu Display

31-24




CMC Main Menu Display The CMC MAIN MENU is displayed when selected using either the display controller or the CCD. 1. Menu Title—Displayed at the top of every menu for easy identification 2. Maintenance Message Menu Selection • The CMC Maintenance Message menu provides access to the associated member systems fault information in various formats. The CMC stores this information in the fault history database.

NOTE There is no fault reset associated with the CMC. Downloading the FHDB will reset the percentage full indication and CAS message. This procedure does not reset the faults stored in the FHDB. They will be over-written as the database reaches 100%, over-writing the oldest information first.

NOTES

3. System Diagnostics Menu Selection • The CMC System Diagnostics menu provides a means to select ground initiated built in tests and system parameter pages. 4. Extended Maintenance Menu Selection • The CMC Extended Maintenance menu provides status information on reporting member systems. Also, file transfers, system configuration pages, and report generations are accessed from this selection. 5. Data Loader Menu Selection • Through this menu selection, software and database loading features are enabled. 6. Fault History Database Percent Full Indication • A percent relative to the FHDB size is indicated here. When the FHDB reaches 90% full, the “Check CMC” message will be triggered and displayed on the ground. To reset this indication, the FHDB must be downloaded to a FHDB report file.


31-25



1.

2.

3.

4.

5.

6.

Figure 31-17. CMC Maintenance Messages Display

31-26




CMC MAINTENANCE MESSAGES Display

NOTES

The CMC MAINTENANCE MESSAGES menu is displayed when the MAINTENANCE MESSAGES bar is selected from the CMC main menu. 1. Menu Title—Displayed at the top of every menu for easy identification 2. ACTIVE maintenance message display selection • This menu will display, in ATA format, currently active faults only. 3. PRESENT LEG maintenance display selection • This menu will display, in ATA format, faults that are associated with the current flight leg. 4. HISTORICAL BY DATE maintenance display selection • This menu will display all faults in the FHDB, associated and displayed by date. 5. HISTORICAL BY ATA maintenance display selection • This menu will display all faults in the FHDB, associated and displayed by ATA chapter. 6. RETURN TO MAIN MENU Selection • To return to the main menu, this menu selection must be utilized.


31-27



1. 2. 5. 7.

3.

4. 6.

Figure 31-18. CMC Active Maintenance Messages Display

31-28




CMC ACTIVE Maintenance Messages Display The CMC ACTIVE maintenance messages menu is displayed when the “ACTIVE” bar is selected from the maintenance messages menu display. 1. Menu Title - Displayed at the top of every menu for easy identification 2. Main Menu Selection • To return to the main menu, this menu selection must be utilized. 3. Previous Display Selection • Only active when the maintenance messages details page is displayed. Selecting will bring back the active messages menu. 4. F l i g h t D e c k E ff e c t ( F D E ) D i s p l a y Selection • FDE is only active on the displayed maintenance messages menu. Used to correlate the Flight Deck Effect CAS message to the current fault. 5. Active Maintenance Messages

b.Once the selected ATA has been entered, the subchapter of that ATA that has faults will be displayed. The item can then be highlighted by the green box and entered to view specific faults associated with that system. c.Once the selected ATA subchapter has been entered, the fault will be displayed detailing fault name and status “ACTIVE”. d.Selecting enter will bring up the details associated with that fault. Details included are: • Fault name and fault type Possible LRUs at fault • A brief description of the fault • A hyperlink to the fault isolation procedure stored on the AMM CDROM. (Only available through the remote terminal and viewable using Adobe Acrobat)

• This area of the display will list all active faults associated by ATA and ATA nomenclature. 6. Green Bar • Indicates the position on the active menu. If more than one menu is available, this indicates the active displayed menu. 7. Green Box a.The green box is scrolled to the desired ATA fault listing by the rotary knob on the CCD. The enter button is then selected to bring up the list of faults in that ATA.


31-29



1.

3. 2.

8.

4.

6.

7. 5.

8.

