Falcon 10/100

SimuFlite

August 2010

Falcon 10/100Initial Pilot Training Manual

Copyright © 2010, CAE SimuFlite, Inc.All Rights Reserved.

Printed in the United States of America.

Excerpted materials used in this publication have been reproduced with permission of

the Dassult Falcon Jet Corp.

NOTICE: This Falcon 10/100 Initial Pilot Training Manual is to be used for aircraft familiarization and training purposes only. It is not to be used as, nor considered a substitute for, the manufacturer’s Pilot or Maintenance Manual.

Falcon 10/100 For SimuFlite training only 1-1April 2000

Welcome toSimuFlite

Welcome to SimuFlite!

Our goal is a basic one: to enhance your safety, proficiencyand professionalism within the aviation community. All of us at SimuFlite know that the success of our companydepends upon our commitment to your needs. We strivefor excellence by focusing on our service to you.

We urge you to participate actively in all training activities.Through your involvement, interaction, and practice, thefull value of your training will be transferred to theoperational environment. As you apply the techniquespresented through SimuFlite training, they will become“second nature” to you.

Thank you for choosing SimuFlite. We recognize that youhave a choice of training sources. We trust you will findus committed to providing responsive, service-oriented training of the highest quality.

Our best wishes are with you for a most successful andrewarding training experience.

The Staff of SimuFlite

1-2 For SimuFlite training only Falcon 10/100April 2000


Introduction

This manual is a stand-alone document appropriate for variouslevels of training. Its purpose is to serve as an informationalresource and study aid.

The Quick Referencesection provides limitation and otherdata for quick review.

The Operating Proceduressection contains chapters that providea pictorial preflight inspection of the aircraft, normal proceduresin an expanded format, standard operating procedures, maneuvers,and other information for day-to-day operations.

The Flight Planning chapter covers weight and balance andperformance; a sample problem is included.

The Systemssection is subdivided by aircraft system. Eachsystem chapter contains a discussion of components, preflightand servicing procedures, and abnormal and emergencyprocedures. At the beginning of the Systems section, a list ofsystems is cross referenced to ATA codes to facilitate furtherself study, if desired, with the manufacturer’s manuals.

The graphics shown at the right direct your attention to a specificlocation in the cockpit. A shaded area locates various instrumentsor switches shown in the adjacent photographs.

The graphic at right (top) represents the overhead panel.

The graphic at right (middle) represents the cockpit forwardpanels.

The graphic at right (bottom) represents the center pedestal andside consoles.

Using thisManual

Falcon 10/100 For SimuFlite training only 2-1March 2001

QuickReference

Chapter 2

This chapter contains the aircraft’s operating limits and require-ments as well as system by system charts summarizing powersources, distribution, controls, and monitors. All limitations areprinted in gray.

It also contains sections on supplement-directed limitations andoperations.

This chapter is intended to serve as a convenient reference.


Quick Reference

Table ofContents

General Limitations . . . . . . . . . . . . . . . . . . 2-5

Authorized Operations . . . . . . . . . . . . . . . . . 2-5

Baggage and Cargo . . . . . . . . . . . . . . . . . . 2-5

Certification . . . . . . . . . . . . . . . . . . . . . . . 2-7

Maneuvers . . . . . . . . . . . . . . . . . . . . . . . 2-7

Maximum Number of Passenger Seats . . . . . . . . 2-7

Minimum Flight Crew . . . . . . . . . . . . . . . . . . 2-7

Noise Levels . . . . . . . . . . . . . . . . . . . . . . 2-8

Operational Limits . . . . . . . . . . . . . . . . . . 2-9

Weight Limits . . . . . . . . . . . . . . . . . . . . . . 2-9

Speed Limits . . . . . . . . . . . . . . . . . . . . . 2-10

Takeoff and Landing Operational Limits . . . . . . . 2-11

Enroute Operational Limits . . . . . . . . . . . . . . 2-17

Systems Limitations . . . . . . . . . . . . . . . . . 2-19

Avionics and Communications . . . . . . . . . . . . 2-19

Electrical (and Lighting) . . . . . . . . . . . . . . . . 2-20

Flight Controls . . . . . . . . . . . . . . . . . . . . 2-20

Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . 2-21

Hydraulics . . . . . . . . . . . . . . . . . . . . . . . 2-23

Ice and Rain Protection . . . . . . . . . . . . . . . . 2-23

Landing Gear . . . . . . . . . . . . . . . . . . . . . 2-24

Oxygen . . . . . . . . . . . . . . . . . . . . . . . . 2-24

Pneumatic and Pressurization . . . . . . . . . . . . 2-24

Powerplant . . . . . . . . . . . . . . . . . . . . . . 2-24

Thrust Reversers . . . . . . . . . . . . . . . . . . . 2-26


Systems Data Summaries . . . . . . . . . . . . . . 2-29

Electrical Systems . . . . . . . . . . . . . . . . . . 2-29

Fire Protection Systems . . . . . . . . . . . . . . . 2-30

Flight Controls . . . . . . . . . . . . . . . . . . . . 2-31

Fuel System . . . . . . . . . . . . . . . . . . . . . . 2-33

Hydraulic Systems . . . . . . . . . . . . . . . . . . 2-34

Ice and Rain Protection . . . . . . . . . . . . . . . . 2-35

Landing Gear/Brakes/Steering Systems . . . . . . . 2-36

Lighting Systems . . . . . . . . . . . . . . . . . . . 2-38

Oxygen System . . . . . . . . . . . . . . . . . . . . 2-38

Pneumatic Systems . . . . . . . . . . . . . . . . . . 2-39

Powerplant . . . . . . . . . . . . . . . . . . . . . . 2-41

Thrust Reversers . . . . . . . . . . . . . . . . . . . 2-42


Quick Reference

Authorized Operations� The aircraft may be used for the following types of operation:

– carriage of passengers– carriage of cargo and mail– icing conditions– extended overwater– ditching– day and night VFR– IFR.

Baggage and Cargo� Do not use aft fuselage equipment bay as stowage for baggage

and spare parts unless an approved installation is provided in thiscompartment. Refer to Figure 2-1 and Table 2-A .

GeneralLimitations

2-1

2-6 For SimuFlite training only Falcon 10/100March 2001

NOTE: SB 0015; Rear Bay –Container for Ground EquipmentKit and Crew Baggage.

SB 0224; Installation of BaggageCompartment in the Rear Bay.

Aircraft Area Maximum Load

Baggage Compartment A 500 lbs

Baggage Compartment B1 220 lbs

Baggage Compartment C 500 lbs

Baggage Compartments A and C Simultaneously 750 lbs

Baggage Compartment DAircraft with SB 0224

200 lbsWithout Thrust Reversers

150 lbsWith Thrust Reversers

Baggage Compartment EAircraft with SB 0224

100 lbsWithout Thrust Reversers

50 lbsWith Thrust Reversers

Ground Equipment Kit and Crew Baggage – Aircraft with SB 0015 55 lbs

Cabin Forward Zone(Frames 12 to 15)

33.6 lb/ftLinear Load on Left or Right Side Floors and Center Aisle

260.0 lbsMaximum Total Weight for Left/Right Side Floors andCenter Aisle

Special Approved Installation2 1,078.0 lbs

Cabin Center Zone (Frames 15 to 20) 375.0 lbs

Cabin Aft Zone (Frames 20 to 22) 500.0 lbs

Table 2-A; Baggage, Cargo, and Cabin Load Limits1 Per change M7, resistance of coat compartment partition during emergency landing.2 Includes stowings and appropriate wing loads.


Quick Reference

Certification� FAR 25 through Amendment 25-2D� FAR 36 through Amendment 36-1� FAA Special Conditions 25-49-EU-14, dated April 16, 1973

Maneuvers� The following maneuvers are not authorized:

– acrobatics–spins.

Maximum Number of Passenger Seats� A maximum of nine seats is approved if the approved associated

interior accomodation is installed.� A maximum of eight seats is approved:

– seven in the cabin if the approved associated interior accom-modation is installed

– one jump seat modified per SB 0067 for authorization at take-off and landing.

� The jump seat occupant must be acknowledged by the pilot asable to use the jump seat and he must be instructed before theflight on its use and on the application of the safety instructions pro-vided on the “Jump Seat” sheet.

Minimum Flight Crew� Pilot and Copilot

NOTE: SB 0067; Jump Seat:Utilization at Takeoff andLanding.


Noise Levels� In compliance with FAR 36, the measured noise levels are as

shown in Table 2-B . No determination has been made by the FAAthat the noise levels of this aircraft are or should be acceptable orunacceptable for operation at, into, or out of any airport.

NOTE: SB 0052; Upgrading ofWeight Structural LimitationsAllowing Takeoff with 18,740 lb.

SB 0238; Upgrading of WeightStructural Limitations AllowingTakeoff with 19,300 lb.

Table 2-B; Effective Percevied Noise Levels1 Takeoff configuration for this noise level is slats + flaps 15° at 18,740 lbs.2 Approach configuration for this noise level is slats + flaps 52° at 17,640 lbs.3 Takeoff configuration for this noise level is slats + 15° flaps at 19,300 lbs.

MeasuringPoint Aircraft Not Modified

Aircraft with SB 0052Aircraft WithSB 0238

Flyover, WithoutNoise Reduction

83.41 84.23

Flyover, WithNoise Reduction

80.61 81.63

Approach 95.42 95.4

Noise Level (EPNdB)

Sideline 86.4 86.2


Quick Reference

Weight LimitsAircraft Not Modified:

Maximum Ramp Weight . . . . . . . . . . . . . 18,300 LBS

Maximum Takeoff Weight . . . . . . . . . . . . 18,300 LBS

Maximum Landing Weight . . . . . . . . . . . . 17,200 LBS

Maximum Zero Fuel Weight . . . . . . . . . . . 12,460 LBS

Minimum Flight Weight . . . . . . . . . . . . . . 9,920 LBS

Aircraft with SB 0052:Maximum Ramp Weight . . . . . . . . . . . . . 18,740 LBS





Aircraft with SB 0238:Maximum Ramp Weight . . . . . . . . . . . . . 19,400 LBS





� The takeoff weight is limited by the most restrictive of the followingrequirements:– climb gradients– balanced field length or distances connected to takeoff– landing weight at destination– maximum brake energy– maximum tire speed.

� The landing weight is limited by the more restrictive of thefollowing requirements:– one-engine go-around climb gradient– landing distance.

OperationalLimits




Speed LimitsVA, Maneuvering . . . . . . . . . . . . . . . . . . . . 220 KTSVFE, High Lift Devices Extended:

Slats Extended . . . . . . . . . . . . . . . . . . . . 200 KTS

Slats + Flaps 15° . . . . . . . . . . . . . . . . . . . 190 KTS

Slats + Flaps 30° . . . . . . . . . . . . . . . . . . . 165 KTS

Slats + Flaps 52° . . . . . . . . . . . . . . . . . . . 165 KTS

VLE, Gear Extended . . . . . . . . . . . . . . . . . . . 220 KTS

VLO, Gear Operating . . . . . . . . . . . . . . . . . . . 190 KTS

VMCA, Minimum Control – Air . . . . . . . . . . . . . . 97 KTSVMCA was established with a maximum 5° bank angle.

VMCG, Minimum Control – Ground . . . . . . . . . . 100 KTSVMO/MMO, Maximum Operating:

Sea Level to 10,000 Ft . . . . . . . . . 350 KTS TO 370 KTS

There is a straight line variation between altitudes.

10,000 Ft to 25,000 Ft . . . . . . . . . . . . . . . . 370 KTS

Above 25,000 Ft . . . . . . . . . . . . . . . . . . . . 0.87 M

Maximum Copilot Pitot/Static Pressure Abnormal . . . . . . . . . . . . . . . 260 KTS/0.76 M

Maximum Cracked Windshield . . . . . . . . . . . . . 230 KTS

Maximum Direct Vision Window Open . . . . . . . . . 200 KTS

Maximum Flaps Extend:15° . . . . . . . . . . . . . . . . . . . . . . . . . . 190 KTS

30° . . . . . . . . . . . . . . . . . . . . . . . . . . 165 KTS

52° . . . . . . . . . . . . . . . . . . . . . . . . . . 165 KTS

Maximum Hydraulic System 1 or 2 Failure . . . 260 KTS/0.76 M

Maximum Slats Extend . . . . . . . . . . . . . . . . . 200 KTS

Maximum Tire Groundspeed:Tires Approved for 180 MPH . . . . . . . . . . . . 157 KTS

Tires Approved for 190 MPH . . . . . . . . . . . . 165 KTS

Maximum Turbulent Air Penetration . . . . . . 250 KTS/0.75 M

Maximum Windshield Wiper Operation . . . . . . . . . 190 KTS

CAUTION: Full application ofrudder and aileron controls aswell as maneuvers that involveangles-of-attack near the stallmust be confined to speedsbelow VA.

CAUTION: Do not exceed themaximum operating speed (VMO/MMO) in any regime of flight(climb, cruise, or descent) unlessa higher speed is authorized forflight tests or pilot training.


Quick Reference

Takeoff and Landing Operational LimitsAirbrakes� The airbrakes must not deliberately be kept extended below a

height of approximately 500 ft.

Altitudes

Airport Pressure Altitude . . . . . . . . . . -1,000 TO 10,000 FT

Ambient Temperature Limits� Observe ambient temperature limits shown in Figure 2-2 .

Autopilot (Collins AP 105/APS 80)� Must not be used during takeoff.

Center of Gravity Limits� Observe limits shown in Table 2-C and Figures 2-3 , 2-4, and 2-5,

following pages.

2-2

Ambient Temperature Limits




Table 2-C; Center of Gravity Limits

Weight (lbs)

Aircraft NotModified

Takeoff 15,72018,300

31.0

Forward Aft

CG Limits (% MAC)

Cruise13,23015,72018,300

14.016.025.0

31.0

16.025.0

Landing13,23015,72017,200

14.016.021.2

31.0

AircraftWithSB 0052

Takeoff 15,72018,740

31.0

Cruise13,23015,72018,740

14.016.025.0

31.0

16.025.0

Landing13,23015,72017,640

14.016.021.8

31.0

AircraftWithSB 0238 Takeoff 15,720

19,30031.0

Cruise13,23015,72019,300

14.016.026.6

31.0

16.026.6

Landing13,23015,72017,640

14.016.021.8

31.0

Taxi 19,400 31.027.0


Quick Reference

2-3

Center of Gravity LimitsBasic Aircraft Without SB

2-4

Center of Gravity LimitsBasic Aircraft With SB 0052


2-5

Center of Gravity LimitsAircraft With SB 0238


Quick Reference

Demonstrated Crosswind

Maximum . . . . . . . . . . . . . . . . . . . . . . . 90°/25 KTS

Elevation

Maximum Field Elevation for Takeoff(Pressure Altitude) . . . . . . . . . . . . . . . . 10,000 FT MSL

Fuel Computers� Must be operative for takeoff.

Jet B or JP-4 Takeoff Envelope� Observe fuel temperature limits in Figure 2-6 .

2-6

Fuel Temperature Limits (Jet B or JP-4)


Operation on Unpaved (Dry) Runways –Supplement No. 1� Aircraft authorized for a 19,300 lb maximum takeoff weight are

limited to 18,740 lbs on an unpaved runway.� Unpaved runways must be dry.� The nosewheel must be equipped with a “wide-chine” tire.� Anti-skid must be in good operable condition and on for takeoff

and landing on grass runways.� The runway bearing strengths are:

– load classification number (LCN) not below 10.– California bearing ration (CBR) not below 10 for an aircraft

weight of 11,023 lbs and 12 for an aircraft weight of 18,740 lbs(linear variation between the two weights).

Tailwind – Maximum

Velocity . . . . . . . . . . . . . . . . . . . . . . . . . . 10 KTS

Runway

Runway Slope . . . . . +2% (UPHILL) OR -2% (DOWNHILL)� The runway must be smooth and hard-surfaced, except when

operating in accordance with Supplement No. 1.

X-Feed and Rear Tank Intercom Valves� These valves must be closed during takeoff.

CAUTION: If the emergencybrake is used or the anti-skidsystem fails, the increase in thelanding distance depends on therunway condition. Expect anincrease of 2,000 ft on gravel orsoil runways.


Quick Reference

Enroute Operational LimitsAirspeed Limits� Observe limits depicted in Figure 2-7 for flight envelope altitudes.

Altitudes

Maximum Flaps Extended . . . . . . . . . . . . 20,000 FT MSL

Maximum Gear Extended . . . . . . . . . . . . 20,000 FT MSL

Maximum Operating . . . . . . . . . . . . . . . 45,000 FT MSL

Ambient Temperature Limits� Observe the ambient temperature limits shown in Figure 2-2 , page

2-11.

Airspeed Limits

2-7




Elevation

Maximum Elevation for Terrain Clearance . . . . . . 30,000 FT

Load Factors

Aircraft Without SB 0052 or 0238:Flaps Retracted . . . . . . . . . . . . . . . . . . . . +3 TO -1

Flaps Extended . . . . . . . . . . . . . . . . . +2 TO ZERO

Aircraft With SB 0052 or 0238:Flaps Retracted . . . . . . . . . . . . . . . . . . . +2.9 TO -1

Flaps Extended . . . . . . . . . . . . . . . . . +2 TO ZERO� These accelerations limit the angle of bank in turns and limit the

severity of pull-up maneuvers.


Quick Reference

Avionics and CommunicationsAutopilot – General� Do not use the autopilot during takeoff.� Hold the control wheel while commanding all autopilot disen-

gagements.� With single engine operation, the following apply.

– After disengaging the autopilot, manually trim the aircraft beforere-engaging the autopilot.

– In consequence of thrust or speed changes, the aircraft must bemanually trimmed with aileron and rudder trims.

– Autopilot operation is authorized for single engine approach.� When using the autopilot, one pilot must be in his seat with his belt

and shoulder straps fastened so that if any malfunction of theautopilot occurs, he can recover control of the aircraft immediately.

Collins AP 105

Maximum Speed . . . . . . . . . . . . . . . . . . . . VMO/MMO

Minimum Height for Operation Enroute . . . . . . . . 1,000 FT

Minimum Height for Operation:IMC Conditions . . . . . . . . . . . . . . . . . . . . 200 FT

VMC Conditions . . . . . . . . . . . . . . . . . . . . . 80 FT� The autopilot is authorized for Category I precision approaches.

Collins APS 80

Maximum Utilization Speed . . . . . . . . . . . . . . VMO/MMO

Minimum Height for Operation Enroute . . . . . 1,000 FT AGL

Minimum Height for Operation:No Visual Reference . . . . . . . . . . . . . . . . . . 200 FT

No Radar Altimeter/With Visual Reference . . . . . . 100 FT

With Radar Altimeter/With Visual Reference . . . . . . 50 FT

Copilot Pitot/Static Pressure – Panel Only

Airspeed . . . . . . . . . . . . . . . . . . . . . 260 KTS/0.76 M

GNS 1000� The GNS 1000 (if installed) must not be used as the primary

means of navigation during approach.

SystemsLimitations


Electrical (and Lighting)Maximum Battery Temperature for Start . . . . . . . . . . 120°F

Maximum Continuous Generator Output . . . . . . . . . . 300A

Maximum Generator Output . . . . . . . . 350A (5 MINUTES)

Maximum Voltage . . . . . . . . . . . . . . . . . . . . . . 31V

Recognition Light� Limit the use of recognition lights on the ground to 5 minutes,

followed by 15 minutes off to ensure sufficient cooling.

Flight ControlsAirbrakes� The pilot must keep a hand on the airbrake control handle until the

normal airbrakes extension or retraction is ascertained.� Do not hesitate to use the airbrakes if the flight envelope limits

(VMO/MMO) are accidentally exceeded.� During approach and below 500 ft AGL, the airbrakes must not be

deliberately kept extended.� Do not use airbrakes below 500 ft AGL.� In manual flight conditions, accompany airbrake extension with

a nose-up maneuver of the pitch trim control to maintain longi-tudinal attitude. In automatic flight conditions, a momentary changeof the longitudinal attitude in the nose-down direction appearsupon airbrake extension.

Flaps� In flight, limit each actuation of the flap control handle to the next

detent position.

Yaw Damper/Q Unit

Q UNIT Annunciator Illuminated . . . . . 260 KIAS/0.76 M

On Approach If Q UNIT Annunciator Illuminates:

RUD Q . . . . . . . . . . . . . 140 KTS MINIMUM UNTIL . . . . . . . .LANDING ASSURED, ADD 10 KTS TO VREF

AIL Q . . . . . . . . . . . . . . . . . ADD 10 KTS TO VREF

� If aileron Q unit is not operating, limit speed to 260 KIAS (0.76 M).

