B737 NG GEN FAM

SECTION

TITLE

1

Introduction

2

Structures

3

Equipment Centers

4

Flight Compartment

5

Common Display System

6

Communication and Recording

7

Navigation

8

Autoflight

9

Electrical Power

10

Fuel

11

Auxiliary Power Unit

12

Power Plant

13

Hydraulics

14

Landing Gear

15

Flight Controls

16

Environmental Systems

17

Fire Protection

18

Ice and Rain

19

Cabin Systems

20

Lights

21

Airplane Access Abbreviations and Acronyms

Table of Contents

Introduction About This Book

Principal Characteristics

•

Operating Experience

This document presents a general technical description of the Boeing 737. It is based on the standard airplane, but also includes details of some of the most popular options.

The main characteristics and major structural differences of the 737-600, -700, -800, and -900 are shown on the subsequent pages. Each airline selects different options. All options are not shown in this manual. Most of the systems are similar between models; only the major differences are covered.

•

Self-Sufficiency

•

Eye-Level Maintenance

For detailed information, or information on a specific customer airplane, refer to these publications: • • • • •

Airplane Flight Manual Operations Manual Airplane Maintenance Manual Configuration Specification Document Configuration Control Document.

If the information in this book does not agree with the information in any of these publications, the publications should be used. OVERVIEW To understand the airplane, it is necessary to understand the airplane systems. This document gives an introduction to the systems in the 737. Each system is shown from a component, installation, and operational view. Flight compartment instruments and panel data help show system operation.

November 2000

1-1

112 ft 7 in (34.4m) (Ref)

117 ft 2 in 117 ft 7 in (35.7m) (35.8m) (With Winglets) (With Winglets) (Unloaded) (Loaded) 47 ft 1 in (14.3m)

12ft 4 in (3.8m) 18 ft 9 in (5.7m)

Airplane Dimensions Airplane Dimensions The wing span and horizontal tail span is the same for all models in the next generation family. The dimensions are shown above.

1-2

November 2000

Introduction

41 ft 3 in (12.5m)

737-600

13 ft 5 in (4.0m) (Typ)

36 ft 10 in (11.2m)

102 ft 6 in (31.2m)

41 ft 2 in (12.5m)

737-700

41 ft 3 in (12.6m)

110 ft 4 in (33.6m)

21 ft 1 in (6.4m) (With Winglets)

41 ft 2 in (12.5m)

737-800

51 ft 2 in (15.6m)

129 ft 6 in (39.5m)

21 ft 1 in (6.4m) (With Winglets)

41 ft 2 in (12.5m)

737-900

56 ft 4 in (17.17m)

138 ft 2 in (42.11m)

Airplane Dimensions November 2000

1-3

Oslo 737-600/700 Reykjavik

London Paris Madrid

Seattle

737-800/900

San Francisco Denver Los Angeles

New York

737-300 Mexico City

Miami

737-400

Bogota 737-500

Airplane Ranges Operating Experience The 737 models now in service have a high dispatch reliability. The 737 flies a large number of short length flights. It can also fly longer range flights. The airplane use rate is very high. The use rates are shown this way: • • •

Average daily utilization Average flight length Schedule reliability

The range map above shows the longer range of the 737-600/700/800 /900 over the 737-300/400/500 aircraft. The ranges are for aircraft in these conditions: • • •

Normal flights Full passenger payload 85% annual winds.

7.6 hr 1.4 hr 99.3%.

The 737-600/700/800/900 design improves on the 737/300/400/500 model design. These are the improvements: • • • • • •

Larger payload Higher service ceiling More range Improved fault isolation Improved systems design Flight compartment common display system.

1-4

November 2000

Introduction

12 ft 4in (3.76m) (typ)

737-600

737-700

737-700C

133 500 to 153 500 (69 626) (60 555)

171 500 (77 791)

156 000 to 173 000 (60 780) (70 762)

737-800

737-900

Maximum Gross Weight, Pounds (Kilograms) Taxi

124 500 to 144 000 (65 318) (56 473)

Brake Release

124 000 (56 246)

143 500 (65 092)

133 000 (60 328)

153 000 (69 400)

171 000 (77 564)

155 500 (70 535)

172 500 (60 550)

164 000 (74 380)

174 200 (78 240)

Landing

120 500 (54 659)

120 500 (54 659)

128 000 (58 060)

128 000 (58 060)

134 000 (60 781)

144 000 (65 318)

144 000 (65 318)

146 300 (66 360)

146 300 (66 360)

Zero Fuel

113 500 (51 484)

113 500 (51 484)

120 500 (54 658)

120 500 (54 658)

126 000 (57 153)

136 000 (61 690)

136 000 (61 690)

138 300 (62 732)

138 300 (62 732)

164 500 to 174 700 (79 243) (74 615)

Engine Thrust, lb Basic Option Option

CFM56-7B20 20 600 CFM56-7B22 22 700 CFM56-7B24 24 200

CFM56-7B18 19 500 CFM56-7B20 20 600 CFM56-7B22 22 700

CFM56-7B24 CFM56-7B26 CFM56-7B27

24 200 26 300 27 300

CFM56-7B24 24 200 CFM56-7B26 26 300 CFM56-7B27 27 300

Fuel capacity, U.S. Gallons (liters) 6878 (26033) Passengers Mixed Class All Tourist, 32-in Pitch All Tourist, 30-in Pitch Lower Hold Volume, ft3 (m3)

108 123 130 Fwd

Aft

Fwd

Total

6878 (26033)

6878 (26033)

128 140 148

160 175 189

177 189 189

Aft

Total

386 596 982 (10.9) (16.9) (27.8)

257 488 745 (7.3) (13.8) (21.1)

Speed Capacity Maximum Operating Airspeed, Knots (KCAS) Maximum Operating Mach Number

6878 (26033)

Service Ceiling 340 0.82

Fwd

Aft

Total

899 1566 667 (18.9) (25.5) (44.3)

Fwd

Aft

Total

840 1012 1852 (23.8) (28.7) (52.4)

41 000 feet 12 497 meters

Principal Characteristics November 2000

1-5

Lengthened Vertical Stabilizer

Redesigned Airfoil and Increased Wing Span Additional Flight Spoiler Dorsal Fin

Strengthened Stabilizer and Center Section

New APU and Firewall Increased Horizontal Stabilizer Span

Redesigned Krueger Flaps

Revised Aft Wing-Body Fairing Shape

Redesigned Main Landing Gear

Flap Track Fairings CFM56-7 Engine

Redesigned Nose Landing Gear Redesigned EE Compartment

Additional Slat Tip Extension

Winglet Option Major 737 Changes (Typical)

From 737-500 t0 737-600 (No Difference in Fuselage Dimensions)

737 Changes 1-6

November 2000

Introduction

From 737-300 to 737-700 (No Difference in Fuselage Dimensions)

64-inch (163-cm) Body Extension

46-inch (117-cm) Body Extension

Tail Skid Expanded Environmental Control System (ECS)

From 737-400 to 737-800 (Additional Changes) 42-inch (107-cm) Body Extension

62-inch (157-cm) Body Extension

From 737-800 to 737-900 (Additional Changes)

737 Changes November 2000

1-7

Fueling

Brake and Hydraulic Service Baggage Handling

Baggage Handling

Galley Service Galley Service

Optional Airstair Potable Water Service Lavatory Service

Air Conditioning

Ground Air

Airplane Servicing Self-Sufficiency

Eye-Level Maintenance

The airplane can operate at improved and unimproved airports. These are the systems that make the airplane self sufficient:

This design permits eye-level maintenance access at ground level. This permits easy access to systems and unit modules. Many major systems are grouped together.

• • •

• • •

•

•

The APU supplies on-ground or in-flight electrical power The APU supplies air for engine starting The APU maintains an airconditioned cabin during ground operations The APU is started from the airplane battery The airplane has large fuel and water capacity Alternative systems allow dispatch with inoperative systems Two or more systems which have equal function but one system operates at a time which allows more rapid fault dispatching Self-contained airstair (option).

1-8

The operator benefits from minimum maintenance cost and low ramp turnaround times. Less money is spent to purchase maintenance equipment. Engines are changed at ground level. Hand operated hoists attach to the airplane engine struts to remove and install engines. Component and System Access The major hydraulic components, are in the main landing gear wheelwell and can be maintained at ground level.

The air-conditioning units are easy to reach. They are behind two doors under the wing center section. Access doors, forward and aft of the nose wheelwell, give access to electronic equipment compartments. Extension of forward and aft wing high lift devices permit access to other system components. Potable water and lavatory systems are easy to maintain because of ground access to their service panels. Built-in-Test Built-in test (BITE) and checkout function for systems simplify fault isolation. Many airframe/engine modules and most avionics modules include BITE. BITE access is at the face of the module or through the flight compartment control display units (CDU). November 2000

Structures Features

•

Fuselage

BASIC STRUCTURAL DESCRIPTION

•

Wing

•

Stabilizers

•

Composites

The airplane is a low wing twin engine design. The engines are below the wings on struts. It has full cantilever wings and tail surfaces. The fuselage is a semi-monocoque design. HIGH-FATIGUE DESIGN LIFE The design service objective is 75,000 flight cycles. For typical airline operations, the aircraft reaches this objective after 25 years of service. CORROSION PREVENTION Years of extensive in-service experience lead to an optimum airframe design. This knowledge along with new material technology gives the operator an airframe that results in: • • •

Minimal corrosion Longer in service periods Less maintenance costs.

November 2000

2-1

Tail Cone

Floor Beams

Aft Pressure Bulkhead

Frames

Stringers

Wing-To-Body Fairing

Nose Radome Forward Pressure Bulkhead

Fuselage Fuselage The fuselage is a pressurized semimonocoque structure. The primary materials for the fuselage are aluminum alloys. Pressure bulkheads at the forward and aft ends of the fuselage form a pressure vessel. These auxiliary structures attach to fuselage: • • •

Nose radome Wing-to-body fairing Tail cone.

2-2

November 2000

Structures

Leading Edge Flap

Engine Nacelle/Pylon

Rib Access Cutout

Slat Fuel Tank Access Panel (Typical) Ground Spoiler Wing Tip

Flaps Flight Spoiler

Aileron

Aileron Balance Tab

Wing Wing The wing is a cantilever structure. The basic wing structure is aluminum. The wing has these features: • • •

Stores fuel Contains fuel system components Attach points for the engine strut, landing gear, and flight control surfaces.

Fuel tank access panels on the bottom wing skin permit access to the fuel tanks.

November 2000

2-3

Removable Leading Edge

Rear Spar

Rudder Front Spar Center Section Truss Structure Dorsal Fin Front Spar Rear Spar

Vertical Stabilizer

Elevator Horizontal Stabilizer

Horizontal/Vertical Stabilizer Stabilizers The horizontal and vertical stabilizers are made of aluminum alloys. The elevator and rudder are made of graphite.

2-4

November 2000

Structures Trailing Edge Panels Rudder Tail Cone

Aileron Dorsal Fin

Aileron Tab

Elevator

Flap Track Fairings GRAPHITE

FIBERGLASS Wing-To-Body Fairing

FIBERGLASS/GRAPHITE

Inboard and Outboard Fixed Trailing Edge

Radome

Nose Landing Gear Door (Graphite)

Thrust Reverser Inboard Fixed Leading Edge Lower Skin Panel (Fiberglass)

Outboard Fixed Leading Edge (Fiberglass)

Composites Composites Some airplane structure and parts are made from composite materials. These are some advantages of composite materials: • • • •

High strength Corrosion resistant Increased fatigue life Light weight.

November 2000

2-5

Equipment Centers Features

•

Electronic Equipment Compartment

ACCESSIBLE LOCATION Most electronic equipment is in a compartment below the cabin floor aft of the nose wheel well. This compartment is easily accessible from ground level. TRANSVERSE RACK The electronic equipment compartment includes 3 equipment racks. The main equipment rack is a transverse rack across the aft end of the compartment. Equipment removal and installation is easy due to the rack design. Interconnecting wiring, mounts, and accessory boxes are accessed through panels in the forward cargo compartment.

November 2000

3-1

Electronic Equipment Compartment

Optional E6 Rack (Aft Cargo Compartment)

E4 Optional E8-1 Shelf (Not Shown) E3

E1 E5

E2

Electronic Equipment 3-2

November 2000

Equipment Centers

E1-1 Shelf (Blow Through Cooling)

E1-3 Shelf (Blow Through Cooling)

E1-5 Shelf (Draw Through Cooling)

IFSAU Gnd Auto Throttle Prox Comp

Flap/ Slat Elex Mod

AntiPA VHF Skid Amp Com Auto 1 Brake Control

VHF Comm 2

ATC 2

DME 2

Flight Control Computer Channel A

DME 1

VOR 2

TCAS Computer

ATC 1

MMR 2

VOR/ MB 1

MMR 1

Flight Control Over heat Computer Cont Channel B Mod

E1-2 Shelf (Blow Through Cooling)

E1-4 Shelf (Draw Through Cooling)

Electronic Equipment Rack E1 (Typical Arrangement)

Electronic Equipment Compartment Electronic Equipment Compartment Electronic equipment is in a compartment below the main cabin floor aft of the nose wheel well. On the ground, you enter this electronic equipment (EE) compartment through a door in the bottom of the fuselage.

Most equipment rack shelves are cooled with air. Air is blown through or drawn through the equipment racks. There is a drip shield over the racks to protect the equipment from moisture condensation.

There are five standard equipment racks. These are the E1, E2, E3, E4, and E5 racks. More equipment racks may be needed on airplanes with optional systems. Shelf assemblies have equipment mounts, interconnected wiring, and accessory boxes. They are easy to remove. This makes troubleshooting and modification easier. Equipment that installs boxes on shelves is adjustable. Access to the most frequently used boxes is improved. November 2000

3-3

E4 Rack

CAB AIR CAB CON TEMP REU PRS SELCONT CAL RLY CONT 2

AUTO SPEED BRAKE ACCESS MOD

E3 Rack

SATCOM

WINDOW WINDOW T/R CONT CONT 3 3 4

ADF 2

SATCOM

T/R 2

BUS PWR CONT UNIT

GCU 2

RAD ALT 2

MUX POWER DISTRIBUTION PANEL 2

CDS DEU 2

DFDAU

AUDIO ENT PLAYER

E2 Rack

ADF 1

CDS DEU 1

ENG VIB SIG COND

ACARS MU

ENG ACC

RAD ALT 1

SYMD SYMD 2 1

VHF 3

WINDOW CONTROL 2

APU GCU

GCU 1

1

FIRE CAB & PRESS APU CONT DET 1 CONT MOD

APU START CONVERTER UNIT

T/R 1

BATTERY PRIMARY CHARGER BATTERY CHARGER

STATIC INVERTER

APU START POWER UNIT

AUX BATTERY CHARGER POWER DISTRIBUTION PANEL 1

BATTERY

J9 PANEL

Electronic Equipment Racks E2, E3, and E4 (Typical Arrangement) GCU

L ADIRU

FMC 1

FMC 2

R ADIRU

Electronic Equipment Rack E5 (Typical Arrangement)

Electronic Equipment Compartment 3-4

November 2000

Equipment Centers Voice Recorder SATCOM HF 1 HF 2 APU ECU

E6 Rack (No Cooling) (Aft Cargo Compartment)

HF-1

FWD E6 Rack (No Cooling) (Aft Cargo Compartment)

Weather Radar Transceiver

HF-2

E8-1 (Blow Through Cooling) (EE Compartment)

Optional Electronic Equipment Racks (Typical Arrangement) Optional Electronic Equipment Racks Two optional equipment racks are available when more space is necessary. The E8-1 shelf is above the E1 rack in the EE compartment. The E8-1 shelf has blow-through cooling. The E8-1 shelf option is available when airstairs are not installed. The E6 rack is in the aft cargo compartment. The E6 rack is not cooled. Airplanes that do not have the optional E6 rack have a mount in the aft cargo compartment for the voice recorder and the APU electronic control unit (ECU). This mount location is the same as the E6 location.

November 2000

3-5

Flight Compartment Features

•

Flight Compartment Panels

DESIGN PHILOSOPHY

•

Glareshield Panel

The flight deck maintains the same type-rating as all previous 737s, while integrating refinements proven on the 737 with innovative 777 technologies.

•

Instrument Panels

•

Center Forward Panels

•

Control Stand

FLAT PANEL LIQUID CRYSTAL DISPLAY UNITS

•

Aft Aisle Stand Panels

•

Overhead Panels

•

Other Flight Compartment Components

The flat panel liquid crystal display units need less power and have a larger display area than conventional CRT display units. MODE CONTROL PANEL The mode control panel uses integrated LED light switch assemblies to increase reliability.

November 2000

4-1

P5 Aft Overhead Panel P5 Forward Overhead Panel P7 Glareshield Panel P2 Center Instrument Panel P3 First Officer Instrument Panel

P1 Captain Instrument Panel

P9 Forward Electronic Panel Control Stand P8 Aft Electronic Panel

Flight Compartment Panels Flight Compartment Panels

These are the aisle stand panels:

The main instrument panel has three panels. These are the panels:

• • •

• • •

P1 P2 P3.

The P7 glareshield panel is above the main instrument panels.

4-2

P8 aft electronic panel P9 forward electronic panel Control stand.

These are the two overhead panels: • •

P5 forward overhead panel P5 aft overhead panel.

November 2000

Flight Compartment

COURSE

P7 Glareshield Panel

Glareshield Panel Glareshield Panel The glareshield panel is the P7 panel. The P7 panel contains these panels: • • • •

Mode control panel (MCP) EFIS control panels Master caution annunciations Fire warning light.

The MCP uses integrated LED light switch assemblies. This design improves the reliability and maintainability of the mode control panel. The EFIS control panels are on the glareshield panel for easier access by the pilots. These control panels are similar to the Boeing 747-400 and 777 EFIS control panels.

November 2000

4-3

P1 Captain Instrument Panel

Main Instrument Panels Captain Instrument Panel The captain instrument panel is the P1 panel. The P1 panel has these features: • • • • • • • • •

Left outboard display unit Left inboard display unit Display switching module Clock Autoflight status annunciator Conditioned air outlet controls Lighting controls for the captain Master dim and test switch Yaw damper indicator.

The display switching modules let the pilots show different display formats on the inboard and lower display units.

4-4

November 2000

Flight Compartment

P3 First Officer Instrument Panel

Main Instrument Panels First Officer Instrument Panel The first officer instrument panel is the P3 panel. The P3 panel is almost the same as the P1. The P3 panel also has these features: • •

Ground proximity module Brake pressure indicator.

November 2000

4-5

P2 Center Instrument Panel

Center Instrument Panel Center Instrument Panel The center instrument panel is the P2 panel. The P2 panel has these items: • • • • •

Engine control module Antiskid and autobrake switches and lights Landing gear lever and position indicators Upper center display unit Standby instruments.

4-6

November 2000

Flight Compartment

P9 Forward Electronic Panel

Forward Electronic Panel Forward Electronic Panel The forward electronic panel is the P9 panel. The P9 panel contains these displays: • •

Lower center display unit Control display units.

November 2000

4-7

Forward Thrust Levers Reverse Thrust Levers P9 Forward Electronic Panel

Takeoff/Go-Around Switches Flap Lever Horn Cutout

Stabilizer Trim Indicator STAB TRIM

STAB TRIM

Speed Brake Lever

1

2

2

5 10 UP

15 25

Start Levers

Parking Brake Lever

P8 Aft Electronic Panel Horizontal Stabilizer Manual Trim Wheels

30 40

1

2

Stabilizer Trim Cutout Switches

Control Stand Control Stand The control stand has controls that are easy to reach by either pilot. The control stand has these controls: • • • • • • • • • •

Forward thrust levers Reverse thrust levers Takeoff/go-around switches Speed brake handle Horizontal stabilizer manual trim wheels Parking brake lever and indication light Flap lever Stabilizer trim cutout switches Horn cutout Engine start levers.

4-8

November 2000

Flight Compartment

P8 Aft Electronic Panel

Aft Electronic Panel Aft Electronic Panel The aft electronic panel is the P8 panel. The P8 panel has these features: • • • • • • • • • •

VHF and HF radio control panels Nav control panels Audio control panels ADF control panel ATC control panel Cargo fire panel ACARS interactive display unit Aileron and rudder control Lighting controls Weather radar control panel.

Radio tuning panels can replace individual radio control panels. The radio tuning panels tune the VHF and HF radios from one control panel.

November 2000

4-9

GPS

P5 Aft Overhead Panel

Aft Overhead Panel Aft Overhead Panel

Forward Overhead Panel

The P5 aft overhead panel has controls that are seldom used in flight.

Because of its central location, either pilot can reach any of the systems controls. The panel has controls for these systems:

The P5 aft overhead panel has these controls and Indications: • • • • • • • • • • • • • •

Leading edge devices annunciator panel Inertial system display unit IRS mode selector unit Service interphone switch Dome light switch Observer’s audio control panel Reverser fault lights Engine control lights Electronic engine control (EEC) alternate mode light/switches Passenger and crew oxygen system Flight recorder test panel Mach/airspeed warning panel Stall warning test panel Landing gear down & locked indicator lights.

4-10

• • • • • • • • • • •

• • •

Exterior lights APU Engine start.

The primary system control panels, fuel, electrical, hydraulic, and air conditioning, are light grey.

Flight controls Instrument switching Fuel Electrical Window and air data probe heat Engine and wing anti-ice Hydraulics Door warning Voice recorder Air-conditioning Pressurization.

The forward overhead panel has switches for these functions: • • • • •

Overhead panel lights Equipment cooling Emergency exit lights Passenger signs Rain removal November 2000

Flight Compartment

P5 Forward Overhead Panel

Forward Overhead Panel November 2000

4-11

P61

P6

P18

Other Flight Compartment Components Other Flight Compartment Components The main circuit breaker panels are behind the first officer and captain. The P6 and P18 have the component load circuit breakers. Circuit breakers are organized by airplane systems. Emergency equipment is placed within easy reach of the crew. Emergency equipment includes these items: • • •

A fire extinguisher on the P6 panel A crash axe on the P18 panel Escape lanyards above the sliding windows.

The data loader control panel is on the P61 panel. The bulkhead and sidewalls have provisions for stowing crew luggage, flight manuals, coats, and hats. 4-12

November 2000

Common Display System Features

•

Common Display System

INTEGRATED FUNCTIONS

•

EFIS layout

The common display system shows flight and engine information to the flight crew. It does the functions of the electronic flight instrument system (EFIS), the engine instrument system (EIS) and most of the mechanical primary flight instruments that are used on the 737-300/400/500 models.

•

EFIS

•

PFD/ND

•

Engine Display

•

System Display

•

BITE

RELIABILITY The system is designed for maximum reliability. The software design gives a high level of efficiency and accuracy. The software functions are separated to increase safety. REDUNDANCY There are two display electronic units (DEUs). Either DEU can drive all six of the display units. DISPLAY MANAGEMENT The display units are interchangeable. The displays automatically switch for any failure to provide safe displays. Display select panels let the flight crew manually reconfigure the display system. They let the pilots change the type of information that shows on the inboard, outboard and lower center display units. BITE Fault information shows on the flight management computer control display unit (FMC CDU.) The DEU can store faults in memory for sixty-four flight legs.

November 2000

5-1

Coax Coupler

Coax Coupler

Coax Coupler

Coax Coupler

Display Electronic Unit

Display Electronic Unit Control Panels

Control Panels

Common Display System Common Display System The common display system (CDS) shows information on six liquid crystal displays (LCDs). These display units (DUs) show primary flight, navigation, and engine information to the flight crew. The CDS has these components: • • • • • • •

Display electronic units (2) Coax couplers (4) Display units (6) EFIS control panels (2) Engine display control panel Display select panels (2) Display source selectors.

The CDS can show the primary flight and navigation data in two optional formats. These are the options: • •

The EFIS format shows information that looks similar to the 737-300/ -400/-500 flight compartment. The PFD/ND format shows information that looks similar to the 747 and 777 flight compartment displays. The CDS can show engine information in two optional formats. These are the options: • •

Side-by-side display Over/under display.

You load display electronic unit (DEU) software to change the CDS display formats.

The DEU gets input from the avionics and airframe systems. It calculates the graphics data in the correct display format and sends the information to two of the coax couplers. Each coupler sends information to all six DUs. This is for redundancy. If one DEU fails, the other can calculate all the necessary graphics data for all the DUs. If one coupler fails, a DU can use graphics data from any one of its other three inputs. The DUs are liquid crystal display units. The LCD is a sharper display over a wide range of lighting conditions than a cathode-ray tube unit.

EFIS PFD/ND.

5-2

November 2000

Common Display System

DEU

RADIO

MINS

FPV

BARO

DEU

MTRS

IN

BARO

RST VOR APP OFF

VOR 1

MAP

PLN

CTR

20 40 80 160 10 5

TFC

320

VOR 2

STA

WPT

ARPT

DATA

POS

ENG PRI

ND

MFD

TERR

EFIS Control Panels

Display Select Panels

N1 SET

SPD REF

AUTO 1

SOURCE AUTO

DISPLAYS CONTROL PANEL ALL ON 2

BOTH ON 1

BOTH ON 2

NORMAL

Display Source Selectors

2

VR WT V REF

FUEL FLOW

BOTH ON R

NORMAL

AUTO BRAKE DISARM a

V1

2

BOTH ON L

AUTO BRAKE

AUTO

BOTH

BOTH ON 2

NORMAL

ALL ON 1

ADF 2

IRS

VHF NAV BOTH ON 1

PFD INBD

OFF

640

LOWER DU NORM

MAIN PANEL DUs NORM ENG OUTBD PRI PFD

STD

ADF 1 WXR

NAVIGATION

HPA

1

3 MAX

OFF

RESET RTO

MFD RATE

ENG USED

SYS

ANTI SKID ANTI SKID INOP a

L E FLAPS TRANSIT a

LE FLAPS EXT g

Engine Control Module

Common Display System Common Display System CONTROL PANELS The control panels allow the pilots to select the type of information and the location for the displays. The EFIS control panel controls the primary flight and navigation information. They let the pilots control this information: • • • • • • • •

Radio and barometric minimum altitude Barometric reference Metric altitude display Weather radar TCAS traffic display Navigation display modes Navigation display information VOR/ADF display.

November 2000

The engine display control panel controls the engine display and the mach air speed indication (MASI) on the primary EFIS or PFD display. It lets the pilot select N1 reference bugs, see fuel flow or fuel used data, and select the speed reference bugs on the MASI. The display select panels let the pilots change the information that shows on a particular display unit to another display unit. For example, the pilots can select the engine display to show on the lower center display unit.

The display source selectors let the pilots select the source of data for the displays. They can select the source for inertial data and VOR. The instrument switching module also has a switch to select the DEU to send information to the display units. An EFIS control panel switch allows the pilots to select the EFIS control panel to use.

5-3

Capt Primary EFIS

Capt Secondary EFIS

F/O Secondary EFIS

F/O Primary EFIS

EFIS Display EFIS Layout The outboard display units show various displays including attitude and heading. The inboard display units also show various displays including the NAV display.

5-4

The indications on the captain and first officer display units are in different positions. They are different to keep this information in the basicT configuration: • • • •

For example, the mach airspeed indication (MASI) is on the outboard display unit for the captain. The MASI is on the inboard display unit for the first officer.

Airspeed Altitude Attitude Heading.

November 2000

Common Display System

Capt Primary EFIS

Capt Secondary EFIS

EFIS Display EFIS The captain outboard EFIS display has displays of these functions: • • • •

Mach airspeed indicator Radio distance magnetic indicator Attitude direction indicator Horizontal situation indicator, 210 degree format.

The inboard EFIS display has displays of these functions: • • •

Vertical speed indicator Baro altimeter indicator Navigation display.

November 2000

The first officer outboard EFIS display has displays of these functions:

The navigation display is one of seven formats. These are the formats:

• •

• • • • • • •

• •

Attitude direction indicator Horizontal situation indicator, 210 degree format Vertical speed indicator Baro altimeter indicator.

The inboard EFIS display has displays of these functions: • • •

Expanded VOR Centered VOR Expanded approach Centered approach Expanded map Centered map Plan.

Mach airspeed indicator Radio distance magnetic indicator Navigation display.

5-5

Centered VOR

Expanded VOR

VOR Display EFIS

This VOR information shows on the display:

VOR The VOR mode shows in a centered or expanded display format. The centered VOR mode shows 360 degrees of the compass rose with the airplane symbol and the lateral deviation bar in the center.

• • • • • • •

VOR deviation TO/FROM annunciation System source annunciation Station identification and frequency Station bearing Selected course DME distance.