Figure 31-19. CMC Present Log Maintenance Messages Display

31-30




CMC PRESENT LEG Maintenance Messages Display

c. Once the selected ATA sub-chapter has been entered, details of the fault will be displayed.

The CMC PRESENT LEG maintenance messages are displayed when the PRESENT LEG bar is selected from the maintenance messages menu display.

d. Selecting enter will bring up the details associated with that fault, which are: • Fault name and fault type

1. Menu Title—Displayed at the top of every menu for easy identification 2. Date and Flight Leg Title Area

• Possible LRUs at fault • A brief description of the fault

3. CMC Main Menu Selection • To return to the main menu, this menu selection must be utilized. 4. Previous Display Selection • Only active when the maintenance messages details page is displayed. Selecting will bring back the Present Leg Maintenance Messages menu.

• A hyperlink to the fault isolation procedure stored on the AMM CDROM. • Maintenance messages occurrences box, which lists all occurrences of this fault by date, time, activity (inactive and active), leg number, and current flight phase if known.

5. Flight Deck Effect Display Selection • See previous page. 6. Stored Active Maintenance Messages • This area displays a list of all active faults associated by ATA for the current flight leg. 7. Green Bar • See previous page. 8. Green Box a. The green box is scrolled to the desired ATA fault listing by the rotary dial on the CCD. The enter button is then selected to bring up the list of faults in that ATA. b. Once the selected ATA has been entered, the sub-chapter of that ATA that has faults will be displayed. The item can then be highlighted by the green box and entered to view specific faults associated with that system.


31-31



1.

2. 8.

4. 3. 6.

5.

6.

Figure 31-20. CMC Historical by Date Maintenance Messages Display

31-32




CMC HISTORICAL BY DATE Maintenance Messages Display The CMC HISTORICAL BY DATE maintenance messages are displayed when the HISTORICAL BY DATE bar is selected from the Maintenance Messages menu display.

NOTE The Historical by ATA menu selection works exactly the same as previous menu selections, but faults are associated by ATA first.

1. Menu Title—Displayed at the top of every menu for easy identification

NOTES

2. Main Menu Selection • To return to the main menu this menu selection must be utilized. 3. Previous Display Selection • Only active when the maintenance messages details page is displayed. Selecting will bring back the Historical by Date messages menu. 4. Active Maintenance Messages • This area of the display will list all active faults associated by year, month, date, day of week, and flight leg. 5. Green Bar • See previous page. 6. Green Box a. The green box is scrolled to the desired monthly fault listing by the rotary dial on the CCD. The enter button is then selected to bring up the list of faults in that month. b. Once the selected month has been entered, the date of that month that has faults will be displayed. The item can then be highlighted by the green box and entered to view flight legs on that date. c. Once the selected flight leg has been entered, the stored maintenance messages page will appear. This works exactly the same as the Stored Maintenance Messages display in the Present Leg menu.


31-33



CMC SYSTEM DIAGNOSTICS Menu Display The CMC provides a means to initiate ground initiated built in tests and view system parameter pages through the SYSTEM DIAGN O S T I C S m e n u . To v i ew t h e s u b m e n u s

associated with the SYSTEM DIAGNOSTICS menu, position the green box on the SYSTEM DIAGNOSTICS menu item and select the ENTER button on the CCD.

Figure 31-21. CMC System Diagnostics Menu Display

31-34




Once the SYSTEM DIAGNOSTICS menu has been selected, the display will show the SELECT A SYSTEM menu. This menu will display: • Systems by ATA, and once selected, the

ATA sub-chapters that are reporting diagnostic (parameter pages, rigging pages, and IBITs) information to the CMC. • Ground maintenance test switch status • A means to return to the main menu

Figure 31-22. CMC Select a System Menu Display


31-35



Rigging pages are also available through the C M C S Y S T E M D I AG N O S T I C S m e n u . Currently, only the visual guidance system has rigging pages available in the CMC. To view these pages: 1. Select the SYSTEM DIAGNOSTICS Menu display from the main menu.