CAUTION: To maintain a suf-ficiently high engine setting inicing conditions, approach maybe done in landing configurationwith airbrakes extended to 500 ftabove the ground. At 500 ft, theairbrakes must be retracted andthe engine synchronizer, ifinstalled, must be switched off.


Quick Reference

FuelApproved Additives� See Table 2-D , following page, for approved additives.� Anti-icing additives are approved under the following circumstances:

– conform to U.S. MIL-I-27686 specifications or equivalent– homogeneous mixture– concentration not in excess of 0.15% by volume.

� Shell ASA-3 anti-static additive, or equivalent, is approved in anamount to bring the fuel up to 300 conductivity units (300 picomhosper meter) if concentration does not exceed 1 mg by liter or 1 ppm.

� SOHIO Biobor JF biocide additive, or equivalent, is approved foruse in the fuel at a concentration not to exceed 20 ppm of elementalBoron.

Approved Fuels� See Table 2-D , following page, for approved fuels.

Fuel Asymmetry� Do not deliberately fuel only one wing.

Fuel Pressure

Minimum Fuel Pressure Warning . . . . . . . . . . . . 4.5 PSIG

Jet B or JP-4 Takeoff Envelope� See Takeoff and Landing Operational Limits, this chapter.

LO FUEL Light� The LO FUEL annunciator illuminates when one feeder tank con-

tains no more than 300 lbs of fuel.� Do not attempt a go-around with feeder tank level below 100 lbs.� If the wing tanks are empty, do not exceed 14° attitude for longer

than 20 seconds.

Pressure Refueling

Maximum Pressure . . . . . . . . . . . . . . . . . . . . . 50 PSI

Unusable Fuel� The fuel remaining in the tanks when the fuel quantity indicators

read zero is not usable in flight.� The amber area of the fuel quantity indicators, 0 to 300 lbs, is a LO

FUEL caution range.


X-Feed and Rear Tank Intercom Valves� These valves must be closed during takeoff.� If the type of fuel used is Jet B or JP-4, the X-FEED valve must

be opened whenever the flight is made at an altitude for whichthe operation of the engine with a failed booster pump is notguaranteed.

Table 2-D; Approved Fuels* Check information with the fuel supplier.

** JP-4 and JP-5 fuels are approved, but not recommended; if these fuels are in thesystem, the fuel heater may cause a vapor lock following a rapid power reduction.JP-4, JP-5, and JP-8 fuels have anti-icing additive preblended by the refinery. Seethe SimuFlite Technical Manual Fuel System chapter for more information.

EMS 53111 (Kerosene)ASTM D 1655 – Jet ACAN 2-3.23 – Jet A

-40-40

**

––

*With

Fuel SpecificationFreezingPoint(°C) Anti-Ice

NATOCodeAnti-Static

Additives

EMS 53112 (Kerosene)ASTM D 1655 – Jet A1CAN 2-3.23 – Jet A1DERD 2494 – AVTURDERD 2453 – AVTUR/FSiiMIL-T-83133 – JP-8**AIR 3405CAIR 2405CCAN 3GP23

-47-47-50-50-50-50-50-50

**

WithoutWithWithWithoutWith*

––F35F34F34F35F34–

*With******

EMS 53113 (Wide Cut Type)ASTM D 1655 – Jet BCAN 2-3.22 – Jet BMIL-T-5624 – JP-4**AIR 2407BDERD 2486 – AVTAGDERD 2454 – AVTAG/FSiiCAN 2-3.33

-50-50-58-58-58-58-58

**WithWithWithoutWithWith

––F40F40–F40F40

*With***

WithoutWith

EMS 53116 (High FlashPoint Type)AIR 3404CAIR 3404CDERD 2497 – AVCATMIL-T-5624 – JP-5**DERD 2452 – AVCAT/FSiiCAN 3GP24CAN 3GP24

-46-46-46-46-46-46-46

WithoutWithWithoutWithWithWithoutWith

F43F44F43F44F44F43F44

**

WithoutWithoutWithout**


Quick Reference

HydraulicsAirspeed Limitation

Hydraulic System 1 or 2 Failure . . . . . . . . 260 KTS/0.76 M

Approved Hydraulic Fluids� The only hydraulic fluid approved for use must conform to the

MIL-H-5606B/C specification.

Ice and Rain ProtectionEngine Anti-Ice (Ground or Flight)� Engine anti-ice must be on when total air temperature is below

+5°C and icing conditions are anticipated.� Engine anti-ice must be off when total air temperature is above

+10°C.

N1 Settings� For adequate performance of wing and engine anti-icing systems,

the N1 speed of the operative engines must not be lower than theminimum values shown in Table 2-E .

Windshield Electrical Heating� The PILOT and COPILOT WINDSHIELD control switches must

not be selected to the MAX position except in flight and then onlyif the NORM position is not sufficient in icing conditions.

Wing Anti-Ice (Flight Only)� Must be on in flight when total temperature is below +5°C and

when icing conditions are anticipated.� Wing anti-ice must be off when total temperature is above +10°C.� Do not operate on the ground.

Table 2-E; Engine Settings for Anti-Ice

CAUTION: If ice has alreadystarted to accumulate, turn onthe ignition system (start selectorswitches to AIRSTART) prior toselecting the engine anti-icingsystems on; turn on the engineanti-ice systems one after theother with a time delay of no lessthan 30 seconds. Wing anti-icingcan then be switched on.

In Flight –MinimumRecommended

7678

7375

6264

6769

Approach 69 69 6469

One Engine 88 85 7479

Phase of Flightat -30°COAT

at -20°COAT

at -5°COAT

at -10°COAT

% N1


Landing GearBrake Pedals

Observe the Caution at left.

Non-Latching of Gear in Icing Conditions

Observe the Caution at left.

Nosewheel Tire� The nose wheel must be equipped with a chined tire.

Oxygen� Above FL410, each passenger must wear an oxygen mask secured

to the face and connected to the oxygen system.

Pneumatic and PressurizationMaximum Cabin Pressure Altitude . . . . . . . . . . . 8,000 FT

Maximum Differential Pressure . . . . . . . . . . . . . . 9.1 PSI

PowerplantApproved Oils� AiResearch Manufacturing Co. of Arizona EMS 53110, Type II.

Engine Synchronizer� Engine synchronization, if installed, must not be used during take-

off, go-around, and landing.

Fuel Computers� Must be operative for takeoff.

Oil Consumption

Maximum . . . . . . . . . . . . . . . . . . . . . . . . 0.05 GPH

Oil Pressure Range (Minimum to Maximum)

Takeoff Thrust . . . . . . . . . . . . . . . . . . . 38 TO 46 PSI

Maximum Continuous Thrust . . . . . . . . . . . 38 TO 46 PSI

Idle . . . . . . . . . . . . . . . . . . . . . . . . . 25 TO 46 PSI

Transient (Maximum) . . . . . . . . . 55 PSI FOR 3 MINUTES

CAUTION: It is strictly for-bidden to depress the brakepedals prior to touch-down.

CAUTION: If red lights of mainlanding gear remain on whenretracting during icing con-ditions, ice may hinder the up-latching of main landing gear.Attempt gear retraction threetimes to clear ice; if unsuccess-ful, speed must not exceed VLO

(190 kts) and fuel consumptioncan be increased by 33% at max-imum.

CAUTION: If a new type of fuelor a mixture of fuel is used,adjust the engine computeraccordingly to preserve both thestarting characteristics and theacceleration and decelerationperformance characteristics ofthe engine.

NOTE: The OIL 1 and OIL 2annunciators illuminate for oilpressures below 25 PSI or ifchips are detected.


Quick Reference

Oil Temperature

Maximum Sea Level to 30,000 Ft . . . . . . . . . . . . . 113°C

Maximum Above 30,000 Ft . . . . . . . . . . . . . . . . 132°C

Maximum Transient (All Altitudes) . . 140°C FOR 2 MINUTES

Minimum for Start . . . . . . . . . . . . . . . . . . . . . -54°C

Operating Limits� Observe the limits shown in Table 2-F .

Thrust Ratings (Uninstalled, Sea Level, 86°F)

Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . 3,230 LBS

Maximum Continuous . . . . . . . . . . . . . . . . . 2,966 LBS

Table 2-F; Powerplant Operating Limits

Starting — — —860

ConditionN2 %RPM

N1 %RPM

TimeLimit

ITT(°C)

Windmilling Airstartfrom Windmilling N 2

to 60% N2

— — 25 Seconds—

Takeoff 100 100 ——

— — 5 Minutes860

Maximum Continuous 100 100 30 Minutes832

Transient 103 103 1 Minute—

105 105 5 Seconds—

Ground Start/StarterAssist Airstart from10% N2 to Lightoff

— — 10 Seconds—

Ground Start fromLightoff to Idle — — 50 Seconds—


Reduced Thrust Takeoff – Supplement No. 4� Restrict the use of reduced thrust to those conditions where the

actual takeoff weight is lower than the maximum takeoff weightlimited by the maximum takeoff thrust performance.

� N1 speed reduction must not exceed 5% below the rated takeoffthrust N1 speed (such speed reduction corresponds to a thrustreduction of approximately 20%).

� The assumed temperature computed with this procedure must notexceed the upper limit normally observed and must not be lessthan the flat rating temperature of the takeoff thrust setting normallyobserved.

� Reduced thrust N1 values must be based on the assumed temper-atures (Figure 2-8 ).

� Use reduced thrust only on smooth, hard-surfaced runways, with-out snow, ice, slush, or water.

� Do not use reduced thrust if:– the anti-skid system is inoperative– the engine or engine and wing anti-icing system is on during

takeoff– the aircraft is operated under an MEL or CDL condition related

to powerplant installation or having an effect on performance.� The use of clearways and stopways is not permitted.

Thrust Reversers� Thrust reversers are approved for ground use only.� The fuel computer must be on and operational for thrust reverser

use.� At power levels above idle, limit thrust reverser use to 30 seconds

maximum with a 10-minute cooling period.� Limit thrust reverser use to 60 seconds maximum in any

30-minute period.� Asymmetric thrust reverse is limited to dry, hard-surfaced run-

ways with operational steering.� Do not use the thrust reversers to back up the aircraft.� For a ground check of the throttle snatch mechanism, observe a

maximum N1 of 50%.� If the throttle snatch mechanism is inoperative, the reverser must

be bolted into the stowed position.� Reverser life is limited to 20,000 hours.


Quick Reference

Reduced Thrust TakeoffSupplement No. 4

2-8


Quick Reference

DataSummaries

Electrical SystemDC and AC Electrical Systems

Power Source Batteries (2)24/26V DC, 23 amp hour (20-cell)22/24V DC, 23 amp hour (19-cell)

Engine generators (2)28.5V DC, 350A, 10.5 kW

Inverters (3) – 750VA115V AC, 400Hz26V AC, 400 Hz

Ground power28.5V, 1,000A maximum

Distribution DC busesBatteryStartMainA,B,C,D

AC busesW bus (pilot) – 115V AC, 400 HzY bus (pilot) – 26V AC, 400 HzX bus (copilot) – 115V AC, 400 HzZ bus (copilot) – 26V AC, 400 Hz

Control DC system switchesPower selectorGenerator controlBattery controlEmergency powerAuxiliary bus (load shed)Battery reset

AC system switchesInverter controlStandby inverter

Monitor DC systemVoltmeter/ammeterBattery temperature

AnnunciatorsGENE. 1/2BATTHOT BAT

AC systemVoltmeterAC FAIL NO. 1/2 annunciators

Protection Circuit breakers

Current limiters

Relays

Generator control units


Fire Protection SystemEngine Fire Protection System

Power Source A busEngine 1 detection circuit (FIRE 1 PULL light)

B busEngine 2 detection circuit (FIRE 2 PULL light)

A busExtinguisher shot 1 to either engine

Battery busExtinguisher shot 2 to either engine

Teleforce cables on FIRE PULL handlesFuel shutoff valves

Distribution Left and right engines

Control TEST button

HORN SILENCE button

FIRE 1/2 PULL handles

FIRE XTING 1/2 switches

Monitor FIRE 1/2 PULL handles (red lights)

Protection Fire warning test system

Fuel firewall valves

Fire bottle discharge device


Quick Reference

Flight ControlsPrimary Flight Controls

Power Source Hydraulic System 1Flight controlsRudder Arthur Q

Hydraulic System 2Flight controlsYaw damper

Primary A busAileron and elevator Arthur Q unitsNormal stabilizer trimYaw damper

C busAileron trim

D busEmergency stabilizer trimRudder trim

Control Control yoke

Rudder pedals

Trim switches

Arthur Q low speed switch

Monitor AnnunciatorsQ UNITRUDAILSTAB TRIM (takeoff warning)GUST

Hydraulic pressure gages

Trim gages

Protection Arthur Q system overstress protection

Electric Arthur Q – automatic shutdown uponelevator/aileron Q failure

Torque bar (secondary AFU) centers flight controlsupon primary control linkage failure

Yaw damp limiting to 1.5° at high speed

Pitch trim limiting to 4° nose-up above 250 kts

Rudder actuator pressure limiting (1,850 PSIpermitted)


Secondary Flight Controls

Power Source Primary A bus – normal trim

Auxiliary D bus – emergency trim

Load-Shed D bus – rudder trim

Control Control yoke trim switches

AP DISC switch on control yoke

STAB EMG switch on pedestal

Monitor Trim gage

STAB TRIM annunciator

Audio warning (clacker)

Protection 4° cruise stop

Trim limit switches

Slats/Flaps System

Power Source Hydraulic System 1 (normal mode)Inboard/outboard slats

Hydraulic System 2 (emergency slats mode [standby])Outboard slats extension only

Hydraulic Systems 1/2 (stall protection mode [automatic]Outboard slats automatic extension

Primary B busEmergency slats

Auxiliary C busFlap position indicator system

Auxiliary D busFLAP CONTROL CBFlaps motor

Control SLATS/FLAPS handle

EMERG SLATS switch

Monitor Slat/flap position indicators

FLAP ASYM annunciator

Flap overspeed aural warning

Stall aural warning

Slat configuration agreement lights (red/green)

Airbrakes

Power Source Hydraulic System 1Primary A bus

Control Airbrake handle

Monitor AIRBRAKE configuration light (amber)AIRBRAKE annunciator (red)


Quick Reference

Fuel System

Power Source A busLeft FUEL QTY gageNo. 1 boost pumpNo. 1 jet pump

B busRight FUEL QTY gageNo. 2 boost pumpNo. 2 jet pump

C busIntercom valveCrossfeed valveFuel counterFuel flow gages

Distribution L/R wing tanks

L/R feeder tanks

NACA scoops (venting system)

Boost pumps

Jet pumps

Intercom valve

Crossfeed valve

Gravity feed

Suction feed

Single point refueling system

Gravity refueling

Pylon valves

Control TRANSFER switches

REAR TANK INTERCOM knob

BOOSTER PUMP switches

X-FEED knob

TOT/REAR TANK switch

Feeder tank high/low level float switches

Fueling valves/float valves

Fueling test valve

Monitor FUEL QTY gages

FUEL �°C gage

FUEL P1/P2 annunciators

LO FUEL annunciator

Protection Feeder tank high/low level switches

NACA vents

Transfer switches


Hydraulic Systems

Power Source Engine-driven pumps (2)

Main busStandby electric pump power

B busStandby electric pump controlHYDR QTY gagesHYDR PSI gages

Primary A busSystem annunciators

Distribution System 1Rudder, elevator, and aileron servo-actuatorsRudder Arthur Q unitNormal (inboard/outboard) slatsAirbrakesLanding gearNormal braking

System 2Rudder, elevator, and aileron servo-actuatorsYaw damperEmergency (outboard) slatsNosewheel steeringEmergency and parking brakesThrust reversers

Control ST-BY PUMP switch (standby pump)

Monitor HYDR QTY/PSI gages

AnnunciatorsHYD 1/2HYD.TK1/TK2ST-BY PUMPP.BRAKE

Hydraulic reservoir sight gages

Parking brake accumulator pressure switch


Hydraulic system pressure relief valves

Reservoir air pressure relief valves


Quick Reference

Ice and Rain ProtectionEngine/Wing Anti-Ice System

Power Source EngineHP bleed airA bus – left engine Pt2Tt2

B bus – right engine Pt2Tt2

WingHP/LP engine bleed airA bus – wing anti-ice valve

Distribution HP bleed airEngine nacelle anti-ice valvesNacelle anti-ice valvesEngine/nacelle pressure switches

HP/LP bleed airWing anti-ice valveWing leading edges

Control ENG 1/ENG 2 ANTI ICE switches

WINGS ANTI ICE switch

Monitor ENG 1/ENG 2 lights

WINGS light


Pitot/Static and Windshield Anti-Ice System

Power Source A busL/R windshield wipersPilot windshield heatPilot pitot/static heat

D busCopilot windshield heatCopilot pitot/static heat

Distribution Windshield temperature regulators

Control PILOT/COPILOT PITOT switches

A/A switch (if installed )

PILOT/COPILOT WINDSHIELD switches

Wiper switches

Monitor Window heat X.FR light

AMP meter

PITOT 1/PITOT annunciators


Windshield temperature regulators


Landing Gear/Brakes/Steering SystemsLanding Gear System

Power Source A bus – gear selector and handle light

B bus – position lights

Hydraulic System 1 – landing gear actuators

Control Landing gear selector

Emergency gear T-handle

Horn silence button

Landing gear handle

Override pushbutton (overrides ground down-lockdevice; its use is discouraged)

Monitor Hydraulic System 1 pressure/quantity gages

HYD.1 annunciator

Down-and-locked lights (3 green)

LDG GEAR MOVING lights (3 red)

Gear handle light (blinking red)

Warning horn

Protection Mechanically controlled emergency gear valve

Ground down-locking device

Nose gear centering after takeoff

Proximity (squat) switches

Proximity (Squat) Switches

Power Source Primary A bus

Distribution Nose gear switchGear retraction protection

Left main switchAirbrake warningBattery blowerGear retraction protectionFreon condenser blowerPressurization controllerStall system 1Thrust reversers

Right main switchConditioning valveHydraulic standby pumpSTAB TRIM annunciatorStall system 2

Control Automatic via extension/compression of three landinggear struts


Quick Reference

Brakes/Nosewheel Steering System

Power Source A bus – nosewheel steering

C bus – anti-skid

Hydraulic System 1 – normal brakes

Hydraulic System 2Emergency brakePark brakeAccumulatorNosewheel steering

Control Brake pedals

PARK BRAKE leverFirst detent – emergency brakesSecond detent – parking brake

Anti-skid switch

Anti-skid test button

Nosewheel steering handwheel

Monitor Hydraulic Systems 1/2 pressure/quantity gages

AnnunciatorsHYD 1/2 (amber)ANTI-SKID BRAKE L/R (amber)BRAKE (red)P. BRAKE (amber)

Protection Steering prevention via gear-extend microswitches


Lighting Systems

Power Source 28V DCA and B buses (non-shedding)C and D buses (load-shedding)

115V AC, 400 HzW bus (non-shedding)X bus (load-shedding)

Control Switches and rheostatsFlight deckPassenger individual lightsEntrance and cabinExterior lightsGear handle (landing light)

Monitor Warning and advisory lights


Oxygen System

Power Source Oxygen cylinder (1,850 PSI at 70°F)

Distribution Crew/passenger oxygen masks

Passenger oxygen system

Therapeutic oxygen mask (EROS)

Control Oxygen cylinder shutoff valve

Crew masks (N/100% PUSH)

Passenger oxygen control unitAUTO/OFF/MANUAL (Robertshaw)NORMAL/OVERRIDE/FIRSTAID/CLOSED (EROS)

Monitor Bottle pressure gage

Cockpit oxygen pressure gage

Bottle safety discharge disc

Protection Bottle safety valve (2,500 to 2,700 PSI)


Quick Reference

Pneumatic SystemsPneumatic System

Power Source LP/HP bleed air – either engine

Distribution Air conditioning via sidewall ducts and cockpitdistributors

Wingshield defog outlets

Cockpit footwarmer

Emergency pressurization

Outflow valve control pressure and vacuum

Control HP BLEED switch

COND’G VALVE switch

Emergency pressurization valve

Cabin temperature regulator

Footwarmer defog valves

Anti-ice switches

Engine power setting

Monitor ENG 1/2 anti-ice lights

WINGS anti-ice light

AnnunciatorsCOND O’HEATHYD. TK. 1/2CABIN (cabin pressure over 10,000 ft)DEICING (some aircraft)