This additional information shows: • • • • •

Ground speed True airspeed Wind speed and direction Weather radar information TCAS information.

The expanded VOR mode shows 60 degrees of the compass rose with the airplane symbol and the lateral deviation bar at the bottom.

5-6

November 2000

Common Display System

Expanded Approach

Centered Approach

ILS Display EFIS

This ILS information shows on the display:

This additional information also shows:

• • • • • •

• • • • •

APPROACH The approach mode shows in a centered or expanded display format. The centered approach mode shows 360 degrees of the compass rose with the airplane symbol and the lateral deviation bar in the center. Glide slope deviation shows on the side of the display.

Localizer deviation Glide slope deviation System source annunciation Station frequency Selected runway heading DME distance.

Ground speed True airspeed Wind speed and direction Weather radar information TCAS information.

The expanded approach mode shows 60 degrees of the compass rose with the airplane symbol and the lateral deviation bar at the bottom.

November 2000

5-7

Centered Map

Expanded Map

Map Display EFIS

This map information shows on the display:

This additional information also shows:

• • • • • • • • • • • •

• • • • •

MAP The map mode shows in a centered or expanded display format. The centered map mode shows 360 degrees of the compass rose with the airplane symbol in the center. The expanded map mode shows 60 degrees of the compass rose with the airplane symbol at the bottom.

5-8

Airplane track and heading Flight plan waypoints Flight plan path lines Active waypoints Airports Navigation aids Distance to go Estimated time of arrival Vertical path deviation Trend vectors FMC/IRU position difference FMC source.

Ground speed True airspeed Wind speed and direction Weather radar information TCAS information.

November 2000

Common Display System

Plan

Plan Display EFIS PLAN The plan mode shows in a centered format. The flight crew uses the plan mode to create, view or change a flight plan. The display is a north up display. The airplane symbol shows present position and FMC track.

November 2000

5-9

Compacted Display

Compacted Display EFIS COMPACTED DISPLAY The compacted display shows automatically when the inboard or the outboard display unit fails. Also, you can use the display select module to show the compacted display. The compacted display has: • • • • • •

Mach airspeed indicator Attitude direction indicator Altimeter indicator Radio distance magnetic indicator Horizontal situation indicator, 360 degree format Vertical speed indicator.

5-10

November 2000

Common Display System

PFD

ND

Primary Flight Display and Navigation Display PFD/ND

NAVIGATION DISPLAY

PRIMARY FLIGHT DISPLAY

The captain and the first officer have a navigation display (ND). The ND normally shows on the inboard display unit.

The captain and first officer have a primary flight display (PFD). The PFD normally shows on the outboard display unit. The PFD can also show on the inboard display unit.

The ND shows flight and navigation information in one of several formats. There are seven formats:

This shows on the PFD: • • • • • • • •

Airspeed parameters Attitude parameters Barometric altitude parameters Heading parameters Vertical speed Radio altitude parameters Flight mode annunciations Landing aids parameters.

November 2000

• • • • • • •

Expanded VOR Centered VOR Expanded APP Centered APP Expanded MAP Centered MAP Plan.

5-11

Expanded VOR

Centered VOR

VOR Display PFD/ND

This VOR information shows on the display:

VOR The VOR mode shows in a centered or expanded display format. The centered VOR mode shows 360 degrees of the compass rose with the airplane symbol and the VOR course deviation bar in the center.

• • • • • • •

VOR course deviation TO/FROM annunciation System source annunciation Station identification or frequency Station bearing Selected course DME distance.

This additional information shows: • • •

Ground speed True airspeed Wind speed and direction.

The expanded VOR mode shows 80 degrees of the compass rose with the airplane symbol and the VOR course deviation bar at the bottom.

5-12

November 2000

Common Display System

Centered APP

Expanded APP

Approach Display PFD/ND The APPROACH (APP) mode shows in a centered or expanded display format. The centered APP mode shows 360 degrees of the compass rose with the airplane symbol and the localizer deviation bar in the center. Glideslope deviation shows on the side of the display.

This ILS information shows on the display: • • • • • •

Localizer deviation Glide slope deviation System source annunciation Station identifier or frequency Selected runway heading DME distance.

This additional information shows: • • •

Ground speed True airspeed Wind speed and direction.

The expanded APP mode shows 80 degrees of the compass rose with the airplane symbol and the localizer deviation bar at the bottom.

November 2000

5-13

Expanded Map

Centered Map

Map Display PFD/ND

This map information shows on the display:

MAP The map mode shows in a centered or expanded display format. The centered map mode shows 360 degrees of the compass rose with the airplane symbol in the center. The expanded map mode shows 80 degrees of the compass rose with the airplane symbol at the bottom.

5-14

• • • • • • • • • •

Airplane track and heading Flight plan waypoints Flight plan path lines Active waypoints Airports Navigation aids Distance to go ETA Vertical path deviation Trend vectors.

November 2000

Common Display System

Plan Mode

Plan Display PFD/ND PLAN The flight crew uses the plan mode to create, view or change a flight plan. The display is a north up display. The airplane symbol shows present position and FMC track.

November 2000

5-15

A/T LIM

CRZ REV

TAI

-12c

96.0

TAI

87.7

10

TAT

87.7

10

0 8

0 8

2

6

6

4

N1

START VALVE OPEN OIL FILTER BYPASS LOW OIL PRESSURE

START VALVE OPEN OIL FILTER BYPASS LOW OIL PRESSURE

2 4

100

50

100

50

0

663

663

0 OIL P

200 EGT

200

100

100

X-BLD START

87.7

0

87.7

0 OIL T

75 N2 12

2 12

3.75

3.76 0

0

2

4

4

FF/FU LB X 1000

5

0

1

24 000

7 600

A 3

2

7 600

FUEL LB

4

2

RF

B 3

4

1

2

0

70

5

0

VIB

4

CTR

3

1

1

8

8

75

OIL Q %

3

HYD P HYD Q %

4

1

60

0 RF

Engine Display (Side-By-Side)

Side-By-Side Engine Display Engine Display

This engine information shows on the side-by-side engine display:

SIDE-BY-SIDE ENGINE DISPLAY The side-by-side engine display normally shows on the upper center display unit. It can also show on the lower center display unit or the inboard display units.

5-16

• • • • • • • • • • • • • • • •

N1 Thermal anti-ice indication EGT N2 Fuel flow/fuel used Fuel quantity Oil pressure Oil temperature Oil quantity Engine vibration Hydraulic pressure Hydraulic quantity Crew alert messages Autothrottle limit message Thrust mode Total air temperature.

November 2000

Common Display System

Engine Primary

Engine Secondary

Engine Display Engine Display

This engine information shows on the primary engine display:

This engine information shows on the secondary engine display:

• • • • • • • •

• • • • • •

OVER/UNDER ENGINE DISPLAY The over/under engine display normally shows the engine information on two displays. The primary engine display normally shows on the upper center display unit. The secondary engine display normally shows on the lower center display unit.

November 2000

N1 Thermal anti-ice indication EGT Fuel quantity Crew alert messages Autothrottle limit message Thrust mode Total air temperature.

N2 Fuel flow/fuel used Oil pressure Oil temperature Oil quantity Engine vibration.

5-17

Compacted Display

Compacted Engine Display Engine Display COMPACTED ENGINE DISPLAY The over/under engine display normally shows the engine information on two displays. The compacted display shows automatically when the upper center or the lower center display unit fails. The engine information that shows on the compacted engine display are all the primary and secondary engine information.

5-18

November 2000

Common Display System

Systems Display

Systems Display Systems Display The systems display shows when you select SYS on the engine control module. The systems display can show on the lower center display unit or either inboard display unit. This information shows on the systems display: • • • •

Hydraulic quantity Hydraulic pressure Brake temperature (option) Flight control surface position (option).

November 2000

5-19

M A I NT

EEC

B ITE

I NDE X

1 / 1

< F MC S

E NG I NE S >

< DF CS

A P U>



< A D I RS

APU

< CDS

Flight Management Computers (2)

< I NDE X

Display Electronics Unit (2)

FMC CDU Display BITE Maintenance data shows on the flight management computer control display unit (FMC CDU.) The maintenance data includes: • • •

CDS LRU faults Engine exceedance information Hardware and software configuration information.

There are several CDS tests. An FMC CDU menu permits access to these tests. The tests include: • • •

DEU self test DU loop status test DU optical test.

The DEU is also the interface for BITE information from the EEC and the APU.

The display electronic unit (DEU) stores sixty-four flight legs of faults.

5-20

November 2000

Communications and Recording Features

•

Flight Interphone

DIGITAL AUDIO CONTROL

•

Service Interphone

The digital audio control is a system that processes all audio information to, from, and in the airplane.

•

Flight and Ground Crew Call

•

Passenger Address

HF COMMUNICATIONS (OPTIONAL)

•

Audio/Video Entertainment

•

VHF Communications

•

HF Communications

•

SELCAL

•

Aircraft Communications Addressing and Reporting System (ACARS)

•

Voice Recorder

•

Flight Data Recorder

•

Aural Warning

•

Electronic Clocks

•

SATCOM

The HF antenna is a new design because of the 737-600/-700/-800 vertical stabilizer redesign. AIRCRAFT COMMUNICATIONS ADDRESSING AND REPORTING SYSTEM (ACARS) (OPTIONAL) ACARS is a digital data link that supplies communication between the airplane and ground operations with the VHF radio.

November 2000

6-1

MIC SELECTOR

w

w

w

w

1-VHF-2-VHF-3

HF

w

SERV INT

w

FLT INT

w

PA

AAU 1-NAV-2

Control Wheel Mic Switch

R/T

MASK

I/C

BOOM

1-ADF-2

V

B

MKR

SPKR

NORM

Audio Control Panel (3)

MOD

EXTERNAL POWER

CONN.

Oxygen Masks

A[]

B[]

REMOTE ELEC AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

Headsets

Headphones

CAPT

HDPH FLT SVR

Flight Interphone Speakers ATTENTION

C[]

D[]

ONICS UNIT

Communication Radios Navigation Receivers Passenger Address Service Interphone Voice Recorder

ENV CAT AAZ

GPWS

INTERPHONE

SERIAL NO. DATE MFD WEIGHT 7.75 LBS

FLIGHT SERVICE PILOT CALL NWW LIGHT ON

Hand Mics

F/O

HDPH FLT SVR

ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

ALT

R

OBS

EXT ATT

SVR INT SVR INT FLT INT DME 2

TCAS DFCS

Remote Electronics Unit NOT IN USE

NORMAL

External Power Panel

Flight Interphone System Flight Interphone System The flight interphone system lets the crew members in the flight compartment communicate with each other. It also connects with the audio communication system and the ground crew members. There are three independent systems, one for each flight crew station and the observer station. The captain system is shown on the graphic. The flight crew selects a system on the audio control panel (ACP) to transmit or receive audio. These are the systems that the pilot can select: • • • • •

Communication radios Navigation receivers Cabin interphone Passenger address Flight and service interphone.

6-2

When the pilot selects a system on the ACP, the remote electronics unit (REU) sends audio from the hand microphones, boom microphones or oxygen mask microphones to that system. The REU also sends the audio from the system to the headsets and speakers. The REU also integrates and sends warning audio to the headsets and speakers. The warning audio comes from these systems: • • •

Ground proximity warning system Traffic alert and collision avoidance system Digital flight control system.

The interphone/radio push-to-talk (PTT) switches are on the control wheels for use with the oxygen mask or boom microphones. The RADIO INT switch on the audio control panel does a similar function.

November 2000

Communications and Recording MIC SELECTOR

w

w

w

w

1-VHF-2-VHF-3

1-NAV-2

Control Wheel PTT Switch (2)

R/T

MASK

I/C

BOOM

HF

1-ADF-2

V

B

w

SERV INT

w

FLT INT

MKR

SERVICE INTERPHONE

w

PA

SPKR

AAU SVR INT SVR INT FLT INT DME 2

ALT

R

OBS

EXT ATT

F/O

HDPH FLT SVR

CAPT

OFF

HDPH FLT SVR

ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

ON

ATTENTION

Service Interphone Switch EXTERNAL POWER

INTERPHONE

NORM

MOD

A[]

B[]

C[]

D[]

Audio Control Panel (3)

CONN.

FLIGHT SERVICE PILOT CALL NWW LIGHT ON

Headsets Headphones Oxygen Masks Hand Mics 1 2 3 4 5 6 7 8 9 * 0 #

REMOTE ELEC AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

ONICS UNIT NOT IN USE

External Power Panel ENV CAT AAZ

LIGHTING

ENTRY

CEILING

WINDOW

Attendant Panel

NORMAL

SERIAL NO. DATE MFD WEIGHT 7.75 LBS

APU Aft Service Door EE Compartment Main Wheel Well Fueling Station

Remote Electronics Unit

Attendant Handset (2)

Service Interphone System Service Interphone The service interphone system is for communication between these personnel: • • •

Flight crew Cabin attendants Maintenance personnel.

They are in different areas around the airplane. Attendants use handsets at each attendant station to communicate on the service interphone system. The flight crew selects service interphone on the audio control panel.

A toggle switch in the flight compartment connects the external service interphone headset jacks to the service interphone. Service interphone headset jacks are in these areas: • • • • • •

External power panel APU Aft service door EE compartment Main wheel well Underwing fueling station.

The handsets send audio to the remote electronics unit. The REU amplifies the audio and sends it back to the handsets.

November 2000

6-3

1 2 3 4 5 6 7 8 9 0 #

*

Exit Locator Sign (2)

LIGHTING

Attendant Handset (2)

ENTRY

CEILING

WINDOW

CALL

Attendant Panel (2)

GRD CALL

ATTEND

b

Ground Crew Call Horn

Passenger Sign Module

EXTERNAL ________ _____ POWER

CONN.

Passenger Address Amplifier

__________ INTERPHONE

FLIGHT SERVICE PILOT CALL NWW LIGHT __________ _________ ON

NOT IN USE

NORMAL

External Power Panel

Aural Warning Module

Flight and Ground Crew Call System Flight and Ground Crew Call System The flight and ground crew call system permits call signals between these areas: • • •

Flight compartment and cabin attendant stations One cabin attendant station to another cabin attendant station Flight compartment and ground crew.

The system tells personnel to use the service interphone.

6-4

The flight compartment can call a cabin attendant station with the ATTEND switch on the forward overhead panel. A pink light at the forward and aft exit locator signs turn on and there is a two-tone chime from the passenger address system. A cabin attendant can call another attendant station with the 5 pushbutton on the handset. The exit locator pink light at the called station comes on and there is a chime from the passenger address system.

A crew member pushes the GRD CALL switch in the flight compartment to call the ground crew. The switch is on the forward overhead panel. A horn in the nose wheel well comes on when the pilots use this switch. The ground crew pushes the PILOT CALL switch on the external power panel to call the flight compartment. A call light on the forward overhead panel comes on and there is a chime from the aural warning module.

A cabin attendant pushes the 2 pushbutton on the handset to call the flight compartment. A call light on the forward overhead panel comes on and there is a chime from the aural warning module.

November 2000

Communications and Recording MIC SELECTOR

w

w

w

w

w

w

w

AAU 1-VHF-2-VHF-3

1-NAV-2

Control Wheel PTT Switch (2)

R/T

MASK

I/C

BOOM

HF

1-ADF-2

V

B

FLT INT

SERV INT

MKR

PA

SPKR

SVR INT SVR INT FLT INT DME 2

OBS F/O CAPT

EXT ATT

HDPH FLT SVR

ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

HDPH FLT SVR

ATTENTION

-1

ALT

R

NORM

Audio Control Panel (3) Headsets Headphones Oxygen Masks Hand Mics

Attendant Area Speakers

REMOTE ELEC AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

0 db

ONICS UNIT

+1

Either Engine On

NORM TEST

CAL

ENV CAT AAZ

Passenger Service Unit Speakers

SERIAL NO. DATE MFD WEIGHT 7.75 LBS

Remote Electronics Unit

Lavatory Speakers 346D-2B

PA Control Stand Mic

NO FASTEN SMOKING BELTS

1 2 3 4 5 6 7 8 9 * 0 #

OFF AUTO ON

ATTEND

CALL

Attendant Handset (2)

Passenger Address Amplifier

GRD CALL

b

Passenger Sign Module

Tape Reproducer

Passenger Address System Passenger Address System The passenger address (PA) system gives announcements and music to the passenger compartment. The pilots or cabin attendants can make announcements. The pilot announcements have the highest priority. The pilots make announcements with the microphones and the PA selection on the audio control panel. Attendant announcements have second priority. An attendant pushes the eight push-button and the PA push-to-talk switch on a handset to make announcements.

The PA amplifier sends the highest priority audio to all the speakers in the passenger compartment. The PA amplifier decreases the output level of the audio when the engines are on. The PA amplifier supplies audio for the attendant area speakers through the remote electronics unit. An attendant announcement mutes this audio to stop microphone feedback. The PA amplifier also supplies the chimes. Chimes are superimposed over any audio.

Prerecorded announcements from the tape reproducer have third priority. Boarding music from the tape reproducer has fourth priority.

November 2000

6-5

videoHi8

VIDEO CASSETTE PLAYER EVP-90B

-1

0 db

+1

Video Monitor

Video Reproducer SEB CTT BRKR

TEST

NORM

CAL A

B

C

Passenger Control Unit 346D-2B

Audio Seat Electronics Unit

Passenger Address Amplifier

Main Multiplexer Audio Reproducer

Audio/Video Entertainment Audio/Video Entertainment (Optional) Audio and video entertainment systems are optional equipment. A typical audio entertainment system has these units: • • • • •

Audio reproducer Main multiplexer Seat electronics boxes Passenger control units Headsets.

A video entertainment system may be added to the audio entertainment system. The typical video system has these units: • •

Audio from the video entertainment system goes through the audio entertainment system. The audio reproducer sends audio to the main multiplexer. The multiplexer changes the analog audio signals to a digital format and sends the digital signals to the seat electronics units. The passenger then selects the audio channel with the passenger control unit. The passenger hears the audio in their headset. The passenger address system sends audio to the main multiplexer. Passenger address audio has priority over all the entertainment audio.

Video tape reproducer Video monitors.

6-6

November 2000

Communications and Recording ACTIVE

STANDBY

AAU OFF

SVR INT SVR INT FLT INT DME 2

TEST

VHF1

VHF2

VHF3

HF1

AM

HF2

Radio Communication Panel (3)

w

w

w

1-VHF-2-VHF-3

HF

w

SERV INT

w

FLT INT

HDPH FLT SVR ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

HDPH FLT SVR

LRU

CONTROL

ATTENTION

ANTENNA

MIC SELECTOR

w

OBS F/O CAPT

EXT ATT

REMOTE ELEC AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

w

ONICS UNIT

PHONE ENV CAT AAZ

PA

MIC SERIAL NO.

1-NAV-2

R/T

MASK

I/C

BOOM

1-ADF-2

V

B

MKR

R

SPKR

DATE MFD WEIGHT 7.75 LBS

VHF-900

FDAU

Remote Electronics Unit

ALT

VHF Antenna (3)

SELCAL Decoder

NORM

Audio Control Panel (3) VHF Transceiver (3)

Headsets Headphones Oxygen Masks Hand Mics ACARS Management Unit

VHF Communication Systems VHF Communication System The very high frequency (VHF) communication system supplies line of sight voice and data communications from air-to-ground or air-to-air. A dual VHF communication system is basic. A third VHF transceiver is available as an option.

The transceiver sends and gets audio and data from the antenna. The remote electronics unit does these functions for transmissions with inputs from the audio control panel: • • •

A radio tuning switch, on a radio communication panel (RCP) selects one of the transceivers. The frequency selectors select the desired frequency. This shows on the liquid crystal display standby frequency window. The frequency transfer switch moves the standby frequency to the active frequency. The RCP sends tuning data to the selected transceiver.

November 2000

Microphone selection Headphone or speaker monitoring PTT.

The aircraft communications addressing and reporting system (ACARS) controls data to an optional third VHF communication system. ACARS is an optional system. When installed, ACARS uses the VHF transceiver to get and send digital data to and from a ground station.

6-7

ACTIVE

STANDBY

OFF

HF Antenna

TEST

VHF1

VHF2

VHF3

HF1

AM

HF2 LRU FAIL

Radio Communication Panel (3)

AAU SVR INT SVR INT FLT INT DME 2

MIC SELECTOR

w

w

w

w

1-VHF-2-VHF-3

1-NAV-2

R/T

MASK

I/C

BOOM

HF

w

SERV INT

1-ADF-2

V

B

w

FLT INT

MKR

R

w

KEY INTERLOCK

OBS F/O CAPT

EXT ATT

HDPH FLT SVR

ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

HDPH FLT SVR

SQL/LAMP TEST ATTENTION

PA

FDAU

CONTROL INPUT FAIL

HFS-700

SELCAL Decoder

PHONE MIC

SPKR REMOTE ELEC AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

ALT

ONICS UNIT

HF Transceiver (2) ENV CAT AAZ

NORM SERIAL NO. DATE MFD WEIGHT 7.75 LBS

Audio Control Panel (3)

Remote Electronics Unit

Headsets Headphones Oxygen Masks Hand Mics

HF Antenna Coupler (2)

HF Communication Systems HF Communication System The high-frequency (HF) communication system is for longrange voice communications. Each HF communication system has these units: • • •

HF transceiver HF antenna coupler Shared antenna.

A radio tuning switch on the radio communication panel (RCP) selects one of the transceivers. The frequency selectors select the desired frequency. This shows on the liquid crystal display standby frequency window. The frequency transfer switch moves the standby frequency to the active frequency. The RCP sends tuning data to the selected transceiver.

6-8

The transceiver sends and gets the audio. The remote electronics unit does these functions with audio control panel selections: • • •

Microphone selection Headphone and speaker monitoring PTT.

The antenna and the antenna couplers are in the vertical stabilizer. The antenna coupler matches the impedance of the antenna to the impedance of the transceiver. The coupler tunes when you first key the HF transmitter.

November 2000

Communications and Recording MIC SELECTOR

w

w

w

w

1-VHF-2-VHF-3

HF

w

SERV INT

w

FLT INT

w

PA

LRU FAIL KEY INTERLOCK 1-NAV-2

CONTROL INPUT FAIL

1-ADF-2

MKR

SPKR

Alert/Reset

SQL/LAMP TEST

HFS-700

R/T

MASK

I/C

BOOM

PHONE MIC

V

B

R

ALT

NORM

Audio Control Panel (3)

HF Transceiver (2) AAU SVR INT SVR INT FLT INT DME 2

LRU CONTROL ANTENNA

MOD

OBS

EXT ATT

F/O

HDPH FLT SVR

ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2

A[]

CAPT

HDPH FLT SVR

ATTENTION

B[]

C[]

D[]

PHONE

REMOTE ELE

MIC

VHF-900

SELCAL Coding Switch

ONICS UNIT

AVTECH P/N

Select Aural Warning Relay

BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

ENV CAT AAZ

Aural Warning Module

SERIAL NO.

SELCAL Decoder

VHF Transceiver (3)

DATE MFD WEIGHT 7.75 LBS

Remote Electronics Unit

SELCAL System SELCAL (Optional) The selective calling (SELCAL) system monitors all communication radios on the airplane. The system alerts the flight crew when it gets a ground call with the correct airplane code. This reduces the flight crew workload because they do not have to continuously listen to the airline communication frequencies.

When the SELCAL gets a call, these things occur: • •

Call light on the audio control panel SELCAL chime from the aural warning module.

When the flight crew selects a PTT for the applicable radio, the light goes off and the system resets.

The SELCAL decoder gets audio signals from the VHF and HF communication systems. The SELCAL decoder gives signals to the flight crew if the signal received has the airplane unique SELCAL code. The airplane unique SELCAL code comes from the SELCAL coding switch.

November 2000

6-9

VHF 3 Antenna Selected Voice Frequency Data Remote Data Line Station

Data Line Control ARINC

ARINC

Telephone System Airline Dispatch Communications

ARINC

Voice Station

Airline Land Lines

FLT DEP FUEL DES A B C D E 1 2 F G H I J 4 5 K L M N O 7 8 P Q R S T + 0 U V W X Y Z

VHF Transceiver Satellite Data Unit

3 6 9 --

MENU CLEAR

FULL

YOX ENTER

EMPTY PAPER

Interactive Display Unit Proximity Switch Electronics Unit Program Switch Modules Program Pins ACARS Management Unit

Multipurpose Printer Flight Management Computer System Flight Data Acquisition Unit Airshow Digital Interactive Unit

Aircraft Communications Addressing and Reporting System (ACARS) (Optional) ACARS (Optional) The aircraft communication addressing and reporting system (ACARS) gives a high-speed digital datalink between the airplane and ground facilities. ACARS reduces flight crew workload by automatically transmitting and receiving data. This is the type of data the ACARS transmits and receives: • • •

• • • •

Airplane identification Flight identification Out of gate, off the ground, on the ground, into the gate (OOOI) reports Delay reports Fuel reports Weather Airplane operating data.

6-10

ACARS also provides voice telephone patch communication between the airplane and ground telephone circuits. It uses VHF radio, airline land lines, ARINC lines, and telephone systems. The main ACARS component is the ACARS management unit (MU). The MU uses program pins and a program switch module for airline identification and the proximity switch electronics unit (PSEU) to define the OOOI event parameters. The MU uses VHF transceiver 3 to receive and transmit data.

The flight crew uses the multipurpose printer to print ACARS reports stored in the MU. A call light in the flight compartment comes on, and there is a chime when ACARS receives a voice request from a ground station. The IDU shows the requested voice communication frequency to the flight crew. The flight crew can use the IDU to tune the VHF transceiver to make a voice call to the ground.

The flight crew controls the system with an interactive display unit (IDU) or the control display unit (CDU).

November 2000

Communications and Recording

COCKPIT VOICE RECORDER MICROPHONE MONITOR

Parking Brake Set and Ground Sensing Relays

Erase Test

ERASE TEST r g

STATUS

Audio Channel 4

HEADPHONE

Voice Recorder Control Panel Clock AAU SVR INT SVR INT FLT INT DME 2

OBS F/O CAPT

HDPH FLT SVR ADJ DME 1 ADJ PA SENS PA GAIN PA ST AUDIO POT 1 AUDIO POT 2 EXT ATT

HDPH FLT SVR

ATTENTION

Audio Channel 1 Audio Channel 2 Audio Channel 3

Headphones

Voice Recorder

Oxygen Masks Hand Mic

REMOTE ELE AVTECH P/N BOEING P/N TSO C50c RTCA 00-16 B2AKXXXXFX

ONICS UNIT

ENV CAT AAZ

SERIAL NO. DATE MFD WEIGHT 7.75 LBS

Remote Electronic Unit

Voice Recorder System Voice Recorder System The voice recorder makes a continuous record of the last 120 minutes of flight crew communications. It erases automatically so that only the last 120 minutes are in memory. The voice recorder records four audio channels. These are the four channels: • • • •

The test switch on the control panel does a functional check of the system. The bulk ERASE switch on the control panel erases all stored audio in the voice recorder. The switch erases the audio only when the airplane is on the ground with the parking brake on. An underwater locator beacon is on the front of the voice recorder.

Captain microphone First officer microphone First observer microphone Area microphone on the voice recorder control panel.

Microphone inputs from the captain, first officer, and first observer go to the remote electronic unit (REU). The REU amplifies the audio and sends it to the voice recorder. The audio from the area microphone is amplified at the microphone and goes directly to the voice recorder. November 2000

6-11

MACH AIRSPEED WARNING TEST NO 1 NO 2

FLIGHT RECORDER TEST

NORMAL OFF

a

Flight Recorder/Mach Airspeed Warning Test Module

Flight Data Recorder

115V AC FDAU Status Relay Control Wheel Position Sensors (2) Control Column Position Sensors (2) Rudder Pedal Position Sensor Rudder Position Transmitter Aileron Position Transmitters (2) Elevator Position Transmitters (2) Accelerometer Program Switch Module Discrete Flight Data Analog Flight Data Digital Flight Data

System Test Plug/Connector DATA LOAD SELECT ACMS

NORM

DFDAU

Data Loader Control Panel Quick Access Recorder Printer Flight Data Acquisition Unit (FDAU)

Flight Data Recorder System Flight Data Recorder System The flight data recorder system (FDRS) records the last 25 hours of flight parameters in a crash proof container. It records parameters that are required by regulatory agencies and requested by the airline. The flight data recorder system operates automatically when either engine is operating or the airplane is in the air. A flight recorder panel shows the status of the flight recorder system. If there is a system fault, an amber OFF light shows. The OFF light also comes on if the system is off.