2. Select ATA 34 Navigation 3. Select the ATA subchapter 26 Visual Guidance System 4. Select the RIG page that you would like to utilize and follow the directions given on the CMC screen.

Figure 31-23. CMC Rigging Pages

31-36




System Test pages are also available through the CMC SYSTEM DIAGNOSTICS menu. To view these pages:

3. Select the ATA that you would like to init i a t e t h e G M T o n ( i . e . , . ATA 3 4 Navigation).

1. Ensure that the maintenance test switch is enabled.

4. Select the ATA sub-chapter you would like (i.e., ATA 34-26 Visual Guidance System).

2. Select the SYSTEM DIAGNOSTICS Menu display from the main menu.

5. Select the test page that you would like to execute and follow the directions given on the CMC screen.

Figure 31-24. System Test Pages


31-37



CMC EXTENDED MAINTENANCE Menu Display The CMC provides status information on reporting systems through this menu selection. File transfers, system configuration pages, and report generations are also accessed from this menu.

To view the submenus associated with the EXTENDED MAINTENANCE menu, position the green box on the EXTENDED MAINTENANCE menu item and select the ENTER button on the CCD.

Figure 31-25. CMC Extended Maintenance Menu Display

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The MEMBER SYSTEM STATUS menu allows the operator to view the status of member systems. When the menu is selected, ATAs that have problems (no status) will be highlighted in blue.When the main ATA chapter is selected, the subchapters will be displayed and the “no status” items will be blue.

Items that are “operational” are white. Operational items indicate that the system is on-line and reporting to the CMC.

Figure 31-26. Member System Status


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Figure 31-27. CMC File Transfer Menu Display

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CMC FILE TRANSFER FROM AIRCRAFT Menu Display

NOTES

The operational aspects of this page have not been fully defined. Data will be added as it becomes available.


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CMC CONFIGURATION Menu Display The CMC CONFIGURATION menu provides hardware, software, and part number information on the member systems. This page will be very useful for the Honeywell Planeview

modules in the MAUs. To view the sub-menus associated with the CONFIGURATION menu, position the green box on the CONFIGURATION menu item and select the enter button on the CCD.

Figure 31-28. CMC Configuration Menu Display

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The SYSTEM CONFIGURATION page will appear when the CONFIGURATION menu is selected. A list of ATA chapters will appear and the operator can select the desired ATA chapter. ATA 31 contains the information on Honeywell Planeview modules. Once in the

ATA 31 submenu, select the desired module. Data presented is hardware part number, software part number, and the module serial number. These items are hidden with the backshell connected to the module.

Figure 31-29. CMC System Configuration Pages


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CMC REPORTS Menu Display The CMC REPORTS menu provides the operator a means to save report files into various locations for viewing later on a personal computer. They can also be printed to the cockpit printer. To view the sub-menus associated with the REPORTS menu, position the green box on the REPORTS menu item and select the ENTER button on the CCD.

Once CMC REPORTS is entered, the menu will show what reports are available. Currently, ACMF and CMC reports are viewable. ACMF reports are based on an initiated report generation from the cockpit event switch. If there are no reports available in that category, the CMC will indicate “No Reports Found”.

Figure 31-30. CMC Reports Menu Display

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CMC reports that are available are Active FDE/Maint Msg, Current Leg FDE/Maint Msg, System Configuration, and Export Fault History. Once a report has been selected, it can be sent to the cockpit printer, local storage (RT hard drive), or DMU PCMCIA Slot 2.

The bottom of the screen will give the report status, when it has been selected for storage. The status line will indicate the path to which the report will be stored and the saved file name.