Cabin temperature gage (if installed )

Protection HP check valves

LP check valves

Check valves – aft pressure bulkheads


Pressurization Systems

Power Source HP/LP bleed air from either or both engines

W bus (115V AC) (aircraft with IDC controller )

A bus (28.5V DC) (aircraft with ABG-SEMCAcontroller )

Distribtion Pressurized fuselage area

Control CABIN ALTITUDE controllerRATE knobPULL FOR BARO knob

AUTO/MAN/DUMP pressurization mode selector

UP/DOWN manual pressurization knob (cherry picker)

HP BLEED switch


Monitor CABIN annunciator (10,000 ft)

IndicatorsCabin rate of climbCabin altitudeCabin differential pressure

Protection Overpressuration controller (9.1 PSI)

Squat switch (landing depressurization)

Freon Air Conditioning System

Power Source Main bus via operating generator or GPU

Distribution Air conditioning ducts

Control SwitchesFreon systemAuxiliary busPressure

Temp regulation valve

Start buttons

30-second timer

Monitor Generator ammeter

Cabin temperature gage

Protection Suction pressure switches

Discharge pressure switches

Starter button shutoff

Generator/GPU power source relay


Quick Reference

Powerplant

Power Source A bus – left engineN1, N2, oil pressure, oil temperature, boost pump, ignition, fire detection, fuel computer,fire XTING switch position 1

B bus – right engineN1, N2, oil pressure, oil temperature, boost pump, ignition, fire detection, fuel computer

Auxiliary busesFuel flow indicators and totalizer

Battery busIgnition backupFire XTING switch position 2

GeneratorsDirect backup power for engine computers (GEN 1, GEN 2)

Control Thrust lever

PRESS TO START buttons

SwitchesIGNITER ONSPRSTART/STOPAIR/GROUND/MOTOR-STARTFuel computer ENGINE 1/ENGINE 2

Monitor Engine operationFuel computer

IndicatorsN1

N2

Oil temperatureOil pressureFuel flow

AnnunciatorsOIL 1/2ENG. C1/C2FUEL P1/P2GENE. 1/2IGNITER ON (2)

Protection Computer ONFlat rating - 3,230 lbs thrustN1, N2, 100%EGT limiting 860°Ultimate overspeed protectionN1 – 109%; N2 – 110%

MechanicalOverspeed 105%


Thrust Reversers

Power Source Hydraulic System 2 pressure

Aircraft without SB 0154A bus (28V DC)No. 1 battery

Aircraft with SB 0154A busMain bus

Control No. 1 battery switch (aircraft without SB 0154 )

HORN TEST button

EMERGENCY STOW switch

Main throttle lever

Thrust reverser throttle lever

Monitor Reverser warning horn

Reverser TRANS/REV lights

Hydraulic System 2 pressure gage

Battery voltmeter (Main bus voltage)

Protection Secondary latch

Left main squat switch

Throttle auto retard (Snatch)

Reverser maximum RPM limit stop


This section presents four individual chapters of flight operations:preflight inspection, expanded normal procedures, standardoperating procedures (SOP), and maneuvers. Although theyare addressed individually in this manual, their smoothintegration is critical to ensuring safe, efficient operations.

The Preflight Inspection chapter illustrates a step-by-stepexterior inspection of the aircraft. Preflight cockpit and cabinchecks are also discussed.

The Expanded Normal Procedureschapter presents check-lists for normal phases of flight. Each item, when appropriate,is expanded to include limitations, cautions, warnings, and lightindications.

The Standard Operating Procedureschapter details PilotFlying/Pilot Not Flying callouts and verbal or physical responses.

The Maneuvers chapter pictorially illustrates common andemergency profiles. Additionally, written descriptions areincluded for most phases of flight with one or both enginesoperating.

OperatingProcedures

Falcon 10/100 For SimuFlite training only 3A-1April 2000

An essential part of the preparations made before any flight is thepreflight inspection. During this inspection, the aircraft’s flightreadiness is verified. A thorough initial preflight is a later benefitin that subsequent inspections that day can be carried out in lesstime.

No detail should be overlooked during the first preflight of theday. Abnormal conditions (e.g., low tire pressure) must becorrected before flight. Even minor discrepancies should befixed before flight to ensure safety.

The preflight inspection begins inside the aircraft where theinitial cockpit setup and essential functions are verified. Theactual exterior inspection follows; it begins at the left side ofthe nose, proceeds clockwise around the aircraft, and ends atthe cabin door. Lastly, the pilot returns to the interior of the air-craft to check the passenger compartment, baggage area, andcockpit for flight readiness.

PreflightInspection

Chapter 3A

3A-2 For SimuFlite training only Falcon 10/100April 2000


Preflight Inspection

Table ofContents

Checklist Usage . . . . . . . . . . . . . . . . . . . 3A-5

Normal Procedures . . . . . . . . . . . . . . . . . . 3A-5

Abnormal and Emergency Procedures . . . . . . . . 3A-5

Cockpit Inspection and Setup . . . . . . . . . . . 3A-7

Pilot’s Circuit Breaker Panel and Side Console . . . 3A-9

Pilot’s Instrument Panel . . . . . . . . . . . . . . . 3A-9

Copilot’s Instrument Panel and Side Console . . . . 3A-10

Copilot’s Circuit Breaker Panel . . . . . . . . . . . 3A-11

Pedestal and Center Instrument Panel . . . . . . . 3A-11

Overhead Panel . . . . . . . . . . . . . . . . . . . 3A-12

Exterior Walkaround . . . . . . . . . . . . . . . . 3A-15

Exterior (General) . . . . . . . . . . . . . . . . . . 3A-15

Left Forward Fuselage . . . . . . . . . . . . . . . 3A-17

Nose Wheel and Wheel Well . . . . . . . . . . . . 3A-19

Right Forward Fuselage . . . . . . . . . . . . . . . 3A-21

Right Fuselage and Engine Inlet . . . . . . . . . . 3A-25

Right Wing . . . . . . . . . . . . . . . . . . . . . . 3A-27

Right Wing Trailing Edge . . . . . . . . . . . . . . 3A-29

Right Main Gear . . . . . . . . . . . . . . . . . . . 3A-31

Right Aft Fuselage . . . . . . . . . . . . . . . . . . 3A-33

Horizontal and Vertical Stabilizer . . . . . . . . . . 3A-37

Rear Compartment . . . . . . . . . . . . . . . . . 3A-39

Left Aft Fuselage . . . . . . . . . . . . . . . . . . 3A-41

Left Main Gear . . . . . . . . . . . . . . . . . . . 3A-45

Left Wing Trailing Edge . . . . . . . . . . . . . . . 3A-47

Left Wing and Engine Inlet . . . . . . . . . . . . . 3A-49

Cabin Inspection . . . . . . . . . . . . . . . . . . 3A-51



ChecklistUsage

Normal ProceduresTasks are executed in one of two ways:

� as a sequence that uses the layout of the cockpit controls andindicators as cues i.e., “flow pattern”

� as a sequence of tasks organized by event rather than panel loca-tion; e.g., After Takeoff, Gear – RETRACT, Yaw Damper –ENGAGE.

Placing items in a flow pattern or series provides organization andserves as a memory aid.

A challenge-response review of the checklist follows execution ofthe tasks. The PNF calls the item, and the appropriate pilot respondsby verifying its condition; e.g., Engine Anti-Ice (challenge) – ON(response).

Two elements are inherent in the execution of normal procedures:

� use of either the cockpit layout or event cues to prompt the cor-rect switch and/or control positions

� use of normal checklists as done lists.

Abnormal and Emergency ProceduresFlow patterns will only be used for immediate recall items; theremaining tasks shall be done using a challenge-response (i.e., “do”list) method.



The cockpit inspection and setup is the first step in the preflightinspection. Except for the battery checks, the procedures in thischecklist do not require the use of electrical power.

CockpitInspectionand Setup

CBPANEL

A CBPANEL

C

B

D

E

F


Cockpit Inspection and Setup



A Pilot’s Circuit Breaker Paneland Side Console

Circuit Breakers . . . . . . . . . . . . . . . . CHECKED/INVerify that all circuit breakers are set. Reset any open circuit breakersafter checking system status in the maintenance log.

Windshield Demist/Floor Heat . . . . . . . . . . . . . COLDNormally these controls are left in the cold position to reduce bleedair demands during engine start.

Oxygen Mask . . . . . . . . . . . . . . . . . CHECKED/SETRemove the mask from the box; perform an operational check of theharness inflating mechanism by squeezing the control plates. Pushthe red oxygen test button and listen for oxygen flow. Verify opera-tion of the mask microphone.

B Pilot’s Instrument Panel

Pitot/Static Selector . . . . . . . . . . . . . . . . . NORMALLeave the pitot/static selector in NORMAL. In EMERGENCY, thesystem uses the static port in the nosecone.

Audio Panel . . . . . . . . . . . . . . . . . . . . . . . . . SETFor normal use, verify that the VHF1 audio, VHF1 microphone, SPK(speaker), and ST (side tone) buttons are depressed. Select HANDor BOOM on the microphone selector (if installed). Set loudspeakervolume control to desired level.

Compass Switch . . . . . . . . . . . . . . . GUARDED/MAGSet the compass switch to MAG so the directional gyro synchronizeswith magnetic heading.

Airspeed Bugs . . . . . . . . . . . . . . . . . . . . . . . SETSet the airspeed indicator bugs to the computed airspeeds on theTOLD card.

Course and Heading Knobs . . . . . . . . . . . PUSHED INOn aircraft with Collins APS 105 system, push the COURSE andHDG knobs on the course indicator in.

Altimeter(s) . . . . . . . . . . SET TO FIELD ELEVATIONSet the altimeter to field elevation. Also set the standby altimeter (ifinstalled). The altimeters should agree within 75 ft.

Altitude Alerter(s) . . . . . . . . . . . . . . . . . . . . . SETSet the altitude preset/alerter to the initial departure altitude.


Standby Horizon . . . . . . . . . . . . . . . . TESTED/OFFTest the horizon, then cage the unit to prevent damage when the air-craft is powered up.

Thrust Reverser EmergencyStow Switches . . . . . . . . . . . . . . . DOWN/GUARDED

Verify that both switches are normal (down) and the guards coverthe switches.

Fuel Quantity Selector . . . . . . . . . . . . . . . . . TOTALSet the fuel quantity indicator selector to TOTAL.

Fire Handles . . . . . . . . . . . . . . . . . . . . . . . . . INVerify that both fire handles are in. Pulling a handle closes the fuelshutoff valve. Exercise the handles periodically.

Fire Extinguisher Switches . . . . . . . . . . . 0 POSITIONVerify that the switches are in the 0 position (down) and safety wired.

C Copilot’s Instrument Paneland Side Console

Fuel Counter . . . . . . . . . . . . . . . . . . . . . ZEROEDZero out the fuel counter by rotating the knob on the lower edge ofthe indicator.

Gear Handle . . . . . . . . . . . . . . . . DOWN/LATCHEDVerify that the landing gear handle is down and the latch is in position.

Airspeed Bugs . . . . . . . . . . . . . . . . . . . . . . . SETSet the airspeed indicator bugs to the values on the TOLD card.

Pressure Controller . . . . . . . . . . . . . . . . . . . . . SETSet the pressure control knob to the center detent (-450 FPM) andthe barometric pressure to 29.92. Set the altitude selector to theplanned flight altitude.

Passenger Oxygen Controller . . . . . . . . . . . . . . AUTOSet the passenger oxygen controller to AUTO. In AUTO, the systemprovides automatic operation based on cabin altitude. Check for aminimum of 965 PSI on oxygen gage (minimum for dispatch – ferryonly). Pressure requirements vary depending on flight profile, pas-senger number, and flight duration.



Audio Panel . . . . . . . . . . . . . . . . . . . . . . . . . SETFor normal use, verify that the VHF1 audio, VHF1 microphone, SPK(speaker), and ST (side tone) buttons are depressed. Select HANDor BOOM on the microphone selector (if installed). Set volume con-trol to desired level.

Compass Switch . . . . . . . . . . . . . . . GUARDED/MAGSet the compass switch to MAG so the directional gyro synchronizeswith magnetic heading.

Pitot/Static Selector . . . . . . . . . . . . . . . . . NORMALSet the selector to NORMAL. In PANEL ONLY, pressure is suppliedto the panel instruments only and all other supplied equipment isinoperative.

Oxygen Mask . . . . . . . . . . . . . . . . . CHECKED/SETRemove the mask from the box; perform an operational check of theharness inflating mechanism by squeezing the control plates. Pushthe red oxygen test button and listen for oxygen flow. Periodicallyverify operation of the mask microphone.

D Copilot’s Circuit Breaker Panel

Circuit Breakers . . . . . . . . . . . . . . . . CHECKED/INVerify that all circuit breakers are in. Reset any open circuit breakersafter checking system status in the maintenance log.

E Pedestal and Center Instrument Panel

Emergency Gear Handle . . . . . . . . . . . . . . STOWEDVerify that the emergency gear handle is stowed.

Autopilot/Yaw Damper . . . . . . . . . DISENGAGED/OFFVerify that the autopilot and yaw damper are disengaged.

Emergency Slats . . . . . . . . . . . . . . . . . . GUARDEDCheck the emergency slat switch to verify that it is off and guarded.

Slat/Flap Handle . . . . . . . . . . . . . . . . . . . . CLEANCheck the slat/flap handle to verify that it is in the full forward posi-tion (CLEAN) and locked in the detent.

Power Levers . . . . . . . . . . . . . . . . . . . . . CUTOFFVerify that the power levers are in the full aft (cutoff) position. Astart cannot be initiated with the power levers out of cut-off.


Airbrake Handle . . . . . . . . . . . . . . . . . . . . . . . INCheck the airbrake handle to ensure that it is in the full forward (IN)position.

Trim Circuit Breaker . . . . . . . . . . . . . . . . . . . . . INVerify that the trim circuit breaker is set.

Park Brake Handle . . . . . . . . . . . . . . . . . . . . . SETThe handle is set to the middle detent to provide sufficient braking for75% engine power settings. The handle contains a spring-loadedrelease that must be squeezed prior to moving from one detent toanother.

Anti-Skid Switch . . . . . . . . . . . . . . . . . . . . . . . ONVerify that the anti-skid switch is on.

Standby Hydraulic Pump . . . . . . . . . . . . . . . . . OFFTurn the standby hydraulic pump off.

Freon/HP Bleed Switches . . . . . . . . . . . OFF/CLOSEDCheck that the Freon switch is OFF. Check that the HP bleed switchis in CLOSED to reduce engine bleed air requirements during enginestart.

Pressurization Selector . . . . . . . . . . GUARDED/AUTOEnsure that the pressurization selector switch is in AUTO and guarded.

Temperature Controller . . . . . . . . . . . . . . . . . . SETSet the controller to the 12 o’clock position. This provides automat-ic temperature control with an approximate temperature of 72°F.

Radios . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

F Overhead Panel

Wiper . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFFCheck that the copilot’s windshield wiper switch is off.

Exterior Lights . . . . . . . . . . . . . . . . . . . . . . . OFFVerify that the anti-collision, nav, strobe, landing, taxi, wing ice, andother optional (logo, recognition, etc.) exterior lights are off.

Interior Lights . . . . . . . . . . . . . . . . . . . . . . . OFFVerify that the exit and cabin lights are off.



Seat Belts/No Smoking Switches . . . . . . . . . . . . PUSHPush both switches so that the lights illuminate when power isconnected.

Inverters (3) . . . . . . . . . . . . . . . . . . . . . . . . OFFCheck that the Standby, No. 1, and No. 2 inverters are off.

Volt/Ammeter Selectors . . . . . . . . . . . . . BAT/STD BYSet the DC volt/ammeter selector to BAT to read battery voltage. Setthe inverter voltmeter selector to STD BY because it is the firstinverter checked.

DC Power Selector . . . . . . . . . . . . . . . . . NORMALThe three position (LOW TEMP START/NORMAL/EXT POWER)switch is set to NORMAL for battery powered engine starts.

Fuel Computer Switches . . . . . . . . . . . . . . . . . . ONThe fuel computer switches are normally left on unless damage froman unreliable power source is anticipated (GPU).

Rear Tank Intercom/Crossfeed . . . . . . . . . . . CLOSEDVerify that the rear tank intercom and crossfeed switches are in theclosed (vertical) position.

Fuel Transfer Switches . . . . . . . . . . . . . . GUARDEDVerify that the fuel transfer switches are in the guarded position.

Booster Pumps . . . . . . . . . . . . . . . . . . . . . . . OFFVerify that both booster pump switches are off.

Start Selector Switches . . . . . . . . . . . . . . GRD STARTCheck that start selector switches are in the GRD START (middle)position. A start cannot be initiated in any other position.

Wiper . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFFCheck that the pilot’s windshield wiper switch is off.

Pitot Heat/AOA . . . . . . . . . . . . . . . . . . . . . . OFFCheck that the pitot heat (pilot and copilot) and angle-of-attack (AOA)heat switches are off.

Side/Windshield Heat . . . . . . . . . . . . . . . . . . . OFFCheck that the side, pilot, and copilot windshield heat switches are off.

Anti-Ice Switches . . . . . . . . . . . . . . . . . . . . . . OFFCheck that the ENG 1, ENG 2, and WINGS anti-ice switches are off.Significant damage can occur if wing anti-ice is activated on theground.


Generators/Auxiliary Bus . . . . . . . . . . . . . . . . . . ONVerify that the generator No. 1, No. 2, and auxiliary bus switches areon. A start cannot be initiated with the appropriate generator switchin the off position.

Batteries . . . . . . . . . . . . . . . . . . . . CHECKED/ON� No. 1 battery ON, 23 volt minimum voltage checked. BATT

annunciator check illuminated.� No. 2 battery ON, BATT annunciator extinguished.� No. 1 battery OFF, 23 volt minimum voltage on battery No. 2

checked. BATT annunciator illuminated.� No. 1 battery ON, 23 volt minimum voltage with both batteries

ON, BATT annunciator extinguished.Do not attempt a normal battery start with less than 23 volts indicated.Leave the battery switches on to complete the following checks.

Fuel Quantity . . . . . . . . . . . . . CHECKED/SET REARCheck that total fuel quantity is adequate for the planned flight ornote the quantity to calculate the fuel order.

P Brake Annunciator . . . . . . . . . . . . EXTINGUISHEDVerify that the parking brake annunciator is extinguished. If not,motor the right engine to raise the accumulator pressure to extinguishthe light or chock the aircraft prior to starting the right engine.

Engine Computer Lights . . . . . . . . . . EXTINGUISHEDCheck that the two engine computer lights are extinguished.Illumination of a light indicates fuel computer malfunction.

Battery Switches . . . . . . . . . . . . . . . . . . . . . . OFFTurn off the batteries.

Clock . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETSet and wind the clock (if installed) on the pilot’s and copilot’s controlyokes.


3A-16


Preflight InspectionWalkaround PathExterior

WalkaroundUnfold the preflight inspection diagram on the following page for easeof reference. Note that each segment of the preflight inspection isidentified by letters A through M. Subsequent pages providesequenced checklists of each preflight inspection segment. Largelocator photos identify the general location of each inspection.Adjacent photos detail the checklist items. Photos read left to right.

Limitations and specifications are noted if relevant to the checklist.

Before starting the exterior inspection of the aircraft, obtain a flash-light, screwdriver, and bucket or other suitable container for disposalof fuel samples.

Exterior (General)Make a general check for security, condition, and cleanliness of theaircraft and components. Check particularly for damage, fluid leak-age, security of access panels, and removal of keys from locks.Remove all covers from the pitot tubes, static ports, probes, andengine inlets and exhausts.


Preflight InspectionWalkaround Path

B

C

D

JK

N

M

A

E

F

H

L G

I

1

5 6 7

2 3 4



A Left Forward Fuselage1. Cabin Door: Check the condition of the cabin door seal formisalignment and pliability. Check the seal for cuts and tears.

2. Static Ports: Check the condition of the ports for cleanlinessand freedom from obstructions.

3. Angle-of-Attack Vane: Check that the AOA vane moves freelyand is free from obstructions. Leave the vane in the horizontal posi-tion.

4. Pilot’s Windshield and Side Window: Check the generalcondition of the windshield and side window. Inspect for chips,cracks, and delamination.

5. Left Side Nosecone Latches: Check that the latches aresecure and flush with the fuselage. Check that the nosecone is cleanand free from damage.