Airplane systems send digital, discrete, and analog data to the flight data acquisition unit (FDAU). The FDAU formats the data. The FDAU sends the data to the flight data recorder (FDR). The FDR records the data in a fire and crash resistant LRU. An underwater locator beacon is on the front of the FDR. An optional FDAU can also collect data for the airplane condition monitoring system (ACMS). The ACMS is an optional system. The FDAU sends the ACMS data to a quick access recorder or to a diskette in its internal disk drive.

The flight recorder panel also has a TEST/NORMAL switch. When the TEST/NORMAL switch is in the TEST position, the recorder receives power on the ground.

6-12

November 2000

Communications and Recording

Takeoff Warning System Cabin Altitude Warning System Landing Gear Warning System Digital Flight Control System Mach Warning System Fire Warning System SELCAL System Crew Call System

Bell Horn Clacker Wailer Chimes Aural Warning Module

FIRE WARN BELL CUTOUT r

MASTER CAUTION PUSH TO RESET a

FLT CONT

ELEC

ANTI-ICE

ENG

IRS

APU

HYD

OVERHEAD

OVHT/DETa

DOORS

FUEL

a

a

AIR COND

MASTER CAUTION

FIRE WARN

PUSH TO RESET a

BELL CUTOUT r

a

Aural Warning System Aural Warning System The aural warning system tells the flight crew of incorrect airplane system conditions with aural indications. The system also provides the aural indication for SELCAL and crew calls. The aural warning module monitors several airplane systems. The airplane systems send a signal to the aural warning module if they detect an unsafe condition.

The aural warning module supplies these aural indications: • •

• • • •

Bell for a fire Intermittent horn for a unsafe takeoff configuration or cabin altitude too high Steady horn for unsafe landing configuration Clacker for overspeed Wailer for autopilot disconnect Chimes for crew call and SELCAL.

There are other independent warning systems. These systems are not part of the aural warning system: • • •

The ground proximity warning system The traffic collision and avoidance system The digital flight control system.

For an unsafe condition, there are also visual indications. The caution lights and the fire warning lights are the visual indications. These lights come on when airplane systems detect an unsafe condition.

November 2000

6-13

CHR

TIME/DATE MAN UTC

60

FMC VALID

SET

TIME

10

50 DAY

Flight Management Computer

MOYR

20

40 ET

PWR ON

FDAU

ET CHR

RUN HLD

30 RESET Clock

Voice Recorder

Electronic Clocks Electronic Clocks There are two clocks in the flight compartment, one on the captain and the other on the first officer instrument panel. Each clock shows this information: • • • •

Time Date Elapsed time in hours and minutes Chronograph time in minutes and seconds.

The captain clock sends time on an ARINC 429 bus to the flight data acquisition unit (FDAU), the flight management computer, and the voice recorder.

6-14

November 2000

Communications and Recording

Satellite Network

Aircraft Earth Station (AES)

Passenger Telephone System

High Gain Antenna

LNA/Diplexer

BSU

Combiner

ACARS MU

High Gain Antenna

SDU

REU

RFU

High Power Amplifier

High Power Relay

LNA/Diplexer

BSU

Ground Earth Station (GES)

SATCOM System Satellite Communication (SATCOM) System (Optional)

AIRCRAFT EARTH STATION DESCRIPTION

point the beam at the desired satellite.

The SATCOM system sends and gets data and voice messages. The system uses satellites as relay stations for long distance communication. SATCOM is more reliable than HF communication because it is not affected by atmospheric conditions.

The basic SATCOM configuration is a high-gain system with high-gain side-mounted antennas.

The AES interfaces with the ACARS MU that send and get data messages. The AES also interfaces with the passenger telephone system and the remote electronics unit for voice call audio and control signals.

The system is the satellite network, the ground earth station (GES), and the aircraft earth station (AES). The satellite network relays radio signals between the AES and the GES. Each GES is a fixed radio station that interfaces with communication networks through ground links and the aircraft earth stations through the satellite. The AES is the SATCOM system on the airplane that interfaces with various onboard communication systems and the ground earth stations. November 2000

The satellite data unit (SDU) is the interface between all other related airplane systems and the SATCOM system. The radio frequency unit (RFU) changes the signal from the SDU to an L-band signal for the high power amplifier (HPA). The HPA supplies adequate radio frequency power to the antenna. The low noise amplifier (LNA) and diplexer are one unit. The diplexer couples transmit signals from the HPA to the antenna. It also couples signals from the antenna to the LNA. The LNA amplifies the low level L-band signal from the antenna. The SDU sends directional control signals to the beam steering unit (BSU). The BSU electronically steers the antenna to

6-15

Navigation Features

•

Air Data Inertial Reference System

•

VHF Omnidirectional Range (VOR) System

•

Marker Beacon System

•

Instrument Landing System

•

Distance Measuring Equipment

•

Automatic Direction Finder System

•

Radio Altimeter System

•

Air Traffic Control System

•

Traffic Alert and Collision Avoidance System

•

Weather Radar System

•

Ground Proximity Warning System

•

Global Positioning System

•

Head-Up Display System (HUD)

NEW AVIONICS The 737-600/700/800/900 uses some new avionics components such as the air data inertial reference unit (ADIRU). The ADIRU puts the air data computer and the inertial reference unit together in a single unit. This saves space and weight. UPGRADED AVIONICS All navigation systems now use ARINC specification 429 for digital communication between units and systems. This results in increased reliability, fewer components, and fewer wires. SATELLITE NAVIGATION The 737-600/700/800/900 offers the newest in navigation systems, the global positioning system (GPS). GPS gives improved navigation accuracy along the flight plan. This saves fuel and improves the airplane on-time performance. COMMON COMPONENTS The 737-600/700/800 uses many navigation system units that are common with the 777, 747-400, 767, 757, and 737-300. This reduces the cost of maintenance and spares. HEAD-UP DISPLAY SYSTEM The head-up display system uses electronics and optics to calculate and display flight and guidance symbols. The symbols project onto a transparent glass screen in the forward field of view of the pilot. The symbols overlay and optically combine with the outside view through the windshield.

November 2000

7-1

F/O Pitot Probe

Capt Static Port

AOA Sensor

Alt Pitot Probe

F/O Static Port

Alt Static Port

To Stby Airspeed Right ADIRU

ADM

ADM

To Stby Inst and Cabin Press

To User Systems ADM

Left ADIRU ADM

Capt Pitot Probe

Capt Static Port

TAT Probe AOA Sensor

F/O Static Port

Alt Static Port

Air Data Inertial Reference System - Air Data Function Air Data Inertial Reference System

The ADIRU uses the ADM, TAT, and AOA data to calculate these values:

The air data inertial reference system (ADIRS) has two separate functions in a single line replaceable unit (LRU). The two functions use the same power supply. All other operations are separate. The air data function is active when electrical power is on. The inertial reference function is active when the pilots select it on.

• • • • • • • • • •

Altitude Computed Airspeed Mach Air temperature Angle-of-attack Baro-corrected altitude Maximum allowable airspeed True airspeed Altitude rate Static air temperature.

The ADIRU sends these values on digital data buses to the systems that use calculated air data values. The alternate pitot probe and the alternate static ports send the pressure to standby instruments. The cabin pressure control system also uses alternate static pressure.

AIR DATA FUNCTION The air data modules (ADMs) give pitot and static pressure data to the air data inertial reference unit (ADIRU). The ADMs convert pressure to digital data. The total air temperature (TAT) probe gives TAT to the ADIRU. The angle-of-attack (AOA) sensors give AOA data to the ADIRU.

7-2

November 2000

Navigation IRS DISPLAY

DSPL SEL HDG

V O R

Inertial System Display Unit

V O R

ADF

ON DC

ALIGN

ON DC

FAULT

DC FAIL

FAULT

DC FAIL

ALIGN

NAV

ALIGN

NAV

OFF

OFF

ATT L

IRS

Flight Data Acquisition Unit

ADF

Radio Magnetic Indicator

HG2050 ADIRU ALIGN

TCAS Computer

Flight Management Computer Autothrottle Computer

ATT

R

HUD Computer

Mode Select Unit Weather Radar Transceiver

ADIRU (2) CDS Display Electronic Unit

CDS Display Units (2)

Stall Management Yaw Damper

IRS Transfer Switch

Flight Control Computer (2)

FMC CDU

Anti-Skid Autobrake Control Unit

Ground Proximity Warning Comp

Air Data Inertial Reference System - Inertial Reference Function Air Data Inertial Reference System

• •

INERTIAL REFERENCE FUNCTION

The inertial reference function has these components:

The ADIRS inertial reference function uses laser gyros and accelerometers to measure airplane movement. The inertial reference function uses the gyro and accelerometer data to calculate these values:

• • •

• • • • • • • • • • • •

Attitude (pitch, roll, yaw) Position (latitude, longitude) True heading Magnetic heading Inertial velocity vectors Linear accelerations Angular rates Track angle Wind speed and direction Inertial altitude Vertical speed Ground speed

November 2000

•

Drift angle Flight path angle.

ADIRUs (2) Mode select unit (MSU) Inertial system display unit (ISDU) IRS transfer switch.

You use the MSU to select the mode of operation for the ADIRU inertial reference function. You can select these modes: • • • •

Off Align Navigate Attitude.

Before the ADIRUs can operate in the navigation mode, they must do an alignment. The airplane must not move during the alignment. You can use the flight management computer (FMC) control display unit (CDU) or

the ISDU to put the airplane present position into the ADIRU. Inertial reference data shows on the common display system (CDS) display units, the HUD, and on the radio magnetic indicator (RMI). Navigation, autoflight and other airplane systems also use inertial reference data. The IRS transfer switch selects the ADIRU that gives inertial reference data to the display system. When the switch is in the NORMAL position, the left ADIRU gives information to the captain displays and the right ADIRU gives information to the first officer displays. You move the switch to BOTH ON LEFT or BOTH ON RIGHT to cause one ADIRU to give information to all displays.

7-3

Flight Management Computer

VOR

Flight Control Computer

MKR DATA IN

HUD Computer

TEST

Navigation Control Panel

VHF NAV BOTH

BOTH

ON 2

ON 1

CDS Display Electronic Unit

NORMAL

VHF NAV Transfer Switch

HDG

AAU OBS

VOR Receiver (2)

F/O CAPT

V O R

ATTENTION

MOD A

B

C

D

DFCS Mode Control Panel

VOR/LOC Antenna

Remote Electronics Unit

V O R

ADF

ADF

Radio Magnetic Indicator

Flight Data Acquisition Unit

VHF Omnidirectional Range (VOR) System VHF Omnidirectional Range (VOR) System

These units use the VOR information:

The VHF omnidirectional range (VOR) system gives bearing information to ground stations. The pilots and the airplane systems use this information.

• •

The pilots use the navigation control panel to tune the VOR receiver. The VOR system receives the signals from the ground station and calculates magnetic bearing to the station.

• • • •

Flight control computer (FCC) Flight management computer (FMC) Common display system display electronic unit (CDS DEU). Radio magnetic indicator (RMI) Flight data acquisition unit (FDAU) Head-up display system (HUD).

Magnetic bearing shows on the radio magnetic indicator (RMI) and on the common display system (CDS) display units.

The VHF NAV transfer switch selects the VOR data that shows on the CDS display unit. In the NORMAL position, VOR 1 data shows on the captain display unit and VOR 2 data shows on the first officer display unit. You move the switch to BOTH ON 1, or BOTH ON 2 to cause one VOR receiver to give data to the captain and first officer. The VOR receivers send audio from the VOR station to the remote electronics unit (REU). The REU sends the audio to the flight compartment speakers and pilot headsets.

The CDS DEUs use the selected course from the DFCS mode control panel to calculate VOR course deviation.

7-4

November 2000

Navigation AAU OBS

F/O CAPT

VOR MKR DATA IN TEST

ATTENTION

MOD A

B

C

D

Marker Beacon Antenna

VOR Receiver 1

Remote Electronics Unit

Display Electronic Unit (2) CDS Display Unit (2)

HUD Computer

Airway Intersection

Marker Beacon Antenna

Flight Data Acquisition Unit

Distance from end of runway different for each location

Marker Beacon System Marker Beacon The marker beacon system gives aural and visual indications in the flight compartment as the airplane flies over a marker beacon transmitter. Marker beacon transmitters are on flight paths and runway approach paths. The transmitter sends a narrow, vertical radio frequency beam with an audio tone. Marker beacon transmitters on runway approach paths send one of three different audio tones.

November 2000

As the airplane flies over the beam, the marker beacon system receives the audio tone. The marker beacon system then sends the aural and visual indications to the common display system, flight data acquisition unit (FDAU), and the head-up display (HUD) system. Each VOR receiver has a marker beacon module. The marker beacon functions only in VOR receiver 1.

7-5

AAU OBS

F/O CAPT

Standby Attitude Indicator

ATTENTION

MOD A

Navigation Control Panel

B

C

D

PWR FMC ON VALID TEST OK MMR FAULT BUS IN FAIL TEST ANT

Flight Management Computer

TEST

Remote Electronics Unit

Glide Slope Antenna

Ground Proximity Warning Computer

MMR RECEIVER

Multi- Mode Receiver (2)

Localizer Antenna

VOR/ Localizer Antenna

Flight Control Computer B

ILS Relay (2)

Integrated Flight Systems Accessory Unit

Flight Control Computer A

Flight Data Acquisition Unit

HUD Computer

CDS Display Electronic Unit

Instrument Landing System Instrument Landing System The instrument landing system (ILS) gives precision approach guidance on instrument approaches. The ILS gives position information to the glidepath and runway center line. Two ILS systems are on the airplane. Each system operates the same.

The multi-mode receivers use the VOR/ localizer antenna on the vertical stabilizer until the flight control computer (FCC) changes the multi-mode receivers to the localizer antenna. This happens on approach. ILS deviation shows on the CDS display units, the head-up display, and on the standby attitude indicator.

The ILS system is active when a pilot selects an ILS frequency on the navigation control panel. The control panel also sends course information to the multi-mode receivers.

These units use ILS information:

The multi-mode receivers calculate up and down glidepath deviation from the signal it receives from the glide slope antenna. They calculate left and right deviation from the signal they receive from the localizer antennas.

•

7-6

ILS audio goes to the remote electronics unit (REU). The REU sends audio to the flight compartment speakers and headsets.

• • •

• • •

Flight control computers (FCCs) Standby attitude indicator Ground proximity warning unit (GPWS) Flight management computer (FMC) Flight data acquisition unit (FDAU) Common display system display electronic units (DEUs) Head-up display (HUD) computer.

November 2000

Navigation AAU

OBS

F/O CAPT

ATTENTION

MOD A

B

C

D

Antenna Flight Control Computer A Navigation Control Panel

Flight Data Acquisition Unit

DME Interrogator 1

Remote Electronics Unit PWR FMC ON VALID

HUD Computer

Flight Management Computer

CDS Display Electronic Unit

CDS Display Unit

Distance Measuring Equipment System Distance Measuring Equipment System The distance measuring equipment (DME) system gives the pilots distance to a DME ground station. Two DME systems are on the airplane. Each system operates the same. These units use the distance information: • • • • •

Flight control computers (FCCs) Flight data acquisition unit (FDAU) Display electronic units (DEUs) Flight management computer (FMC) Head-up display (HUD) computer.

November 2000

The pilot can tune one DME station with the navigation control panel. This DME distance shows on the CDS display units. The FMC, HUD and the FCC also use this distance. The FMC can tune up to four DME stations at the same time. The FDAU and the FMC use these DME distances. DME audio goes to the remote electronics unit. The pilots can hear the audio on the flight compartment headsets and speakers.

7-7

ADF Antenna HDG

V O R

V O R

ADF

Radio Magnetic Indicator

ADF

ADF Antenna

HUD Computer

CDS Display Unit CDS Display Electronic Unit

Ground Radio Station

AAU OBS F/O CAPT

MOD A

B

C

D

ADF Receiver ADF Control Panel

Remote Electronics Unit

Automatic Direction Finder System Automatic Direction Finder System The automatic direction finder (ADF) system is a navigation aid that receives radio signals from a ground station. The ADF calculates bearing information to a ground station. It can also give audio to the pilots.

You can tune the ADF receiver to receive and calculate bearing to any radio transmitter with a frequency between 190 and 1750 kHz. Bearing shows on the radio magnetic indicator (RMI), the HUD, and on the common display system display units. The receiver sends audio to the remote electronics unit (REU).

The audio is a station identifier or a broadcast from a radio station. Some ADF stations also supply weather information. The ADF system has these components: • • •

ADF receiver Control panel Combined loop and sense antenna.

7-8

November 2000

Navigation RADIO

MINS

MTRS

FPV BARO

IN

MAP

VOR 1

VOR APP OFF

PLN CTR

20 10 5

40 TFC

ADF 1 WXR

CDS Display Electronic Unit

Transmit Antenna

Receive Antenna

Flight Control Computer

Radio Altimeter Transceiver (2)

Auto Throttle Computer

STA

BARO HPA STD

RST

WPT

ARPT

VOR 2 80 160 320 OFF 640 ADF 2

DATA

POS

TERR

EFIS Control Panel (2)

TCAS Computer

Ground Proximity Warning Computer

Flight Data Acquisition Unit

HUD Computer

Radio Altimeter System Radio Altimeter System

These units use radio altitude:

The radio altimeter (RA) system gives the pilots and airplane systems the altitude of the airplane above the ground. The RA system operates from 0 to 2500 feet.

• • • •

The system has two transceivers. Each transceiver has a transmit antenna and a receive antenna. The transmit antenna sends a signal to the ground which comes back to the receive antenna. The transceiver uses the time between transmission and reception to calculate the altitude above the ground. Radio altitude shows on the CDS display units when radio altitude is 2500 feet or less.

•

November 2000

• •

Flight control computers (FCCs) Autothrottle computer Display electronic units (DEUs) Traffic alert and collision avoidance system (TCAS) computer Ground proximity warning computer (GPWC) Flight data acquisition unit (FDAU) Head-up display (HUD) computer.

Each pilot can set a radio minimums altitude on the EFIS control panel. The radio minimums shows on the CDS display unit, and on the HUD. When the radio altitude is equal to or less than the radio minimums, the radio minimums and radio altitude change color and size, and momentarily flash.

7-9

Top ATC/Mode S

Bottom ATC/Mode S Antenna Location

MODE S TRANSPONDER

A

Antenna Switch

TCAS Computer

ADIRU (2)

A A ATC Control Panel

Antenna Switch

ATC Transponders (2)

Air Traffic Control System Air Traffic Control System The air traffic control (ATC) system has the airborne components that the TCAS computers and the ground facilities use to track the airplane movement. The ATC transponder replies to interrogation signals from the ground and from TCAS airplanes. The reply to most ground stations is airplane code (mode A) or airplane altitude (mode C). Selective calling (mode S) ground stations enhance the operation of the ATC system because it adds a discrete interrogation capability and a data link feature.

7-10

You use the ATC control panel to set the airplane code, select transponder 1 or 2, and start an identification pulse. The altitude source select switch on the control panel selects the ADIRU that supplies altitude to the transponder. Only one ATC transponder is active at a time. Antenna switches connect the ATC antennas to the active transponder. The ATC transponder also works with the traffic alert and collision avoidance system (TCAS).

November 2000

Navigation AAU OBS F/O CAPT

ATTENTION

MODE S TRANSPONDER

Antenna Switches

Top TCAS (Directional Antenna)

MOD A

B

C

D

PASS TPR

FAIL

UPPER ANT LOWER ANT ALT CTL

Remote Electronics Unit

TEST

ATC Transponder (2)

ATC Control Panel

HUD Computer

FDAU TCAS Computer CDS Display Electronic Unit

Radio Altimeter (2) Ground Proximity Warning Computer

RADIO

Left ADIRU

Bottom TCAS (Directional or Omnidirectional Antenna)

MINS

FPV

MTRS

BARO

IN

MAP

VOR 1

VOR APP OFF

PLN CTR

20 10 5

40 TFC

ADF 1 WXR

STA

BARO HPA STD

RST

WPT

ARPT

VOR 2 80 160 320 OFF 640 ADF 2

DATA

POS

TERR

EFIS Control Panel (2)

Traffic Alert and Collision Avoidance System Traffic Alert and Collision Avoidance System The traffic alert and collision avoidance system (TCAS) gives aural and visual indications to the flight crew. The indications are advisories. The traffic display shows other airplanes and possible collision conditions. TCAS uses the ATC/Mode S transponder system to send TCAS data to other TCAS airplanes. The TCAS system has these units: • • • • •

TCAS computer Top directional antenna Bottom directional or omnidirectional antenna Mode S ATC transponders (2) ATC control panel.

November 2000

TCAS gives two types of advisories to the pilots. One type is the traffic advisory (TA) which tells of other airplanes in the area. The other type of advisory is the resolution advisory (RA). The RA gives the pilots directions to prevent a collision.

The left ADIRU gives heading data to the TCAS computer. TCAS uses radio altitude to change the range limits of advisories at low altitudes. Radio altitude also helps TCAS to know if an airplane is on the ground.

The TCAS computer sends data to the DEUs and the HUD Computer. You push the traffic button on the EFIS control panel to show the location and track of other airplanes on the display units. The display units also show the pilots how to change or maintain vertical speed to prevent a collision. TCAS sends aural alerts to the flight compartment through the remote electronics unit. The ground proximity warning computer (GPWC) alerts have higher priority than TCAS advisories. When both computers give warnings at the same time, you do not hear the TCAS indications.

7-11

RADIO

MINS

FPV

MTRS

BARO

IN

MAP

VOR 1

VOR APP OFF

PLN CTR

20 10 5

40 TFC

ADF 1 WXR

BARO HPA STD

RST

STA

WPT

ARPT

VOR 2 80 160 320 OFF 640 ADF 2

DATA

POS

TERR WX

EFIS Control Panel (2)

TEST

WX RADAR

WX/TURB MAP

TILT 5 10 GAIN

15

UP

CDS Display Unit (3)

CDS Display Electronic Unit (2)

0 DN AUTO

15

MAX 5

10

Weather Radar Control Panel Collins

TEST

CONN

LRU

STATUS

XMIT

ANT FAIL

REC

ANT FAIL

Weather Radar Antenna

PSEU

TEST

Weather Radar Transceiver Radio Altimeter (2)

ADIRU (2) Autothrottle Switch PacK Landing Gear Lever Switch

Ground Proximity Warning Computer

Weather Radar System Weather Radar System The weather radar system shows the weather conditions along the flight path of the airplane. The pilots can change the flight path to fly around bad weather conditions. The pilots can also use the weather radar system as a navigational aid. The weather radar transceiver sends weather data to the common display system display electronic units (CDS DEUs). The DEU shows weather radar on the CDS display unit in four colors. The four colors define these conditions: • • • •

Green - light rainfall Yellow - moderate rainfall Red - heavy rainfall Magenta - rainfall with turbulence.

Push the WXR button on the EFIS control panel to show weather radar on the onside CDS display unit. The pilots use the weather radar control panel to set these functions: • • •

Mode of operation Gain control Antenna tilt angle.

The weather radar uses attitude signals from the ADIRU to stabilize the antenna scan. The weather radar transceiver uses EFIS control panel range for display of weather data on the display units. The pilots can set a different range on each EFIS control panel. The weather radar system can show predictive windshear (PWS) messages on the display units (DUs).

windshear condition. PWS uses inputs from the PSEU, landing gear lever switch, and autothrottle switch pack to determine if the airplane is in a takeoff or approach mode. PWS supplies information to the ground proximity warning system (GPWS) computer for types of messages and message priority. The GPWS computer supplies audio inhibit signals to the PWS function in the weather radar transceiver. These are the different PWS annunciations that show on the DUs: • • • •

Windshear caution-yellow Windshear warning-red Windshear symbol bar-black and red Windshear attention bars-yellow.

PWS uses inputs from the radio altimeter and ADIRUs to detect a 7-12

November 2000

Navigation GROUND PROXIMITY

Gear Lever Switch Flap Switches

INOP

a

FLAP GEAR TERR INHIBIT INHIBIT INHIBIT

SYS TEST

NORMAL NORMAL NORMAL

Ground Proximity Module

CDS Display Electronic Unit

CDS Display unit

BELOW G/S P-INHIBIT a Below G/S Light (2) Radio Altimeter AAU OBS

F/O CAPT

SMYD Left ADIRU

HUD Computer MOD A

PWR FMC ON VALID

B

C

D

Flight Management Computer Ground Proximity Warning Computer

TCAS Computer

Remote Electronics Unit

MMR (2)

Ground Proximity Warning System Enhanced Ground Proximity Warning System

The GPWS uses inputs from these units to calculate warning conditions:

The enhanced ground proximity warning system (enhanced GPWS) gives the pilots aural and visual warnings of unsafe conditions. The warnings continue until the pilots correct the condition. The system operates when the airplane is less than 2450 feet above the ground.

• • • • • • •

The enhanced GPWS displays terrain forward of the airplane, and also alerts the flight crew of early descent when landing. The ground proximity warning computer sends terrain data to the common display system to show on the navigation displays.

November 2000

•

Flap landing gear warning switch Flap takeoff warning switch Landing gear lever switch Radio altimeter (2) Flight management computer (FMC) Left air data inertial reference unit (ADIRU) Stall management yaw damper (SMYD) computer Multi-mode receiver (MMR) (2).

GPWS visual warnings show on the common display system (CDS) display units or on the below glideslope annunciators. GPWS windshear warning shows on the HUD. GPWS aural warnings go through the remote electronics unit to the pilots headsets and speakers. The GPWS prevents TCAS warnings when both systems give a warning at the same time.

The FLAP, GEAR, and TERR INHIBIT switches on the ground proximity module prevent certain warnings and displays. The FLAP INHIBIT switch sends a flaps down signal to the GPWC. The GEAR INHIBIT switch sends a landing gear down signal to the GPWC. The TERR INHIBIT switch stops the terrain information on the navigation displays.

7-13

ADIRU (2)

TEST OK MMR FAULT BUS IN FAIL TEST ANT

GPS

TEST

PWR ON

FMC VALID

DO NOT PAINT

GPS Antenna (2)

MMR RECEIVER

Flight Management Computer

FMC CDU

Multi-Mode Receiver (2)

Global Positioning System Global Positioning System

The MMR calculates these values:

The global positioning system (GPS) is a satellite radio aid for navigation. The GPS is part of a multi-sensor navigation system.

• • • •

GPS uses navigation satellites to give accurate airplane position to the FMC and the flight crew.

The FMC uses GPS data to help calculate airplane present position. The FMC CDU shows GPS present position. The pilots can turn off the use of GPS data in FMC calculations with the FMC CDU.

The air data inertial reference units (ADIRUs) gives inertial reference data to the multi-mode receiver (MMR). The MMR can use this information to find the best satellites during system initialization. It also uses the data for certain modes of operation.

7-14

Airplane latitude Airplane longitude Airplane altitude Accurate time.

November 2000

Navigation

Combiner AIII Drive Electronics Unit

Overhead Unit

APCH WARN

FLARE COMPUTER FAULT

TEST

R

HGS COMPUTER

Airplane Sensors

Annunciator Panel

FLIGHT DYNAMICS

PROTECTED BY US PATENT No.

HUD Computer

COMPUTER

1

2

3

4

5

6

7

8

9

ENTER

0

TEST

MODE

H G S

STBY RWY G/S

FAULT

CLR

BRT + DIM -

Control Panel

Head-Up Display System Head-Up Display System OPERATION The head-up display (HUD) system shows flight and guidance symbols. The flight crew uses the HUD for low visibility takeoffs and CAT IIIa approach and landings. They can also use the system during normal weather conditions. These components are in the HUD system: HUD • • • • • •

Computer Drive electronics unit (DEU) Annunciator panel Control Panel Overhead unit (OHU) Combiner.

The HUD computer uses sensor data from other airplane systems to calculate the symbols and their position on the combiner. November 2000

The DEU amplifies the symbol data and sends it to the overhead unit. The DEU also contains the power supplies for the DEU, OHU, and combiner. There is a cathode ray tube (CRT) and an optics projection assembly in the overhead unit. These components project the symbols onto the combiner. The combiner is a glass plate assembly. The assembly has two ground glass outer pieces with a special thin clear coating between them. The special coating reflects only the green symbol displays from the CRT. The combiner optically combines the symbols with the view through the captain windshield. The combiner also contains brightness controls.