Figure 31-31. CMC Reports Pages


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Figure 31-32. CMC Menu Displays (Air Mode)

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CMC Menu Displays (Air Mode) The CMC is powered and functional inflight, but it has limited functionality to the operator. While in air mode, the CMC provides the following functions: • AC T I V E M A I N T E NA N C E M E S SAGES • S TO R E D M A I N T E NA N C E M E S SAGES • REPORTS T h e AC T I V E M A I N T E NA N C E M E S SAGES and REPORTS screens are identical to the screens generated while on the ground. The STORED MAINTENANCE MESSAGES screens work and are viewed the same way as the historical by date and ATA maintenance messages screens on the ground. Maintenance data is available, but IBITs and rigging functions are locked out while in air mode.

facility. Considerations should also be given when deciding to remove the CMC module for maintenance purposes. Before removing the CMC module, pull and tag the CMC Shutdown Power circuit breaker (LEER H6). This will remove the 24VDC power input from the L E-Batt.

Central Maintenance Computer Function The central maintenance function (CMF) is a software application hosted on the CMC module. It provides the real-time collection of faults and the subsequent storage of associated maintenance messages. Additionally, this application provides a central location from which other maintenance functions (rigging, IBIT) can be performed.

NOTES

CMC Power Up/Down Procedure The CMC operates using WindowsNT technology. As with any Windows product, there is a two minute power up period before it will be available for use after the aircraft is initially powered up. On shutdown, the same applies. The CMC must have two minutes of power to shutdown its functions. To accomplish this, the CMC will utilize the left emergency battery for power to complete it’s shutdown cycle after aircraft power is secured. The power source is a direct line from the L E-Batt to the “CMC Shutdown Power circuit breaker (LEER H6). With this design, it is very important to never have the L E-Batt removed or have its front panel circuit breakers pulled during an aircraft power down cycle. If this occurs, it is possible for data corruption to occur or the CMC can become “locked” and unavailable to the operator. These conditions will require the CMC to be repaired by a Honeywell repair FOR TRAINING PURPOSES ONLY

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CMC COMMUNICATION

CMC Flight Deck Effect and Maintenance Message Correlation

The CMC has no direct communication with any of the member systems that it is monitoring. The member systems transmit their associated data on ARINC 429 or ASCB. An input/output module receives the ARINC 429 and the NIC/PROC receives the ASCB data. The data is then collected from the MAU backplane by the CMC Central Maintenance Function. The CMF processes the received maintenance data through the LDI, through which, the maintenance messages are derived.

The Central Maintenance Function is capable of correlating CAS messages, called flight deck effects, to maintenance messages. To view these, the operator must be on a maintenance message display menu (active, present leg, historical, or stored). Once there, the maintenance message must be highlighted by the green box. The operator can then select the FDE” button in the lower right hand corner of the CMC display.


TLD Data

FHDB




I/O Software


I/O Process


RS-422 Data Bus RIB

ADVANCED GRAPHICS MODULE (AGM)


VIDEO MODULE

Remote Image Bus


Main Test Switch



Aircraft FHDB


Cockpit Printer




Ethernet LAN ASCB-D

Figure 31-33. CMC Communication Diagram

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The screen will change once the selection is made to the ACTIVE CORRELATED FDE page (Active or Stored), which will list the correlated CAS (warnings, cautions, or advisories) associated with the displayed maintenance message.

O n t h e AC T I V E C O R R E L AT E D F D E ” screens, the selected Monitor Warning System will be displayed. This correlates to the MWF that was selected when the failure occurred. This information is not present on any of the STORED CORRELATED FDE screens.

Figure 31-34. CMC Flight Deck Effect and Maintenance Message Correlation


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Figure 31-35. CMC CAS Messages

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CMC CAS Messages

NOTES

There are 7 CAS messages associated with the CMC. They are: • L-R Engine Maintenance STD (Amber Caution) • L-R Engine Maintenance LTD (Blue Advisory) • L-R Engine Maintenance 150 (Blue Advisory) • Check CMC (Blue Advisory) • Event Record (Blue Advisory) • Exceedance Record (Blue Advisory) • CMC Fail (Blue Advisory)

Fault Classification Member Systems identify fault occurrences to the CMC and classify the faults. Engine faults are classified as Long Term Dispatch, Time Limited Dispatch 150, 140, etc., or Short Term Dispatch. The CMC takes this information and presents it through the CAS system. Without the CMC, all engine faults become “Do Not Dispatch”. All other member system faults are classified as “Long Term Dispatch”, according to the CMC, which initiates the “Check CMC” advisory CAS message when fault information is present. The message is inhibited when in the air mode and airspeed is greater than 50 knots.