6. Ram Air Temperature Sensor: Check that the sensor is clearand clean.

7. Left Pitot Tube: Remove the pitot cover and check the tubefor cleanliness and freedom from obstructions.


1

2 3 4

5 6



B Nose Wheel and Wheel Well1. Nose Gear Tire: Check for correct tire pressure. Inspect thetire for chine, irregular wear, delaminations, cuts or tears, and slipmark alignment. Inflate the tire with dry nitrogen only; never useoxygen or air. Normal tire pressure is 94 PSI.

2. Nose Strut: Examine the nose strut for normal extension. Strutextension height varies with ambient temperature. Normal strutextension at 68°F is approximately 2 1/2 inches. Consult the aircraftmaintenance manual strut extension/temperature chart.

3. Taxi Light: Check the taxi light and cannon plug for generalcondition and security. Check the lens for any evidence of crackingor chipping. Check the condition of the bulb.

4. Nosewheel Steering Locking Pin: Verify that the nose-wheel steering locking pin is installed properly.

5. Nosewheel Steering Electrical Connection: Verify thatthe electrical connection is secure.

6. Nosewheel Well Doors: Check the overall condition of thedoors and linkages.


2 3 4

6 7

1

5



C Right Forward Fuselage1. Total Air Temperature Probe: Check the condition of theprobe for cleanliness and freedom from obstruction.

2. Recognition Light: Check that the lens is not cracked or heatdiscolored. Check bulb condition.

3. Right Pitot Tube: Remove the pitot cover and check the tubefor cleanliness and freedom from obstructions.

4. Oxygen Filler Access Door: Verify that the oxygen valve isopen, then open the door on the right side of the nose. Check theoxygen pressure gage for 1,850 PSI or minimum computed for flight.Close the door and verify that it is secure.

5. Oxygen Discharge Disc: Check that the oxygen dischargedisc is present and intact.

6. Nosecone Latches: Verify that the latches are secure andflush with the fuselage.

7. Pressurization Static Port: Check that the static port coveris removed and that the port is clean and free from obstructions.


8 9

11

10

12



C Right Forward Fuselage (continued)8. Angle-of-Attack Vane: Verify that the AOA vane is un-

obstructed and free to move. Leave the vane in the horizontal position.

9. Windshield/Side Windows: Inspect the windshield andwindow for cleanliness. Check for cracking, chipping, and delamination.

10. Right Static Port: Check for condition, cleanliness, and free-dom from obstruction.

11. Angle-of-Attack Probe (optional): Inspect the probe forcleanliness and freedom from obstructions. Verify that it movesfreely.

12. Wing Leading Edge Light (optional): Check the light (wingice light) for lens condition and bulb appearance.


2 3

1



D Right Fuselage and Engine Inlet1. Bottom Fuselage, Antenna, and Anti-Collision Light:Check the fuselage underside for general condition and any sign offluid leaks. Check the VHF 1 antenna for security, condition, andfreedom from dents and cracks. Inspect the anti-collision light forsecurity, lens condition, and bulb condition.

2. Emergency Exit Light: Check the emergency exit light onthe bottom of the wing root for lens condition and bulb appearance.

3. Engine Inlet: Inspect the engine inlet for foreign objects andloose fasteners. Check the condition of the fan blades and inlet guidesfor nicks, cracks, and signs of damage. The Pt2/Tt2 probe should beclean and undamaged.


1

32 4

5 6



E Right Wing1. Emergency Exit: Ensure that the emergency exit is flush withthe fuselage. Check that the red plexiglass cover over the emergencyexit handle is in place.

2. Landing Light: Check the light for security. Check the lensfor cracks and signs of overheating. Check the condition of the bulb.

3. Leading Edge: Inspect the slats for general condition, securi-ty, as well as for dents or other damage. Check the wing fence forsecurity, condition, and damage.

4. Wing Fuel Vent: Check the fuel vent on the underside of thewing for cleanliness, freedom from obstructions, and leaks.

5. Filler Fuel Cap: Check that the filler cap is properly secured andthat the locking tab is facing aft with the tab flush with the wingsurface.

6. Wing Tip and Navigation/Strobe Lights: Inspect the wingtip for damage. Check the navigation and strobe lights for lensdamage, security, and any indication of a burned out bulb. If nightflight is anticipated, perform an operational check of the navigationand strobe lights.


1

32 4



F Right Wing Trailing Edge1. Static Dischargers: Check that the static dischargers on theaileron and wing tip are present, secure, and undamaged and thatthe brushes are equal in length and fullness (if installed). Two differentstatic dischargers are used on the Falcon 10/100. One consists of athree brush assembly; the other consists of a solid plastic unit withsharp points. The number of static dischargers used varies from air-craft to aircraft depending on modification status, operator prefer-ences, and static discharger type.

2. Aileron: Inspect the aileron for damage and security.

3. Flaps: Check the condition and security of the flaps. Verify thatthe position of the flaps corresponds with flap handle position. Checkthe condition of all cables, actuators, and lines forward of the flaps,if extended.

4. Airbrakes: Check the condition and position of the airbrakes.Verify that the position of the airbrakes corresponds with airbrakehandle position.


1

6

2 3 4

5



G Right Main Gear1. Wheel Embellisher: Check that no fasteners are missing fromthe embellisher.

2. Tires and Wheels: Check tires for proper inflation, abnormalwear, slip mark alignment, and other damage. Use only nitrogen toinflate the tires; never use dry air or oxygen. Inflation pressuredepends on aircraft weight and service bulletin status. Check thewheel for dents, metal pickup on wheel surface, and other damage.

3. Brakes: Inspect the brakes for wear pin extension, leaks, dam-age, and signs of overheating. If the wear pin is recessed in the brakereturn box, excessive wear is indicated; maintenance is requiredbefore flight.

4. Strut: Check the strut for leaks and extension. Normal strutextension is 6 inches at 68°F; extension varies with ambient temp-erature. Consult the maintenance manual extension/temperaturechart.

5. Wheel Well: Inspect the wheel well for general condition, secu-rity of hydraulic lines and cables. Check for indications of fluidleaks.

6. Fuselage Underside: Inspect the bottom of the fuselage forsigns of leaks or damage.


6

2 3 4

5

1



H Right Aft Fuselage1. Single Point Refueling Panel: Check for leaks. Verify thatthe panel is closed and secure.

2. Fire Extinguisher Discharge Disc: Check the red fire ex-tinguisher disc to ensure that it is not absent (indicating bottledischarge).

3. Engine Fuel Filter Bypass: Check that the indicator is flushto indicate normal filter operation. If the indicator is protruding, thefilter is bypassing; maintenance is required before flight.

4. Engine Cowl: Check that the cowling is properly latched andthat there are no leaks on the underside of the cowl.

5. Engine Oil Filter Bypass: Check that the red indicator isflush to indicate normal filter operation. If the indicator is protrud-ing, the filter is bypassing; maintenance is required before flight.

6. Oil Sight Gage: Push the spring-loaded access door to visual-ly check engine oil quantity through the sight gage. Total engine oiltank capacity is approximately 1.5 gallons. Oil quantity should bechecked within one hour after engine shutdown.


7 8

9B

9A

9C



H Right Aft Fuselage (continued)7. Engine Tailpipe/Thrust Reverser: Inspect the engine turbineblades for nicks, cracks, or damage from foreign object ingestion.Confirm that the thrust reverser is stowed and that the exhaust ductis clear.

8. Engine Pylon and Static Discharger: Check the enginepylon for general condition and evidence of leaks. Ensure the inboardcowl latches are secure. One static discharger (solid type) is on theengine pylon trailing edge. Check for security and damage.

9. Rear Lower Side Fuselage: Check that the external powerplug is properly connected (9A) or that the door is securely closed.Make sure the air conditioner heat exchanger inlet is unobstructed(9B). Check that the grates over the Freon condenser cooling outletsare clear (9C).


1

6

2 3 4A

54B



I Horizontal and Vertical Stabilizer1. Right Horizontal and Vertical Stabilizer: Check the righthorizontal and vertical stabilizer, elevator, and rudder for generalcondition, security, and damage.

2. Right Feeder Tank Vent: Inspect the tank vent for cleanli-ness and freedom from obstruction.

3. Static Dischargers: Check the static wicks for security and thatall brushes are of equal length and fullness (brush type). Two typesof static dischargers are used: a brush type and a solid plastic type.The number of dischargers varies depending on aircraft modificationstatus, operator preference, and discharger type.

4. Anti-Collision and Strobe/Navigation Light: Visuallyinspect the lights on the vertical stabilizer top (4A). If night flight isanticipated, perform an operational check. Inspect the white nav-igation light on the tailcone for security, lens condition, and bulbcondition (4B).

5. Left Horizontal and Vertical Stabilizer: Inspect the gen-eral condition of the left horizontal and vertical stabilizer, rudder,elevator, and static dischargers for general condition, security, anddamage.

6. Stabilizer Alignment Mark and Left Feeder Tank Vent:Visually check that the horizontal stabilizer is aligned with the cen-ter or takeoff mark on the vertical stabilizer. Check that the tankvent is clean and free from obstructions.


2 3 4

5

1



J Rear Compartment1. Rear Compartment Access Door: Open the door to gainaccess to the rear compartment.

2. Hydraulic Reservoirs: With the sight gage on the reservoirside, check the two hydraulic reservoirs for adequate fluid quantity.Fluid should be between the FILLING LEVEL and BELOW THISLIMIT REFILL TO LEVEL indexes. Inspect all plumbing forsecurity and loose fittings. Inspect the hydraulic system componentsand the general area for evidence of hydraulic leaks.

3. Circuit Breakers: Verify that all circuit breakers on the bat-tery box are in, and that the battery charge switch is in the appropriateposition.

4. Batteries: Inspect the battery cases and area surrounding themfor evidence of leaks. Check both battery connectors for tightness byturning them clockwise.

5. Rear Compartment: Inspect all areas of the rear compartmentfor general condition, evidence of leaks, and equipment security.Close and secure the rear compartment access door.


1

4

5

2 3



K Left Aft Fuselage1. Fuselage Underside: Check the underside of the fuselage forevidence of fluid leaks (streaks, etc.). Examine all antennas for secu-rity and damage.

2. Fire Extinguisher Discharge Disc: Check that the disc isintact. Absence of the disc indicates fire extinguisher bottle over-pressurization and discharge; maintenance is required.

3. Engine Pylon and Static Discharger: Check the enginepylon for general condition and evidence of leaks. Ensure the inboardcowl latches are secure. One static discharger (solid type) is on theengine pylon trailing edge. Check for security and damage.

4. Engine Tailpipe/Thrust Reverser: Inspect the turbine exhaustblades for cracking, nicks, or damage from foreign object ingestion.Confirm that the thrust reverser is properly stowed and that theexhaust duct is clear.

5. Engine Cowl: Check the engine cowl for general condition,and ensure that it is secure and latched.


6 7 8



K Left Aft Fuselage (continued)6. Engine Fuel Filter Bypass: Check that the indicator is flushto signify normal filter operation. If the indicator is protruding, thefilter is bypassing; maintenance is required before flight.

7. Oil Sight Gage: Push the spring-loaded access door to visuallycheck oil quantity through the sight gage. Total engine oil tank capa-city is approximately 1.5 gallons. Oil quantity should be checkedwithin one hour after engine shutdown.

8. Engine Oil Filter Bypass: Check that the red indicator isflush to indicate normal filter operation. If the indicator is protrud-ing, the filter is bypassing; maintenance is required before flight.


1

4

5

2 3



L Left Main Gear1. Wheel Embellisher: Check that no fasteners are missing fromthe embellisher.

2. Tires and Wheels: Check tires for proper inflation, abnormalwear, slip mark alignment, and other damage. Use only nitrogen toinflate the tires; never use dry air or oxygen. Inflation pressuredepends on aircraft weight and service bulletin status. Check thewheel for dents, metal pickup on wheel surface, and other damage.

3. Brakes: Inspect the brakes for wear pin extension, leaks, dam-age, and signs of overheating. If the wear pin is recessed in the brakereturn box, excessive wear is indicated; maintenance is requiredbefore flight.

4. Strut: Check the strut for leaks and extension. Normal strutextension is 6 inches at 68°F; extension varies with ambient tem-perature. Consult the maintenance manual extension/temperaturechart.

5. Wheel Well: Inspect the wheel well for general condition, andsecurity of hydraulic lines and cables. Check for indication of fluidleaks.


1

42 3



M Left Wing Trailing Edge1. Airbrakes: Check the condition and position of the airbrakes.Verify that the position of the airbrakes corresponds with airbrake han-dle position.

2. Flaps: Check the condition and security of the flaps. Verify thatthe position of the flaps corresponds with flap handle position. Checkthe condition of all cables, actuators, and lines forward of the flaps,if extended.

3. Aileron: Inspect the aileron for damage and security.

4. Static Dischargers: Check that the static dischargers on theaileron and wing tip are present, secure, and undamaged and thatthe brushes are equal in length and fullness (if installed). Two differentstatic dischargers are used on the Falcon 10/100. One consists of athree brush assembly, the other consists of a solid plastic unit withsharp points. The number of static dischargers varies depending onaircraft modification status, owner preference, and discharger type.


1

4

6

2 3

5



N Left Wing and Engine Inlet1. Wing Tip and Navigation/Strobe Lights: Inspect the wingtip for damage. Check the navigation and strobe lights for lens dam-age, security, and any indication of a burned out bulb. If night flightis anticipated, perform an operational check of the navigation andstrobe lights.

2. Filler Fuel Cap: Check that the filler cap is properly secured andthat the locking tab is facing aft with the tab flush with the wingsurface.

3. Wing Fuel Vent: Check the fuel vent on the underside of thewing for cleanliness, freedom from obstructions, and fuel leaks.

4. Leading Edge: Inspect the slats for general condition, andsecurity, as well as for dents or other damage. Check the wing fencefor security, condition, and damage.

5. Engine Inlet: Inspect the engine inlet for foreign objects andloose fasteners. Check the condition of the fan blades and inlet guidesfor nicks, cracks, and signs of damage. The Pt2/Tt2 probe should beclean and undamaged.

6. Landing Light: Check the light for security. Check the lensfor cracks and signs of overheating. Check the condition of the bulb.



The cabin inspection completes the preflight checklist for the Falcon10/100.

First Aid Kit . . . . . . . . . . . . . . . . . . . . CHECKEDCheck the first aid kit for adequate supplies.

Crash Axe . . . . . . . . . . . . . . . . . . . . . CHECKEDIf included on the aircraft equipment list, check for presence andproper storage.

Fire Extinguishers . . . . . . . . . . . . . . . . . CHECKEDCheck that the portable fire extinguishers are present and fully charged.Check for valid inspection.

Emergency Exit Safety Pin . . . . . . . . . . . . CHECKEDRemove the emergency exit safety pin.

Survival Equipment . . . . . . . . . . . . . . . . CHECKEDFor overwater flight, there should be a life jacket stowed under eachseat for the passenger and crew and life rafts (with sufficient capac-ity for crew and passengers).

Documents . . . . . . . . . . . . . . . . . . . . . CHECKEDEnsure that the following documents are onboard the aircraft:� registration certificate� airworthiness certificate� radio station license� aircraft flight manual (AFM)� operating manuals (for aircraft and avionics equipment)� performance manual (with weight and balance information).

Seatbelts . . . . . . . . . . . . . . . . . . . . . . CHECKEDCheck the passenger seatbelts for security and condition prior to pas-senger loading.

Passenger Oxygen Masks . . . . . . . . . . . . . CHECKEDEnsure that the masks are secure within their panels.

Galley and Lavatory . . . . . . . . . . . . . . . . CHECKEDCheck the galley for adequate supplies. Check the galley and lavatoryfor cleanliness.

CabinInspection

Falcon 10/100 For SimuFlite training only 3B-1April 2000

This chapter outlines and expands normal operating proce-dures and includes applicable cautions and warnings. Alsopresented are cold and hot weather operations, as well asparking, mooring, and aircraft storage requirements.

ExpandedNormalProcedures

Chapter 3B

3B-2 For SimuFlite training only Falcon 10/100April 2000


Expanded Normal Procedures

Normal Procedures . . . . . . . . . . . . . . . . . 3B-5

Before Starting Engines . . . . . . . . . . . . . . . 3B-5

Systems Checks – GPU Available . . . . . . . . . . 3B-8

Engine Start, Part 1 . . . . . . . . . . . . . . . . . 3B-11

Systems Not Previously Checked . . . . . . . . . . 3B-13

Engine Start, Part 2 . . . . . . . . . . . . . . . . . 3B-15

After Start . . . . . . . . . . . . . . . . . . . . . . 3B-15

Taxi . . . . . . . . . . . . . . . . . . . . . . . . . 3B-17

Before Takeoff . . . . . . . . . . . . . . . . . . . . 3B-18

Line Up . . . . . . . . . . . . . . . . . . . . . . . 3B-18

Climb . . . . . . . . . . . . . . . . . . . . . . . . 3B-19

Cruise . . . . . . . . . . . . . . . . . . . . . . . . 3B-19

Descent . . . . . . . . . . . . . . . . . . . . . . . 3B-19

Approach . . . . . . . . . . . . . . . . . . . . . . 3B-20

Before Landing . . . . . . . . . . . . . . . . . . . 3B-20

After Landing . . . . . . . . . . . . . . . . . . . . 3B-20

Shutdown . . . . . . . . . . . . . . . . . . . . . . 3B-21

Quick Turn . . . . . . . . . . . . . . . . . . . . . . 3B-21

Parking . . . . . . . . . . . . . . . . . . . . . . . 3B-23

Mooring . . . . . . . . . . . . . . . . . . . . . . . 3B-24

Towing/Taxiing . . . . . . . . . . . . . . . . . . . . 3B-25

Towing – Hard Ground . . . . . . . . . . . . . . . 3B-25

Close Up – Hard Ground . . . . . . . . . . . . . 3B-25

Towing – Soft Ground . . . . . . . . . . . . . . . 3B-26

Close Up – Soft Ground . . . . . . . . . . . . . . 3B-27

Taxiing . . . . . . . . . . . . . . . . . . . . . . . . 3B-29

Storage and Restoring . . . . . . . . . . . . . . . 3B-31

Storage . . . . . . . . . . . . . . . . . . . . . . . 3B-31

Aircraft Configuration, General . . . . . . . . . . 3B-31

Aircraft Grounded for One Month,Aircraft Already Outdoors . . . . . . . . . . . . . 3B-32

Table ofContents


Periodic Checks of Storage Conditions,Each Month . . . . . . . . . . . . . . . . . . . . 3B-34

Periodic Checks of Storage Conditions,Every Two Months – Cranking of Dampers . . . . 3B-35

Periodic Checks of Storage Conditions, Every TwoMonths – Cranking of Hydraulic Jacks . . . . . . 3B-35

Periodic Checks of Storage Conditions Every Four/Two Months (Depending on Engine Storage) . . . 3B-35

Restoring . . . . . . . . . . . . . . . . . . . . . . 3B-36

Hot Weather and Desert Operations . . . . . . . 3B-39

Exterior Inspection . . . . . . . . . . . . . . . . . 3B-39

Engine Start . . . . . . . . . . . . . . . . . . . . . 3B-39

Taxi . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39

Takeoff . . . . . . . . . . . . . . . . . . . . . . . . 3B-39

Shutdown and Postflight . . . . . . . . . . . . . . 3B-39

Cold Weather Operations . . . . . . . . . . . . . 3B-41

Preflight Inspection . . . . . . . . . . . . . . . . . 3B-41

GPU Available . . . . . . . . . . . . . . . . . . . . 3B-42

1,000 AMP Rating . . . . . . . . . . . . . . . . . 3B-42

GPU Not Available . . . . . . . . . . . . . . . . . 3B-42

Ambient Temperature -15°C to -35°C . . . . . . . 3B-42

After Engine Start . . . . . . . . . . . . . . . . . . 3B-43

Before Takeoff . . . . . . . . . . . . . . . . . . . . 3B-44

After Takeoff . . . . . . . . . . . . . . . . . . . . . 3B-44

Cruise/Descent . . . . . . . . . . . . . . . . . . . 3B-45

Landing . . . . . . . . . . . . . . . . . . . . . . . 3B-45

Taxi-In and Park . . . . . . . . . . . . . . . . . . . 3B-45



Before Starting EnginesPassenger Briefing . . . . . . . . . . . . . . . COMPLETED

According to Part 91.519 requirements, the pilot-in-command or acrewmember briefs the passengers on smoking, use of safety belts,location and operation of the passenger entry door and emergencyexits, location and use of survival equipment, and normal and emergencyuse of oxygen equipment. For flights over water, the briefing shouldinclude ditching procedures and use of flotation equipment. An excep-tion to the oral briefing rule is if the pilot-in-command determines thepassengers are familiar with the briefing content. A printed card withthe above information should be available to each passenger tosupplement the oral briefing.