Status and warning annunciations show on the annunciator panel. The panel supplies the first officer with changes in HUD status during manual ILS approach and landing operations to CAT IIIa minimums. The flight crew uses the control panel to select and show HUD modes and to enter data. Maintenance personnel use the panel to operate system BITE. BITE To do a system self-test, push TEST on the control panel or push the red reset switch on the front of the computer. There is a fault ball indicator on the front of the computer and the DEU. Normally, the indicator color is black. The color changes to white when there is a fault. The indicator resets automatically.

7-15

Roll Scale and Pointer Pitch Scale Airplane Reference Horizon Line

Selected Altitude Altitude Scale and Index

Selected Speed

Airspeed Scale and Index

Flight Path Guidance Cue

Combiner Display Head-Up Display System COMBINER DISPLAY The graphic shows an example display during approach. The horizon line shows an artificial horizon. When the horizon line and airplane reference symbol overlap, the airplane is in a level, zero degree pitch attitude. Airplane reference shows the projected centerline of the airplane. The function of this symbol is equivalent in operation to the airplane symbol on a standard EADI. Airspeed shows on the left side of the combiner. Barometric altitude shows on the right side of the combiner. A pitch scale shows above and below the horizon line.

7-16

The roll scale and pointer are above the airplane reference symbol. The scale shows bank angle. The flight path symbol shows the flight path vector of the airplane. This symbol gives an instant indication of where the airplane is going. The pilot can operate the airplane and fly a flight path to a desired point. The guidance cue functions the same way as an integrated cue flight director. For the pilot, the objective is to capture the guidance cue inside the flight path symbol circle with pitch and roll control inputs. Selected speed shows the airspeed the pilot sets on the DFCS mode control panel (MCP). It also shows the airspeed command set by the FMCS. Selected altitude shows the altitude set on the MCP.

November 2000

Autoflight Features

•

Flight Management Computer System (FMCS)

FLIGHT MANAGEMENT COMPUTER SYSTEM (FMCS)

•

Dual Flight Management Computers

•

Digital Flight Control System (DFCS)

•

Autothrottle

•

Built-In Test Equipment (BITE)

•

Wheel to Rudder Interconnect System (WTRIS)

The FMCS allows preplanned flight profile control and guidance for best performance and performance management. OPTIONAL DUAL FLIGHT MANAGEMENT COMPUTER SYSTEM A dual system permits navigation under the primary means of navigation criteria. DIGITAL FLIGHT CONTROL SYSTEM (DFCS) The DFCS includes these functions: • • • • •

Autopilot Flight director Mach trim Speed trim Altitude alert.

AUTOTHROTTLE The autothrottle controls the engines independently to get the best performance. BUILT-IN TEST EQUIPMENT (BITE) The BITE system gives fast and accurate troubleshooting of the main flight management system (FMS) components. YAW DAMPER/WHEEL TO RUDDER INTERCONNECT SYSTEM (WRTIS) The yaw damper decreases the yaw rates associated with dutch roll and turbulence. The WTRIS assists manual turns when both A and B hydraulic systems are in standby.

November 2000

8-1

FMCS Flight Management Computer

DFCS Control Display Unit

Flight Control Computers

Mode Control Panel

A/T System

ADIRS

CDS

SMYD

Autothrottle Computer

Air Data Inertial Reference Units

Common Display System

Yaw Damper WTRIS Functions

Flight Management System Flight Management System (FMS) The flight management system is a group of systems that operate together to decrease flight crew workload. The design of the system permits fast, accurate troubleshooting and maintenance. You do the built-in test (BITE) for most of the FMS components from a common location. The BITE procedures for each system are almost the same. FLIGHT MANAGEMENT COMPUTER SYSTEM (FMCS)

systems of the FMS to follow the flight plan. The BITE for other FMS systems goes through the FMC and shows on the CDU. DIGITAL FLIGHT CONTROL SYSTEM (DFCS) The digital flight control system (DFCS) includes the autopilot, flight director, and other functions. For a fully automatic flight, the FMC commands the autopilot to fly the route. AUTOTHROTTLE (A/T) SYSTEM

The central part of the flight management system is the flight management computer system (FMCS). The flight crew uses the control display unit (CDU) to enter the route and performance data for the flight. The FMC calculates the lateral and vertical components of the flight. It then sends commands to the other 8-2

The autothrottle (A/T) system controls the engine thrust levers. The FMC sends thrust and speed targets to the autothrottle for best overall flight performance.

AIR DATA INERTIAL REFERENCE SYSTEM (ADIRS) The air data/inertial reference system (ADIRS) supplies attitude, air data, and navigation information to the other FMS systems. COMMON DISPLAY SYSTEM (CDS) FMS information shows on the common display system (CDS). YAW DAMPER/WTRIS FUNCTION The yaw damper and WTRIS functions are part of the stall management yaw damper (SMYD). The yaw damper gives commands to the rudder to decrease yaw rates from dutch roll and turbulence. The WTRIS assists manual turns when on the standby hydraulic system. The BITE for the SMYD is on the front of the SMYD. November 2000

Autoflight

Top Of Climb PWR ON

Step Climb

FMC VALID

Optimized Speeds And Altitudes

Climb Takeoff Guidance Control Display Unit (CDU) (2)

Flight Management Computer (FMC)

CDS

Origin Cruise Waypoint

Flight Control Computer (2) Stall Management Yaw Damper (2) DME Interrogator(2) Display Electronic Unit (2) Navigation Control Panel (2) Autothrottle Computer Flight Data Acquisition Unit Ground Proximity Warning Computer Aircraft Communication and Reporting System Head-Up Display System

Waypoint Course Alteration

Top Of Descent

Waypoint ILS Outer Marker

Approach

Destination

Flight Management Computer System Flight Management Computer System (FMCS) The flight management computer (FMC) is the main component of the flight management computer system. The flight crew uses the control display unit (CDU) to schedule the route and performance for the flight. With this flight plan data and inputs from airplane sensors, the FMC does these operations: • • •

Navigation Performance Guidance.

An optional connection to the aircraft communication and reporting system (ACARS) is available. It lets the FMC receive flight plan data from a ground station and send flight status to the ground station. NAVIGATION A navigation data base stored in the FMC memory includes the November 2000

navigation data for the area of operation. You update the data base every 28 days. The pilot can preselect the entire flight plan from the navigation data base with standard air traffic control language. Also, the flight crew can have the FMC fly an offset from the route. Required time of arrival (RTA) is also available. The FMC calculates the airplane position as the flight continues. It uses the inertial reference function and navigation aids, if available, to calculate the position. The FMC compares the calculated position with the planned position. Any deviation shows on the common display system. PERFORMANCE A performance data base in the FMC contains a performance model of the airplane and the engines. The flight crew enters gross weight, cruise

altitude, and cost index for the flight. The FMC uses this data to calculate the economy speeds, best flight altitude, and top of descent point. The CDU and the flight displays show target speeds and altitudes. GUIDANCE The FMC sends commands to the digital flight control system (DFCS) and the autothrottle (A/T). The DFCS and the A/T use these signals to control the airplane in the lateral (LNAV) and vertical (VNAV) parts of the flight plan. BUILT-IN TEST EQUIPMENT (BITE) The maintenance technician uses the CDU to operate BITE for the FMC. The full self tests of the FMC and the ability to see the status of the inputs and outputs permits fast trouble-shooting.

8-3

FMC BOTH ON R

BOTH ON L

Normal/ Both On L

P5

FMC 1

CDU 1

NORMAL

PWR FMC ON VALID

Both On R

Transfer Switch and Relays

Digital Outputs

PWR FMC ON VALID 1

A/P

A/T

FMC

P/RST

P/RST

P/RST

2 TEST

Autoflight Status Annunciator (2) CDU 2

FMC 2

Dual FMC Option Dual Flight Management Computer (FMC) The dual FMC configuration is a flight management computer system (FMCS) option. It permits navigation under the primary means of navigation criteria. This lets the airplane fly a more direct route which saves fuel and time. The main components are two flight management computers and two control display units (CDUs). Each FMC makes a position calculation with different sets of navigation aids and ADIRU inputs. If navigation aids are not available, the FMCs only use separate ADIRU inputs for the calculation. The FMCs combine their calculations to calculate a best position. Both FMCs transmit this best position to the CDS and other user systems.

8-4

One FMC is the primary FMC and the other is the secondary. A transfer switch controls the system configuration through transfer relays. When the transfer switch is in NORMAL or BOTH ON LEFT, FMC 1 is primary and FMC 2 is secondary. When the transfer switch is in BOTH ON RIGHT, FMC 2 is primary and FMC 1 is secondary. When the transfer switch is in NORMAL, FMC 1 sends data to all single systems and to the number 1 system of dual systems. The data from FMC 2 goes to the number 2 system of dual systems. When the transfer switch is in BOTH ON LEFT, FMC 1 is the source for all outputs. When the transfer switch in BOTH ON RIGHT, FMC 2 is the source for all outputs.

The output of each CDU connects to both FMCs. The primary FMC processes the CDU inputs and sends the data to the secondary FMC. The display on both of the CDUs is from the primary FMC. The two FMCs compare inputs and outputs at all times. A large difference between the two causes the secondary FMC to conditionally fail and to alert the flight crew.

November 2000

Autoflight A/T ARM

COURSE

IAS/MACH

V NAV

HEADING

L NAV

ALTITUDE

VERT SPEED

C/O

DN

VOR LOC MA F/D

CWS A

SPEED

LVL CHG

HDG SEL

APP

ALT HLD

MA DISENGAGE

V/S UP

OFF

COURSE

CWS B

SEL

OFF N1

A/P ENGAGE CMD B CMD A

F/D

OFF

Mode Control Panel (P7)

Altitude Alert Tone

Remote Electronic Unit

Sensors Autopilot Aileron and Elevator Actuators

A/P

A/T

FMC

P/RST

P/RST

P/RST

Flight Control Computer (2)

1 2 TEST

Autoflight Status Annunciator (2) Aural Warning Module

Common Display System

Stabilizer Trim Electric Actuator

Mach Trim Actuator

Digital Flight Control System Digital Flight Control System (DFCS) Two independent flight control computers (FCCs) control these functions: • • • • •

Autopilot Flight director Altitude alert Mach trim Speed trim.

To do these functions, each FCC does these things: • • •

Receives flight crew requests and airplane sensor inputs Does the calculations for each function Controls the indicators and actuators for each function.

AUTOPILOT/FLIGHT DIRECTOR The flight crew uses the mode control panel to control the autopilot and the flight director. The crew November 2000

engages the autopilot, turns on the flight director, selects the modes, and selects the targets on the MCP. For the autopilot function, the FCC sends commands to the autopilot aileron and elevator actuators to control the surfaces. For the flight director function, the FCC sends pitch and roll guidance commands to the common display system. The status and the mode of operation of the autopilot and the flight director show on the common display system. MACH TRIM AND SPEED TRIM Mach trim and speed trim functions operate with no flight crew input. For mach trim, the FCC controls the mach trim actuator. It moves the elevator to increase airplane stability at high air speeds. For speed trim, the FCC controls the stabilizer trim electric actuator. It

moves the stabilizer to increase airplane stability at low air speeds. ALTITUDE ALERT The altitude alert function uses the altitude selected on the mode control panel. The FCC alerts the flight crew when the airplane approaches or departs the selected altitude. There is an aural alert from the remote electronics unit and a visual alert on the common display system. DISENGAGE WARNING There is a visual and an aural warning to alert the flight crew if the autopilot disengages. A red light on the autoflight status annunciator and the wailer from the aural warning module give the warning. BITE The flight control computer gives accurate, reliable, and fast built-in test (BITE) capability. 8-5

A/T Disconnect Switches TO/GA Switches A/T ARM

IAS/MACH

C/O OFF N1

Engine (2)

SPEED

Mode Control Panel (P7)

Autothrottle Servomotor (2)

Sensor Autothrottle Computer

1

A/P

A/T

FMC

P/RST

P/RST

P/RST

2 TEST

Autoflight Status Annunciator (2)

Flight Control Computer (2)

Thrust Mode Annunciator

Flight Management Computer Common Display System EICAS Display

Autothrottle System Autothrottle • The full-range autothrottle system can control the thrust levers from takeoff to touchdown. It gives maximum fuel conservation through smooth, precise thrust control. Like other flight management subsystems, the autothrottle design gives maximum operational and cost benefits. There is a single autothrottle computer that operates the thrust levers through two independent servomotors. The autothrottle controls the engines independently to get the best performance from each engines. These are the auto throttle controls in the flight compartment: •

•

Arm switch on the mode control panel (MCP) arms the autothrottle Takeoff/go-around (TO/GA) switches select the takeoff or

8-6

•

go-around modes Autothrottle (A/T) disconnect switches disconnect the autothrottle Mode select push-buttons on the MCP select thrust or speed control.

The flight crew can select the autothrottle mode with the mode select push-buttons on the MCP. Usually the digital flight control system selects the correct autothrottle mode for the flight phase. The active autothrottle mode shows on the common display system.

control to give improved performance. A red light on the autoflight status annunciator gives a visual warning if the autothrottle disengages. The autothrottle system has built-in test. This lets you do fast accurate maintenance. You do the tests of the autothrottle system with the CDU. From the CDU, you can select these functions: • • • •

Current status Inflight faults Interactive Ident/Config.

The autothrottle moves the thrust levers to control thrust or airspeed. For thrust control, the FMC calculates the correct thrust setting for the flight phase. For airspeed control, the autothrottle accepts mach and airspeed commands from the FMC or MCP. The autothrottle operates with the electronic engine November 2000

Autoflight A/T ARM

COURSE

IAS/MACH

V NAV

HEADING

L NAV

ALTITUDE

C/O

DN

VOR LOC MA F/D

SPEED

CWS A

LVL CHG

HDG SEL

APP

ALT HLD

COURSE

CWS B MA

SEL

OFF N1

A/P ENGAGE CMD B CMD A

VERT SPEED

DISENGAGE

V/S UP

OFF

F/D

OFF

Mode Control Panel (P7) TO/GA Switches

Autothrottle Mode

Autopilot/Flight Director Modes

Status

Primary Flight Display

Autoflight Autoflight Modes And Mode Display Mode selection and system engage switches for the autopilot, flight director, and autothrottle are on the mode control panel (MCP). The TO/GA switches select takeoff and go-around modes. They are on the thrust levers. The engage status and mode of operation show on the display units. The takeoff mode is a combined flight director and autothrottle mode. The autothrottle sets engine thrust to the target value calculated by the flight management computer (FMC). The flight director gives commands to control the rate of climb and then to control the selected airspeed set on the MCP.

November 2000

Vertical navigation (VNAV) and lateral navigation (LNAV) are the normal cruise modes. When these modes are selected, the FMC sends commands to the autopilot, flight director, and autothrottle. The airplane flies the FMC route at the airspeed and altitude for the selected performance. The autopilot, flight director, and autothrottle work together in these modes. Other modes are available at the option of the crew. The crew can select modes to change and hold the airplane altitude and to fly a radio beam or heading. The crew uses the mode control panel to select the desired autopilot and flight director mode. The autopilot or flight director then sets the autothrottle mode that gives the best combined performance.

The approach mode (APP) is for landing. In the approach mode, the flight director and the autopilot use localizer and glideslope radio signals. When the approach mode is selected, you can engage both autopilots for an automatic landing. The autothrottle holds the MCP selected airspeed in approach. In go-around, the autothrottle sets the thrust levers to go-around thrust. The flight director and, if available, the autopilot, control rate of climb, airspeed, and track. The flight mode annunciator (FMA) is on the primary flight display. Autopilot flight director status shows in the status field of the FMA. Autothrottle, pitch, and roll modes show in the other positions of the FMA.

8-7

DF CS B I T E T E S T * * E X T E N D E D MA I N T E N A N C E * * < B I T E L I B RA RY T E S T

MAI NT B I T E T EST < FM CS ENGINES > < DFCS

APU >

< A/ T

FQIS >

< R I G G I NG < MC P T E S TS < S E N S OR V A L U E S

* F L T XXXXX * 10: 26Z * AL T 0 3 1 0 0

< ADI RS E NU>

< CDS

D F CS B I T E T E S T MA I N ME NU < F A UL T R E V I E W

< IND EX

DF CS B I T E L E G 0 2 * A / P DI CHA N- A * R/ A MONI T OR * S US P E CT L RU( S ) L RRA - 1

< L RU T E S T S

T EST SC * 0 1 / 0 3 * APPR * *

1 1 MA R 9 5 * A/ S 2 2 0 KT P RE V ME NU>

In-Flight Faults

< L RU RE P L A CE ME NT T E S T < L A ND V E R I F Y

Main BITE Index

< E X T E NDE D M A I NT E NA NCE < I D E N T I F I C A T I ON C O N F I GU R A T I O N < EXI T

AND

DFCS BITE Index DF CS B I T E T E S T * S T AT U S * 11. 00 * * S U M M A RY * - A */ B * 1 / X * T EST F A I L ED * * S US P E CT L RU( S ) * I RS S W

DF CS B I T E T E S T * L R U R E P L A C E ME N T * < A DI RU LNAV > < A/ T < DME < LRRA < F CC

< C ON T I N U E

PREV MENU >

Current Status

< EX I T

MCP > FMC > SMYDC > VOR > P R E V ME N U >

LRU Interface

D F CS B I T E TE S T EL EV R I G - A/ - B 5 1 . 0 1 SET ST AB " B " D I ME N ( I N)

NOM= UL = LL=

S T A B P OS ( V A C)

NOM= 0. UL = 0. LL= - 0. B = + - XX.

A = + - XX. XX < C ON T I N U E

39. 89 39. 90 39. 88 00 07 07 XX

P R E V ME N U >

Rigging Tests

Built-In Test Equipment Built-In Test Equipment (BITE) The CDU gives access to selfcontained in-flight monitoring and ground test capabilities for these systems or components: • • • • • • • •

Flight management computer system - FMCS Digital flight control system DFCS Autothrottle - A/T Air data inertial reference system - ADIRS Common display system - CDS Electronic engine control ENGINES Electronic control unit - APU Fuel quantity indicating system FQIS.

A technician uses the CDU to control these BITE functions. All system BITE displays are on the CDU in English. The technician selects test options from a menu and supplies interactive operator responses through the CDU. BITE does the tests and supplies data to the technician to do trouble-shooting. The technician uses the verification tests to do a test of the interfaces after replacement of a line replaceable unit (LRU).

The ENGINES prompt gives access to engine exceedance data for present and past flights. The APU prompt gives access to APU fault and system information. The FQIS prompt gives access to fuel quantity system fault and system information.

The system allows fast, line maintenance fault isolation to a single line replaceable unit. Fast and accurate diagnostics mean fewer delays and smaller spares inventories.

Each FMS component contains tests for itself, its sensor inputs, and other interfaces. Also, each component stores in-flight fault data for analysis on the ground.

8-8

November 2000

Autoflight

YAW DAMPER YAW DAMPER

a

Command

OFF

Feedback

ON

Warning Light and Engage Switch (P5 Flight Control Panel)

Rudder Power Control Units

ADIRU Control Wheel Position Sensor

X XXX XX XX XXXX XXXXXXX XXXX X X X X X X X XX X X X

YAW DAMPER

Stall Management Yaw Damper (2)

Yaw Damper Indicator (P2 Center Instrument Panel)

Yaw Damper/WTRIS System Yaw Damper/WTRIS The yaw damper moves the rudder to decrease yaw rates that are related to dutch roll. The WTRIS assists manual turns when the standby hydraulic system is on. The yaw damper system connects to the main and standby rudder power control units to do the yaw damper function. SMYD 1 connects to the main rudder power control unit. SMYD 2 connects to the standby rudder power control unit. The WTRIS function is only in SMYD 2. These are the yaw damper components in the flight compartment: • • •

The yaw damper is available for the full flight. You engage the yaw damper and the WTRIS with the yaw damper switch on the overhead panel. The yaw damper warn light goes off when the yaw damper is engaged. The yaw damper uses inputs from the inertial reference function of the ADIRU to get airplane yaw rate and lateral acceleration. The yaw dampers send commands to the rudder power control units to move the rudder and stop the dutch roll. The rudder pedals do not move when the yaw damper moves the rudder. The yaw damper indicator shows yaw damper movement of the rudder.

The WTRIS senses control wheel movement and sends commands to the standby rudder power control unit to move the rudder. Monitor circuitry in the computer disengages the yaw damper/WTRIS for problems with the position feedback or with the servo. You do a test of the yaw damper and WTRIS functions at the stall management yaw damper. A menu lets you select these BITE functions: • • • •

Existing faults Fault history Ground tests Other functions.

Yaw damper switch Yaw damper warn light Yaw damper indicator.

November 2000

8-9

Electrical Power Features THREE ELECTRICAL GENERATORS Two generators are engine-driven and one is APU-driven. AUXILIARY POWER UNIT (APU) The APU makes the airplane selfsufficient. It supplies electrical power on the ground and in flight. AUTOMATIC BUS TRANSFER SYSTEM This system increases electrical power reliability and reduces pilot workload.

BATTERY FOR STANDBY DC AND AC STANDBY POWER The airplane battery supplies power to the standby system when normal power fails. The battery supplies power to essential systems for a minimum of thirty minutes. BUILT-IN TEST AND TROUBLESHOOTING

•

Electrical Power System

•

Power Control and Protection

•

Controls and Indications

•

Control, Indication, and BITE

•

System Power Distribution, DC Standby Power Control

•

Electrical Component Location

Operational information shows in the flight compartment. Built-in test features are in the protection and control units in the electronic equipment (EE) compartment and on the P5 panel in the flight compartment.

ACCESSIBLE ELECTRONIC EQUIPMENT RACKS You have access to all electronic racks from the ground.

November 2000

9-1

External Pwr AC

GCU

APU

SCU

GEN

AGCU

GCB1

GEN

EPC

APB

GCB2

GCU

115V AC Transfer Bus 2

115V AC Transfer Bus 1 BTB2

BTB1

BPCU

115V AC Gnd Service Bus 1

115V AC Main Bus 1

115V AC Main Bus 2

Galley

Aux Battery Charger

Transformer Rectifier 1

Aux Battery

28V DC Battery Bus

28V DC Bus 1

115V AC Standby Bus 1

Transformer Rectifier 2

28V DC Standby Bus

Galley

Transformer Rectifier 3

115V AC Gnd Service Bus 2 Battery

28V DC Hot Battery Bus

28V DC Bus 2

Battery Charger

28V DC Switched Hot Battery Bus

Static Inverter

Electrical Power Distribution Electrical Power System • The electrical power system supplies 115v ac and 28v dc electrical power to the airplane. AC Power Distribution The electrical power distribution system normally operates as two separate systems. IDG 1 supplies ac power to transfer bus 1. IDG 2 supplies ac power to transfer bus 2. The APU can supply power to both transfer buses on the ground and in flight. AC power sources include these components: •

•

Two IDG units, each driven by an engine, supply115v ac. Each IDG supplies up to 90 KVA A starter-generator is a 90 KVA (to 32000 ft, 66 KVA above) generator and an electric starter

9-2

for the APU External power receptacle rated at 90 KVA.

External power supplies the airplane in these two ways. •

•

From the external power receptacle directly to the ground service buses. The ground service buses supply cabin lighting, electrical outlets, and the battery chargers From the external power receptacle through the external power contactor (EPC) and bus tie breakers (BTBs), which supply power to all buses.

controlled load shed relays shed loads as necessary in this condition. Bus transfer allows continuation of normal flight operations. The pilot starts and uses the APU. The BTBs open when the bus has power again. Bus Protection There are four AC system control units; three GCUs and the BPCU. The BPCU provides overall system management and protection. In addition to a GCU, the APU has a start converter unit (SCU) for excitation and voltage regulation. The engine GCUs supply generator voltage regulation and protection.

Bus Transfer System

There is protection for these faults:

If either transfer bus loses power, the bus power control unit (BPCU) uses engine generator control units (GCUs) to close the bus tie breakers (BTB) to supply power from the opposite transfer bus. BPCU

• • • • •

Differential current Over/under voltage Over/under frequency Overcurrent Unbalanced phase current. November 2000

Electrical Power

BPCU FAULT

GCU FAULT

EPC FAULT

GCB/APB FAULT

EP DIST/ BUS FAULT

BPCU IDG FAULT

PASS

FEEDER FAULT DIST/BUS FAULT GCU PASS

BTB FAULT

BPCU

GCU

TEST

TEST

Bus Power Control Unit

Generator Control Unit

Power Control and Protection DC Power Distribution The 28v dc system has these components: •

•

•

Three transformer rectifiers (TRs). The TRs convert 115v ac to 28v dc One 36 ampere-hour (AH) battery or optional, two 40 AH or 48 AH batteries Battery charger(s).

DC BUSES The main battery or the main battery charger supplies power to the hot battery bus and the switched hot battery bus. Whichever has the higher output voltage becomes the source. TR 3 normally energizes the battery bus. Transfer bus 2 or transfer bus 1 energizes TR 3. If TR 3 fails, the battery charger or the battery supplies power to the battery bus. November 2000

TR 1 and TR 2 and alternately TR 3 normally supply power to dc buses 1 and 2. DC bus 1 is the normal source for the dc standby bus. The battery bus alternately supplies the dc standby bus. Standby Power The auxiliary battery and/or the main battery are the standby source of power. The standby system supplies ac and dc power to primary flight instruments, communication, navigation, and other equipment. If all ac generators fail, ac transfer bus 1 does not have power. This condition results in these functions: • •

The batteries energize the dc standby bus The static inverter energizes the ac standby bus. The inverter uses battery power to create single phase 115v ac.

Automatic control of the standby and dc system is in the standby power control unit (SPCU). The SPCU is in the P6 panel in the flight compartment. The SPCU controls the contactors in the system and monitors the status of the buses and relays. The SPCU also supplies information to the flight compartment. Ground Power The APU generator is a convenient power source for ground operation. In addition, external 115v ac power connects through a receptacle on the right side above the nose gear wheelwell. A switch on the forward overhead panel connects this power to the airplane system. Another switch on the forward attendant panel connects this power to cabin related service circuits and the battery charger.

9-3

DC AMP

CPS FREQ

DC VOLTS / AC AMPS / AC VOLTS

BAT DISCHARGE

ELEC

TR UNIT a

a

a

MAINT

APU GEN

AUX BAT

P5 Forward Overhead Panel

TR1

BAT BAT BUS STBY PWR

GEN1

TR2

TR3

GEN2

GRD PWR

INV

STBY PWR

TEST

TEST

OFF OFF

BAT

GALLEY ON

ON

DC

AC

Control, Indication, and BITE Controls and Indications These electrical power controls and indications are on the P5 forward overhead panel. The DC AMPS meter shows the current of TR1, TR2, TR 3, and BAT. The selector controls the indication. The DC VOLTS meter shows the voltage of the source selected by the dc meter selector (all positions). The DC meter selector selects the dc source for the dc voltmeter and ammeter indications. The TEST position connects the dc meters to a BITE in the panel. The BAT switch switches the hot battery bus. The switched hot battery bus connects to the battery with the switch on and guard down. The battery is then available for backup to the ac and dc standby buses.

9-4

The CPS FREQ meter shows the frequency of the source selected by the AC meter selector.

• • •

The AC VOLTS meter shows the voltage of the source selected. The AC AMPS meter shows load current (phase B) of the generator source selected by the ac meter selector. The AC meter selector selects the ac source for ac volt and frequency indications. The TEST position connects the AC meters to a BITE in the panel. The GALLEY switch supplies electrical power to the galleys. BITE The ELEC light shows that there is a BITE message. This is the procedure to see a BITE message:

Airplane on ground Put the AC and DC meter selectors to TEST Push the MAINT button.

Messages show on the display. These are examples of BITE messages: • • • • • • •

BAT CHGR INOP AUX BAT CHGR INOP SPCU INOP PANEL FAILURE INTERFACE FAILURE STAT INV INOP TR3 XFR RLY INOP.

The amber TR UNIT light comes on when any of the TR units fail on the ground. It also comes on when TRU1 or TRU2 and TRU3 fail in the air. The amber BAT DISCHARGE light comes on when a large current flow occurs in a short time.

November 2000

Electrical Power

P5 Forward Overhead Panel

Control and Indication Controls and Indications The STANDBY PWR OFF light comes on amber when the standby ac or dc bus does not have power.The light also comes on when the battery bus does not have power and the battery switch is in the on position. The STANDBY PWR switch has these three positions: •

• •

AUTO - Normal position. The battery or batteries automatically connect to supply the dc standby and ac standby buses with loss of all ac power in flight OFF - Turns off power to the standby power buses BAT - The battery or batteries supply power to the battery bus, dc standby bus, and ac standby bus.