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Figure 31-36. CMC Parameter Page Viewing

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CMC Parameter Page Viewing From the SYSTEM DIAGNOSTICS menu, aircraft systems parameter pages can be viewed. Parameters are displayed in real-time. This can be a valuable tool when evaluating faults associated with the engines or APU.

Aircraft Condition Monitoring Function The ACMF compliments the fault storage capabilities of the CMC by providing storage and analysis of additional data that is not necessarily associated with a fault. This stored data can be downloaded as a report file utilizing the remote terminal and the LAN. The reports can then be viewed using Microsoft Excel. There will also be future capabilities to downlink these reports via the Communications Management Function. This is not yet implemented.

The ACMF provides exceedance monitoring/recording, standard recordings/reports, and customized recordings/reports. These recordings/reports are stored on the CMC module. They can then be downloaded using the REPORTS display menu. The system will overwrite older reports with newer ones as the memory disc gets full. The ACMF shall monitor the defined parameters that appear on ASCB, collecting information for pre-trigger measures and and evaluating trigger expressions. When a trigger expression becomes true, the ACMF shall create a report of the condition, storing all defined measures and for the function. The ACMF has the ability to record 100 seconds before and up to 100 seconds after the trigger expression becomes true. This time can be varied through the software at the request of Gulfstream.

Figure 31-37. Aircraft Condition Monitoring Function


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Figure 31-38. ACMF Reports Viewing

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ACMF Reports

• Main AC Volts

In the event of an exceedance (red/amber) being recorded, a report will be generated by the ACMF and an Exceedance Record advisory will be presented in CAS. If the FDR/CMC event switch is utilized to generate the report, the ACMF will produce the report and an “Event Record” advisory will appear in CAS.

• Main DC Volts

Red/Amber Exceedances that are monitored by the ACMF are:

• Oil Pressure Low • Oil Temp High • Oil Temp Low • Overweight Landing • Takeoff • TAT Split • Throttle Config

• Flap/Stab Exceedance

• WOW Fault Advisory

• APU Exceedance

• WOW Fault Caution

• Engine Exceedance • Altitude Exceedance

NOTES

• Gear Exceedance • Engine Hot Some examples of ACMF report capabilities (Figure 31-38) are:

• • • • • • • • • • • • • • • • •

Aircraft Config Airspeed Miscompare Autopilot Fail Autothrottle Fail Bleed Config DU Blinking End of Flight Engine Backup Air Data Event Record EVS External DC Power Overload Flap Asymetry Fuel Tank Temp Generator Overload Ground Spoiler HMG Overload Landing


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Figure 31-39. Communication Management Displays

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Communication Management Function

NOTES

The Aircraft Diagnostic and Maintenance System will have the capability to transfer the following maintenance data from the aircraft to a ground station while in-flight: • Active Maintenance Messages • Stored Maintenance Messages • Aircraft Condition Monitoring Function Reports The ADMS transfers the data to the Communications Management Function (CMF), which in turn formats the data for the appropriate protocol. The CMF then forwards the information to the appropriate datalink function such as VHF COM or SATCOM. The datalink function transmits the information to a ground station using a service like Global Data Center. At the ground station, the file is converted to an e-mail and then e-mailed to the subscriber’s e-mail list (Figure 31-39).

Dispatch Considerations The central maintenance computer is not required for dispatch, but it must be understood that any engine fault will become a “Do Not Dispatch” without the CMC in the aircraft and operating normally. The central maintenance computer is NOT used to determine the airworthiness of the aircraft.


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Figure 31-40. Remote Terminal Setup Displays (Sheet 1 of 2)

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REMOTE TERMINAL REQUIREMENTS

• The laptop shall include a PS/2-style Mini-DIN 6-pin connector for an external keyboard or mouse.