Cockpit Check/Preflight . . . . . . . . . . . . COMPLETED

Standby Horizon . . . . . . . . . . . . . . . . TESTED/OFF

Generators 1/2 . . . . . . . . . . . . . . . . . . . . . . . . ON

Auxiliary Bus . . . . . . . . . . . . . . . . . . . . . . . . ON

Batteries . . . . . . . . . . . . . . . . . . . . CHECKED/ON

DC Power Selector . . . . . . . . . . . . . . . . . . . . . SETSet the DC power selector switch for the type of engine start to beaccomplished.

Battery Start . . . . . . . . . . . . . . . . . . . . NORMALGPU Start . . . . . . . . . . . . . . . . . . . . EXT POWERCold Weather (below 5°C) . . . . . . . . . . . . LOW TEMP

Engine Computers . . . . . . . . . . . . . ON/LIGHTS OUT

Cabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . ON

Overhead Warning Lights . . . . . . . . . . . . . . TESTEDPress to test the following lights:� engine and wing anti-ice (three)� windshield X.FR (one)� left and right ignition (two)� No. 1 and 2 inverters (two)� optional cabin emergency (one).

Fuel Quantity . . . . . . . . . . . . . . . CHECKED/REARSelect TOTAL on the switch between the fuel quantity gages to checktotal fuel quantity. Select REAR to determine fuel quantity in the feed-er tanks. Check for imbalance.

Annunciator Panel Lights . . . . . . . . . . . . . . TESTEDPress the TEST button on the center annunciator panel. Observe thatall annunciators illuminate.

NormalProcedures


Fire Warnings . . . . . . . . . . . . . . . . . . . . . TESTEDFIRE 1 System TEST Button . . . . . . . . . . . . . PRESS

Observe that the light in the FIRE 1 PULL handle illuminates; thefire warning horn sounds.

Fire Warning Horn . . . . . . . . . . . . . . . . . SILENCEObserve that the light in the FIRE 1 PULL handle remainsilluminated.

FIRE 1 System TEST Button . . . . . . . . . . . . RELEASEObserve that the light in the FIRE 1 PULL handle extinguishes.

FIRE 2 System TEST Button . . . . . . . . . . . . . PRESSObserve that the light in the FIRE 2 PULL handle illuminates; thefire warning horn sounds.

Fire Warning Horn . . . . . . . . . . . . . . . . . SILENCEObserve that the light in the FIRE 2 PULL handle remainsilluminated.

FIRE 2 System TEST Button . . . . . . . . . . . . RELEASEObserve that the light in the FIRE 2 PULL handle extinguishes.

Landing Gear Panel . . . . . . . . . . . . . . . . . TESTEDLanding Gear TEST Button . . . . . . . . . . . . . . PRESS

The gear warning horn sounds and the following illuminate:� gear panel lights� flight director annunciator panels� thrust reverser lights� anti-skid panel annunciator� transfer panel lights.

Landing Gear Warning Horn . . . . . . . . . . . . SILENCEObserve that the following lights remain illuminated:� gear panel lights� flight director annunciator panels� thrust reverser lights� anti-skid panel annunciator� transfer panel lights.

Landing Gear TEST Button . . . . . . . . . . . . RELEASEObserve that following lights extinguish:� gear panel lights� flight director annunciator panels� thrust reverser lights� anti-skid panel annunciator� transfer panel lights.



Battery Temperature Warning . . . . . . . . . . . . TESTEDBATTERY TEST Button . . . . . . . . . . . . . . . . PRESS

Observe the temperature indicator needles rising. When the leftneedle indicates 120°, the amber warning light illuminates. When theright needle indicates 150°, the red warning light illuminates.

BATTERY TEST Button . . . . . . . . . . . . . . RELEASE

Hydraulic Quantity . . . . . . . . . . . . . . . . . NORMAL

Parking Brake . . . . . . . . . SET/P BRAKE LIGHT OUT

Cabin Altitude Warning . . . . . . . . . . . . . . . TESTEDCABIN Test Button . . . . . . . . . . . . . . . . . . PRESS

The CABIN annunciator illuminates, and the cabin warning hornsounds.

Cabin Warning Horn . . . . . . . . . . . . . . . . SILENCEObserve that the CABIN annunciator remains illuminated.

CABIN Test Button . . . . . . . . . . . . . . . . . RELEASEObserve that the CABIN annunciator extinguishes.

VMO Warning . . . . . . . . . . . . . . . . . . . . . TESTEDVMO/MMO Test Button . . . . . . . . . . . . . . . . . PRESS

The overspeed warning horn sounds and cannot be silenced.

VMO/MMO/Test Button . . . . . . . . . . . . . . . RELEASEThe warning horn silences.


Systems Checks – GPU AvailableInverters . . . . . . . . . . . . . . . . . . . CHECKED/OFF

Both inverter lights illuminate to signify that the 115V AC buses arenot powered.

AC Voltmeter Selector . . . . . . . . . . . . . . . . STD BYStandby Inverter Switch . . . . . . . . . . . . NO. 1 AC BUS

Observe that the No.1 AC FAIL light extinguishes.

AC Voltmeter . . . . . . . . . . . . . . . . CHECK/115V ACNo. 1 Inverter . . . . . . . . . . . . . . . . . . . . . . . . ONStandby Inverter Switch . . . . . . . . . . . . . . . . . . OFF

Observe that the No. 1 AC FAIL light remains extinguished andvoltage reads zero.

Standby Inverter Switch . . . . . . . . . . . . NO. 2 AC BUSObserve that the No. 2 AC FAIL light extinguishes and AC volt-meter reads 115V AC.

No. 2 Inverter . . . . . . . . . . . . . . . . . . . . . . . . ONStandby Inverter Switch . . . . . . . . . . . . . . . . . . OFF

Observe that the voltmeter reads zero and No. 2 AC FAIL lightremains extinguished.

AC voltmeter Selector . . . . . . . . . . . . . . . . . . NO. 2Observe 115V AC.

AC Voltmeter Selector . . . . . . . . . . . . . . . . . . NO. 1Observe 115V AC.

No. 1 and No. 2 Inverters . . . . . . . . . . . . . . . . . OFF

No. 1 Power Lever . . . . . . . . . . . . . . . . . ADVANCE

Hydraulic Standby Pump . . . . . . . . . . . ON/NORMALHydraulic Standby Pump . . . . . . . . . . . . . . . . . . ON

Observe that the hydraulic pressure in the No. 1 system rises to2,150 PSI, then cycles between 1,600 and 2,150 PSI.

Hydraulic Standby Pump . . . . . . . . . . . . . . NORMALObserve that the hydraulic pressure in the No. 1 system rises to2,150 PSI, then cycles between 1,600 and 2,150 PSI.



Trims . . . . . . . . . . . . . . . . . . . . . CHECKED/SETWith the stabilizer trim in the green range and the No. 1 power leveradvanced to takeoff, check the stabilizer trim.Aircraft with Split Trim Switches:

Trim . . . . . . . . . . . . . . . . . . . . . . . . . CHECKNo movement unless both switches are pressed.

One Stab Trim Switch onPilot’s Control Wheel . . . . . . NOSE-DOWN/NOSE-UP

Check that normal trim does not move.

Other Stab Trim Switch onPilot’s Control Wheel . . . . . . NOSE-DOWN/NOSE-UP


One Stab Trim Switch onCopilot’s Control Wheel . . . . . NOSE-DOWN/NOSE-UP


Other Stab Trim Switch onCopilot’s Control Wheel . . . . . NOSE-DOWN/NOSE-UP


Pilot’s Stab Trim . . . . . . . . . . . . . . . . NOSE-DOWNObserve movement nose-down.

Copilot’s Stab Trim . . . . . . . . . . . . . . . . . NOSE-UPCheck that movement of the stabilizer trim stops.

Stab Trim Switches . . . . . . . . . . . . . . . . . RELEASEPilot’s Stab Trim . . . . . . . . . . . . NOSE-DOWN UNTIL . . . . . . . . . . . . . .WARNING LIGHT ILLUMINATES

The red STAB TRIM annunciator illuminates at 4°.

Pilot’s Stab Trim . . . . . . . . . . . . . . . . . . NOSE-UPObserve movement nose-up. The STAB TRIM annunciatorextinguishes.

Copilot’s Stab Trim . . . . . . . . . . . . . . NOSE-DOWNCheck that movement of the stabilizer trim stops.

Stab Trim Switches . . . . . . . . . . . . . . . . . RELEASEEmergency Trim . . . . . . . . . . . . . SELECT NOSE UP

Observe that the normal trim CB opens. Use emergency trim untilred STAB TRIM annunciator illuminates which is at 8° nose-up.

Emergency Trim . . . . . . . . . . . . NOSE-DOWN UNTIL . . . . . . . . . . . . . . . . . . . . .ANNUNCIATOR OUTStabilizer Normal Trim . . . . . . . . . . . . . . CB RESETStab Trim . . . . . . . . . . . . . . . . SET FOR TAKEOFF

Use the normal system.

CAUTION: Whenever the sta-bilizer trim is in motion, an auralwarning sounds.

3B-10 For SimuFlite training only Falcon 10/100March 2001

Aileron/Rudder Trim . . . . . . . . . . . . . CHECKED/SETMonitor the trim indicator, control wheel movement, and the rudderpedals for proper operation.

Airbrakes . . . . . . . . . . . . . . . . . . . . . . . EXTENDObserve the illumination of the amber AIRBRAKE annunciator onthe landing gear panel and the red AIRBRAKE annunciator on thefailure warning panel.

No. 1 Power Lever . . . . . . . . . . . . . . . . . . CUTOFFObserve that the red AIRBRAKE annunciator extinguishes.

Airbrakes . . . . . . . . . . . . . . . . . . . . . . RETRACTObserve that the amber AIRBRAKE annunciator extinguishes.

No. 1 Stall Warning . . . . . . . . . . . . . . . . . . TESTEDNo. 1 STALL Test Button . . . . . . . . . . . . . . . . PUSH

The stall warning horn sounds and cannot be silenced. Observe thatthe slat indicator illuminates red, then green when the outboard slatsare extended. Observe that the igniter lights on the overhead panelilluminate.

No. 1 STALL Test Button . . . . . . . . . . . . . RELEASEObserve that the slat indicator light illuminates red, then extinguisheswhen the outboard slats retract. The igniter lights extinguish after 10seconds.

Slats/Flaps . . . . . . . . . . . . . . . . . . . . . . CYCLEDSlats Only . . . . . . . . . . . . . . . . . . . . . . . SELECT

Observe that the slat indicator turns green when the slats are extended.

Flaps . . . . . . . . . . . . . . . . . . . SELECT 15°/30°/52°Observe that the flaps stop and indicate appropriately at each setting.

Slats/Flaps . . . . . . . . . . . . . . . . . SELECT CLEANObserve that the red slat indicator does not extinguish before theflaps reach the clean position.

Hydraulic Standby Pump . . . . . . . . . . . . . . . . . OFF

Batteries/External Power . . . . . . . . . . . . . . . . . OFF

Airbrake Limitations� The pilot must keep his hand

on the airbrake control handleuntil the normal extension orretraction of the airbrakes isconfirmed.

� Do not hesitate to use theairbrakes in the event the flightenvelope limits (VMO/MMO)are accidentally exceeded.

� In approach and below a heightof 500 ft AGL, the airbrakesmust not deliberately be keptextended.

� Do not use below 500 ft AGL.� In manual flight conditions,

accompany airbrake extensionwith a nose-up maneuver of thepitch trim control to maintainlongitudinal attitude. Inautomatic flight conditions, amomentary nose-downmovement occurs on airbrakeextension.

Falcon 10/100 For SimuFlite training only 3B-11March 2001


Engine Start – Part 1Direct Vision Window . . . . . . . . . . . . . . . . . . OPEN

Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . ON

DC Power Selector . . . . . . . . . . . . . . . . . . . . . SET

Emergency Exit Lights . . . . . . . . . . TESTED/ARMEDCabin Emergency Light Switch . . . . . . OFF/ARMED/ON

Observe that the exit lights illuminate in the cockpit dome fixture.

Cabin Emergency Light Switch . . . . . . . . . . . ARMEDCabin Lights . . . . . . . . . . . . . . . . . . . . . . . . ON

Anti-Collision Light . . . . . . . . . . . . . . . . . . . . . ON

Freon/HP Bleed . . . . . . . . . . . . . . . . . OFF/CLOSED

No. 1 Radio . . . . . . . . . . . . . . . . . . . . . . . . . . ONAt least one radio should be on for engine start.

Parking Brake . . . . . FULL AFT/P BRAKE LIGHT OUTVerify that the parking brake handle is set full aft and that the P BRAKEannunciator is out.

No. 2 Booster Pump . . . . . . . . . . . . . ON/LIGHT OUTTurn the No. 2 booster pump on; confirm that the FUEL P2 annunci-ator extinguishes.

No. 2 Engine . . . . STARTED/PARAMETERS CHECKEDAbort the start if any of the following occurs:� engine oil pressure does not rise within 10 seconds after lightoff� EXH. TEMP. does not rise within 10 seconds with SPR switch

energized� EXH. TEMP. is rising rapidly and approaching the 860°C limit� N1 rotation is not observed by 20% N2� N2 speed does not increase smoothly and rapidly to 24% after light-

off idle speed is not reached in time limit shown in Table 3B-A .

Oil LimitationSee the Quick Reference chapterfor oil consumption, pressure,and temperature limitations.

Powerplant LimitationSee the Quick Reference chapterfor powerplant operating limits.

Condition Time Limit

Ground Start/Starter Assist Airstart from 10%N2 to Lightoff

10 seconds

Ground Start from Lightoff to Idle 50 seconds

Windmilling Airstart from Windmilling N2 to60% N2

25 seconds

Table 3B-A; Start Limitations


No. 2 Start Button . . . . . . . . . PRESS MOMENTARILYPress the button for no more than two seconds. Check that N2 rises.

At 10% N2 and N1 Rotation:No. 2 Power Lever . . . . . . . . . . . . . . . . . . . . IDLE

Observe that:� ignition light illuminates� ITT rises� N2 increases� N1 rotates by 20% N2� oil pressure rises.

ITT . . . . . . . . . . . . . . . . . . . . . . . . . MONITOR

At 50% N2:Igniter Light . . . . . . . . . . . . . . . . . . . CHECK/OUT

The light extinguishes when the starter relay disengages.

Oil Pressure Annunciator . . . . . . . . . . . . CHECK/OUTAll Engine Instruments . . . . . . . . CHECK/STABILIZED

No. 2 Generator . . . . . . . . . . . . . . . . . . CHECKEDDC Voltmeter Selector . . . . . . . . . . . . . . . . . GEN 2

Observe a positive ammeter reading and a voltage of 27V for threeminutes.

No. 2 Stall . . . . . . . . . . . . . . . . . . . . . . . TESTEDNo. 2 STALL Test Button . . . . . . . . . . . . . . . . PUSH

The stall warning horn should sound and cannot be silenced. Observethat the slat indicator illuminates red, then green when the outboardslats are extended. Observe that the igniter lights on the overheadpanel illuminate.

No. 2 STALL Test Button . . . . . . . . . . . . . RELEASEObserve that the slat indicator illuminates red and remains red. Theigniter lights extinguish after 10 seconds.

Generator LimitationsMaximum ContinuousAmperage . . . . . . . . 300A

MaximumOutput . . 350A (5 MINUTES)

Maximum Voltage . . . . . 31V



Systems Not Previously CheckedNo. 1 Power Lever . . . . . . . . . . . . . . . . . ADVANCE

Hydraulic Standby Pump . . . . . . . . . . . ON/NORMALHydraulic Standby Pump . . . . . . . . . . . . . . . . . . ON

Observe that the hydraulic pressure in the No. 1 system rises to2,150 PSI, then cycles between 1,600 and 2,150 PSI.

Hydraulic Standby Pump . . . . . . . . . . . . . . NORMALObserve that the hydraulic pressure in the No. 1 system rises to2,150 PSI, then cycles between 1,600 and 2,150 PSI.

Slat Indicator . . . . . . . . . . . . . . . . . . . . . CHECKIndicator extinguishes after outboard slats retract.Hydraulic pressure should cycle between 1,650 and 2,150 PSI.

No. 1 Stall . . . . . . . . . . . . . . . . . . . . . . . TESTEDNo. 1 STALL Test Button . . . . . . . . . . . . . . . . PUSH

The stall warning horn should sound and cannot be silenced. Observethat the slat indicator illuminates red, then green when the outboardslats are extended. Observe that the igniter lights on the overheadpanel illuminate.

No. 1 STALL Test Button . . . . . . . . . . . . . RELEASEObserve that the slat indicator illuminates red, then extinguisheswhen the outboard slats retract. The igniter lights extinguish after 10seconds.

Trims . . . . . . . . . . . . . . . . . . . . . CHECKED/SETWith the stabilizer trim in the green range and the No. 1 power leveradvanced to takeoff, check the stabilizer trim.Aircraft with Split Trim Switches:

Trim . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKThere is no movement unless both switches are pressed.

One Stab Trim Switch onPilot’s Control Wheel . . . . . . . NOSE-DOWN/NOSE-UP


Other Stab Trim Switch onPilot’s Control Wheel . . . . . . . NOSE-DOWN/NOSE-UP


One Stab Trim Switch onCopilot’s Control Wheel . . . . . . NOSE-DOWN/NOSE-UP


Other Stab Trim Switch onCopilot’s Control Wheel . . . . . . NOSE-DOWN/NOSE-UP


CAUTION: Whenever the sta-bilizer trim is in motion, an auralwarning sounds.


Pilot’s Stab Trim . . . . . . . . . . . . . . . . NOSE-DOWNObserve movement nose-down.

Copilot’s Stab Trim . . . . . . . . . . . . . . . . . NOSE-UPCheck that movement of the stabilizer trim stops.

Stab Trim Switches . . . . . . . . . . . . . . . . . RELEASEPilot’s Stab Trim . . . . . . . . . . . . NOSE-DOWN UNTIL . . . . . . . . . . . . . .WARNING LIGHT ILLUMINATES

The red STAB TRIM annunciator illuminates at 4°.

Pilot’s Stab Trim . . . . . . . . . . . . . . . . . . NOSE-UPObserve movement nose-up. The STAB TRIM annunciatorextinguishes.

Copilot’s Stab Trim . . . . . . . . . . . . . . NOSE-DOWNCheck that movement of the stabilizer trim stops.

Stab Trim Switches . . . . . . . . . . . . . . . . . RELEASEEmergency Trim . . . . . . . . . . . . . SELECT NOSE UP

Observe that the normal trim CB opens. Use emergency trim untilred STAB TRIM annunciator illuminates at 8°.

Emergency Trim . . . . . . . . . . . . NOSE-DOWN UNTIL . . . . . . . . . . . . .WARNING LIGHT EXTINGUISHESNormal Trim . . . . . . . . . . . . . . . . . . . . CB RESETStab Trim . . . . . . . . . . . . . . . . SET FOR TAKEOFF

Use the normal system.

Aileron/Rudder Trim . . . . . . . . . . . . . CHECKED/SETMonitor the trim indicator, control wheel movement, and the rudderpedals for proper operation.

Airbrakes . . . . . . . . . . . . . . . . . . . . . . . EXTENDObserve the illumination of the amber AIRBRAKE annunciator onthe landing gear panel and the red AIRBRAKE annunciator on thefailure warning panel.

No. 1 Power Lever . . . . . . . . . . . . . . . . . . CUTOFFObserve that the red AIRBRAKE annunciator extinguishes.

Airbrakes . . . . . . . . . . . . . . . . . . . . . . RETRACTObserve that the amber AIRBRAKE annunciator extinguishes.

Airbrake Limitations� The pilot must keep his hand

on the airbrake control handleuntil the normal extension orretraction of the airbrakes isconfirmed.

� Do not hesitate to use theairbrakes in the event the flightenvelope limits (VMO/MMO)are accidentally exceeded.

� In approach and below a heightof 500 ft AGL, the airbrakesmust not deliberately be keptextended.