June 1998

The amber DRIVE light comes on when IDG 1 or IDG 2 oil pressure is low. The DISCONNECT switch deenergizes IDG 1 or 2 and disconnects the input shaft. The engine start levers must be in the idle position for the disconnect to operate. The GRD POWER AVAILABLE light comes on blue when ground power is connected to the receptacle. The quality of the ground power must also be good for the light to come on. The TRANSFER BUS OFF lights come on amber when a transfer bus does not get power by any source. The SOURCE OFF light comes on amber when a transfer bus does not get power by the selected engine, APU generator, or external power.

The GEN OFF BUS light comes on blue when the IDG does not supply power to the transfer bus on the same side. The APU GEN OFF BUS light comes on blue when the APU is in operation but the generator does not supply power to one or both transfer buses. The GEN1, APU GEN, and GEN2 switches are three-position switches, momentary on/off and spring loaded to center. The BUS TRANSFER switch has these two positions: •

•

AUTO (guard down) - Transfer bus automatically transfers to opposite power source if normal source is inoperative OFF - Isolates left and right sides and prevents bus transfer.

9-5

First Observer Seat (Stowed)

Second Observer Seat (Optional)

P6-1

AUTO FLT B ENG 2 FLT CONT HYD SYS INPH PWR FIRE PROT

R ADIRU RADIO NAV B FMCS COMM 2 WXR SMYD 2 CDS

RADIO NAV 1 AUTO FLT A L ADIRU AUTO THROTTLE P6-2

STANDBY POWER CONTROL UNIT

P18-2 P6-3 AC/DC BUS IND

AC/DC POWER RECP

P18-1 ENGINE 1 CDS COMM 1 SMYD 1 FLT RECORDER

FUEL SYSTEM ENGINE FUEL LANDING GEAR MASTER CAUTION

AIR COND PRESS

ANTI-ICE & RAIN INTERIOR/EXTERIOR LIGHTING EQUIP COOLING OXYGEN LAVATORIES

APU EQUIP COOLING ELEC

P18-3

P6-4

PASS ACCOM

P6-5 WINDOW HEAT POWER P6-12

WINDOW HEAT POWER

P18-4

P6-11

Flight Compartment P6, P18 (Looking Aft)

System Power Distribution, DC Standby Power Control Aircraft Installation

Wiring

Automatic Load Shedding

Two power distribution panels (PDPs) in the EE compartment and two circuit breaker panels in the flight compartment contain the electrical power buses. The main buses and individual high power load controls are in the two PDPs. Lower current buses and individual load controls are in the PDPs and the circuit breaker panels in the flight compartment.

Wire bundles have numbers for quick identification to wiring diagrams and drawings. They are in raceways that run the length of the airplane. The raceways minimize radio interference, reduce wire splicing, and permit wire bundle removal and replacement.

Several methods protect the electrical system from damage. These are the methods:

Electronic Equipment Racks Electronic and electrical equipment are on racks in the EE compartment below the cabin floor, aft of the nose gear wheel well. You can easily get access from the ground through the access door in the lower body.

9-6

•

•

•

•

The BPCU sheds noncritical loads when there is a loss of an ac power supply Failure of components and associated wiring disconnect from the bus by circuit breakers The GCU trips the generator breaker when there is a left or right system problem Automatic electrical isolation of left and right sides when landing.

November 2000

Electrical Power

TRU 3

TRU 2

Bus Power Control Unit

GCU 2

APU Start Converter Unit

AGCU

GCU 1

Main Battery Charger

TRU 1 Static Inverter

Batteries PDP 1 (P91)

PDP 2 (P92)

J9 Junction Box

Auxiliary Battery Charger

E2, E3, E4 Equipment Racks

EE Access Door

Electrical Component Location EE Compartment Electrical Equipment ELECTRICAL SYSTEM BITE The GCUs collect and show maintenance data from the GCU, GCB, IDG, and power feeders. A front panel display on each GCU shows any failures. The BPCU collects and shows maintenance data from the BPCU, EPC, and ground power feeders. A front panel display on the BPCU shows failures. The APU start converter unit (SCU) collects maintenance data from the start power unit (SPU) and voltage regulator (VR). Fault data goes to the APU electronic control unit (ECU). This information and other APU data goes to the control display units (CDUs) in the flight compartment.

June 1998

COMPONENT LOCATION The E2 equipment rack has these components: • • • • • •

APU SCU AGCU GCU 1 TRU 1 Main battery charger Static inverter.

Power distribution panel 1 (PDP 1), P91, below the E2 rack, has these components in it: • • • • •

BTB 1 GCB 1 AGB 1 Transfer bus 1 Main bus 1.

The J9 junction box has these components in it: • •

The E3 equipment rack has the auxiliary battery charger. The auxiliary battery, and/or the main battery are below the E3 equipment rack. The E4 equipment rack has these components: • • • •

BPCU GCU 2 TRU 2 TRU 3.

Power distribution panel 2 (PDP 2), P92, below the E4 rack, has these components in it: • • • • •

BTB 2 GCB 2 EPC Transfer bus 2 Main bus 2.

Static inverter RCCB Auxiliary battery RCCB. 9-7

Fuel Features

AUTOMATIC CENTER TANK SCAVENGE

FUEL CAPACITY Each of the two main tanks holds 8,630 pounds (3,915 kg) of fuel. The center tank holds 28,803 pounds (13,066 kg) of fuel.

The automatic center tank scavenge system transfers residual center tank fuel to main tank 1. This increases the quantity of usable fuel.

UNDERWING REFUEL STATION

FUEL QUANTITY INDICATING SYSTEM (FQIS)

The refuel station is in the right wing. The maximum refuel rate is 2025 pounds (918 kg) per minute. The maximum refuel pressure is 55 psi.

The FQIS uses a variable capacitance principle and an advanced microprocessor to measure fuel quantity.

FUEL TANK COMPONENT REPLACEMENT WITHOUT DEFUELING

JETTISON SYSTEM NOT NECESSARY

Defueling is not necessary for removal of many fuel system components that are on the front and rear spar.

•

Fuel Tanks and Vent System

•

Pressure Refuel System

•

Defuel System

•

Engine and APU Fuel Feed System

•

Fuel and Water Scavenge System

•

Fuel Quantity Indicating System

•

Fuel System Control

Because the maximum take off weight is not substantially higher than the maximum landing weight, a fuel jettison system is not necessary.

AUTOMATIC WATER SCAVENGE SYSTEM An automatic water scavenge system decreases water accumulation in the center and main tanks.

November 2000

10-1

Surge Tank Main Tank 2 1,288 US Gal (4,876 Liters) 8,630 Lbs (3,915 Kgs)

Main Tank 1 1,288 US Gal (4,876 Liters) 8,630 Lbs (3,915 Kgs) Center Tank 4,299 US Gal (16,273 Liters) 28,803 Lbs (13,066 Kgs)

Surge Tank

Note: Total Volume = 6,875 US Gallons (26,025 Liters) 46,063 Lbs (20,896 Kgs)

Fuel Tanks and Vent System Fuel Tanks

Fuel Vent System

The fuel system has these fuel tanks:

The fuel vent system keeps the fuel tanks at near ambient pressure during all phases of airplane operation. Each tank vents to the surge tanks through channels in the wing.

• • •

Main tank 1 Main tank 2 Center tank.

There is a surge tank at the outer end of main tank 1 and 2. The surge tanks are part of the wing structure. Most fuel system components are inside the tanks. The boost pumps and scavenge pumps are on the front and the rear spar. You can remove these components without defueling.

10-2

The vent channels also permit fuel overflow into the surge tank if necessary.

November 2000

Fuel VALVE POSITION LIGHTS

OPEN

OPEN

OPEN

CLOSED

CLOSED

CLOSED

TEST GAUGES OFF FUEL DOOR SWITCH BYPASS

FUEL QTY LB

TANK 2

FUEL QTY LB

CENTER TANK

FUEL QTY LB

TANK 1

P15 Refuel Panel Center Tank Refuel Manifold

Main Tank 2 Main Tank 1 Float Switch (3)

Refuel Manifold

Refuel Manifold

Pressure Refuel System Pressure Refuel System The refuel station is on the right wing. It has a single refuel receptacle. The refuel station has these components: • • • • •

Power for the refuel system comes from the battery bus, the hot battery bus, and external power through the bus power control unit (BPCU) transformer rectifier.

Single refuel receptacle Individual fuel quantity indicators Fueling valve control switches Fueling valve open lights Fueling power control switch.

There is one fueling valve for each tank. The fueling valve control switches on the refuel panel control the fueling valves. A fueling float switch in each tank closes the fueling valves when the fuel quantity in each tank is at capacity. The fueling valves also operate manually. Fueling can occur for each tank individually or for all tanks at the same time.

November 2000

10-3

Forward Boost Pump Tank 1

Forward Boost Pump Tank 2

Bypass Valve

Bypass Valve

Fuel Spar Valve

Fuel Spar Valve Defuel Valve Refuel Station

Aft Boost Pump Tank 1

Aft Boost Pump Tank 2

Center Tank Boost Pumps

Crossfeed Valve

Defuel System Defuel System The defuel system permits the removal of fuel from each tank. It also permits the transfer of fuel between tanks on the ground. Use the fuel boost pumps to get fuel out of the tanks and into the fuel feed manifold. When the defuel valve is open, fuel transfers to the refuel station. The defuel valve operates manually. It is on the right front spar near the refuel station. Use these components to transfer fuel between tanks: • • •

Boost pumps Defuel valve Fueling valves.

A bypass valve in main tank 1 and main tank 2 permits suction defueling from those tanks. 10-4

November 2000

Fuel

Forward Boost Pump Tank 2

Forward Boost Pump Tank 1

To Engine

To Engine

Bypass Valve Fuel Spar Valve

Fuel Spar Valve

Defuel Valve Refuel Station

To APU

Bypass Valve

Aft Boost Pump Tank 2

Aft Boost Pump Tank 1 APU Fuel Shutoff Valve

Crossfeed Valve Center Tank Boost Pumps

Engine and APU Fuel Feed System Engine Fuel Feed System

APU Fuel Feed System

There are two boost pumps each for main tank 1, main tank 2, and the center tank. The boost pumps supply fuel from the fuel tanks to the engines.

The APU can receive fuel from any tank with the use of the applicable boost pump and crossfeed valve.

A crossfeed valve connects the left and right fuel feed manifold. This lets any tank supply fuel to any engine with the use of the applicable boost pumps. The center tank boost pumps have a higher output pressure than the boost pumps in the main tanks. Because of this, the engines receive center tank fuel first. With only residual fuel in the center tank, the boost pumps in the main tanks supply fuel to the engines.

November 2000

When the boost pumps are off, the APU gets fuel from main tank 1. An APU fuel shutoff valve lets fuel flow from the engine fuel feed manifold to the APU fuel manifold. The APU master switch, on the forward overhead panel, controls the APU fuel shutoff valve. An optional APU DC fuel pump supplies fuel from main tank 1.

10-5

Motive Flow Inlet

Boost Pump Inlet Motive Flow Discharge

Motive Flow Discharge

Motive Flow Pump Motive Flow Inlet

Forward Boost Pump Tank 1

Motive Flow Pump Boost Pump Inlet Motive Flow Pump

Aft Boost Pump (2)

Water Scavenge System

Center Tank Boost Pump (2)

Float Valve

Motive Flow Inlet Fuel Scavenge System

Fuel and Water Scavenge System Fuel Scavenge System

Water Scavenge System

The fuel scavenge system removes remaining fuel in the center tank and transfers it to main tank 1. This increases the usable fuel quantity in the center tank.

The water scavenge system removes water from the low points in each tank to help prevent corrosion.

The forward boost pump in main tank 1 supplies motive flow to a jet pump. The jet pump removes fuel from the center tank and transfers it to main tank 1. A float valve controls fuel sent to main tank 1.

There is a jet pump in the main tanks. There are two jet pumps in the center tank. The jet pumps in the main tanks use motive flow from the aft boost pumps. The jet pumps in the center tank use motive flow from both center tank boost pumps. Each jet pump removes fuel and water from its related tank and discharges it to the boost pump inlet. The water mixes with the fuel and vaporizes during combustion.

10-6

November 2000

Fuel Compensator (3) Fuel Temperature Sensor Fuel Measuring Sticks (16)

Tank Unit (32)

429 Data Bus

To Flight Management Computer System

FQIS C

CROSS

FEED

VALVE POSITION LIGHTS

FUEL PUMPS L CTR AFT

FWD

FWD

1

AFT

FUEL PUMPS

8500 LOW

1

OPEN

OPEN

OPEN

CLOSED

CLOSED

CLOSED

TEST GAUGES

R

2

CTR

28 000 CONFIG

FUEL LB

2

OFF FUEL DOOR SWITCH BYPASS

8500

FUEL QTY

FUEL QTY

FUEL QTY

LB

LB

LB

IMBAL

TANK 2

Fuel Control Panel (P5)

Upper Display Unit

CENTER TANK

TANK 1

P15 Refuel Panel

Fuel Quantity Indicating System Fuel Quantity Indicating System The fuel quantity indicating system (FQIS) measures fuel quantity, calculates fuel weight, and shows fuel weight. FQIS components include these components: • • • •

Tank units Compensators Fuel quantity processor unit Densitometer (optional).

The tank units supply a capacitance signal that is equal to fuel depth. This signal goes to the FQIS processor. The processor uses the ARINC 429 data bus to send a fuel weight signal to the common display system (CDS), the refuel panel, and the flight management computer system (FMCS). There are 32 variable capacitance tank units in the three tanks.

November 2000

The compensators supply a capacitance signal that is equal to fuel quality and fuel temperature. This signal goes to the processor to calculate fuel density. There is one compensator in each tank.

Fuel measuring sticks in each fuel tank supply a direct indication of fuel quantity. The measuring stick is a calibrated, flat, bendable stick that is attached to the bottom surface of the wing.

The fuel quantity processor unit does these functions:

The fuel quantity indicators on the display units (DUs) in the flight compartment can show these messages:

• • • •

Calculates total fuel weight Calculates fuel weight in each tank Monitors the fuel system for faults Sends fault data to the control display units.

The fuel quantity processor also sends fuel weight information to the common display system and the P15 refuel panel. Fuel weight can show in either pounds or kilograms. An optional densitometer supplies a signal that is equal to fuel density. There is one densitometer in each tank.

•

•

•

LOW - when fuel quantity in a main tank is less than 2000lbs (907 kg) IMBAL - when the fuel quantity between the two main tanks is different by more than 1000lbs (453 kg) CONFIG - when there is more than 1600 lbs (725 kg) in the center tank, both center tank boost pumps are off, and either or both engines are on.

10-7

ENG VALVE CLOSED b

ENG VALVE CLOSED b SPAR VALVE CLOSED b

-20

0 FUEL 20 TEMP

-40

SPAR VALVE CLOSED b

40

C

FILTER a BYPASS

FILTER a BYPASS

VALVE OPEN b

CROSS

FEED

LOW a LOW a PRESSURE PRESSURE

FUEL PUMPS OFF

L

P5 Forward Overhead Panel

R ON

LOW a LOW a PRESSURE PRESSURE

AFT

FWD

CTR

LOW a LOW a PRESSURE PRESSURE

FWD

ON

1

AFT OFF

OFF

FUEL PUMPS

ON

2

Fuel Control Panel

Fuel System Control Fuel System Control The fuel system control is from the forward overhead panel. Forward and aft fuel pump switches control boost pump operation in main tank 1 and main tank 2. Amber LOW PRESSURE lights come on when the output pressure of the boost pump is low. Center tank fuel pump switches control boost pump operation in the center tank. LOW PRESSURE lights come on when the output pressure of the pump is low and the fuel pump switch is ON.

crossfeed valve is closed and the selector is in the OFF position. The blue SPAR VALVE CLOSED light comes on bright when the valve is in transit. The SPAR VALVE CLOSED light is on dim when the valve is closed. The SPAR VALVE CLOSED light is off when the valve is closed and the fuel pump switch is OFF. The amber FILTER BYPASS lights come on when the filter is going to clog. The fuel filter is on the engine fuel pump housing. The fuel temperature indicator shows the fuel temperature in main tank 1.

A crossfeed selector controls crossfeed valve operation. The blue VALVE OPEN light comes on brightly when the valve is in transit. The VALVE OPEN light comes on dimly when the valve is open. The VALVE OPEN light is off when the

10-8

November 2000

Auxiliary Power Unit Features

OPERABLE DURING REFUELING

•

Auxiliary Power System

OPERATES ON THE GROUND AND IN FLIGHT

The APU can supply electrical and pneumatic power during refueling.

•

APU Fuel System

•

APU Pneumatic System

The auxiliary power unit (APU) is an electrical and pneumatic power source for aircraft systems on the ground and in flight.

FULL AUTHORITY DIGITAL ELECTRONIC CONTROL

•

APU Lubrication System

•

APU Ignition and Starting System

•

APU Control and Indication

•

APU BITE and Maintenance Indications

ELECTRICAL POWER A 90 KVA starter generator supplies electrical power up to 32,000 feet altitude. Above 32,000 feet to 41,000 feet, the generator rating decreases to approximately 66 KVA.

The electronic control unit (ECU) is a full authority, digital electronic control unit that controls APU operation. ENHANCED HISTORICAL RECORDING A data memory module records and keeps APU operational data.

PNUEMATIC POWER The load compressor supplies pneumatic power up to 17,000 feet. APU STARTING A starter-generator can start the APU at altitudes up to 41,000 feet. EDUCTOR COOLING SYSTEM An eductor cooling system cools the APU compartment and the APU oil. It is highly reliable and its parts do not move.

November 2000

11-1

Electronic Control Unit

Data Memory Module

Auxiliary Power System Auxiliary Power System

ELECTRONIC CONTROL UNIT

DATA MEMORY MODULE

The auxiliary power unit (APU) supplies electrical and pneumatic power to the airplane. The APU can start at all altitudes up to 41,000 feet. Electrical power is available up to 41,000 feet. Pneumatic power is available up to 17,000 feet.

The electronic control unit (ECU) controls and monitors all phases of APU operation. It also stores system and fault information. System and fault information shows on the captain and first officer control display units (CDUs).

A data memory module (DMM) records APU hours and starts. The DMM also keeps various APU component part and serial numbers. Programmed and recorded information shows on the captain and first officer CDUs.

The APU is an AlliedSignal 131-9(B). The APU has these features:

The ECU also causes an APU protective shut down to prevent damage to the APU.

The DMM is on the left side of the APU compressor inlet plenum.

• • • •

Single-stage centrifugal compressor Two-stage axial turbine Separate single-stage centrifugal load compressor Modular design.

The ECU is in the aft cargo compartment, on the right side aft of the cargo door.

The APU is in the aft section of the fuselage.

11-2

November 2000

Auxiliary Power Unit Oil Cooler

Accessory Gearbox

Fuel Manifold Oil Fill Port

CompressorInlet Plenum

Fuel Control Unit Starter/Generator Bleed Control Valve APU - Right Side

Compressor Inlet Plenum Oil Cooler Accessory Gearbox

Igniter

Starter/Generator

Data Memory Module Oil Cooler Temperature Control Valve

Fuel Control Unit Lubrication Module Oil Fill Port APU - Left Side

APU Components November 2000

11-3

Engine Compressor

Turbine

Exhaust

FWD

Gearbox

Load Compressor

Air Inlet

Combustor

SCV Air Exit Duct

APU Engine - Introduction APU Engine - Introduction PURPOSE The APU engine supplies power to operate the load compressor and the APU starter-generator. GENERAL DESCRIPTION

The engine operates at a constant speed to provide 400 Hz generator output. The APU engine also supplies air for airplane systems. An inlet screen prevents foreign object damage (FOD) to the APU compressors.

The APU engine has these main sections: • • • • •

Accessory gearbox Single-stage load compressor Single-stage engine compressor Combustor chamber Two-stage axial flow turbine.

All the components in the engine that turn are on a common shaft. The shaft turns the accessory gearbox and the load compressor. The accessory gearbox turns the APU generator and other components.

11-4

November 2000

Auxiliary Power Unit

APU

Fuel Feed

Electronic Control Unit

Fuel Control Unit Fuel Flow Divider Fuel Manifold

APU Fuel System APU Fuel System

FUEL FLOW DIVIDER

The APU fuel system receives fuel from any fuel tank. The APU fuel system has these major components:

The flow divider separates metered fuel to the primary and secondary fuel manifolds. The primary fuel manifold always supplies fuel when the APU is in operation. The flow divider supplies fuel to the secondary manifold when metered fuel flow pressures increases to a certain level. The ECU controls a solenoid on the flow divider to change the secondary manifold fuel supply with reference to altitude and APU speed.

• • •

Fuel control unit Fuel flow divider Primary and secondary fuel manifolds.

FUEL CONTROL UNIT The fuel control unit pressurizes, filters, and meters fuel. The electronic control unit sends fuel control signals to the fuel control unit. The fuel control unit uses these signals to meter fuel. The electronic control unit controls fuel flow from start to shutdown.

November 2000

The primary and secondary fuel manifolds supply fuel to the fuel injector nozzles. INTERFACE Regulated fuel pressure, from the fuel control unit, operates the inlet guide vanes and the surge control valve actuator.

11-5

Inlet Plenum Load Compressor

FWD Delta P Transducer Bleed Air Valve

Total Pressure Transducer

Bleed Air Valve

Surge Control Valve

Surge Control Valve Inlet Guide Vane Actuator

FWD Electronic Control Unit

APU Pneumatic System APU Pneumatic System The APU supplies pneumatic power for the environmental control system and for main engine start. The air inlet door directs air into the inlet plenum. The APU takes inlet plenum air for use in two APU sections. The APU has two separate compressors, a power compressor and a load compressor. The power compressor supplies compressed air to the combustor. The load compressor supplies compressed air to the airplane pneumatic system. The engine section operates the load compressor.

11-6

Inlet guide vanes control the amount of air that enters the load compressor. The ECU controls the inlet guide vanes with the inlet guide vane actuator. The inlet guide vane actuator uses fuel pressure to operate the inlet guide vanes. A surge control valve lets enough air go through to the load compressor to prevent a compressor surge. The surge control valve sends excess pressurized air to the APU exhaust. This protects the load compressor from a surge. The ECU controls the surge control valve. Regulated servo fuel pressure, from the fuel control unit, operates the surge control valve actuator.

The ECU senses the APU bleed switch position on the P5 panel in the flight deck. When the bleed switch is in the ON position, the ECU energizes a solenoid on the bleed air valve. When the solenoid energizes, APU bleed air duct pressure causes the bleed air valve to move from the closed to the open position. The ECU monitors the position of these components: • • •

Inlet guide vane actuator Surge control valve actuator Bleed air valve.

The position of these components can be seen in the APU maintenance pages on the control display unit (CDU).

November 2000

Auxiliary Power Unit

Air/Oil Separator

Gearbox StarterGenerator

Turbine Load Engine Compressor Compressor Bearing Compt

Oil Cooler Lube Module Temperature Control Valve

Magnetic Chip Detector/Drain Plug Temperature Control Valve

Oil Cooler Legend: Pressure Scavenge Vent

Differential Pressure Switch Lubrication Assembly

Oil Fill Port and Sight Glass

Oil Level Sight Gage and Oil Level Sensor

APU Lubrication System APU Lubrication System The lubrication system lubricates and cools the gears, bearings, and shafts of these major components: • • •

Power compressor Load compressor Gearbox.

The lubrication system also lubricates the APU starter-generator. The lubrication module pressurizes, filters, and scavenges the oil. The lubrication module contains these components: • • •

Pressure and scavenge pumps Pressure and scavenge filters High oil temperature sensor.

November 2000

These are other lubrication system components: • • • • • •

Low oil pressure switch Starter-generator scavenge oil filter differential pressure switch Oil cooler Temperature control valve Magnetic chip detector Air/oil separator.

The gearbox sump is the oil reservoir. An oil fill port supplies oil to the reservoir. A sight glass shows oil quantity. The gearbox vents to the APU exhaust.

Scavenge pumps send oil to the scavenge filter before the oil returns to the reservoir. The APU exhaust gas operates an eductor. The eductor pulls APU compartment air through the oil cooler to cool the oil and APU compartment. A temperature control valve regulates oil flow to and from the oil cooler. If the starter-generator scavenge oil filter clogs, the differential pressure switch operates. If the airplane is on the ground and the engines are shut down at the time of the clog, the ECU causes an APU protective shutdown.

From the reservoir, the oil is pressurized, filtered, and cooled before it goes to the major components and the startergenerator.

11-7

MAINT

8 6

StarterGenerator

b

LOW OIL OVER PRESSURE FAULT a SPEED a a

10 EGT c x 100 0 4 2

APU OFF ON START

APU Start Switch and Indicators

Ignition Unit and Igniter Plug

Start Power Unit

Start Converter Unit Electronic Control Unit

APU Ignition and Start System APU Ignition and Start System

AC power supply to increase the speed of the starter motor.

The ignition and start system supplies the combustion spark and starts the APU acceleration. The system includes these components:

The starter-generator uses AC power from the start converter unit to turn the APU gearbox and APU.

• • • • •

The ignition unit supplies electrical power to the igniter unit. The igniter supplies spark to the combustion chamber. The ECU controls the power to the ignition unit.

Starter-generator Start power unit Start converter unit Ignition unit Igniter.

The APU generator is also the APU starter. The start power unit (SPU) changes AC and low voltage DC to high voltage DC power. The SPU then sends this high voltage DC power to the start converter unit (SCU). The start converter unit receives the high voltage DC electricity and changes it to AC. The SCU then matches phases and sends the AC power to the starter-generator. The SCU increases the frequency of the 11-8

November 2000

Auxiliary Power Unit MAINT b

LOW OIL PRESSURE a

FAULT

a

OVER SPEED a

10 8 EGT 6 o

4

c x 100

100% 0

100% Governed Speed

95%

2

APU Switch (P5)

Electric and Pneumatic Loads are Available

APU OFF ON

1 Starter-Generator Deenergizes

70%

START

60%

Ignition Unit Deenergizes

30%

Low Oil Pressure Light (P5) Goes Out

• Battery Switch On • APU Switch to the START Position & Release to ON • APU Fuel Shutoff Valve Open • Air Inlet Door Open • Ignition Unit Energizes • Starter-Generator Energizes 7%

Fuel Solenoid Valve Opens 60 Seconds (Std Day)

1

When you put the APU switch to ON or START, the APU fuel shutoff valve and air inlet door opens.

APU Engine - Operation - Start Airborne Auxiliary Power Operation - Start GENERAL You can start the APU up to an altitude of 41,000 feet (12,500 meters). The APU electronic control unit (ECU) controls these components: • • • • •

APU inlet door APU fuel shutoff valve APU fuel Ignition APU start system.

PRESTART The battery switch must be ON before you can start and operate the APU. If AC power is available, operate one or more of the fuel boost pumps. This gives pressurized fuel to the APU. Pressurized fuel makes the APU start better. November 2000

APU START Move the APU switch to the START position and release it. This sends a signal to the electronic control unit (ECU). The ECU then opens the APU fuel shutoff valve and the APU air inlet door. The ECU also causes the low oil pressure light to come on. When the air inlet door is fully open, the door switch sends a door fully open signal to the ECU. APU SEQUENCE The ECU controls this APU start sequence: •

•

•

At 0 percent speed and before the start system is energized, the ECU energizes the ignition unit At 0 percent speed for start or 7 percent speed for restart, the ECU energizes the startergenerator At 7 percent speed, the fuel solenoid valve opens

•

• • •

•

At approximately 30 percent speed, the low oil pressure light (P5) goes out. At 60 percent speed, the ignition unit deenergizes At 70 percent speed, the startergenerator deenergizes At 95 percent speed, the APU can supply electrical power and air The APU accelerates to, and stays at 100 percent speed.

Note: The inlet guide vanes (IGVs) close to 15 degrees with the APU bleed air valve closed. This keeps the load compressor cool when it does not have a load. The BAT DISCHARGE light on the electrical meters, battery, and galley power module comes on when the APU start uses DC power. The BAT DISCHARGE light does not come on when the APU uses AC power to start.

11-9

MAINT b

LOW OIL PRESSURE

FAULT a

a

OVER SPEED

a

10

EGT

8 6

C X 100

4

0

2

APU

OFF

P5 Forward Overhead Panel ON START

APU Switch and APU Indicators

APU Control and Indication APU Control and Indication

INDICATION

CONTROL

The indicator on the P5 panel shows exhaust gas temperature.