Remote Terminal is software that connects to the Central Maintenance Computer in the airplane via 10BASE-2 Ethernet. The software is provided by Honeywell. Remote Terminal needs to be loaded on a laptop (Provided with the A/C) that meets these minimum requirements:

• For connection to the aircraft LAN for troubleshooting, the laptop must interface with a 10BASE-2 (ThinLAN coax Ethernet). This requires the following equipment (or equivalent):

• The processor shall be at least equivalent in speed to an Intel Pentium 3 at 1.2GHz clock speed. • The laptop display shall be at least 13.3 in diagonal with at least 1024x768 resolution. A larger display and higher resolution are desired. The display should be TFT Active Matrix. Emphasis must be placed on sunlight readability. • The laptop shall include a 15-pin highdensity D-sub connector for connecting an external monitor. • Disk storage shall be at least 20GB. • The laptop must be capable of accepting a 1.44MB floppy drive, either builtin, through a removable drive bay device, or externally through a parallel port, disk drive connector, or USB port. • Random Access Memory (RAM) shall be at least 128MB.

1. Linksys EC2T PCMCIA Ethernet adapter with 10Base-2 coax adapter 2.Belkin 50 ohm coax Cable kit with B N C Te e ( R 6 C 0 7 3 - E ) , B N C Terminator (R6C043-E), and 10 ft ThinLAN cable (F3K101-10-E).

Remote Terminal Setup Remote Terminal, when initialized, looks exactly like the CMC except for the menu bar at the top of the screen. For it to work properly, the laptop must be configured to run Remote Terminal, if it hasn’t been in the past. • A network connection needs to be created using an IP address between 192.168.201.1 and 192.168.201.254 with a subnet mask of 255.255.0.0 To get to this setup menu: 1.Select “Start” from the Task bar

• The operating system must support software certified for use with Microsoft Wi n d o w s 2 0 0 0 o r Wi n d o w s X P Professional. Limiting operations to software created for Windows 95, 98, 98SE, NT, or Me are not acceptable.

2.Select “Settings” from the Start Menu

• T h e l a p t o p s h a l l h a v e 1 0 BA S E - T Ethernet capability at 100Mbps with a built-in RJ-45 jack. Required for interface to the aircraft for CMC operation.

5.The “Local Area Connection Status” Wi n d o w w i l l a p p e a r. C l i c k o n “Properties”.

3.Select “Network and Dial Up Connections” from the Settings Menu 4.Select “Linksys Combo PCMCIA Ethernet Card” and Click

• The laptop shall have a built-in (or removable drive bay) CD-RW, or combination CD-RW/DVD-ROM drive. • The laptop shall include at least one USB port. Two USB ports are preferred.


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Figure 31-40. Remote Terminal Setup Displays (Sheet 2 of 2)

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6.The “Properties for this connection” will a p p e a r. S e l e c t “ I n t e r n e t P r o t o c o l (TCP/IP)” and click on “Properties”. 7.The “TCP/IP Internet Protocol Properties” Menu will appear. The “Obtain IP Address automatically” will be selected.

Remote Terminal Connection When the laptop is connected and remote terminal has been initiated, the menu bar at the top of the display can be used to make some changes to the remote terminal configuration.

8.Select “Use the following IP address”, and enter an address that corresponds to the number addresses above. Click OK. The setup procedure is now complete.

Figure 31-41. Remote Terminal Connection Display (Sheet 1 of 2)


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Figure 31-41. Remote Terminal Connection Display (Sheet 2 of 2)

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The second menu selection is “Settings”. Under this menu, the user can:

NOTES

• Connection Information—Set the component that Remote Terminal is connected to. • Display Size—Laptop or Desktop • Download/Report Directory—Designate the directory that reports are written to the local storage drive. • Administrative Tools—Requires password. For future use. A third menu selection, “Edit ACMF Data”, will be available in the future. Under this menu, the user will be able to input Aircraft Hours, Engine Hours, and Engine/APU Serial Numbers.

Figure 31-42. Edit ACMF Data Display


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G500-G550 MTM V1

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