� Do not use below 500 ft AGL.� In manual flight conditions,

accompany airbrake extensionwith a nose-up maneuver of thepitch trim control to maintainlongitudinal attitude. Inautomatic flight conditions, amomentary nose-downmovement occurs on airbrakeextension.



Engine Start, Part 2No. 1 Booster Pump . . . . . . . . . . . . . ON/LIGHT OUT

No. 1 Booster Pump . . . . . . . . . . . . . . . . . . . . ONFuel P1 Annunciator . . . . . . . . . . . . CHECKED/OUT

No. 1 Engine . . . . STARTED/PARAMETERS CHECKEDStart the No. 1 engine using the same procedures for starting the No.2 engine.

No. 1 Generator . . . . . . . . . . . . . . . . . . CHECKEDDC Voltmeter Selector . . . . . . . . . . . . . . . . . GEN 1

Observe a positive ammeter reading and a voltage of 27V for threeminutes.

DC Power Selector . . . . . . . . . . . . . . . . . NORMAL

Ground Power . . . . . . . . . . . . . . . . . . . REMOVEDIf ground power was used for start, place the DC power selector switchto NORMAL prior to disconnecting the GPU. The GENE. 1/2 annun-ciators do not extinguish until the GPU is disconnected.

After StartInverters . . . . . . . . . . . . . . . . . . . . CHECKED/ON

Exterior Lights . . . . . . . . . . . . . . . . . . . . . . . SET

Engine Computers/Anti-Ice . . . . . . . . . . . . CHECKEDEngine Anti-Ice Switches . . . . . . . . . . . . . . . . . ONNo. 1 Engine Computer Switch . . . . . . . . . . . . . . OFF

Guard the No. 1 engine power lever at the same time; an enginerunaway is possible. Engine RPM increases slightly.

No. 1 Power Lever . . . . . . . . . . ADVANCE SLIGHTLYEngine RPM follows power lever movement. Check for illuminationof the left engine’s ANTI-ICE light as the power lever is advancedduring the computer check.

If Engine Anti-Ice Is Not Required:

ENG 1 ANTI-ICE Switch . . . . . . . . . . . . . . . . . OFFObserve the ENG ANTI-ICE light extinguishes.

No. 1 Power Lever . . . . . . . . . . . . RETARD TO IDLENo. 1 Engine Computer Switch . . . . . . . . . . . . . . ON

Guard the No. 1 engine power lever at the same time; an enginerunaway is possible.

No. 2 Engine Computer Switch . . . . . . . . . . . . . . OFFGuard the No. 2 engine power lever at the same time; an enginerunaway is possible. Engine RPM increases slightly.

Generator LimitationsMaximum ContinuousAmperage . . . . . . . . 300A

MaximumOutput . . 350A (5 MINUTES)

Maximum Voltage . . . . . 31V

Wing Anti-Ice Limitations� Must be on in flight when total

temperature is below +5°C andwhen icing conditions areanticipated.

� Must be off when total temper-ature is above +10°C.

� Do not operate on the ground.

Engine Anti-Ice Limitations� Must be on when total air

temperature is below +5°C andicing conditions are anticipated.

� Must be off when total airtemperature is above +10°C.


No. 2 Power Lever . . . . . . . . . . ADVANCE SLIGHTLYEngine RPM follows power lever movement. Check for illuminationof the right engine’s ANTI-ICE annunciator as the power lever isadvanced during the computer check.

If Engine Anti-Ice Is Not Required:

ENG 2 ANTI-ICE Switch . . . . . . . . . . . . . . . . . OFFObserve the ENG ANTI-ICE light extinguishes.

No. 2 Power Lever . . . . . . . . . . . . RETARD TO IDLENo. 2 Engine Computer Switch . . . . . . . . . . . . . . ON

Guard the No. 2 engine power lever at the same time; an enginerunaway is possible.

Engine Anti-Ice/Pitot Heat . . . . . . . . . . AS REQUIRED

Windshield Heat . . . . . . . . . . . . . . . . . . . NORMAL

Emergency Power/Standby Horizon . . . . . ON/UNCAGED

Q Unit/Yaw Damper . . . . . . . . . . . . . . . TESTED/ONYaw Damper . . . . . . . . . . . . . . . . . . . . . ENGAGEAnnunciator Panel Test Button . . . . . . . . . PRESS/HOLD

Observe that the yaw damper disengages with all annunciators,except the Q UNIT annunciator, which illuminates after severalseconds along with the RUD/AIL lights on the copilot’s CB panel.

Annunciator Panel Test Button . . . . . . . . . . . RELEASEYaw Damper . . . . . . . . . . . . . . . . . . . . . ENGAGE

Radios/Avionics . . . . . . . . . . . . . . . . . . . . . . . ON

Battery Temperatures . . . . . . . . . . . . . . . CHECKED

Standby Pump/Freon/HP Bleed . . . . . . ON/AUTO/OPEN

Circuit Breakers . . . . . . . . . . . . . . . . . . CHECKED

Galley Power . . . . . . . . . . . . . . . . . . . . . . . . . ON

Direct Vision Windows . . . . . . . . . . . . . . . . CLOSED

TOLD Card/Bugs . . . . . . . . . . . . . COMPUTED/SET

Yaw Damper/Q UnitLimitationsQ UNIT AnnunciatorON . . . . . 260 KIAS/0.76 M

On Approach if Q UNIT Light On:

RUD Q . 140 KTS MINIMUMUNTIL LANDING ASSURED

AIL Q . . . ADD 10 KTS to VAP

If Not Operating:

Airspeed . . 260 KIAS (0.76 M)MAXIMUM)

Windshield ElectricalHeating LimitationThe PILOT and COPILOTWINDSHIELD control switchesmust not be selected to the MAXposition, except in flight and onlyif the NORM position is notsufficient in icing conditions.

Battery LimitationMaximum battery temperature forstart is 120°F.



TaxiBecause nosewheel steering receives power from the right engine,taxiing with only the right engine operating is authorized. Taxiingwith only the left engine operating is not recommended.

Normal Brakes/Steering . . . . . . . . . . . . . . CHECKEDCheck operation of normal brakes and steering when leaving the park-ing area at a slow speed.

Anti-Skid . . . . . . . . . . . . . . . . . . . . . . . TESTEDAfter clearing the ramp area, test the anti-skid for the left and rightbrakes individually.While Holding Pressure on Left Brake:

Anti-Skid Test Button . . . . . . . . . . PRESS/RELEASEBrake pressure is momentarily released (amber anti-skid lightilluminates) and then restored (amber anti-skid light extinguishes).

While Holding Pressure on Right Brake:

Anti-Skid Test Button . . . . . . . . . . PRESS/RELEASEBrake pressure is momentarily released (amber anti-skid lightilluminates) and then restored (amber anti-skid light extinguishes).

Emergency Brakes . . . . . . . . . . . . . . . . . CHECKEDSmoothly move the emergency brake handle aft.

Thrust Reversers (if installed) . . . . . . . . . . CHECKEDReverse Levers . . . . . . . . . . . . . . . IDLE REVERSE.

Observe the amber TRANS lights illuminate and then extinguishas the GREEN REV indicators illuminate.

Reverse Thrust . . . . . . . INCREASE APPROXIMATELY . . . . . . . . . . . . . . . . . . . . .5% N2 ABOVE IDLE

Emergency Stow Switches . . . . . . . . . . EMERG STOWObserve that the power levers automatically return to reverse idle.Observe that the GREEN REV indicators extinguish, followed by theAMBER TRANS indicators extinguishing.

Reverse Levers . . . . . . . . . . . . . . . . . . . . . STOWEmergency Stow Switches . . . . . . . . . . . . . . . NORMT/R Test Button . . . . . . . . . . . . . . . . . . . . . PRESS

This tests the warning horn, which sounds if a reverser door positiondoes not agree with the respective power lever.

Thrust Reverser Limitation� Thrust reversers are approved

for ground use only.� Maximum 50% N1 for power

lever retarder check.


Slats/Flaps . . . . . . . . . . CYCLED/SET FOR TAKEOFFSlats Only . . . . . . . . . . . . . . . . . . . . . . . SELECT

Observe that the slat indicator turns green when the slats are extended.

Flaps . . . . . . . . . . . . . . . . . . . SELECT 15°/30°/52°Observe that the flaps stop and indicate appropriately at each setting.

Slats/Flaps . . . . . . . . . . . . . . . . . SELECT CLEANObserve that the red slat indicator does not extinguish before theflaps reach the clean position.

Slats/Flaps . . . . . . . . . . . . . . . . SET FOR TAKEOFF

Airbrakes . . . . . . . . . . . . . . . . . . . . RETRACTED

Trims . . . . . . . . . . . . . . . . . . . . . . . THREE SET

Flight Controls/Hydraulic Systems . . . . FREE/CHECKED

Before TakeoffIn case of crosswind or tailwind, set takeoff N1 speed progres-sively when reaching approximately 30 kts.

Flight Instruments . . . . . . . . . . . . . . NO FLAGS/SET

Pressurization . . . . . . . . . . . . . . . . . . . . . . . . SETSet the pressure to 29.92, the altitude to the flight plan altitude, and therate knob to the center detent.

Crew Briefing . . . . . . . . . . . . . . . . . . COMPLETED

Line UpParking Brake . . . . . . . . . . OFF/BRAKE LIGHT OUT

Radar/Transponder . . . . . . . . . . . . . . . . . . . . . ON

Pitot/AOA Heats . . . . . . . . . . . . . . ON/LIGHTS OUT

Start Selectors . . . . . . . . . . . . . . . . . . . AIRSTART

Strobes/Recognition Lights (if installed) . . . . . . . . . . ON

Warning Lights . . . . . . . . . . . . . . . . . . . . . CLEAR

F.A.T.S. Check . . . . . . . . . . . . . . . . . COMPLETEDFlaps, airbrakes, trims, and speeds are safety-checked for takeoff.

Engine SynchronizerLimitationEngine synchronization, ifinstalled, must not be used duringtakeoff, go-around, and landing.

Fuel Computer LimitationFuel computers must be operativefor takeoff.

X-Feed and Rear TankIntercom Valves LimitationThese valves must be closedduring takeoff.

Recognition Light LimitationUse of recognition lights on theground is limited to 5 minutes,followed by 15 minutes off toensure sufficient cooling.

Flap LimitationIn flight, each actuation of the flapcontrol handle must be limited tothe next detent position.



ClimbLanding Gear . . . . . . . . . . . . . . . . . . . . . . . . UP

Slats/Flaps . . . . . . . . . . . . . . . . . . . . . . . CLEAN

Anti-Ice Panel . . . . . . . . . . . . . . . . . . . . . . . . SET

Start Selectors . . . . . . . . . . . . . . . GROUND START

Landing/Taxi Lights . . . . . . . . . . . . . . . . . . . . OFF

Cabin Signs . . . . . . . . . . . . . . . . . . AS REQUIRED

Pressurization . . . . . . . . . . . . . . . . . . . CHECKED

Standby Pump . . . . . . . . . . . . . . . . . . . . NORMAL

Altimeter (At Transition Altitude) . . . . . . . . . . . . SET

Recognition Lights . . . . . . . . . . . . . . . . . . . . . OFF

CruiseElectrical Systems . . . . . . . . . . . . . . . . . CHECKED

Fuel Panel . . . . . . . . . . . . . . . . . . . . . CHECKED

Engine Instruments . . . . . . . . . . . . . . . . CHECKED

Hydraulic Systems . . . . . . . . . . . . . . . . . CHECKED

Pressurization . . . . . . . . . . . . . . . . . . . CHECKED

Oxygen . . . . . . . . . . . . . . . . . . . . . . . CHECKED

DescentPressurization . . . . . . . . . . . . . . . . . . . . . . . . SET

Fuel Quantity/Panel . . . . . . . . . . . . . . . . CHECKED


TOLD Card . . . . . . . . . . . . . . . . . . . COMPUTED

Crew Briefing . . . . . . . . . . . . . . . . . . COMPLETED

Altimeter (At Transition Level) . . . . . . . . . . . . . . SET

Recognition Lights (if installed) . . . . . . . . . . . . . . ON

Pressurization LimitationsMaximum CabinPressure Altitude . . . 8,000 FT

MaximumDifferential Pressure . . 9.1 PSI

Operating LimitationMaximumOperating Altitude . . . FL 450

Oxygen LimitationAbove FL 410, all passengersmust wear an oxygen masksecured to the face and connectedto the oxygen system.

X-Feed Valve LimitationIn flight, if Jet B or JP-4 fuel isused, the X-FEED valve must beopened whenever the flight ismade at an altitude for which theoperation of the engine with afailed booster pump is notguaranteed.


ApproachSlats/Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . SET

Standby Pump . . . . . . . . . . . . . . . . . . . . . . . . ON


Avionics . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

Cabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . ON

Before LandingLanding Gear . . . . . . . . . . DOWN/3 GREEN/NO RED

Hydraulic Pressure/Quantity . . . . . . . . . . . CHECKED

Anti-Skid . . . . . . . . . . . . . . . . . . . . . . . TESTED

Start Selectors . . . . . . . . . . . . . . . . . . . AIRSTART

Slats/Flaps . . . . . . . . . . . . . . . . SET FOR LANDING

Landing Lights . . . . . . . . . . . . . . . . . . . . . . . SET

After LandingAnti-Ice/Pitot/Windshield Heat . . . . . . . . . . . . . . OFF

Start Selectors . . . . . . . . . . . . . . . GROUND START

Strobe/Recognition Lights . . . . . . . . . . . . . . . . . OFF

Landing/Taxi Lights . . . . . . . . . . . . . . . . . . . . SET

Slats/Flaps . . . . . . . . . . . . . . . . . . . . . . . CLEAN

Airbrakes . . . . . . . . . . . . . . . . . . . . RETRACTED

Trims . . . . . . . . . . . . . . . . . . . SET FOR TAKEOFF

Radar/Transponder . . . . . . . . . . . . . . . . . STANDBY

Flap LimitationIn flight, each actuation of the flapcontrol handle must be limited tothe next detent position.

Engine SynchronizerLimitationEngine synchronization, ifinstalled, must not be used duringtakeoff, go-around, and landing.

Brakes LimitationIt is strictly forbidden to depressthe brake pedals prior totouchdown.



ShutdownEmergency Power/Standby Horizon . . . . . . OFF/CAGED

Parking Brake . . . . . . . . . . . . . . . . . . . . . . . SET

Standby Pump/Freon/HP Bleed . . . . . OFF/OFF/CLOSED

Power Levers . . . . . . . . . . . . . . . . . . . . . CUTOFFOperate engines at idle for two minutes after landing before shutdown.

Exterior/Emergency Exit Lights . . . . . . . . . . . . . OFF

Inverters . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

Booster Pumps . . . . . . . . . . . . . . . . . . . . . . . OFF

Radios . . . . . . . . . . . . . . . . . . . . . . . . . ALL OFF

Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

Auxiliary Bus . . . . . . . . . . . . . . . . . . . . . . . . OFF

Chocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IN

Parking Brake . . . . . . . . . . . . . . . . . . . . . . . OFF

Quick TurnBatteries . . . . . . . . . . . . . . . . . . . . . . . . . . . ON

DC Power Selector . . . . . . . . . . . . . . . . . . . . . SET

Cabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . ON

Emergency Exit Lights . . . . . . . . . . . . . . . . ARMED

Anti-Collision/Navigation Lights . . . . . . . . . . . . . . ON

Fire Warning . . . . . . . . . . . . . . . . . . . . . TESTED

No. 1 Radio . . . . . . . . . . . . . . . . . . . . . . . . . . ON

Parking Brake . . . . . FULL AFT/P BRAKE LIGHT OUT

No. 2 Booster Pump . . . . . . . . . . . . . . . . . . . . . ON

No. 2 Engine . . . . STARTED/PARAMETERS CHECKED

No. 2 Generator . . . . . . . . . . . . . . . . . . CHECKED

No. 1 Booster Pump . . . . . . . . . . . . . . . . . . . . . ON

No. 1 Engine . . . . STARTED/PARAMETERS CHECKED

No. 1 Generator . . . . . . . . . . . . . . . . . . CHECKED

DC Power Selector . . . . . . . . . . . . . . . . . NORMAL

Ground Power . . . . . . . . . . . . . . . . . . . REMOVED


FEEDERTANK VENTS

AIR SCOOP

ENGINE AIRINTAKE

VENTURI ANGLE-OF-ATTACK VANE

PITOTHEADS

CABINPRESSURIZATIONSTATIC PORT

STATICHEADS

FUEL NACAVENTS

ENGINEEXHAUSTS

RUDDERLOCK

3B-1



ParkingAircraft . . . . . . . . . . . . . TURNED INTO THE WIND

Lower Fuselage/Tires . . . . . . . . . . . . . . . CHECKEDCheck that nothing can damage the airframe if a tire deflates or a shockstrut is depressurized.

All Servicing Materials . . . . . . . KEPT AT A DISTANCE . . . . . . . . . . . . . . . . . . . . . . .FROM AIRCRAFT

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . OFF

Parking Brake . . . . . . . . . CHECKED/NOT ENGAGED

Opening Windows . . CHECKED/CLOSED AND LOCKED

Wheel Chocks . . . . . . . . . . . . . . . . . . INSTALLEDPlace in front and behind main landing gear wheels.

Protective Covers or Caps (Figure 3B-1 ) . . . . INSTALLEDInstall covers and caps on the following elements:� engine air intakes� engine exhausts� pitot heads� static heads� cabin pressurization static port� air scoops� wing air vents� collector tank air vents� pitch sensor.

Rudder Lock . . . . . . . . . . . . . . . . . . . INSTALLED

Gear Safety Devices (Figure 3B-1 ) . . . . . . . . INSTALLED


MooringAircraft . . . . . . . . . . . . . . . . . . . . . . . . PARKED

Main Gear Mooring Fitting Clamp . . . . . . . INSTALLEDUse TF10A 10201 fittings.

Nose Gear Towing Spool/Locking Device . . . . INSTALLEDUse TF10A 10202 fittings.

Aircraft . . . . . . . MOORED WITH RINGS (Figure 3B-2 )

Mandatory Safety Precautions . . . . . . . . . . OBSERVEDIf Snow is Anticipated:Snow Fittings . . . . . . . . . . . . . . . . . . INSTALLED

Use TF10A 07 106 fittings.

TF10A 07 107 Fitting . . . . . . . . . . . . . . TIGHTENED1,000 KG Load . . . . . . . . . . . . . . . INSTALLED ON . . . . . . . . . . . . . . . . . . .TF10A 07 107 SHACKLE

Red Flag . . . . INSTALLED ON TF10A 07 107 SHACKLE

TOWINGSPOOL

RING

FITTING CLAMP

LOCKINGDEVICE

RING

LOCKING PIN

3B-2



Towing/TaxiingFor all towing procedures, disconnect the steering control. Duringtowing operations, check that wheel chocks have been removed; anassistant must be in the cockpit to release the emergency/parkingbrake.

Use the maneuvering fork to guide the aircraft on the ground overshort distances; use the towing bar to pull or push the aircraft.

The nose wheel can rotate in the following arcs:� 55° with steering control engaged� 120° with steering control disengaged.

Towing – Hard Ground

Steering Control . . . . . . . . . . . . . . DISCONNECTED

Parking Brake . . . . . . . . . . . . CHECKED/RELEASED

Battery Switches . . . . . . . . . . . . . . . . . BAT 1/BAT 2Make sure the aircraft configuration green lights are illuminated.

Hydraulic System 1Pressure . . . . . . . . . . CHECKED/2,465 PSI (170 BAR)

If pressure is less than 2,465 PSI (170 bar), use the electropump topressurize the system.

If Maneuvering Fork Is Used:Maneuvering Fork . . . . . . . . . . . . . . . ASSEMBLED

Attach maneuvering fork to nose gear and guide aircraft with fork.

If Towing Aircraft:Tow Bar . . . . . . . . . . . . . . . . . . . . ASSEMBLED

Attach tow bar to nose gear. Tow aircraft slowly; avoid jerkingmotions.

Close Up – Hard Ground

Aircraft . . . . . . . . . . . . . . . . . . . . . . . STOPPED

Nose Wheel . . . . . . . . . . . . . . . STRAIGHT AHEAD

Wheel Chocks . . . . . . . . . . . . . . . . . . INSTALLED

Maneuvering Fork or Towing Bar . . . . . . . . REMOVED

Steering Control . . . . . . . . . . . . . . . . CONNECTED

CAUTION: For operations re-quiring nosewheel to be turnedby more than 120°, disconnectbonding braid and electrical con-nector (102). Use only the Falcon10 towing bar with torque andtensile stress shear pins designedto prevent damage to the nosegear. Never use other towing bars(Falcon 20, etc.).