The ECU controls these APU functions: • • • • • • • • •

Start and ignition Fuel control Surge control Inlet guide vane control Normal shutdown Protective shutdown APU indications BITE/fault recording Data storage.

The APU switch on the P5 panel controls normal APU start and shutdown.

11-10

The MAINT light comes on when there is an APU maintenance problem. Maintenance problems do not always require immediate action. The LOW OIL PRESSURE light comes on when the APU oil pressure is low. Low oil pressure causes a protective shutdown. The FAULT light comes on when the APU has a fault that causes a protective shutdown. The OVERSPEED light comes on when the APU speed is too high and a protective shutdown occurs. The light also comes on when the overspeed protection test fails and overspeed protection is lost.

November 2000

Auxiliary Power Unit FAULT LIGHT - APU FUEL VALVE SHUTDOWN - DC POWERLOSS SHUTDOWN - ECU SHUTDOWN - FIRE SHUTDOWN - INLET DOOR SHUTDOWN - INLET OVERHEAT SHUTDOWN - LOSS OF EGT SHUTDOWN - LOSS OF SPEED SHUTDOWN - NO ACCELERATION SHUTDOWN - NO APU ROTATION SHUTDOWN - NO FLAME SHUTDOWN - OIL FILTER SHUTDOWN - OIL TEMPERATURE SHUTDOWN - OVERTEMPERATURE SHUTDOWN - REVERSE FLOW SHUTDOWN - SENSOR FAILURE - UNDERSPEED SHUTDOWN

LOW OIL PRESSURE

MAINT

FAULT

OVER SPEED

10

8 6

EGT o

4 cx100 2

0

APU Indicator Panel (P5)

OVERSPEED LIGHT - FUEL CONTROL UNIT SOLENOID FAILURE - OVERSPEED PROTECTION - OVERSPEED SHUTDOWN

APU BITE TEST MAIN MENU

1 / 1

< CURRENT STATUS < FAULT HISTORY MAINTENANCE HISTORY > < IDENT / CONFIG < INPUT MONITORING < INDEX

LOW OIL PRESSURE LIGHT - OIL PRESSURE SHUTDOWN

OIL QUANTITY >

CDU Display (P9)

PROTECTIVE SHUTDOWN CONDITIONS ECU

APU - Protective Shutdown Airborne Auxiliary Power Protective Shutdown GENERAL

PROTECTIVE SHUTDOWN These are the conditions that cause a protective shutdown and a fault light:

A protective shutdown prevents damage to the APU or the airplane.

•

The ECU controls the automatic protective shutdown of the APU. If the ECU finds a fault, it does a protective shutdown.

• • • •

There are three different protective shutdown indications in the flight deck. These are the flight compartment protective shutdown indications: • • •

Fault light Overspeed light Low oil pressure light.

The cause for the shutdown shows on the control display unit (CDU) on the P9 panel.

November 2000

• • • • • • • • • • • •

Fuel shutoff valve not in commanded position Loss of dc power ECU failure APU fire Inlet door not in commanded position APU inlet overheat Loss of both EGT signals No speed signal No acceleration No APU rotation No flame (no ignition) Generator filter clogged High oil temperature Overtemperature (EGT) Reverse flow (load compressor) Oil temperature or inlet air temperature sensor failure Underspeed.

These are the conditions that cause a protective shutdown and an overspeed light: • • •

Fuel control unit solenoid failure Loss of overspeed protection Overspeed.

Low oil pressure for 20 seconds causes a protective shutdown and a low oil pressure light. When a protective shutdown occurs, the ECU removes electrical power from these components: • • • • • •

Fuel solenoid Ignitor SCU start signal Bleed air valve (BAV) Fuel control unit (FCU) Surge control valve (SCV).

11-11

APU BITE TEST MA IN MENU < CURRE NT STATUS < FAUL T

APU BITE TEST 1 / 1

D ATE

HI STORY

< I DENT /

NO

1 3 0 1

MESS#

MANUAL

< I NDEX

< I NDEX

OIL QUA NTI TY >

BITE Test Main Menu

HI S T ORY

G MT 1 3 0 1

MA I N T

M ESS #

BL EED

AIR

49 - X X X X X CYCLE

4 9 - XXXXX

VALVE

HOURS

APU

P O S IT I O N

DI S A GRE E S

W IT H

< I NDEX

CURRENTS TA T U S >

CO M M AN D

APU APU OI L

TO

A P U B I T E T E ST

/

OIL

1 / 2

OL D 0

T O ECU

< INDEX

A PU B I T E T E S T BITE REAL TIME DISPLAY

CYCL ES

POWER

Current Status Display

1 / 6 /

QUANTITY

CONFIRM REPAIR

CURRE NT S T A TU S >

I DENT DAT E OCT 0 4

OIL

49 - X X X X X

( LATCHED )

Fault History Display

A PU B I TE TEST MA I N T E N A N C E

LOW

MESS#

SHUT DOWN

SEE FAULT ISOLATION

< I NP UT MONI TORI NG

1 / 3

CURRENT STATUS MAINT

CYCLES OLD 1

ACCELERATION

MAINT

CONFI G

1 / 6

GMT

OCT 04

MAINTENANCE HISTORY >

APU BITE TEST

F AULT HISTORY

S/ N

1 0 1

HOURS

8 3 6

CYCL ES

1 1 6 5

L EVEL

F UL L



< DF C S

APU >

< A/ T

FQIS >

2

ENGINE

1

< AD I RU < CDS < I NDE X

D S P Yw F A I La

INIT REF

RTE

CLB

CRZ

DES

DIR INTC

LEGS

DEP ARR

HOLD

PROG

N1 LIMIT

FIX

PREV PAGE

1

A

NEXT PAGE

2

3

B

C

REVERSER

REVERSER

a

a

EEC

BRT

EEC

EXEC

D

ENGINE CONTROL

E

F

G

H

I

J

K

L

M

N

O

4

5

6

P

Q

R

S

T

7

8

9

U

V

W

X

Y

0

+/ -

Z

DEL

/

CLR

M S G w

ON a

ALTN

w a

ON ALTN

w

ENGINE CONTROL a

a

Engine Panel (P5)

O F S Tw

Control Display Unit

Fault Isolation SYSTEM FAULTS AND BITE

ENGINE PANEL

CONTROL DISPLAY UNIT

An ENGINE CONTROL Indication light comes on for a major engine fault. This light, on the aft P5 panel, is amber. The airplane must be on the ground for the light to come on. The engine may or may not be running.

The CDU shows exceedance data for each engine. A red exceedance box on the engine display tells maintenance personnel to get BITE data from the CDU. BITE data shows the most current exceedance event and the flight leg. Historical data shows on other CDU pages. In addition to exceedance data, engine fault data is available on the CDU. All fault messages have a number that is in the fault isolation manual (FIM). Past and present flight fault data may be recalled. The CDU also supplies this data: • •

Air data faults cause the ALTN amber light to come on. For airplane dispatch, both EEC switches are put to off. This makes both engines run in the hard-alternate mode. This procedure prevents unequal thrust and throttle stagger. The crew must be careful to not exceed the rated engine thrust in this mode.

Engine configuration Functional test data on engine control components and the ignition system.

12-12

November 2000

Power Plant

Drag Link

Forward Thrust

Drag Link

Translating Sleeve

Translating Sleeve Blocker Door (10)

Blocker Door (10)

Cascades

Reverse Thrust

Deployed

Stowed

Thrust Reverser Operation Thrust Reversers The thrust reversers (TRs) use electrical control. The TRs operate with hydraulic power. Both TRs are interchangeable between the two engines except for the cascades. Reverse thrust occurs with a change of fan air direction. When the thrust reversers deploy, the blocker doors close and fan air goes out radially and forward. When the translating sleeves extend, these events occur: • • • •

Cascades uncover Blocker doors deploy Blocked fan air goes out through the cascades Cascades direct the fan air forward.

November 2000

12-13

System A Hydraulics

Fire Warning Switch

STBY Hydraulics Deploy Eng Accy Unit

Control Valve Module

28V DC

Stow

Air Stow Deploy

Thrust Lever Angle Resolver

Interlock Actuator Electronic Engine Control

Ground Stow

Deploy

T/R Sleeve Position Reverser Deployed Engine 1

Thrust Reverser Actuation System Schematic Thrust Reverser Operation The TR system may be armed only during these conditions: • •

Airplane on the ground or less than 10 feet from the ground The fire warning switches are in the normal position.

thrust lever. This movement increases reverse thrust power. STOW

An automatic restow system activates when the engine accessory unit (EAU) senses a sleeve out of the stow position.

Movement of the reverse thrust lever to the stow position changes the direction of hydraulic fluid. The actuators move the sleeves to the stow position.

DEPLOY The TR system arms when movement of the reverse thrust levers sends power to the control valve modules. This energizes solenoids which open valves. The valve modules are in the main landing gear wheel wells. The valves send pressurized hydraulic fluid to the TR actuators. The actuators move the sleeves to the deploy position. When the sleeves deploy, the electronic engine control releases an interlock that permits further movement of the reverse

12-14

The TR hydraulic system pressurizes when you select reverse thrust. It also pressurizes when the automatic restow system energizes. Each engine TR system gets power from separate hydraulic systems. Hydraulic system A supplies power to the left engine TR. System B supplies power to the right engine TR. The standby hydraulic system supplies power to the left, right, or both engine TRs if loss of the normal hydraulic system pressure occurs. The standby system reduces the TR deployment and stow rate. November 2000

Power Plant

P2 Center Instrument Panel P5 Aft Overhead Panel

REVERSER

REVERSER

a

a

EEC ENGINE CONTROL

Engine Display

2

ENGINE

1

ON a

ALTN

EEC w a

ON ALTN

w

ENGINE CONTROL a

a

Engine Panel

Thrust Reverser Indication Thrust Reverser Indication The EEC must be on to sense thrust reverser sleeve position. The EEC sends a signal to the DEUs for thrust reverser sleeve position. A REV indication shows above the N1 indication on the engine display when the thrust reverser operates. The indication is amber if the thrust reverser is not in the commanded position. The indication is green when the thrust reverser is in the fulldeploy position. The indicator is off when the reverser is in the stow position.

November 2000

The engine accessory unit (EAU) operates a REVERSER fault light. The lights are on the engine panel on the P5 aft overhead panel. When this fault light comes on, the master caution light also comes on after 15 seconds delay. When the light comes on, do a test of the system. BITE is on the EAU in the EE compartment for fault isolation.

12-15

Hydraulics Features

CONNECTION

•

Hydraulic Systems

TRIPLE REDUNDANCY

A hydraulic service connection in the main landing gear wheel well makes it possible to service all three reservoirs from one location.

•

System Distribution

•

System Controls

•

Servicing

•

Maintenance

•

Pressure Module

System A has one engine-driven pump (EDP) and one electric motordriven pump (EMDP) for these systems: • • • • • • •

Flight controls Flight spoilers Landing gear Nose gear steering Alternate brakes Left thrust reverser Ground spoilers.

LEAK CONTROL FUSES The system pressure lines have fuses to protect the hydraulic system from fluid loss if a major component fails or if a line leaks.

System B has one engine-driven pump and one electric motor-driven pump for these systems: • • • • • •

Flight controls Flight spoilers Normal brakes Trailing edge flaps Leading edge devices Right thrust reverser.

The standby system has one electric motor-driven pump that supplies the third power source for the rudder control system. The standby system also supplies the second power source for these systems: • •

Thrust reversers Leading edge devices.

FLIGHT CONTROLS Hydraulic systems A and B supply power for all three axes of the flight control system. MULTIPLE FILTRATION The hydraulic fluid goes through filters while servicing and during operation of the system to increase reliability. MODULAR COMPONENTS Modules reduce the number of components, connections, and fittings. This makes the system more reliable and easier to service. SINGLE-POINT SERVICE November 2000

13-1

System A

System B

Standby System

A Reservoir

B Reservoir

Standby Reservoir

Legend: A Accumulator Backup E Electrical Backup

EDP

EMDP

EDP Normal

EMDP

Ailerons

Normal

Aileron Autopilot

Normal

Elevator

Normal

Normal

Elevator Autopilot

Normal

Normal

Rudder

Normal

Normal

Elevator Feel

Normal

No. 2, 4, 9 & 11

Flight Spoilers

No. 3, 5, 8 & 10

No. 1, 6, 7 &12

Ground Spoilers

Left Engine

Thrust Reversers

M

Normal Normal

M

PTU Alternate T

EMDP

NG, MG Actuation NG Steering

Normal

LE Slats and Flaps

Normal

Brakes

Normal

Landing Gear

T Landing Gear System Transfer Valve (For Landing Gear Retraction And NG Steering)

M

M

Manual Extension

Rudder

Right Engine

TE Flaps

M Manual Backup

EMDP (Electric MotorDriven Pump) EDP (Engine-Driven Pump) PTU (Power Transfer Unit)

Thrust Reversers E LE Devices (Extension Only) A

Main Deck Cargo Door (737-700C)

Hydraulic System Hydraulic Systems

SYSTEM B

STANDBY SYSTEM

The three hydraulic systems operate independently at 3000 psi nominal pressure. The three systems are system A, system B, and the standby system. Each system has a reservoir, pumps, and filters.

System B uses one EDP and one EMDP. System B supplies hydraulic power to these systems:

The standby system, which has a separate electric motor-driven pump, is an auxiliary source of hydraulic power. To operate the standby system, move either of the FLT CONTROL switches to STBY RUD or the ALTERNATE FLAPS switch to ARM. The switches are on the overhead panel. The standby system also operates automatically. Standby hydraulic power supplies pressure to these systems:

The hydraulic fluid is fire resistant. SYSTEM A System A uses one engine-driven pump (EDP) and one electric motordriven pump (EMDP). The system supplies hydraulic power to these systems: • • • • • • •

Flight controls Landing gear Nose gear steering Alternate brakes Flight and ground spoilers Left thrust reverser Power transfer unit (PTU).

13-2

• • • • • • • •

Flight controls Normal brakes Trailing edge flaps Leading edge flaps and slats Right thrust reverser Flight spoilers Alternate nose gear steering Alternate gear retraction.

Regulated air from the environmental control system pressurizes system A and system B reservoirs. The air gives a positive supply of hydraulic fluid to each pump. The standby system reservoir gets pressure from fluid in the system B reservoir.

• • •

Rudder control system Either or both thrust reversers Leading edge flaps and slats (full extend only) in the alternate flap mode.

For normal operation, system A and system B are on and the standby system is off.

November 2000

Hydraulics System Distribution

LANDING GEAR AND BRAKES

THRUST REVERSERS

FLIGHT CONTROLS

System A normally supplies pressure for operation (extension and retraction) of the landing gear.

System A supplies power to the left thrust reverser and system B supplies power to the right thrust reverser. If there is no primary system (A or B) pressure, the standby system supplies hydraulic power to operate one or both of the thrust reversers.

Although system A and system B supply hydraulic power for the primary flight controls, either system alone will operate the ailerons, elevators, and rudder. As a third backup, the ailerons and elevators can operate by manual reversion and the rudder operates by the standby hydraulic system. Trailing edge flaps normally receive power from system B. They also receive power from an electrical backup system for extension or retraction. The standby system supplies a secondary means to fully extend the leading edge flaps and slats. The power transfer unit (PTU) supplies a backup source of hydraulic power for normal and/or autoslat operation. The PTU has a hydraulic motor and pump on a common shaft. The PTU receives pressure from system A to turn the motor. The pump of the PTU receives fluid from the system B reservoir. The PTU operates automatically when these conditions are true: • • •

Airplane in the air Trailing edge flap position between up and 15 Engine-driven pump in system B has low pressure.

Some of the flight spoilers continue to function if either system A or system B pressure is not available.

If engine 1 does not operate for takeoff, system B supplies pressure to retract the landing gear. The landing gear transfer valve changes the pressure supply for the landing gear from system A to system B. All of these conditions are necessary for the landing gear transfer valve to automatically move to the alternate, system B position: • • • •

Airplane in the air Landing gear lever not down One or both main landing gear not up and locked Left engine N2 speed less than 50%.

The nose wheel steering system receives pressure through the nose gear and main gear actuation. System A normally supplies pressure for nose wheel steering. The landing gear transfer valve moves manually to the alternate (system B) position when these conditions are true: • • •

Alternate nose wheel steering switch to the alternate position Airplane on the ground Normal quantity in the system B reservoir.

If system A hydraulic pressure is not available for extension, the landing gear can be extended manually. The manual release, in the flight compartment, lets the landing gear free-fall to the down and locked position. The normal brake system gets power from hydraulic system B and the alternate is system A. Braking automatically transfers to the alternate system if hydraulic system B pressure is not available. The normal brake system has a brake accumulator with fluid capacity for braking when system B and system A do not have pressure.

November 2000

13-3

System Controls Controls and indicators for the hydraulic system are on the forward overhead panel and the center main panel. ENGINE-DRIVEN PUMP SWITCHES The ON position is the normal position for each engine pump switch. The depressurizing valve solenoid does not operate and the valve closes.

OFF—Hydraulic system pressure for either system A or B stops to these components: • • • •

SPOILER SHUTOFF VALVE SWITCHES These valves control hydraulic system pressure to the flight spoilers. •

ELECTRIC MOTOR-DRIVEN PUMP SWITCHES

Ailerons Elevators Rudder Elevator feel system.

•

ON—System A or B will supply power to the flight spoilers OFF—No system pressure from A or B is available to the flight spoilers.

• •

Turns on the standby pump Trailing edge flap bypass valve moves to the bypass position Arms the alternate flaps control switch.

The amber lights come on for system A and B EMDPs to show an overheat condition.

STANDBY HYD LOW QUANTITY LIGHT

FLT CONTROL SWITCHES

This light comes on amber when the standby hydraulic reservoir fluid quantity is less than 50% full.

• • •

The standby pump Pressurizes the standby rudder power control unit Closes the system A or B flight control shutoff valve.

ON—Normal pressure for either system A or B to flight controls.

13-4

HYDRAULIC BRAKE PRESSURE INDICATOR

• •

•

Either switch A or B in this position turns on these components:

Low pressure: 2000 psi Normal pressure: 3000 psi Maximum pressure: 3500 psi.

The ARM position operates these functions:

ELECTRIC MOTOR-DRIVEN PUMP OVERHEAT LIGHTS

STBY RUD

• • •

ALTERNATE FLAPS SWITCH PUMP LOW PRESSURE LIGHTS

The flight control switches control hydraulic system pressure to ailerons, elevators, rudder, and elevator feel system.

These pressure indicators show hydraulic system A and system B pressure. When both pumps for a system are off, the indicator shows zero pressure. These are the gage pressure ranges:

The indicator shows brake accumulator pressure. It also shows the accumulator precharge pressure when the accumulator completely bleeds off.

The ON position sends power to the electric motor-driven pump (EMDP).

Amber low pressure lights come on when the pump pressure is less than normal. All low pressure lights operate at the same pressure value.

HYDRAULIC SYSTEM PRESSURE INDICATORS

STANDBY HYD LOW PRESSURE LIGHT This light comes on amber when the standby pump pressure is too low. The light arms when any one of these conditions are true:

SYSTEM RESERVOIR QUANTITY INDICATION The reservoir quantity for system A and system B show digitally in percent of full. The refill level is 76%. When the quantity is 76% or less, an RF message shows adjacent to the quantity indication. System A (center main panel): • •

• •

Either FLT CONTROL switch in the STBY RUD position The ALTERNATE FLAPS switch in the ARM position The standby system operates automatically.

100%-Full: 5.7 U.S. gallons (21.6 liters) 76%-Refill: 4.7 U.S. gallons (16.4 liters).

System B (center main panel): •

•

Normal pressure: 3000 psi Normal precharge pressure: 1000 psi.

•

100%-Full: 8.2 U.S. gallons (31.1 liters) 76%-Refill: 6.9 U.S. gallons (23.6 liters).

The standby system reservoir has a capacity of 3.5 U.S. gallons (13.3 liters). The standby hydraulic low quantity light (P5 overhead panel) comes on when the quantity is 1.8 gallons or less (50% of full).

November 2000

Hydraulics Flight Control Switches

FLT CONTROL A B STBY RUD

Flight Control Low Pressure Lights

STANDBY HYD LOW QUANTITY

STBY RUD

OFF

OFF

A ON

B ON

Standby Hydraulic Low Quantity Light

a

LOW PRESSURE a

ALTERNATE FLAPS

Standby Hydraulic Low Pressure Light

OFF

Flight Spoiler Shutoff Valve Switches

P5 Forward Overhead Panel

LOW PRESSURE

a

LOW PRESSURE

UP a

OFF

SPOILER A B

ARM DOWN

Alternate Flaps Master Switch System A Hydraulic Overheat Light System A Pump Low Pressure Lights System A Hydraulic Pump LOW PRESSURE Switches a

System B Overheat Light

OVERHEAT a LOW PRESSURE a

ELEC 2

ENG 1

OVERHEAT a LOW LOW PRESSURE PRESSURE a a

ELEC 1 OFF

ON

ON

OFF

ON

ON

FEEL DIFF PRESS a

Hydraulic System B Pump Low Pressure Lights

SPEED TRIM FAIL a MACH TRIM FAIL a

YAW DAMPER YAW DAMPER

ENG 2

OFF

OFF

AUTO SLAT FAIL a

a

OFF ON

A

HYD PUMPS

Hydraulic Panel

B

Flight Control Panel System B Hydraulic Pump Switches

Hydraulic Brake Pressure Indicator

3

2

4 BRAKE PRESS

1 0

Hydraulic System B Quantity

Hydraulic System A Quantity

Refill Message

Hydraulic System A Pressure

Systems Display

Hydraulic System B Pressure

Hydraulic Systems Controls and Indications November 2000

13-5

Servicing

Maintenance

All three hydraulic reservoirs fill from a convenient single-point service connection in the right wheel well. These are the main components:

Hydraulic system supply and power components are easy to reach. Most of the hydraulic components are in the main wheel well. These are the components in the wheel well:

• • •

Hand pump with suction hose Connection for ground cart pressure fill Selector valve.

Electrical power is not necessary to read reservoir fluid quantity. System A and B reservoirs have mechanical quantity gauges that are visible from the servicing location. To fill the hydraulic reservoir, maintenance personnel use either a ground cart that connects to the pressure fill connection or the hand pump and suction hose. The system B reservoir fills through the standby reservoir. When the system B quantity indicator shows full, both system B and standby reservoirs are full. Hydraulic system reliability is better because of filtration. System A and B have a pressure, return, and case drain filters. Individual filters in the system supply additional filtration for critical areas such as flight control power units. The standby system has pressure and case drain filters.

13-6

• • • • •

System A and B EMDP Three reservoirs Standby pressure module Filters Hydraulic servicing station.

High reliability comes from modular units which help fluid flow and reduce the number of necessary fittings. The system A pressure module shows as an example. The pressure module includes these components: • • • •

Two pressure filters Two check valves Two pressure switches Pressure relief valve.

Use of this modular assembly removes many extra tubes and hydraulic connections. You can replace the entire pressure module or replace individual components without removal of the module.

November 2000

Hydraulics Reservoir Pressurization Module

System A Reservoir

System B Pressure Module

Reservoir Fill Selector Valve

System B Reservoir

EDP Supply Shutoff Valve (2)

Reservoir Fill Pressure Filter

System A Pressure System A Module Return Filter Module

Manual Fill Hand Pump (and Hose) PTU Pressure Filter

System A EMDP Power Transfer Unit (PTU)

System B Return Filter Module

System B EMDP

Main Landing Gear Wheel Well (Looking Forward)

Engine Accessory Gearbox

Engine-Driven Pump Standby System Reservoir

Standby Pressure Module

Standby System EMDP

FWD

Aft Wing-to-Body Fairing

FWD

Main Landing Gear Wheel Well

Hydraulic Power Systems Component Locations November 2000

13-7

System Pressure Transmitter (To Upper Center Display Unit)

EDP Low Pressure Switch EMDP Low Pressure Switch To Alternate Brake Source Selector Valve and Spoiler Control Valve

To Landing Gear Transfer Valve

EDP Check Valve

Pressure Line From EMDP 2

EMDP Check Valve

Note: EDP (Engine-Driven Pump) EMDP (Electric Motor-Driven Pump) PTU (Power Transfer Unit)

System Relief Valve

FWD

EMP Pressure Filter To PTU

System A Pressure Module (System B Similar) Pressure Module The pressure module supplies and filters EDP and EMDP pump pressure to user systems. The module is on the forward bulkhead of the left main gear wheel well. The pressure module includes these features: • • • • •

Pressure filters Pump low pressure switches System pressure transmitter Pressure relief valve Check valves.

A check valve downstream of each pressure filter and each low pressure switch isolates them from the pressure output of the opposite pump. The system pressure transmitter connects to the system pressure module downstream of both one-way check valves. Because of this, the pressure transmitter shows system pressure and not individual pump pressure. The signal from the pressure transmitter goes to the display electronics unit. You can replace each component on the module without module removal.

A non-bypass cartridge type filter in the pressure line from each pump filters the fluid before it goes to the user systems. The filter has a noncleanable filter element inside a metal bowl.

13-8

November 2000

Landing Gear Features MAIN GEAR WHEEL DOORS ARE NOT NECESSARY A blade seal fits around the outboard wheel. The only doors are small segment doors attached to the shock strut. Complicated linkages and actuators are not necessary. DUAL INDEPENDENT HYDRAULIC BRAKE SYSTEMS Hydraulic system B operates the normal brake system. Hydraulic system A is the primary alternative, and the brake accumulator is the secondary alternative.

November 2000

PRESSURE-MODULATED ANTISKID SYSTEM

•

Main Landing Gear

•

Nose Landing Gear

Provides maximum brake force for different runway conditions.

•

Tires, Wheels, and Brakes

AUTOBRAKE SYSTEM

•

Brake System

The autobrake system applies the brakes on landing or for a refused takeoff (RTO).

•

Antiskid System

•

Autobrake System

NOSE WHEEL STEERING

•

Air/Ground System

The pilot’s steering wheel and rudder pedals control the nose wheel steering.

•

Indication and Warning

•

Landing Gear Controls

•

Brake Controls

14-1

Beam Hanger

Walking Beam

Main Gear Actuator

Reaction link Downlock Springs

Outer Door

Main Gear in Retracted Position

Downlock Link Uplock Assembly

Side Strut

Center Door

Outer Door

Gas Charging Valve Oil Charging Valve (Aft Side Not Shown)

Inner Door

Uplock Roller

Shock Strut

FWD

Main Gear Actuator

Inner Door Downlock Link

Center Door

Side Strut

INBD

Axle Jacking Points (Not Shown) Torsion Links

INBD

Main Gear Components Main Landing Gear The main landing gear is a dualwheel, conventional landing gear. It has high operational reliability and a low maintenance design. The main gear absorbs landing impacts with a nitrogen-oil strut. It also absorbs vibrations while the airplane moves on the ground. The shock and side struts transmit loads from the gear to the airplane structure. The only doors on the main gear installation are small segmented doors attached to the shock strut and hinged to the wing. The doors close when the struts retract. The outer surface of the outboard tire aligns with the contour of the airplane to form the aerodynamic fairing for the wheel well opening.

14-2

Wheel well blade seals reduce noise and drag. The main gear design offers the operator these advantages: • • •

Minimum spare inventory Mechanical gear doors eliminate sequencing valves and actuators Easy access to strut servicing valves.

MANUAL EXTENSION The manual extension system permits landing gear extension if hydraulic system A has no pressure. A manual gear release from the flight compartment starts gear free fall to the down and locked position. The forces that pull down the gear are gravity and wind loads.

NORMAL OPERATION The main landing gear operates hydraulically. Extension uses system A. Retraction uses system A, or system B if necessary. Overcenter mechanical and hydraulic locks hold the gear in these positions: • •

Full extension Full retraction.

November 2000

Landing Gear Upper Drag Brace Link

Nose Gear Actuator Nose Gear Door Mechanism

Lock Link Up/Down Lock Actuator

Nose Wheel Well Door (2)

Trunnion Oil Charging Valve

Up/Down Lock Actuator

Upper Torsion Link Lower Drag Brace Link Lower Torsion Link

Jacking Point

Upper Drag Brace Link

Lock Link

Shock Strut

Steering Actuator (2)

Gear Extended

Nose Gear Door Mechanism

Gas Charging Valve

Dynamic Load Damper

Trunnion

Lower Torsion Link Towing Lever Steering Actuator

Nose Wheel Well Door (2)

Tow Fitting

Nose Gear Components Nose Landing Gear

MANUAL EXTENSION

The nose gear is a dual-wheel type which retracts forward and up into the wheel well.