Towing – Soft GroundThe aircraft can be towed from soft ground forward or rearwardusing an assembly on the main gear; generally the aircraft is towedfrom the rear.

Gear Struts . . . . . . . . . . CHECKED/DOWNLOCKED

Steering Control . . . . . . . . . . . . . . DISCONNECTED

If Towing Rearward:Nose Gear Strut Bonding Braid . . . . . . DISCONNECTEDNose Gear Strut Electrical Connector . . . DISCONNECTED

Disconnecting the electrical cable and fitting prevents damage tothese components due to nosewheel castering while towing rearward.Angular limitation for rear towing is indicated by friction of cableon wheels. For rear towing, maneuvering fork can be used to guidenose wheel.

Parking Brake . . . . . . . . . . . . . . . . . . . NEUTRALIf the wheel brakes are operational, station an assistant in the cockpitto apply brakes as necessary. If the wheel brakes are not operational,take all precautions (e.g., clearing of areas, chocks) to prevent accidents.Brake accumulator capacity permits 10 good braking applications per-formed by completely pulling out emergency/parking brake controlhandle.

Towing Apparatus . . . . . . . . . . . . . . . ASSEMBLEDAttach to tow vehicle.

Aircraft . . . . . . . . . . . . . . . . . . TOWED SLOWLYAvoid jerking motions. Tow along the longitudinal axis.

If Stress Shear Pin Breaks:Stress Shear Pin . . . . . . . . . REPLACED WITH SPARE

A stress shear pin breaks under a load of approximately 5,600 lbs(2,500 daN).

As Soon As Aircraft is on Hard Ground:Towing Apparatus . . . . . . REMOVED FROM AIRCRAFTTowing . . . . . . . . . CONTINUED/NORMAL TOWING . . . . . . . . . . . . . . . . . . . . . .PROCEDURE USED



Close Up – Soft Ground

Wheels and Bearings . . . . . . . REMOVED/INSPECTED/ . . . . . . . . . . . . . . . . . . . . . . . . . . . .SERVICED

Brake Units . . . . . REMOVED/INSPECTED/SERVICED

Landing Gear and Doors . . . . CLEANED/LUBRICATED/ . . . . . . . . . . . . . . . . . . . . . . . . . .PROTECTED

Aircraft . . . . . . . . . . . . . . . . . CLEANED OF MUD

Landing Gear . . . . . CHECKED/PENDULUM MOTION

Gear Struts . . . . . . . . . . . . . . . . . . . . . CHECKED

Landing Gear Functional Test . . . . . . . . PERFORMED


28' 1

1"S

TE

ER

ING

CO

NT

RO

LD

ISC

ON

NE

CT

ED

37' 0

"S

TEE

RIN

G C

ON

TRO

L C

ON

NE

CTE

D

13' 2"

3B-3

Towing and Taxiing Radius



TaxiingThe minimum turning radius (with steering control disengaged) isalmost within the radius of the aircraft (Figure 3B-3 ); the nosewheel is able to turn 90°.

For rotation around the main landing gear, turning radius (withsteering control disconnected) is approximately 28.7 ft (8.60 m).

For maximum nosewheel rotation, the turning radius is approxi-mately 37 ft (11 m).



The aircraft may be stored for up to eight months inside or out-side of a hanger.

StorageAircraft Configuration, General

Aircraft . . . . . . . . . . . . . . . . . . . . . ON WHEELS

Wheel Chocks (if outdoors) . . . . . . . . . . . INSTALLED/ . . . . . . . . . . . . . .FRONT AND BACK OF WHEELS

Aircraft . . . . . . . . . . . . . . . . . . . . . . . MOOREDFront mooring should be accomplished if snow conditions are present.

Tire Pressure . . . . . . . . . . . . . . . . . . . . CHECKED

Parking Brake . . . . . . . . . . . . . . . . . . . RELEASED

Wheel Positions (Figure 3B-4 ) . . . . . . . . . MARKED FOR . . . . . . . . . . . . . . . . . . . .MONTHLY ROTATION

1

34 2

3B-4

Storage andRestoring


Wheels . . . . . . . . . . . . . . . . . . . . . . . COVERED

Passenger Door Seal/Other Rubber Seals . . . . . . . POWDERED WITH TALC

All Doors . . . . . . . . . . . . . . . . . . . . . . . CLOSED

All Covers and Blanking Caps . . . . . . . . . INSTALLED

Rudder . . . . . . . . . . . . . . . . . . . . . . . . LOCKED

Window Panes/Emergency Exit . . . . . . . . . . . . TAPED

Junctions Between Components . . . . . . . . . . . . TAPEDThis prevents water seepage between different sections of slats, flaps,and boxes.

Wing Center Box . . . . . . . . . . . . . . . . . . . . . . . 0°

Visible Rod of Microswitch . . . . . . LIGHTLY GREASEDUse a low-temperature grease.

Aircraft Grounded for One Month,Aircraft Already Outdoors

All Water Drains . . . . . . . . . . . . . . . . . OPERATED

All Airspeed Indicating Systems . . . . . . . . . . . BLEED

Tank Drains . . . . . . . . . . . . . . . . . . . . OPERATEDLift aircraft nose for satisfactory draining of wing center box.

Sound-Proofing Blankets . . . CHECKED/NO MOISTURE

Stub/Drain Compartment . . . CHECKED/NO MOISTURE

Rear Cargo Compartment . . CHECKED/NO MOISTURE

Wheels . . . . . . . . . . . . . . . . . . . . . . . REMOVED

Axles and Bearings . . . . . . . . . . . . . . LUBRICATED

Wheels . . . . . . . . . . . . . . . . . . . . . . INSTALLED

Anti-Corrosion Varnish . . APPLIED ON TIRE SIDEWALLS


Aircraft . . . . . . . . . . . . . . . . . . . . . . GROUNDED



Aircraft Cell . . . . . . . . . . . . . . . . . . . PROTECTEDCorrosion Points . . . . . . . . . . . . . . . . . TOUCH UPScrew/Huck Rivet Heads . . CLEAN/TOUCH UP VARNISHUnpainted Nuts/Pipe Ends . . . . CHECK/NO CORROSION

Pay particular attention to the wheel wells.

Electrical Connectors . . . . . . . CHECK/NO CORROSIONPay particular attention to open structures and wheel wells.

Greasing . . . . . . . . . . . . . . . . . . . . . . PERFORMDo not remove excess grease from around tecalemit grease fittings.

Track Assembly . . . . . . . . . INSPECT/CLEAN/GREASESlat Rollers . . . . . . . . . . . INSPECT/CLEAN/GREASEMicroswitches . . . . . . . . . INSPECT/CLEAN/GREASECover Mechanisms . . . . . . . INSPECT/CLEAN/GREASEFlap Rollers and Rails . . . . . INSPECT/CLEAN/GREASEUplatch Olives/Catches/Bolts/Latches/Visible Pin . . . . . . . . . . . . . . GREASEAuxiliary Elevator Servo-Control AFS Pin . . . . . GREASEUnpainted Aircraft . . . . . . PAINT WITH CELOPROTECT

Regardless of storage location, the entire cell skin including theleading edge should be covered with a coat of average green CELO-PROTECT if unpainted.

Fuel System . . . . . . . . . . . . . . . . . . . . PREPAREDSystem . . . . . . . . . . . . . . . . INSPECT FOR LEAKSFuel Tanks . . . . . . . . . . . . . . . . . . . . . . . DRAINWing Fuel Tanks . . . . . . . . . . FILL TO CAPACITY OR . . . . . . . . . . . . . . . . . . . . . . .HALF-CAPACITY

Collector Tanks . . . . . . . . . . . . . FILL TO TRANSFER . . . . . . . . . . . . . . . . . . . .REGULATION LEVEL

Hydraulic System/Components . . . . . . . . . PREPAREDGas Chromatography Analysisof Hydraulic Fluid . . . . . . . . . . . . . . . . . PERFORM

If analysis shows contamination of hydraulic fluid, replace the fluid.

Hydraulic Reservoirs . . . . . . . . . . . . . . . . . . . FILLSlide Rods of Dampers/Lifting Jacks/Door Mechanism/Strut . . . . . . . . . . CLEAN/MOISTEN

Use a dry cloth to clean the rods; lightly apply hydraulic fluid tothe edges of each.

Damper Pressure . . . . . . . . . . . . . . . . . . . CHECK


Equipment and Furnishings . . . . . . . . . . . PREPAREDAll Batteries . . . . . . . . . . . . . . . . REMOVE/STORE

This includes the DC generator starter, standby horizon, and emer-gency lighting batteries. Store batteries according to manufacturerdirections. Clean battery areas.

Window Silicagel . . . . . . . . . . . CHECK CONDITIONRear Glass PanelDemisting System . . . . . . . . . . . CHECK CONDITIONSilicagel Bags . . . . . . . . . . . . . . . . . . . . . HANG

Hang in radio and electrical racks of cockpit and nosecone. Alsohang bags in the passenger compartment.

Oxygen Cylinder . . . . . . . . . . . . LOWER TO 100 PSIBrake Fluid Level . . . . . . . . . . . . . . . . . . . CHECKFreon Compressor Oil Level . . . . . . . . . . . . . CHECK

Periodic Checks of Storage Conditions, Each Month

Nosecone Condition . . . . . . CHECKED/NO MOISTURE

Visible Slide Rods of Dampers . . CLEANED/LUBRICATEDUse a dry cloth to clean the rods; lightly apply hydraulic fluid.

Outer Condition of Dampers . . . . . . . . . . . CHECKED

Visible Slide Rods of Jacks . . . CLEANED/LUBRICATEDUse a dry cloth to clean the rods; lightly apply hydraulic fluid.

Aircraft Drains . . . . . . . . . . . . . . . . . . OPERATED

Tank Drains . . . . . . . . . . . . . . . . . . . . OPERATED

Silicagel . . . . . . . . . . . . . . . . CHANGED (IF BLUE)Silicagel is pink when new and blue in humidity.

Skin and Wheel Wells . . . . . . . . . . . . . . INSPECTED/ . . . . . . . . . . . .CORROSION POINTS ELIMINATED

Engine Storage Conditions . . . . . . . . . . . . CHECKEDSee manufacturer’s directions for engine storage.

Aircraft Wheels . . . . . . . . . . . . . ROTATED 1/4 TURN


Fuel/Hydraulic Leaks . . . . CHECKED/NONE PRESENT

Freon System . . . . . . . . . STARTED/RAN 5 MINUTESCheck reservoir indicator light for absence of bubbles; clean collectionpan if necessary.

NOTE: Bags of Silicagel mustnot contact sheet metal plate.

NOTE: Refer to MaintenanceManual chapter 5-20 forequipment that have periodicmaintenance procedures.



Periodic Checks of Storage Conditions,Every Two Months – Cranking of Dampers

Charging Pressure . . . . . . . . . . . . . . . . . CHECKED

Aircraft . . . . . . . . . PLACED ON HYDRAULIC JACKSOperate very slowly so that rods move very slowly.

Visible Slide Rods of Dampers . CLEANED/LUBRICATEDUse a dry cloth to clean the rods; lightly apply hydraulic fluid.

Periodic Checks of Storage Conditions,Every Two Months – Cranking of Hydraulic Jacks

Visible Slide Rods . . . . . . . . CLEANED/LUBRICATEDUse a dry cloth to clean the rods; lightly apply hydraulic fluid.

Hydraulic Pressure . . . . . . . . . . . . . . . . . . RAISEDRaise the pressure progressively until there is no friction and there isa slow movement.

Gear/Airbrakes/Slats/Flaps . . . . . . . . . . . EXERCISED

Visible Slide Rods . . . . . . . . CLEANED/LUBRICATEDUse a dry cloth to clean the rods; lightly apply hydraulic fluid.

Periodic Checks of Storage Conditions,Every Four/Two Months (Depending on Engine Storage)

Extinguisher Cartridges . . . . . . . . . . . . . . CHECKED

Engines . . . . . . . . . . . . . . . . . . . . . . RESTOREDRudder . . . . . . . . . . . . . . . . . . . . . . . . UNLOCKServo Control Slide Rods . . . APPLY HYDRAULIC FLUIDEngine . . . . . . . . . . . . . . . . . . . . . . . . RUN UPFlight Servo Controls . . . . . . . . . . . . . . . . . CHECKPressurization . . . . . . . . . . . . . . . . . . . . . CHECKFeeder Tank Drain and Wing Fuel Samples . . . . . OBTAIN/ . . . . . . . . . . . . . . . . . . . .CHECK/NO BACTERIA


RestoringAll Parking Protective Covers . . . . . . . . . . REMOVED

Unpainted Aircraft . . . . . . CELOPROTECT REMOVED

Aircraft . . . . . . . . . . . . . . . . . PLACED ON JACKS

Nosewheel Bearings . . . . . . . CHECKED/LUBRICATED

Engines . . . . . . . . . . . . . . . . . . . . . . RESTORED

Batteries . . . . . . . CAPACITY CHECKED/INSTALLED

Hydraulic/Electric Rigs . . . . . . . . . . . . CONNECTED

Hydraulic System Tests . . . . . . . . . . . . PERFORMEDThese include tests of the gear, airbrakes, flaps, slats, and servo controls.Perform the first tests very slowly to avoid hydraulic leakage.

All Hydraulic Equipment . . . . . TIGHTNESS CHECKED/ . . . . . . . . . . . . . . . . . . . . . . . .NO CORROSION

Extinguishers . . . . . . . MAINTENANCE PERFORMED

Extinguisher Cartridges . . . . . . . . . CHECKED VALID

Gage Insulation Measurements . . . . . . . . PERFORMED

Wing Collector Tank Expansion Valves . . . . . CHECKED

Hydraulic Reservoir Expansion Valves . . . . . . CHECKED

Yaw Damper . . . . . . . . . . . . . . . . . . . . . . TESTED

Brake System Pressure/Operation . . . . . . . . CHECKED

Accumulator Pressures . . . . . . . . . . . . . . CHECKED

Slat Tracks and Rollers/Flap Rollers . . . . . . . . . . . . CLEANED/LUBRICATED

Aircraft . . . . . . . . . . . . . . . . . . . . . LUBRICATED

Airspeed Indicating System . . . . . . . . . . . . . TESTED

Radio Tests . . . . . . . . . . . . . . . . . . . PERFORMED

Electrical Systems . . . . . . . . . . . . . . . . . . . TESTED

Emergency Beacon (if installed) . . CURE DATE CHECKED

Oxygen . . . . . . . . . . . . . . . . . SERVICED/TESTEDFill the oxygen cylinder. Test the system for leakage; check maskoperation.



Silicagel Bags . . . . . . . . . . . . . . . . . . . REMOVED

Pressurization Valves . . . . . CHECKED/SATISFACTORY

Overpressure Valves . . . . . . . . . . . . . . . . CRANKED

Pressurized Refueling . . . . . . . . . . . . . PERFORMED

Freon Motor-Compressor Belt . . . . . . . . . STRETCHED

Freon Circuit . . . . . . . . . . . . . . . . . REPLENISHED

Freon Level . . . . . . . . . . . . CHECKED/CORRECTED

Freon Compressor Oil Level . . . . . . . . . . . CHECKED



Observe aircraft performance limitations computed from Section5 of the AFM. Temperature affects engine thrust, braking, takeoffdistance, and climb performance. In very dry areas, protect theaircraft from dust and sand.

Exterior InspectionPreflight Inspection . . . . . . . . . . . . . . CONDUCTED

All Protective Covers . . . . . . . . . . . . . . . REMOVED

Landing Gear Shock Struts . . . . . . . . . . CLEANED OF . . . . . . . . . . . . . . . . . . . . . . . .DUST AND SAND

Gear Doors, Position Switches,and Squat Switches . . . . . . . . . . . . . . . CONDITION/ . . . . . . . . . . . . . . . . . . . .OPERATION CHECKED

Tires and Struts . . CHECKED FOR PROPER INFLATION

Engine Inlet/Exhaust Ductsand Thrust Reversers . . CLEANED OF DUST AND SAND

During the inspection, be particularly conscious of dust and sand accu-mulations on components that are lubricated with oily or greasy lubri-cants. Be careful of personnel and equipment behind the aircraft duringengine starts.

Engine StartITT . . . . . . . . . . . . . . . . . . . . . . . MONITORED

During engine starts at high outside temperatures, engine ITT is higherthan normal but should remain within limits.

TaxiIf the airport surfaces are sandy or dust covered, avoid the exhaustwake and propwash of other aircraft.

TakeoffEnsure takeoff performance is adequate for conditions and run-way length.

Shutdown and PostflightAll Protective Covers . . . . . . . . . . . . . . . INSTALLED

Sunscreens on Glareshield andSide Windows . . . . . . . . . . . . . . . . . . . INSTALLED

This limits the sunlight entering the cockpit.

Window Shades in Cabin . . . . . . . . . . . . . LOWERED

Hot Weatherand DesertOperations



Preflight InspectionInspect areas where surface snow or frost could change or affect nor-mal system operations. In addition, perform the following supple-mental checks.

All Protection Covers . . . . . . . . . . . . . . . REMOVED

Surface . . . . . . . . . . . . . . . . . . . CHECKED FREE . . . . . . . . . . . . . . . . .OF FROST, ICE, AND SNOW

The wing leading edges, all control surfaces, fuselage, wings, and hor-izontal stabilizer must be free of ice, frost, or snow.

Engine Inlets . . . . . . . . . . . . . . . . . . CLEARED OF . . . . . . . . . . . . . . . . .INTERNAL ICE AND SNOWCheck that the inlet cowling is free of ice or snow and that the fanrotates freely.

Fuel Tank Vents . . CHECKED FREE OF ICE AND SNOWCheck the NACA vents. Remove all traces of ice or snow.

Pitot Heads and Static Ports . . . . . . . CLEARED OF ICEWater rundown resulting from snow removal may refreeze immedi-ately forward of the static ports, causing an ice buildup that results indisturbed airflow over the static ports. Erroneous static readings occureven though the static ports themselves are clear.

Landing Gear and Gear Door Linkage . . . . . . CHECKEDBe sure the landing gear and door linkages are free of impacted ice orsnow. Also, check the landing gear uplocks.

Aircraft Deicing . . . . . . . . . . . . . . . . COMPLETEDThe times of protection (i.e., the holdover times) for different de-icingfluids vary considerably. Furthermore, these times depend to a largeextent on the meteorological conditions and methods of application.When temperatures warrant, consider preheating the followingequipment:� engines (less than -35°C)� batteries (less than -15°C)� windshields (less than -15°C)� interior and avionics.Accomplish engine preheat by blowing heated air into the inlet andexhaust ducts.

Cold WeatherOperations

CAUTION: Do not spray de-icing fluid in areas where sprayor fluid may enter the engineinlets. Deicing fluid may be usedto clear these areas, providingthey are thoroughly wiped cleanbefore starting.


GPU Available

1,000 AMP Rating

For Engine Start:

Batteries . . . . . . . . . . . . . PREHEAT IF NECESSARYThermal blankets may be used.

Windshield Heat . . . . . . . . . . . . . . . . . . . NORMAL

DC Power Selector . . . . . . . . . . . . . . . EXT POWER

Auxiliary heat (if installed) . . USE TO WARM INTERIOR

GPU Not Available

Ambient Temperature -15° C to -35° C

Batteries . . . . . . . . . . . . . . . . . . . . . . . PREHEAT

DC Power Selector . . . . . . . . . . . LOW TEMP START

No. 2 Engine . . . . . . . . . . . . . . . . . . . . . . . STARTThe No. 2 engine becomes a power source; 55% N1 is required forcabin heating.

For Second Engine Start:DC Power Selector . . . . . . . . . . . . . . . . . NORMAL

NOTE: It is highly recommended that the pressurization valve ofeach hydraulic reservoir be heatedfor several minutes.



After Engine StartInstruments . . OBSERVED FOR NORMAL OPERATION

The engine instruments should indicate approximately normal short-ly after reaching idle speed.

Engine Oil Pressure . . . . . . . . . . . . . . . . CHECKEDPower settings above idle are not recommended until engine oil tem-perature is 30°C or warmer. During cold starts, oil pressure transientsup to 55 PSI for three minutes are allowed.