The nose gear manual extension operates by manual release from the flight compartment. The nose gear free falls to the down and locked position when you pull the release.

The nose gear uses a nitrogen-oil strut. A folding drag brace transmits loads from the strut to the airplane structure. At full extension or retraction of the nose gear, the overcenter mechanism of the lock links locks the drag braces. The nose wheel well doors operate by mechanical linkages that connect to lugs on the trunnion. The doors stay open while the gear is down. NORMAL OPERATION The nose gear is hydraulically actuated. Extension uses system A. Retraction uses system A, or system B if required.

November 2000

NOSE WHEEL STEERING The captain steering wheel controls the nose wheel movement to a maximum of 78 degrees in each direction. A first officer steering wheel is optional. The rudder pedals control the nose wheel movement to a maximum of 7 degrees in each direction. Nose gear steering operates hydraulically by system A through the landing gear hydraulic extend line.

Centering cams inside the nose gear strut center the gear before retraction. A dynamic load damper (DLD) in the steering system reduces vibration. A towing lever on the steering metering valve permits the airplane to tow throughout the full steering range. ALTERNATE NOSE WHEEL STEERING If hydraulic system A has no pressure, a switch in the flight compartment operates the landing gear transfer valve and permits steering with hydraulic system B.

The steering wheel overrides the rudder pedal input. Rudder pedal steering is not available after the nose gear strut becomes extended. 14-3

FWD Main Landing Gear Brake

Brake Wear Indicator Pin (2)

Torque Pin (2)

Brake Retention Cable Attachment

Wheel Bearing (2)

Rotor Drive Keys

Wheel Tie Bolts

Pressure Port Heat Shield (9)

Piston/ Automatic Adjuster Assembly (6)

Rotors & Stators

Brake Temperature Sensor (Optional)

Torque Takeout Slot

Outer Wheel Half Shear Bolt (4)

Main Landing Gear Brake

Inner Wheel Half Main Landing Gear Wheel

Main Gear and Brake Components Tires, Wheels, and Brakes

Brake System

•

The 737-600/700 main gear wheels use a standard tire size with an optional larger tire. The 737-800 is only available with the larger tire.

The brake system includes multi-disc brakes for each main gear wheel. Hydraulic system B operates the normal brake system. Hydraulic system A is the primary backup and selected by the alternate brake selector valve if system B fails. If both the A and B systems fail, the accumulator isolation valve selects the accumulator as the secondary backup. The brakesystem has these features:

When the gear retracts, the brakes are applied to stop the main gear wheel rotation. Spin brakes in the nose wheel well stop nose wheel rotation when the gear retracts.

The nose gear tire is the same size for all models. The brakes are steel and available from the two wheel suppliers. The 737-600/700 use a standard brake and the 737-800 uses a high capacity brake. Fuse plugs in the main gear wheels act as safety valves if the wheel temperature becomes too high. Excessive heat caused by unusual heavy use of brakes can cause above normal wheel temperature. A pressure relief valve in all wheels releases tire pressure if it becomes too much.

14-4

• • • • •

Hydraulic brake line fuses limit fluid loss if there is a failure (external leak such as a broken hose to a brake) downstream of the fuse. A fuse closes when the volume of hydraulic fluid through the fuse increases more than normal.

Full antiskid protection Autobrakes for landing or RTO Easy brake service and maintenance requirements Identical brake control from either pilot station Directional control through differential braking

November 2000

Landing Gear Normal System B Pressure

System B Return

Alternate System A Pressure

Accumulator

Alternate Source Selector Valve

Accumulator Isolation Valve

Return Line From Right Normal Brake Metering Line

To System (A) Return Normal Brake Metering Valve

Alternate Brake Metering Valve

Autobrake Pressure Module

Normal Brake Metering Valve

To System B Return

Autobrake Shuttle Valve

Autobrake Shuttle Valve

Normal Antiskid Valve

Normal Normal Antiskid Anti-Skid Valve Valve

Alternate Antiskid Valve

Hydraulic Fuse

Hydraulic Fuse

Hydraulic Fuse

Shuttle Valve

Alternate Brake Metering Valve Brake Pressure Switch

Brake Pressure Switch Parking Brake Shutoff Valve

Gear “Up” Pressure (System A)

Shuttle Valve

To System A Return

Legend: System B Pressure System B Return System A Return System A Pressure Gear Up Pressure

From Left Alternate Brake Metering Valve and Antiskid Return

Normal Antiskid Valve

Normal Antiskid Valve

Alternate Antiskid Valve

Hydraulic Fuse

Hydraulic Fuse

Hydraulic Fuse

Shuttle Valve

Shuttle Valve

Left Wheel Brakes

Right Wheel Brakes

Brake and Antiskid Systems Antiskid System The antiskid system supplies safe brake control for all runway conditions. The system protects the airplane from a skid condition caused by a stop of wheel rotation. The system has a speed transducer in each main wheel. A control unit in the EE compartment controls both the antiskid and autobrake systems. The system also has four normal antiskid valves and two alternate antiskid valves. The antiskid system gets input from each wheel speed transducer when the wheels roll on the ground. The system automatically controls brake pressure to each main wheel. If the pilot or autobrake system applies sufficient pressure to stop wheel rotation, the control unit reads wheel sensor speed inputs and sends applicable signals to the November 2000

antiskid valves. The brake pressure reduces to prevent a skid, then reapplies to an optimum pressure. This operation repeats if skid conditions continue as brakes are applied.

Autobrake System

There is antiskid protection for each wheel when the normal brake system operates. When the alternate brake system operates, antiskid protection is for a pair of wheels.

The system has a pressure control module in the main landing gear wheel well. The autobrake selector lets the pilots arm the system and select the level of auto brake. A control unit in the EE compartment controls the pressure control module.

The ANTISKID INOP amber light comes on when a fault in the system occurs during these conditions: • •

Normal flight operation BITE test.

The light tells the pilots that antiskid may not operate when the aircraft lands and brakes are applied.

The autobrake system automatically applies the brakes to stop the airplane after it lands or if a refused takeoff occurs.

The AUTO BRAKE DISARM amber light comes on when the pilot selects autobrakes and the system has been manually disarmed or a malfunction exists in the autobrake or antiskid systems.

Ground test features are in the antiskid/autobrake control unit for maintenance. 14-5

Air/Ground Discretes Air/Ground Relays

NOSE GEAR r NOSE GEAR g LEFT GEAR r LEFT GEAR g

Air-Ground Sensors (6) Position Sensors (12)

RIGHT GEAR r RIGHT GEAR g

Landing Gear Panel (P2) LEFT GEAR g

RIGHT GEAR g

NOSE GEAR g

Landing Gear Control Lever Position Switch

Landing Gear Indicator Lights (P5) BITE Panel

Flight Control Computers

PSEU

a

Trailing Edge Flaps Position Switches

PSEU Light (P5)

Proximity Switch Electronics Unit

Horn Reset Switch Autothrottle Switch Pack Aural Warning Unit

Aural Warning Horn

Air/Ground, Indication, and Warning Systems Air/Ground System The air/ground system supplies air and ground mode signals to airplane systems. Two air/ground sensors on each landing gear monitor the compression of the shock struts. Sensor signals go to the proximity switch electronics unit. The system processes signals from the air/ground sensors and sends air/ground discretes and signals to operate air/ground relays. Indication Position sensors monitor the up and locked or down and locked positions of the nose and main landing gears. Dual sensors at each location on the landing gear improves dispatch reliability. If one sensor does not

14-6

operate, the system will still give correct indication. The proximity switch electronics unit processes all position sensor inputs and sends outputs to the landing gear position lights in the flight compartment. When the landing gear extends to the down and locked position, three primary and three auxiliary green landing gear position lights come on. Warning There are two types of warnings for the landing gear, visual and aural. VISUAL WARNING Three red landing gear position lights come on when the landing gear moves during extension and retraction and during the gear not down warning.

The red lights come on for a gear not down warning when these conditions are true: • • •

Landing gear not down and locked Either thrust lever at idle Altitude below 800 feet or trailing edge flaps at more than 10 units position.

AURAL WARNING The proximity switch electronics unit sends a signal to the aural warning unit when any landing gear is not down and locked and these conditions are true: • • •

Either thrust lever at idle Altitude below 800 feet Trailing edge flaps extended.

The horn reset switch permits reset of the horn unless the flaps are in the landing position and landing thrust is selected. November 2000

Landing Gear In Transit Landing Gear Indicator Lights (Red)

LEFT GEAR g PSEU NOSE GEAR r NOSE GEAR g

Access Door

LEFT GEAR r LEFT GEAR g

Right Main Gear

RIGHT GEAR r RIGHT GEAR g

Down and Locked Landing Gear Indicator Lights (Green)

RIGHT GEAR g

NOSE GEAR g

PSEU Light and Landing Gear Indicator Lights (P5)

UP

Landing Gear Warning Horn Reset Switch

L A N D I N OFF G

Nose Gear

a

Left Main Gear

FWD

Landing Gear Lever

G E A R DN

Override Trigger

LANDING GEAR LIMIT (IAS)

Note: Manual release handles are on the flight compartment floor behind the first officer.

OPERATING EXTEND 270K-.82M RETRACT 235K EXTENDED 320K-.82M

FLAPS LIMIT (IAS) 1-250K 2-250K 5-250K 10-210K 230K

15-195K 25-170K 30-165K 40-156K ALT FLAP EXTEND

Landing Gear Speed Limit Placard Landing Gear Warning Horn

Landing Gear Panel (P2)

Landing Gear Controls and Indication Landing Gear Controls and Indication There are the landing gear controls and indications in the flight compartment:

PSEU LIGHT

MANUAL RELEASE HANDLES

When the proximity switch electronics unit finds a fault, the amber PSEU light comes on.

Three separate handles manually release all landing gear. These handles are in the flight compartment floor behind the first officer.

LANDING GEAR WARNING HORN LANDING GEAR INDICATOR LIGHTS There are one red and two green indicator lights for each gear. The red light comes on for these conditions: •

•

The gear is not down and locked and either throttle is not at the idle position Gear position does not agree with the landing gear lever position.

The green indicator lights come on when the related gear is down and locked.

November 2000

LANDING GEAR LEVER The landing gear warning horn operates when the airplane is in a landing configuration and the main landing gear is not down and locked. The sound from the horn stops when the main landing gear is down and locked. LANDING GEAR WARNING HORN RESET SWITCH

This lever has three positions; up, off, and down. The override trigger overrides the ground lockout in the landing gear control lever assembly. The ground lockout prevents the placement of the lever to the UP position while the airplane is on the ground.

This switch, on the control stand, stops the warning horn with trailing edge flaps and thrust lever(s) in certain positions. The horn stops automatically when the landing gear moves to the down and locked position. 14-7

NOSE WHEEL STEERING A L T

N O R M

Nose Wheel Steering Switch (P1)

Steering Wheel

Rudder Pedals

Landing Gear Controls and Indication Landing Gear Controls and Indication (continued) These are the controls for nose wheel steering in the flight compartment. STEERING WHEEL The captain steering wheel controls the nose wheel steering movement to a maximum of 78 degrees in each direction. A first officer steering wheel is optional. The steering wheel overrides the rudder pedal input.

available after the nose gear strut becomes extended (the airplane in the air). NOSE WHEEL STEERING SWITCH Nose wheel steering normally receives pressure from hydraulic system A through landing gear extension. If hydraulic system A has no pressure, this switch in the flight compartment operates the landing gear transfer valve and permits steering with hydraulic system B.

A pointer on the steering wheel and a placard on the sidewall panel show the amount of steering movement. RUDDER PEDALS The rudder pedals control the nose wheel steering movement to a maximum of 7 degrees in each direction. Rudder pedal steering input backdrives the steering wheel. Rudder pedal steering is not 14-8

November 2000

Landing Gear

1

BRAKE PRESS

THRUST

4 INCREASE

Hydraulic Brake Pressure Indicator

3 2

0 THRUST

Hydraulic Brake Pressure Indicator (P3)

STAB TRIM

FLAP

AUTO BRAKE

Autobrake Disarm Light

AUTO BRAKE DISARM a

Parking Brake Lever

2

Autobrake Selector Switch

1

3

FLAP DOWN

MAX

OFF

Antiskid Inoperative Light

INCREASE

FLAP UP

PULL

Landing Gear Warning Horn Cutout Button

RTO

ANTI SKID ANTI SKID INOP

a

Parking Brake Light (Red)

Auto Brake/Antiskid Panel (P2)

P10 Control Stand

Brake Controls and Indication Brake Controls and Indication HYDRAULIC BRAKE PRESSURE INDICATOR The indicator shows the pressure of the brake accumulator. Normal operating pressure is 3000 psi. ANTISKID INOPERATIVE LIGHT When the antiskid monitoring system finds a fault, the amber ANTISKID INOP light comes on. PARKING BRAKE LEVER To set the parking brake, push on the brake pedals then pull the parking brake lever. The parking brake linkage latches the brake pedal linkage in the pushed down position. The parking brake warning light comes on with the brake set.

November 2000

PARKING BRAKE WARNING LIGHT

As in the landing mode, manual braking overrides RTO. RTO autobrakes disarm at lift-off.

This light comes on red when the parking brake shutoff valve is in the closed position.

AUTOBRAKE DISARM LIGHT

AUTOBRAKE SYSTEM

The amber light comes on when the pilot selects autobrakes and any of these are true:

The autobrake system applies pressure to all the brakes to slow the airplane at the rate selected by the pilot. The pilot can select one of four deceleration levels before landing. The antiskid system operates normally during autobrake operation. Manual braking by the pilot will override and disarm the autobrake system.

• • •

There is a malfunction in the automatic brake system There is a malfunction in the antiskid system The system has been manually disarmed.

AUTOBRAKE SELECTOR SWITCH This switch permits selection of the necessary level of auto brake and arms the system.

The autobrake system also has a refused takeoff (RTO) mode. The pilot selects RTO prior to takeoff. The system applies maximum brake pressure when the pilot refuses a takeoff. 14-9

Flight Controls HYDRAULIC HIGH LIFT SYSTEM WITH ELECTRIC AND STANDBY SYSTEM BACKUP

•

Flight Controls

•

Lateral (Roll) Control

•

Spoiler System and Speedbrakes

•

Directional (Yaw) Control

High-lift devices supply an increase in lift at slower speeds for takeoff and landing.

The trailing edge flap system and the leading edge flap/slat system usually operate from hydraulic system B. The trailing edge flaps have an electric motor backup. The standby hydraulic system is the backup for the leading edge flap/slat system.

•

Longitudinal (Pitch) Control

•

Longitudinal Trim

HYDRAULICALLY POWERED AILERONS AND ELEVATORS

HYDRAULICALLY POWERED SPOILERS

•

High Lift Devices (T.E. Flaps and L.E. Devices)

Hydraulic systems A and B operate the ailerons and elevators. If hydraulic power is lost, manual reversion (manual operation of flight controls) is available.

Hydraulic systems A and B operate the spoilers.

•

Flight Control Switches and Indicators

WHEEL TO RUDDER INTERCONNECT SYSTEM (WTRIS)

•

Control Stand

Features The flight control system has control surfaces to allow the airplane movement about all three axes. HIGH-LIFT DEVICES

HYDRAULICALLY POWERED RUDDER Hydraulic systems A and B operate the rudder. Backup power comes from the standby hydraulic system.

The rudder moves automatically for roll coordination during manual reversion of the ailerons (standby hydraulic system supplies power to the rudder).

ELECTRIC STABILIZER TRIM WITH MANUAL BACKUP The horizontal stabilizer trim comes from an electric motor. Manual trim wheels, on the control stand, are backup.

November 2000

15-1

Rudder

Aileron (Typical) Aileron Balance Tab (Typical) Flight Spoilers (Typical) Yaw Axis

Leading Edge Slats (Typical)

Elevator Tab (Typical) Elevator (Typical)

Stabilizer

Inboard Flap (Typical) Ground Spoiler (Typical) Outboard Flap (Typical) Ground Spoiler (Typical)

Roll Axis

Pitch Axis Leading Edge Flap (Typical) (Not Shown)

Flight Controls Flight Controls The flight control system controls airplane movement around the roll, pitch, and yaw axes. These are the primary flight controls: • • •

Ailerons Elevators Rudder.

Primary flight controls usually get power form hydraulic system A and B. Either hydraulic system can supply power to all primary control surfaces. If A and B systems lose hydraulic pressure, aileron and elevator control changes to a mechanical backup system (manual reversion). The backup for the rudder is the standby hydraulic system.

The ailerons control airplane movement around the roll axis. The elevators control airplane movement around the pitch axis. The rudder controls airplane movement around the yaw axis. These are the secondary flight controls: • • • •

Spoilers Horizontal stabilizer Leading-edge slats and flaps Trailing-edge flaps.

15-2

November 2000

Flight Controls Leading Edge Devices Transit Lights

LE DEVICES a g g

1

FLAPS a

a

g g

g g

2

3

1 a

a

g g

g

4

2 a

3 TRANSIT EXT FULL EXT

g

4

a

g

a

a g

g

g

5

a g g

6

a g g

7

Leading Edge Slat Extend Light

a g g

Leading Edge Slat Full Extend Light

8

Leading Edge Flap Extend Light

SLATS

SLATS TEST

Leading Edge Devices Annunciator Panel (P5)

Test Switch

Speedbrakes Extended Light SPEEDBRAKES EXTENDED a

2

5

1 UP

Speedbrake Do Not Arm Light Speedbrake Armed Light

10

Trailing Edge Flap Position Indicator

Alert/Annunciator Lights (P3)

15 25

SPEEDBRAKE DO NOT ARM a SPEEDBRAKE ARMED

Leading Edge Devices Transit Light

FLAPS 40

LE FLAPS TRANSIT a

g

Alert/Annunciator Lights (P1)

30

LE FLAPS EXT g

Leading Edge Devices Extended Light

Flap Position Indicator (P2)

Flight Control Switches and Indicators Flight Control Switches and Indicators

SPEEDBRAKES EXTENDED LIGHT

LEADING EDGE DEVICES ANNUNCIATOR PANEL

This amber light comes on when the spoilers deploy and all of these conditions occur:

An annunciator panel, on the aft overhead panel, shows the leading edge flaps and slats position. The lights come on when the leading edge devices are in these positions: •

• •

TRANSIT (amber)—Leading edge devices are in transit, or not in the selected position EXT (green)—Leading edge slats and flaps in extend position FULL EXT (green)—Leading edge slats are in the full extend position.

All lights go off when the leading edge devices are in the retract position. All leading edge lights come on when you push the test switch. November 2000

LE FLAPS EXT LIGHT The green light comes on when the leading edge devices extend to the position selected by the flap lever. LE FLAPS TRANSIT LIGHT

• • •

Airplane in the air Speedbrake lever is more than arm position Trailing edge flaps at or more than 15 units or altitude is below 800 feet.

The amber light also comes on when the speedbrake lever is in the down position when the airplane is on the ground and the ground spoilers get hydraulic pressure. TRAILING EDGE FLAP POSITION INDICATOR The indicator is on the P2 center instrument panel. Separate pointers show position of left (L) and right (R) wing trailing edge flaps.

When the flap lever moves, this amber light comes on if the leading edge devices are in transit or one or more is not in the selected position. SPEEDBRAKE DO NOT ARM LIGHT The amber light comes on when the speedbrake lever is in the armed position and there is a malfunction in the auto speedbrake system. SPEEDBRAKE ARMED The green light tells the pilot the auto speedbrake system is ready to operate.

15-3

Flight Control Switches

FLT CONTROL B A STBY RUD

Flight Control Hydraulic System Low Pressure Lights

Standby System Low Quantity Light Standby System Low Pressure Light

STANDBY HYD STBY RUD

OFF

OFF

A ON

B ON

LOW QUANTITY

a LOW PRESSURE a

ALTERNATE FLAPS

Alternate Flaps Arm Switch

OFF

System A Flight Spoiler Switch P5 Forward Overhead Panel

Alternate Flaps Control Switch

UP

LOW LOW PRESSURE PRESSURE a a

OFF

SPOILER A B

System B Flight Spoiler Switch OFF

ARM DOWN

OFF FEEL DIFF PRESS a

ON

ON

Feel Differential Pressure Light Speed Trim Fail Light

SPEED TRIM FAIL a

Mach Trim Fail Light

YAW DAMPER

MACH TRIM FAIL a

Auto Slat Fail Light

YAW DAMPER

AUTO SLAT FAIL a

Yaw Damper Light Yaw Damper Switch

a

OFF ON

Flight Control Panel

Forward Overhead Panel Flight Control Switches and Indicators (continued) The flight control panel on the forward overhead panel has these lights and switches:

•

• • • • • • • • • •

•

Flight control switches Standby hydraulic lights Spoiler switches Alternat flaps switches Feel differential pressure light Speed trim fail light Mach trim fail light Auto slat fail light Yaw damper switch Yaw damper light.

standby rudder valve. This pressurizes the standby rudder power control unit. ON - In this position there is normal hydraulic system pressure to the primary flight controls. OFF - In this position hydraulic system pressure is shut off to the ailerons, the elevators, the rudder, and the feel computer.

STANDBY HYD LIGHTS The LOW PRESSURE light comes on amber when the standby system pressure is low. The light arms when the standby system operates.

FLIGHT CONTROL HYDRAULIC SYSTEM LOW PRESSURE LIGHTS A LOW PRESSURE light comes on amber when hydraulic pressure to the primary flight controls is low. SPOILER SWITCHES These switches control the position of the spoiler shutoff valves: •

•

ON - In this position the spoiler shutoff valve opens to supply hydraulic pressure to the flight spoilers. OFF - In this position the spoiler shutoff valve closes.

FLT CONTROL SWITCHES These switches control hydraulic system pressure to the primary flight controls. These are the switch positions for system A and B: •

The LOW QUANTITY light comes on amber when the standby system hydraulic fluid quantity is half full or less.

STBY RUD - Move either switch to this position to turn on the standby pump and open the

15-4

November 2000

Flight Controls ALTERNATE FLAPS SWITCHES

FEEL DIFF PRESS LIGHT

YAW DAMPER SWITCH

When the alternate flap arm switch is in the ARM position, these functions occur:

The FEEL DIFF PRESS light comes on amber when the feel computer pressure for system A and system B are different by more than a given amount. When the flaps extend, this light does not come on.

The YAW DAMPER switch controls the yaw damper. A solenoid holds the switch on. The switch spring loads to OFF.

• • • •

The standby pump supplies pressure The TE flap bypass valve moves to bypass The alternate flaps control switch arms The standby system low pressure light arms.

If the alternate flaps arm switch is in the ARM position and you move the alternate flaps control switch, these functions occur: •

•

•

DOWN - Electric motor extends the trailing edge flaps and hydraulic power extends the leading edge devices. OFF - Stops the movement of the trailing edge flaps but not the leading edge devices. The switch spring loads to OFF from the DOWN position. It remains in the UP position until you put it back to OFF. UP - Electric motor retracts the trailing edge flaps. The leading edge devices do not retract with alternate operation.

November 2000

SPEED TRIM FAIL LIGHT The amber SPEED TRIM FAIL light comes on to show a fault of the speed trim function. See the autoflight chapter for more information about the speed trim function.

When you move the switch to ON after loss of hydraulic system A and B, the wheel to rudder interconnect system (WTRIS) operates. YAW DAMPER LIGHT The amber YAW DAMPER light comes on when the yaw damper is not engaged.

MACH TRIM FAIL LIGHT The amber MACH TRIM FAIL light comes on to show a fault of the mach trim function. See the autoflight chapter for more information about the mach trim function. AUTO SLAT FAIL LIGHT The AUTO SLAT FAIL light comes on to show a fault in the autoslat system.

15-5

Speedbrake Lever (With Indicator) Manual Stabilizer Trim Wheel (2) Flap Position Lever (Detented Positions with Indicator)

Stabilizer Position Indicator (2)

Aileron Trim Switches

Rudder Trim Switch Stabilizer Trim Cutout Switches

RUDDER TRIM

LEFT

RIGHT

AILERON NOSE LEFT LEFT WING DOWN

RIGHT WING DOWN

R U D D E R

NOSE RIGHT

STAB TRIM

CAB DOOR

OVERRIDE

CAB DOOR UNLOCKED a

NORMAL

Top View Aileron and Rudder Trim Module

Stabilizer Trim Switching Module Override Switch (Column)

Control Stand Control Stand These are the controls for the flight controls on the control stand. SPEEDBRAKE LEVER The speedbrake lever moves flight spoilers up in air or on ground and ground spoilers up on ground. • • • •

ARMED - Arms the speedbrake control system for landing FLIGHT DETENT - Maximum position of flight spoilers in air UP - Maximum position of flight spoilers on ground DOWN - All spoilers stow.

STABILIZER POSITION INDICATOR This shows the position of the horizontal stabilizer. The green band shows the permitted takeoff stabilizer trim positions. If you start a takeoff with the stabilizer out of the green band range, a takeoff warning horn operates. STABILIZER TRIM CUTOUT SWITCHES

The rudder trim switch moves the rudder neutral position. AILERON TRIM SWITCHES The aileron trim switches move the aileron neutral position. You can see the amount of trim input on the top of the control column. STABILIZER TRIM COLUMN OVERRIDE SWITCH

These switches are on the control stand.

This switch is on the control stand.

•

•

NORMAL - The main electric trim motor has power available CUTOUT - Removes power to the main electric trim motor.

MANUAL STABILIZER TRIM WHEELS

•

Foldout cranks on each trim wheel permit either pilot to manually move the stabilizer.

FLAP LEVER

15-6

RUDDER TRIM SWITCH

•

NORMAL - Permits stabilizer trim through the column switching module OVERRIDE - Permits stabilizer trim in both direction if the column switching module fails.

The flap lever controls the position of the leading edge devices and trailing edge flaps. It has nine detent positions. November 2000

Flight Controls Control Wheel (2)

Trim Indicator (Top of Both Columns)

Transfer Mechanism

Flight Spoilers

Flight Spoiler Control Input (First Officer Cable Only)

Normal Lateral Control Input (Captain Cable Only)

Flight Spoilers

Balance Tab (Typical) Spoiler Mixer & Ratio Changer Aileron cable Quadrants (2) Autopilot Actuator System B

Aileron (Typical)

Autopilot Actuator System A

Aileron Trim Aileron Power Control Unit (2) Actuator Aileron Feel and Centering Unit

Aileron and Flight Spoilers (Roll) Control System Roll Control An aileron on each wing supplies primary control around the airplane roll axis. Two independent hydraulic power control units (PCUs) move the ailerons through cables. One PCU receives hydraulic power from system A and the other receives hydraulic power from system B. Either PCU can operate both ailerons to supply roll control. There is manual reversion for aileron control with both hydraulic system A and B off. Aileron balance tabs and balance panels keep the control forces to a minimum during manual reversion. These are the inputs that move the ailerons: • • •

Pilot command Autopilot command Aileron trim.

November 2000

Pilot input to the power control units is from the control wheels through a cable system. The captain cable system is the normal input path. Movement of the power control units operates a wing cable system which sets the position of the ailerons. A mechanical feel and centering unit with a centering cam, roller, and spring supplies control wheel feel force for the pilots. The autopilot, when engaged, controls the ailerons through autopilot actuators. These actuators supply input to the power control units and backdrive the control wheels.

Four flight spoilers on each wing operate with the ailerons. When the control wheel turns, the spoilers operate to help the roll movement of the airplane. Hydraulic system A operates flight spoilers 2, 4, 9, and 11. System B operates flight spoilers 3, 5, 8, and 10. The flight spoilers also supply lateral control if there is a malfunction in the aileron system. If a malfunction occurs, the first officer spoiler cable system controls the flight spoilers through an transfer mechanism. The flight spoilers can also operate as speed brakes. This function is in the spoiler system description.

The aileron trim switches on the aft of the P8 aft electronic panel control the aileron trim. The trim switches command an electrical linear actuator which moves the feel and centering unit.

15-7

12

11 10 9 8 7

Ground Spoiler Control Valve (Operated By Spoiler Mixer)

Speedbrake Quadrant 6

Spoiler Mixer Ratio Changer 5 4

3 2

Ground Spoiler Interlock Valve (Operated By Push-Pull Cable-Right Gear)

1

Flight Spoiler Quadrant (Typical)

Speedbrake Lever

Speedbrake Quadrant and Electric Actuator

Note: Flight Spoilers - 2, 3, 4, 5, 8, 9,10, and 11 Ground Spoilers - 1, 6, 7, and 12

Spoiler System Spoiler System and Speedbrakes There are four flight spoilers and two ground spoilers on each wing. The flight spoilers operate in the air and on the ground. They help the ailerons with lateral control and also operate as speedbrakes to increase drag and decrease lift. The ground spoilers operate only on the ground to decrease lift and increase drag. The speedbrake lever controls the flight spoilers in the air. The amount that the spoilers move depends on both the control wheel position and the speedbrake lever position. A mechanical spoiler mixer and a spoiler ratio changer give the correct spoiler extension on each wing from the two inputs. Aerodynamic forces can override actuator hydraulic pressure and limit spoiler panel extension to an amount in proportion to airspeed.

15-8

There are two ground spoilers on each wing. One ground spoiler is outboard of the flight spoilers the other is inboard of the nacelle. All the ground spoilers receive hydraulic power from system A. A ground spoiler interlock valve, operated by the right main gear strut, permits ground spoiler use only on the ground. The ground and flight spoilers operate together on the ground. The spoilers extend to reduce lift and increase aerodynamic drag. This helps stop the airplane in a shorter distance.

The spoilers operate either automatically or manually on the ground. Use the speedbrake lever to operate the spoilers manually. The spoilers extend automatically under these landing conditions: • •

Speedbrake lever is in the armed position Airplane on the ground or wheel rotation is more than 60 knots.

All spoilers automatically retract, after automatic extension, when either thrust lever advances. The flight crew can manually move the speedbrake lever to override the automatic spoiler system. During a refused takeoff (RTO), the spoilers extend if these two conditions occur: • •

One of the two reverse thrust levers operates The airplane speed is more than 60 knots. November 2000

Flight Controls

Rudder

Rudder Control Cables

Standby Rudder PCU

Trim Control and Indicator (Aft Section of the Center Aisle Stand)

Rudder Pedal Adjustment Crank (Typical) Captain’s Rudder Pedals

Feel and Centering Unit

Input Rod From Rudder Aft Control Quadrant

Main Rudder PCU (Both A and B Systems) Rudder Trim Actuator

Aft Control Quadrant

Rudder (Yaw) Trim and Control System Rudder (Yaw) Control The rudder gives control of the airplane around the yaw axis. The rudder is a single conventional rudder without tabs. The normal movement of the rudder is from the main rudder power control unit which uses hydraulic systems A and B. A separate power control unit, which uses standby hydraulic power, supplies backup movement. Any of the three hydraulic systems supply rudder control. Either pilot's rudder pedals operate the power control units through cables. A mechanical feel and centering unit gives the pilot feel forces and centers the rudder.

November 2000

An electric actuator on the feel and centering unit supplies rudder trim. A trim control switch on the aisle stand operates the trim actuator. Trim actuator movement gives an input to the power control unit (PCU) to move the rudder. The yaw damper system moves the rudder to prevent dutch roll. This system operates through the hydraulic system B control section of the main rudder power control unit. The yaw damper operates independently of the rudder control system and does not give feedback to the rudder pedals.

The wheel to rudder interconnect system also controls the rudder through the standby power control unit. It is active only when hydraulic systems A and B do not have pressure. Movement of the control wheel sends a signal to the standby power control unit to move the rudder. This gives rudder assist to help turn the airplane when control of the ailerons is through manual reversion.

15-9

Autopilot Actuator System B

Output Torque Tube Mach Trim Actuator

Autopilot Actuator System A

Right Elevator Input

Feel and Centering Unit Control Column Stick-Shaker (2)

Elevator PCU System B

Left Elevator Input Control Column (2)

Elevator PCU System A Breakout Mechanism

Forward Control Quadrant (2)

Input Torque Tube Captain’s Cable First Officer’s Input Quadrants Input Quadrants

Balance Weight (2)

Elevator (Pitch) Control System Elevator (Pitch) Control The elevators supply primary control about the airplane pitch axis. Two elevators connect to the aft end of the left and right horizontal stabilizer sections. Two independent hydraulic power control units (PCU) move the elevators. One PCU receives hydraulic power from system A and the other receives hydraulic power from system B. Either power control unit can operate both elevators to supply pitch control. These are the inputs that move the elevators: • • • •

Pilot command Autopilot command Neutral shift Mach trim.

15-10

Pilot input is from the control columns through a dual cable system and an input torque tube. The input torque tube connects to each PCU with two input rods. An output torque tube connects both power control units to both elevators. A hydraulic feel system supplies control column forces proportional to airspeed and stabilizer position. An elevator feel computer gets input of airspeed and stabilizer position and supplies the appropriate feel force. There is manual reversion for elevator control with hydraulic system A and B off.

Neutral shift moves the elevators to give a different elevator neutral position for different stabilizer positions. Neutral shift rods connect the stabilizer to the elevator feel and centering unit. The elevator’s neutral position changes as the stabilizer moves.

The autopilot, when engaged, controls the elevators through autopilot actuators. These actuators move the input torque tube which supplies input to the power control units.

A stick-shaker stall warning system gives the pilot the positive indication that the airplane is close to a stall. An electric motor attached to each control column shakes the column when the airplane comes near a stall condition.

The mach trim system commands the elevators at high speeds. The mach trim actuator moves the elevator feel and centering unit. This changes the elevator neutral position.

November 2000

Flight Controls

Stabilizer Support Structure (Center Section)

Stabilizer Drive Gearbox

Trim Position Wheels

Aft Cable Drum

Column Switching Modules (2)

Forward Cable Drum

Stabilizer (Pitch) Trim and Control System Stabilizer Trim The moveable horizontal stabilizer gives pitch trim to the airplane. The horizontal stabilizer is a three piece assembly. A jackscrew assembly attaches to the center section. The jackscrew moves the stabilizer assembly. These are the inputs that control the jackscrew: • • •

Main electric trim inputs to the stabilizer trim motor Autopilot and speed trim inputs to the stabilizer trim motor Manual trim wheels through cables.

The autopilot and speed trim systems also give commands to the stabilizer trim motor. The autopilot trim motor speeds are slower than the thumb switch control speeds. The electric trim system includes a column switch module to stop uncommanded stabilizer trim. If the pilot moves the control column opposite to the direction of the uncommanded trim, switches in the module stop the electric trim. An override switch on the aisle stand bypasses the column switch module if it malfunctions.

The manual stabilizer trim control wheels connect to the stabilizer gearbox with a forward and aft cable drum. Foldout handcranks on the trim wheels allow either pilot to manually trim the stabilizer. The cable system also operates trim position indicators next to the trim wheels on the control stand.

Thumb switches on either control wheel command the stabilizer trim motor for the main electric trim system. The motor operates at low speed with flaps up and high speed with flaps not up.

November 2000

15-11

Flap Lever Settings Flap Position Transmitter (Typical)

Input Cables Flap Control Lever

0

0

Cruise

1-40

Flap Control Unit

1 Slats 5-8

LE Flaps 3 and 4 (Not Shown)

LE Flaps 2 LE Slat Extend

LE Flaps 1 and 2 (Not Shown)

5

LE Slat Full Extend

Takeoff 10

Slats 1-4 15

Outboard Flap Track Fairing (4)

25 Flap Power Unit

30

Drive System Torque Tube Outboard Flap Transmission (4)

Landing

40

High Lift Devices Range of Position

High Lift Devices Range of Position High Lift Devices The high lift devices improve wing performance at low speeds. The high lift system includes leading edge flaps and slats and double-slotted trailing edge flaps. Hydraulic System B supplies power to the leading edge devices and the trailing edge flaps. Trailing edge flaps have doubleslotted inboard and outboard assemblies on each wing. Each assembly includes two mechanically linked segments that extend and separate to form a double-slotted surface for added lift. A hydraulic motor drives a flap power drive unit (gearbox) to operate all trailing edge flaps. A torque tube drive system transfers movement from the flap power drive unit to the flaps. Leading edge devices have two leading edge flaps inboard of each engine and four leading edge slats outboard of each engine. To extend 15-12

the leading edge devices, move the flap control lever on the control stand. The leading edge flaps and slats retract with the flap lever in the 0 position. The leading edge flaps extend with the flap control lever in any position from 1 to 40. The leading edge slats move to an extend position with a flap lever position from 1 to 5. The slats move to a full extend position with a flap lever position from 10 to 40. Autoslat operation automatically moves the slats from extend to the full extend position if the airplane approaches a stall. Normal operation of the leading edge flaps and slats comes from hydraulic system B. However, if the engine driven pump for system B has low pressure and the trailing edge flaps are in takeoff position, the power transfer unit (PTU) automatically supplies a backup source of hydraulic system B

power for normal and/or autoslat operation. The flap alternate operation uses electric power to drive the flap system if a failure prevents normal hydraulic operation. The hydraulic pressure shuts off to the hydraulic motor if any of these conditions occur: • •

• •

Flaps become asymmetric Flaps become skewed (inboard end of a flap does not align with outboard end of flap) Flaps have an uncommanded motion (UCM) Flaps operate with the alternate drive.

Takeoff flap positions supply high lift with low drag. Landing flaps produce high lift and high drag which help to decrease approach speeds.

November 2000

Flight Controls The flap load relief system protects the trailing edge flaps from excessive airloads. The flaps move up one position for these conditions: • •

Flaps are at 30 or 40 units Airspeed exceeds a set speed.

The flaps return to the selected flap position when airspeed reduces. The flap design helps with durability and maintainability. The heavy-gage lower surface skin improves damage tolerance. Flap Operation

HIGH LIFT DEVICES INDICATION An indicator on the center main instrument panel shows the trailing edge flap position. A flap position indicator has a right wing and left wing flap pointer. A green LE FLAPS EXT light on the center main instrument panel comes on when all leading edge devices are in the selected position. An amber LE FLAPS TRANSIT light comes on when any of the leading edge devices are in transit or not in the selected position. An annunciator panel, on the aft overhead panel, shows each leading edge flap or slat position.

For normal operations, the pilot selects the desired position with the flap lever on the control stand. Both leading edge devices and trailing edge flaps travel to the position selected. The possible flap lever positions are: UP, 1, 2, 5, 10, 15, 25, 30 and 40 units. An alternate system operates when the normal hydraulic source (system B) is not available for leading edge and trailing edge operation. The standby hydraulic system supplies power during alternate operation to move the leading edge devices to the fully extended position. An electric motor supplies power during alternate operation to move the trailing edge flaps. The alternate flap arming switch and the alternate flaps control switch on the forward overhead panel operate the trailing edge flaps and leading edge devices.

November 2000

15-13

Environmental Systems Features

CARGO COMPARTMENT HEATING

•

Pneumatic

•

Air-Conditioning

These sources supply the pneumatic manifold:

The forward cargo compartment heating is by exhaust air from the equipment cooling system.

•

Conditioned Air Distribution

•

Equipment Cooling

• • •

The aft cargo compartment heating is by outflow air from the passenger cabin.

•

Cargo Compartment Heating

•

Pneumatic and AirConditioning Control Panels

•

Cabin Pressure Control

•

Cabin Pressure Control Panels

PNEUMATIC

Engine bleed air APU bleed air Ground source.

The system controls and indications reduce crew workload. AIR CONDITIONING The air conditioning system is a dual air cycle pack design. The ram air system produces minimum drag. The pack air-cycle machines have air bearings. These bearings require no regular servicing. Pack temperature control is either automatic or manual. Automatic overtemperature protection reduces crew workload.

PRESSURIZATION The cabin pressure control system uses dual, automatic, digital pressure controllers. This increases reliability and reduces crew workload. Pressure controllers have BITE. There is a manual backup pressure control system. Independent, mechanical safety relief valves protect the airplane structure in any mode of pressure control.

System maintenance does not require ladders or special stands. A cabin air recirculation system reduces fuel consumption. EQUIPMENT COOLING The airplane uses two equipment cooling systems. Both systems have backup fans. The equipment cooling system automatically configures for ground and flight operations.

November 2000

16-1

Pneumatic Ground Service Connector Isolation Valve Wing Anti-ice Valve

M 2

Pressure Transmitter

To LE Slats To Starter Precooler Pressure Regulating and Shutoff Valve High Stage Valve

Precooler Control Valve Fan

S

S Pack Valve

Fan

To Left Air Conditioning System

S

5th

Check Valve

S 1

Relief Valve

9th

Apu S Bleed Valve

M Motor Operated Valve S Solenoid Controlled Valve

5th

2

From APU

9th

1

Potable Water Pressure Tap

2

Hydraulic Reservoir Pressure Tap

Pneumatics Pneumatics The pneumatic system supplies pressurized air to these systems and components: • • • • •

Engine starters Air conditioning packs Thermal anti-ice systems Hydraulic reservoirs Potable water system.

These are the sources of pneumatic power: • • •

External ground source APU load compressor Engine bleed air.

The APU regulates bleed air pressure from the APU load compressor. The APU is a primary source of bleed air on the ground. It eliminates the need for ground support equipment. The APU is a backup source of bleed air in flight.

16-2

Engine bleed air comes from the 5th or 9th stage of the high pressure compressor. The change from 5th to 9th is automatic. The pressure regulating and shutoff valves (PRSOVs) regulate engine bleed air pressure.

(e.g. engine starting operations). The isolation valve operation can be automatic or manual.

The precooler system cools the engine bleed air. The precooler is an air-to-air heat exchanger. It cools engine bleed air with engine fan air as the heat sink. The precooler control valve controls the flow of fan air.

The pneumatic system control is from the P5 panel. Improvements of controls and indications decrease crew work load.

The isolation valve isolates the pneumatic manifold into a left and right side when closed. This separates the pneumatic system into two systems. A single duct failure can be isolated. It will not effect the entire system. When open, the valve gives continuity to both sides of the pneumatic manifold. This allows a single source to power systems on one or the other side of the manifold

Pressure transmitters and a gage on the P5 panel show right and left manifold pressures.

Automatic overtemperature and overpressure protection systems protect the airplane from system malfunctions. Overheat sensing elements near the pneumatic ducts monitor the system for duct leaks.

November 2000

Environmental Systems Electronic Control Control Signals P5 Module

Overheat Switches

Temp Sensors

Check Valve (Typical)

Air Cond Relays

Temp Regulator

Mix M Valve

S

Flow Control and Shutoff Valve

M Pack Discharge

Ram Air Exhaust

Mix Muff 35F Control System Water Separator Air Cycle Machine Ram Air System S Solenoid M Motor

Ram Air Duct Ram Air Inlet

Water Spray Nozzle

Heat Exchanger (Typical)

Air Conditioning 737-600/700 Air Conditioning 737-600/700 The air conditioning system uses two independent air-cycle cooling packs, a cabin temperature control system, an air distribution system, and a recirculation system. The system can maintain safe cabin conditions with any one subsystem inoperative. The air conditioning packs are under the wing center section. The air conditioning packs discharge into the mix manifold of the distribution system. Air conditioning pack discharge is used for these purposes: • • • •

Supply fresh air to the cabin at a comfortable temperature Pressurize the airplane Cool the electronic equipment Heat the cargo compartments.

November 2000

The pneumatic manifold supplies compressed air to the air conditioning packs. The flow control and shutoff valves control the air flow through the packs. Heat exchangers and expansion through an air-cycle machine (ACM) cools pack air. The ACM is a refrigeration turbine and has air bearings. No scheduled maintenance is necessary. On the 737-600/700, the mix valve controls pack output temperature. This valve mixes cooled and uncooled pack air to produce discharge air at the proper temperature. The 35F control system prevents freezing temperatures downstream of the ACM. This protects the system from ice damage.

stream. The water sprays into the inlet of the ram air system. The ram air system supplies a cooling flow of ambient air through the heat exchangers. This airstream is the heat sink for the pack system. The ram air system inlet panels move to keep drag to a minimum. During ground operations, the ACM ram air fan pulls air through the system. Improvements of controls and indications on the P5 panel decrease crew work load. There is manual and automatic control of pack flow rates. Temperature control is also automatic or manual. An automatic overheat protection system is active in all modes of operation. The temperature regulator and 35F control system have BITE.

A water separator removes condensation from the cooled air

16-3

Trim Air Valve (Typical) Control Signals

M Trim Air Press Reg Valve

Flight Compartment Zone M

From Other Pack S

Flow Control and Shutoff Valve

Forward Pass Zone M Stby TCV M Water Extractor Overheat Switches

S

Aft Pass Zone

Air Cond. Relays

ACM

TCV

M

Check Valve (Typical) Ram Air Exhaust

Pack Discharge

Digital Pack/Zone Controllers

Condenser Temperature Sensors

Ram Air Fan

Reheater

P5 Module

Ram Air Duct Ram Air Inlet

Ram Air Actuator

Heat Exchanger (Typical)

Water Spray Nozzle

Air Conditioning 737-800/900 Air Conditioning 737-800/900 The air conditioning system uses two independent air-cycle cooling packs, a digital 3-zone cabin temperature control system, an air distribution system, and a recirculation system. The system can maintain safe cabin conditions with any one subsystem inoperative. The location, purpose, function, and integration of the 737-800/900 system is similar to the 737-600/700 system. The flow control and shutoff valves control air flow through the packs. Heat exchangers and expansion through an air cycle machine (ACM) cools pack air. The ACM has air bearings. Scheduled maintenance is not necessary.

16-4

System temperature control is automatic by the pack temperature control valves and trim air valves. Two digital pack/zone controllers operate these components: • • •

Temperature control valves Trim air valves The ram air actuators.

On the 737-800/900, the normal pack output temperature control is by the temperature control valve (TCV). The standby temperature control valve can control pack output temperature if the normal system fails. The temperature control valves control the amount of pack air that does not flow through the cooling components of the pack. This produces the proper discharge temperature. The cabin zone that requires the coolest air sets the pack output temperature. Hot trim air is then added to the ducts for the other two zones.

The pack has a high pressure water extractor system. This system has a reheater/condenser/extractor assembly. This removes the water before it enters the ACM turbine. The ram air system supplies a cooling flow of ambient air through the heat exchangers. The ram air system inlet panels move to keep drag to a minimum. Improvement in controls and indications on the P5 panel reduce crew work load. There is manual and automatic control of pack flow rates. Temperature control and overheat protection is automatic. The digital pack/zone controllers have LRU BITE.

November 2000

Environmental Systems

Passenger Compartment Overhead Distribution Duct

Sidewall Risers (Three on 737-800/900)

Air Conditioning Controls

Diffuser Outlet (Typical)

Flight Deck Distribution System Air Conditioning Packs Electronic Equipment Racks

Mix Manifold, Recirculation Fan and Filter (Two Fans and Filters on 737-800/900)

737-600/-700 Shown 737-800/900 Similar

Conditioned Air Distribution Conditioned Air Distribution

•

GENERAL

FLIGHT COMPARTMENT

The conditioned air distribution system combines the air conditioning pack outputs with recirculated air. It then distributes the air to the flight compartment and the passenger compartment.

The flight compartment receives conditioned air from the left pack discharge.

The mixing manifold and recirculation components are aft of the forward cargo compartment. The737-600/700 distribution system has these two independent temperature control zones: • •

Flight compartment Passenger compartment.

The 737-800/900 distribution system has these three independent temperature control zones: • •

Flight compartment Forward passenger compartment

November 2000

Aft passenger compartment.

If the left pack is off, the flight compartment receives air from the right pack and the mix manifold. Outlets and controls in the flight compartment supply conditioned air for these functions: • • • • • •

Windshield defogging Foot warming Seat warming Shoulder warming Control panel gaspers Ceiling panel gaspers and anemostats.

PASSENGER COMPARTMENT Conditioned air from the mix manifold moves in sidewall ducts to an overhead distribution duct above the center isle. The air comes out through these devices: • • • • •

Overhead duct nozzles Window diffuser outlets Passenger gasper outlets Galley ceiling gasper outlets Lavatory gasper outlets.

Passenger compartment air then moves through air return grills. This air then goes through a filtered recirculation system or goes overboard through the outflow valve. Galley air goes overboard through fuselage vents.

Flight compartment air then moves through vents into the electronic equipment compartment. 16-5

P5 Panel

Supply System Duct (Typical)

P6 Panel Supply Fans

Flow Detectors

Exhaust System Duct (Typical)

Supply Air Filter

Exhaust Fans E4 E5 Rack

Overboard Exhaust Valve

E3 E2 Pilot Control Stand

E1 Rack

Transverse Rack (E2,E3,E4)

Pilot Primary Display (Typical)

Equipment Cooling System Equipment Cooling GENERAL Electronic equipment is air cooled. These systems supply air to the equipment cooling system: • •

The supply system The exhaust system.

Cooling for the most critical electronic equipment is from both cooling systems. This causes a double (push-pull) cooling system. Cooling for less critical equipment is by one system. Electronic equipment that does not require active cooling is not included in the cooling system. Each system has two parallel fans (normal and alternate). Flow detectors monitor the quality of cooling air flow and give an indication of a failure. If the normal fan fails, the flight crew can select the alternate fan. 16-6

The fans, air filter, and the overboard exhaust valve are in the electronic equipment compartment. The flow detectors are in ducts in the forward equipment compartment. Connecting ducts, equipment rack channels, headers, and plenums complete the cooling circuits. Controls and indications are on the P5 forward overhead panel. SUPPLY SYSTEM The supply system pulls cooling air through a cleaner and pushes it over these components: • • • •

Pilot primary displays (3) Pilot control stand E1 and E5 racks Transverse rack.

EXHAUST SYSTEM The exhaust system pulls cooling air over these items: • • • • • •

The exhaust system air then discharges through an overboard exhaust valve, overboard. In flight, differential pressure causes this valve to close. These things happen when the valve closes: •

• The supply system pushes air over the equipment and into the exhaust system or into the electronic equipment compartment.

Pilot primary displays (3) Pilot control stand P6 panel P5 panel E1 and E5 racks Transverse rack.

Diverts warm equipment exhaust air around the forward cargo compartment for cargo heating Increases airplane pressurization control.

November 2000

Environmental Systems A

B

A

B

Forward Cargo Compartment

Aft Cargo Compartment

Diffuser Outlets (Typical)

Air Return Grille (Typical)

Overboard Exhaust Valve

Outflow Valve A-A

B-B

Cargo Compartment Heating Cargo Compartment Heating GENERAL The cargo compartments are not ventilated. There are sealed, fire resistant liners that prevent oxygen from sustaining a fire in a cargo compartment. The volume of air in the cargo compartments is sufficient to sustain the life of animals with these conditions met: • • •

The biomass is not too great The flight duration is not too long The cargo volume does not displace too much air space.

Conditioned air circulated around the cargo liners warms the cargo compartments. This keeps the compartments warm enough to sustain life.

November 2000

The cargo heat system is passive. It uses the differential pressures and heat energies of the air conditioning and pressurization systems. Cargo heating is automatic and controls, indications, or servicing are not necessary. FORWARD CARGO COMPARTMENT The forward cargo compartment is heated only when the airplane is in the air. AFT CARGO COMPARTMENT The aft cargo compartment is heated when the airplane is in the air and on the ground.

16-7

737-600/700 CONT CABIN AIR MIX VALVE

AIR TEMP

PASS CABIN

PASS CABIN

SUPPLY DUCT

AIR MIX VALVE

120 TEMP.160 DUCT OVERHEAT

DUCT OVERHEAT

80 200 40 F

AUTO

AUTO DUAL BLEED

RAM DOOR FULL OPEN

RAM DOOR FULL OPEN

COOL

RECIRC FAN OFF

P5 Forward Overhead Panel

COOL

WARM

WARM WARM OFF MANUAL

WARM OFF MANUAL

COOL

COOL

AUTO 40

737-800/900

60

OVHT 20

EQUIP COOLING

0

L PACK

EXHAUST

SUPPLY

NORM

OFF AUTO HIGH

ALTN

AIR TEMP

80 PSI 100

TEST R PACK

ISOLATION VALVE

S U P P L Y

60 40 TEMP 80 20 100

OFF AUTO HIGH AUTO

ZONE CAB FWD

AFT

D U C T

AFT

F W D CONT CAB

C

R L

TRIM AIR

P A C K

OFF OFF

OFF

a

WING ANTI ICE

a

Equipment Cooling Panel

PACK TRIP OFF

OPEN

PACK TRIP OFF

WING-BODY OVERHEAT

TRIP

WING-BODY OVERHEAT

BLEED TRIP OFF

WING ANTI ICE

ON ZONE TEMP a

BLEED TRIP OFF

1

ON

ZONE TEMP a

FWD CAB

AFT CAB

AUTO

AUTO

AUTO

RESET OFF

ZONE TEMP a

CONT CAB

OFF APU BLEED

ON

C

2

W OFF

Air Conditioning Panel

C

W

W

C OFF

OFF

Cabin Temperature Panel

Pneumatic and Air Conditioning Control Panels Pneumatic and Air Conditioning Control Panels

System indication lights show these conditions:

GENERAL

• •

The control panels are on the P5 forward overhead panel. These are control panel features: • • •

Lighted gages Light plates Positive position toggle switches and selector knobs System condition and caution lights.

• •

Bleed trip off Wing-body overheats (duct leaks) Dual bleed Loss of equipment cooling.

Gages show the mix valve position. A temperature gage and source selector show the system temperatures. System lights show these conditions:

Push-button switches control:

• • •

Pack trip off Duct overheats Ram door position.

• •

Resets of trip off conditions Wing-body overheat tests.

AIR CONDITIONING CONTROLS 737-800/900

AIR CONDITIONING CONTROLS 737-600/700

PNEUMATIC CONTROLS

Switches control these functions:

The higher capacity three-zone temperature control system of the 737-800/900 uses these additional controls:

Toggle switches control these functions:

• • •

• • •

•

• •

Bleed air sources Pneumatic manifold isolation.

A dual needle pressure gage shows right and left duct pressures.

16-8

Pack flow scheduling The right recirculation fan The equipment cooling fans.

Temperature selectors give automatic or manual pack temperature control for the two zones.

•

Left recirculation fan switch Three temperature selectors Three zone temp lights for overheat and fault indication Trim air system control switch.

November 2000

Environmental Systems P5 Forward Overhead Panel

Blowout Panels (Typical)

Pressure Equalization Valves (Typical) Digital Cabin Pressure Controllers (2)

Negative Relief Valve Outflow Valve

Positive Pressure Relief Valves (2)

Manual Signal OFF SCHED AUTO FAIL a DESCENTa

ALTN

g

MANUAL

g

MANUAL

AUTO

V A L V E

FLT ALT C L O S E

O P E N

Auto Signal

ALTN AUTO MAN LAND ALT

Outflow Valve Digital Cabin Pressure Controller (2)

Pressurization Control Panel Pressurized Area

Cabin Pressure Control Cabin Pressure Control NORMAL OPERATION The pressurization system controls the rate of air released from the cabin. The position of the outflow valve controls this rate. The cabin pressurization system maintains a safe, comfortable cabin pressure altitude at all times. Under normal operations, cabin pressure altitude is never more than 8,000 feet. The pilots can control airplane pressurization in these modes: • • •

Automatic mode Alternate mode Manual mode.

Controls for pressurization and indication are on the P5 forward overhead panel.

November 2000

Two digital controllers are in the EE compartment. The controllers have LRU BITE. They use inputs from these to control cabin pressure: • • • •

P5 panel settings Stall management computers Air data computers Aft outflow valve position transducer.

In the automatic modes (auto and alternate), the controllers automatically schedule cabin pressurization for all phases of flight. If both controllers fail, the pilot can control the valve manually. The outflow valve is in the aft, lower right area of the airplane. PRESSURE EQUALIZATION

FAIL-SAFE DEVICES If the systems fail, these valves protect the airplane structure from excessive pressure differentials: • • •

Positive pressure relief valves (2) Negative pressure relief valve Cargo compartment blowout panels.

ALTITUDE WARNING A cabin altitude warning system tells the crew when the cabin pressure altitude goes to 10,000 feet. This system activates by a switch on the ceiling of the lower nose compartment. It operates an aural warning horn on the control stand. The horn cutout button is on the P5 forward overhead panel.

Independent mechanical pressure equalization valves are in the bulkheads of the cargo compartments to allow for pressure changes in the cargo compartments. 16-9

AUTO FAIL

a

OFF SCHED DESCENT a

ALTN

MANUAL g

AUTO

g

MANUAL

0 1

10 0

9

CABIN ALT

40

30

C O L P O E S N E ALTN AUTO MAN

3

8

10

X 1000 FEET

25

20

15

4

7 5

6

LAND ALT PRESS DIFF LIMIT: TAKEOFF & LDG .125 PSI

1 .5

V A L V E

FLT ALT

5

35

P5 Forward Overhead Panel

ALT HORN CUTOUT

2

50

2 3

UP

0

4 DN

3

.5 1

CAB ALT

2

FLT ALT

LAND ALT