Flight Controls . . . . . . . . . . . . . . . . . . . CHECKEDIn extremely cold weather, the flight controls may be sluggish untilthe hydraulic fluid in the reservoir and lines warms. Operate the controlsthroughout their full range until the sluggishness disappears.Verify that the stabilizer, slats, and flaps operate normally. Attemperatures of -30°C, the FLAP MOTOR CB may open during flapretraction. As the aircraft warms, this situation should remedy itself.

Anti-Ice . . . . . . . . . . . . . . . . . . . . AS REQUIREDDuring operation from snow-covered runways, turn on engine anti-ice during taxi and takeoff. Precede takeoff by a static engine run-upto as high a power level as practical to ensure stable engine operationprior to brake release.If severe icing conditions are present, turn on engine anti-icingimmediately after engine start. During prolonged ground operation,periodic engine run-up may be performed to reduce the possibility ofice buildup. However, be aware that ingestion of ice or hardened snowparticles can cause engine damage.

Slats/Flaps . . . . . . . . . . . . . . . . . . . . RETRACTEDThe slats and flaps should be retracted during taxi. Extend slats and/orflaps when the aircraft is on the runway and in takeoff position.

Brakes . . . . . . . . . . . . . . . . . . . . USE MINIMIZEDUsing brakes on surfaces covered with snow or slush can result inmoisture freezing in the brake assemblies that in turn can cause wheeljamming on the subsequent landing. This problem may be lessenedby using the minimum practical braking force and by cycling the gearafter takeoff.


Before TakeoffContaminated Runway Operations . . . . . . . COMPUTED

Service Letter 34 presents formulas for computing recommendedequivalent water depths, takeoff distances, and techniques. See FlightPlanning chapter for details.

Slats/Flaps . . . . . . . . . . . . . . . . SET FOR TAKEOFFJust prior to takeoff, extend the slats and/or flaps as planned.

Engine Operation During Acceleration . . . . . CHECKEDEnsure proper engine operation during takeoff roll. Abort takeoff ifmalfunctions or engine instrument variations are observed.

After TakeoffLanding Gear . . . . . . . . . CYCLED BELOW 190 KIAS

After takeoff in snow or slush, cycle the gear three times to removesnow and slush accumulations on the gear and brakes.

Slats . . . . . . . . . . . . . . . CYCLED BELOW 190 KIAS

If Slats Fail to Fully Retract (red slat-in-transit light on):Wing Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . ONSlats . . . . . . . . . . . . . . . . CYCLE TO REMOVE ICE

If slats fail to fully retract, land as soon as possible.

CAUTION: Flight test data hasshown that engine flameout orcompressor stalls can occurbetween 50 and 60 kts on takeoffin water or on snow.



Cruise/DescentSatisfactory operation of the anti-icing system requires that theengine speed (N1) be not less than shown in Table 3B-B .

Normal flight operations call for the airbrakes to be retracted at500 ft AGL on the approach.

LandingContaminated Runway Operations . . . . . . . COMPUTED

Service Letter 34 presents formulas for computing recommendedequivalent water depths, landing distances, and techniques. See FlightPlanning chapter for details.

Taxi-In and ParkEngine Anti-Ice . . . . . . . . . . . . . . . . AS REQUIRED

It may be necessary to leave the engine anti-ice on while taxiing in.

Wheel Chocks . . . . . . . . . . . . . . . . . . . . IN PLACE

Parking Brake . . . . . . . . . . . . . . . . . . . . . . . OFFRelease the parking brake to eliminate the possibility of the brakesfreezing.

Protective Covers . . . . . . . . . . . . . . . . . INSTALLED

Water Storage Containers . . . . . . . . . . . . . DRAINED

Toilet . . . . . . . . . . . . . . . . . . . . . . . . . DRAINED

Batteries . . . . . . . . . . . . . . . . . . . . . . REMOVEDIt is recommended that the batteries be removed and stored indoors atroom temperature.

CAUTION: If ice has alreadyformed, the ignition switchesmust be set to AIRSTART andanti-ice turned on as follows:Ignition Switches – AIRSTARTEngine 1 Anti-Ice – ON

30 Seconds Later:Engine 2 Anti-Ice – ON

30 Seconds Later:WING ANTI-ICE – ON

CAUTION: In order to main-tain a sufficiently high enginesetting in icing conditions,approach may be done in thelanding configuration with air-brakes extended to 500 ft AGL.At 500 ft AGL, retract the air-brakes and turn off the enginesynchronizer, if installed.

Phase of Flight% N1

at -30°COAT

at -20°COAT

at -10°COAT

at -5°COAT

In Flight – Min. 76 73 67 62

Recommended 78 75 69 64

Approach 69 69 69 64

One Engine 88 85 79 74

Table 3B-B; Engine Settings for Anti-Ice

Falcon 10/100 For SimuFlite training only 3C-1April 2000

SimuFlite strongly supports the premise that the disciplineduse of well-developed Standard Operating Procedures (SOP)is central to safe, professional aircraft operations, especiallyin multi-crew, complex, or high performance aircraft.

If your flight department has an SOP, we encourage you to useit during your training. If your flight department does notalready have one, we welcome your use of the SimuFlite SOP.

Corporate pilots carefully developed this SOP. A product oftheir experience, it is the way SimuFlite conducts its flightoperations.

The procedures described herein are specific to the Falcon 10and apply to specified phases of flight. The flight crew memberdesignated for each step accomplishes it as indicated.

StandardOperatingProcedures

Chapter 3C

3C-2 For SimuFlite training only Falcon 10/100April 2000


Standard Operating Procedures

General Information . . . . . . . . . . . . . . . . . 3C-5

Definitions . . . . . . . . . . . . . . . . . . . . . . 3C-5

Flow Patterns . . . . . . . . . . . . . . . . . . . . . 3C-5

Checklists . . . . . . . . . . . . . . . . . . . . . . . 3C-5

Omission of Checklists . . . . . . . . . . . . . . . 3C-5

Challenge/No Response . . . . . . . . . . . . . . 3C-5

Abnormal/Emergency Procedures . . . . . . . . . . 3C-6

Time Critical Situations . . . . . . . . . . . . . . . 3C-6

Aborted Takeoffs . . . . . . . . . . . . . . . . . . 3C-6

Critical Malfunctions in Flight . . . . . . . . . . . . 3C-6

Non-Critical Malfunctions in Flight . . . . . . . . . 3C-7

Radio Tuning and Communication . . . . . . . . . . 3C-7

Altitude Assignment . . . . . . . . . . . . . . . . . 3C-7

Advising of Aircraft Configuration Change . . . . . . 3C-7

Transitioning from Instrumentto Visual Conditions . . . . . . . . . . . . . . . . . 3C-8

Pre-Departure Briefings . . . . . . . . . . . . . . . 3C-8

Phase of Flight SOP . . . . . . . . . . . . . . . . . 3C-9

Holding Short . . . . . . . . . . . . . . . . . . . . . 3C-9

Takeoff Roll . . . . . . . . . . . . . . . . . . . . . 3C-10

Climb . . . . . . . . . . . . . . . . . . . . . . . . 3C-11

Cruise . . . . . . . . . . . . . . . . . . . . . . . . 3C-13

Descent . . . . . . . . . . . . . . . . . . . . . . . 3C-14

Precision Approach . . . . . . . . . . . . . . . . . 3C-16

Precision Missed Approach . . . . . . . . . . . . . 3C-21

Precision Approach Deviations . . . . . . . . . . . 3C-22

Non-Precision Approach . . . . . . . . . . . . . . 3C-23

Non-Precision Missed Approach . . . . . . . . . . 3C-28

Non-Precision Approach Deviations . . . . . . . . 3C-30

Visual Traffic Patterns . . . . . . . . . . . . . . . . 3C-31

Landing . . . . . . . . . . . . . . . . . . . . . . . 3C-33

Table ofContents



GeneralInformation

DefinitionsLH/RH – Pilot Station. Designationof seat position for accomplishing agiven task because of proximity to therespective control/indicator. Regardlessof PF or PNF role, the pilot in that seatperforms tasks and responds to check-list challenges accordingly.

PF – Pilot Flying. The pilot respon-sible for controlling the flight of theaircraft.

PIC – Pilot-in-Command. The pilotresponsible for the operation and safe-ty of an aircraft during flight time.

PNF – Pilot Not Flying. The pilot whois not controlling the flight of theaircraft.

Flow PatternsFlow patterns are an integral part ofthe SOP. Accomplish the cockpit setupfor each phase of flight with a flowpattern, then refer to the checklist toverify the setup. Use normal check-lists as “done lists” instead of “dolists.”

Flow patterns are disciplined proce-dures; they require pilots who under-stand the aircraft systems/controls andwho methodically accomplish the flowpattern.

A standardized flow pattern for thecockpit setup before starting enginesappears in the Expanded Normalschapter.

ChecklistsUse a challenge-response method toexecute any checklist. After the PF ini-tiates the checklist, the PNF challengesby reading the checklist item aloud.The PF is responsible for verifying thatthe items designated as PF or his seatposition (i.e., LH or RH) are accom-plished and for responding orally to

the challenge. Items designated on thechecklist as PNF or by his seat posi-tion are the PNF’s responsibility. ThePNF confirms the accomplishment ofthe item, then responds orally to hisown challenge. In all cases, theresponse by either pilot is confirmedby the other and any disagreement isresolved prior to continuing thechecklist.

After the completion of any checklist,the PNF states “ checklist is com-plete.” This allows the PF to maintainsituational awareness during checklistphases and prompts the PF to contin-ue to the next checklist, if required.

Effective checklists are pertinent andconcise. Use them the way they arewritten: verbatim, smartly, and pro-fessionally.

Omission of ChecklistsWhile the PF is responsible for initi-ating checklists, the PNF should askthe PF whether a checklist should bestarted if, in his opinion, a checklist isoverlooked. As an expression of goodflight deck management, such prompt-ing is appropriate for any flight situa-tion: training, operations, or check-rides.

Challenge/No ResponseIf the PNF observes and challenges aflight deviation or critical situation,the PF should respond immediately. Ifthe PF does not respond by oral com-munication or action, the PNF mustissue a second challenge that is loudand clear. If the PF does not respondafter the second challenge, the PNFmust ensure the safety of the aircraft.The PNF must announce that he isassuming control and then take thenecessary actions to return the aircraftto a safe operating envelope.

3C-6 For SimuFlite training only Falcon 10/100March 2001

Abnormal/EmergencyProceduresWhen any crewmember recognizes anabnormal or emergency condition, thePIC designates who controls the air-craft, who performs the tasks, and anyitems to be monitored. Following thesedesignations, the PIC calls for theappropriate checklist. The crew-member designated on the checklistaccomplishes the checklist items withthe appropriate challenge/response.

The pilot designated to fly the aircraft(i.e., PF) does not perform tasks thatcompromise this primary responsibil-ity, regardless of whether he uses theautopilot or flies manually.

Both pilots must be able to respond toan emergency situation that requiresimmediate corrective action withoutreference to a checklist. The elementsof an emergency procedure that mustbe performed without reference to theappropriate checklist are called mem-ory or recall items. Accomplish allother abnormal and emergency proce-dures while referring to the printedchecklist.

Accomplishing abnormal and emer-gency checklists differs from accom-plishing normal procedure checklistsin that the pilot reading the checkliststates both the challenge and theresponse when challenging each item.

When a checklist procedure calls forthe movement or manipulation ofcontrols or switches critical to safety offlight (e.g., throttles, engine fireswitches, fire bottle discharge switch-es), the pilot performing the actionobtains verification from the other pilotthat he is moving the correct controlor switch prior to initiating the action.

Any checklist action pertaining to aspecific control, switch, or equipmentthat is duplicated in the cockpit is read

to include its relative position and theaction required (e.g., “Left Throttle –OFF; Left Boost Pump – NORMAL”).

Time Critical SituationsWhen the aircraft, passengers, and/orcrew are in jeopardy, remember threethings.� FLY THE AIRCRAFT – Maintainaircraft control.� RECOGNIZE CHALLENGE –Analyze the situation.� RESPOND – Take appropriateaction.

Aborted TakeoffsThe aborted takeoff procedure is a pre-planned maneuver; both crewmembersmust be aware of and briefed on thetypes of malfunctions that mandate anabort. Assuming the crew trains to afirmly established SOP, either crew-member may call for an abort.

The PF normally commands and exe-cutes the takeoff abort for directionalcontrol problems or catastrophic mal-functions. Additionally, any indicationof the following malfunctions prior toV1 is cause for an abort:� engine failure� engine fire� thrust reverser deployment.In addition to the above, the PF usuallyexecutes an abort prior to 80 KIAS forany abnormality observed.

When the PNF calls an abort, the PFannounces “Abort.” or “Continue.”and executes the appropriate pro-cedure.

Critical Malfunctionsin FlightIn flight, the observing crewmemberpositively announces a malfunction.As time permits, the other crewmem-ber makes every effort to confirm/identify the malfunction before ini-tiating any emergency action.

NOTE: “Control” meansresponsible for flight control ofthe aircraft, whether manual orautomatic.

Falcon 10/100 For SimuFlite training only 3C-7March 2001


If the PNF is the first to observe anyindication of a critical failure, heannounces it and simultaneously iden-tifies the malfunction to the PF bypointing to the indicator/annunciator.

After verifying the malfunction, thePF announces his decision and com-mands accomplishment of any check-list memory items. The PF monitorsthe PNF during the accomplishmentof those tasks assigned to him.

Non-Critical Malfunctionsin FlightProcedures for recognizing and veri-fying a non-critical malfunction orimpending malfunction are the sameas those used for time critical situa-tions: use positive oral and graphiccommunication to identify and directthe proper response. Time, however,is not as critical and allows a moredeliberate response to the malfunction.Always use the appropriate checklist toaccomplish the corrective action.

Radio Tuning andCommunicationThe PNF accomplishes navigation andcommunication radio tuning, identifi-cation, and ground communication.For navigation radios, the PNF tunesand identifies all navigation aids.Before tuning the PF’s radios, heannounces the NAVAID to be set. Intuning the primary NAVAID, the PNFcoordinates with the PF to ensure prop-er selection sequencing with theautopilot mode. After tuning andidentifying the PF’s NAVAID, the PNF

announces “(Facility) tuned andidentified.”

Monitor NDB audio output anytimethe NDB is in use as the NAVAID. Usethe marker beacon audio as backup tovisual annunciation for marker pas-sage confirmation.

In tuning the VHF radios for ATCcommunication, the PNF places thenewly assigned frequency in the headnot in use (i.e., preselected) at the timeof receipt. After contact on the newfrequency, the PNF retains the previ-ously assigned frequency for a rea-sonable time period.

Altitude AssignmentThe PNF sets the assigned altitude inthe altitude alerter and points to thealerter while orally repeating thealtitude. The PNF continues to pointto the altitude alerter until the PFconfirms the altitude assignment andalerter setting.

Advising of AircraftConfiguration ChangeIf the PF is about to make an aircraftcontrol or configuration change, healerts the PNF to the forthcomingchange (e.g., gear, speedbrake, and flapselections). If time permits, he alsoannounces any abrupt flight pathchanges so there is always mutualunderstanding of the intended flightpath.

Time permitting, a PA announcementto the passengers precedes maneuversinvolving unusual deck or roll angles.

NOTE: The acronym AWAREstands for the following:� aircraft status� weather� airport information� route of flight� extra.


Transitioning from Instrument toVisual ConditionsIf visual meteorological conditions(VMC) are encountered during aninstrument approach, the PNF nor-mally continues to make callouts forthe instrument approach being con-ducted. However, the PF may requesta changeover to visual traffic patterncallouts.

Pre-Departure BriefingsThe PIC should conduct a pre-depar-ture briefing prior to each flight toaddress potential problems, weatherdelays, safety considerations, and oper-ational issues. Pre-departure briefingsshould include all crewmembers toenhance team-building and set the tonefor the flight. The briefing may be for-mal or informal, but should includesome standard items. The acronymAWARE works well to ensure nopoints are missed. This is also anopportunity to brief any takeoff ordeparture deviations from the SOP dueto weather or runway conditions.


Falcon 10/100 3C-9August 2010

Developed for Training Purposes

Phase of Flight SOP

Holding ShortPF� PPF�

Ccaal:� “Before Takeoff checklist.”cCctOPl:� Complete Before

Takeoff checklist.

Ccaal:� “Before Takeoff checklist complete.”

Takeoff Briefing

cCctOPl:� Brief the following:

� Assigned Runway for Takeoff

� Initial Heading/Course

� Initial Altitude

� Airspeed Limit (If Applicable)

� Clearance Limit

� Emergency Return Plan

� SOP Deviations

Consider the following:

� Impaired Runway Conditions

� Weather

� Obstacle Clearance

� Instrument Departures Procedures

� Abort

Cleared for cakeoff

cCctOPl:� Confirm Assigned Runway for Takeoff and Check Heading Indicator Agreement

Ccaal:� “Assigned Runway Confirmed, Heading Checked”

Ccaal:� “Line-Up checklist”

cCctOPl:� Confirm assigned runway for takeoff and check heading indicator agreement

Ccaal:� “Assigned runway confirmed, heading checked”

cCctOPl:� Complete Line-Up checklist

Ccaal:� “Line-Up checklist Complete

Falcon 10/1003C-10March 2001

Developed for Training Purposes

cakeoff RollPF� PPF�

Setting cakeoff Power

Ccaal:� “_____ set.”

ct 60 KCcS

cCctOPl:� Move left hand to yoke.

Ccaal:� “My yoke.”

Ccaal:� “Airspeed alive.” At 80 KIAS,

Ccaal:� “80 knots crosscheck.”

Ccaal:� “Your yoke.”

ct V1

cCctOPl:� Move hand from throttles to yoke.

Ccaal:� “V1.”

ct VR

cCctOPl:� Rotate to 16° pitch attitude or charted pitch attitude for takeoff.

Ccaal:� “Rotate.”



PF PNF

Climb

At Positive Rate of Climb

Only after PNF’s call,

CALL “Gear – UP.”

CALL “Positive rate.”

CALL “Gear selected UP.”

When gear indicates UP,“Gear indicates UP.”

ACTION Immediatelyaccomplish attitudecorrelation check.� PF’s and PNF’s ADIdisplays agree.

� Pitch and bankangles are acceptable.

CALL “Attitudes check.” Or, if a fault exists,give a concise statement of the discrepancy.

After Gear Retraction

At VFR and 400 Ft Above Airport Surface (Minimum)

CALL “Clean the wing.”

CALL “Clean wing speed.”

CALL “Slats selected.”

When flaps indicate 0°,“Slats indicated.”“Clean wing selected.”

When clean wing is indicated,“Clean wing indicated.”


PF PNF

Climb (cont.)

At Transition Altitude

CALL “29.92 set.” CALL “29.92 set.”

CALL “ (altitude) for(altitude).”

(e.g., “9,000 for10,000.”)


(e.g., “9,000 for10,000.”)

At 1,000 Ft Below Assigned Altitude

CALL “Climb checklist.”ACTION Complete Climb

checklist.

CALL “Climb Checklistcomplete.”

At 15,000 FT (Minimum) Above Airport Surface andWorkload Permitting

CALL “Climb power.”CALL “Climb power set.”

At VENR (Minimum)



PF PNF

Cruise

CALL “Cruise checklist.”ACTION Complete Cruise

checklist.

CALL “Cruise checklist complete.”

CALL “Correcting.”

CALL “Altitude.”


CALL “Course.”

Altitude Deviation in Excess of 100 Ft

Course Deviation in Excess of One Dot


PF PNF

Descent

CALL “Descent checklist.”ACTION Complete Descent

checklist.

CALL “Descent checklistcomplete.”


(e.g., “10,000 for9,000.”)


(e.g., “10,000 for9,000.”)

CALL “Check.Speed 250 kts.”

CALL “10,000 ft.”

Maintain sterile cockpit below 10,000 ft above airport surface.

At 10,000 Ft

At Transition Level

At 1,000 Ft Above Assigned Altitude

CALL “Altimeter set .” CALL “Altimeter set .”



PF PNF

Descent (cont.)

At Appropriate Workload Time

Review the following:� approach to be executed� field elevation� appropriate minimum sector altitude(s)� inbound leg to FAF, procedure turn direction and altitude� final approach course heading and intercept altitude� timing required� DA/MDA� MAP (non-precision)� VDP� special procedures (DME step-down, arc, etc.)� type of approach lights in use (and radio keying procedures, if required)

� missed approach procedures� runway conditions information

REVIEW REVIEW

Accomplish as many checklist items as possible. The Approachchecklist must be completed prior to the initial approach fix.

ACTION Brief the following:� configuration� approach speed� minimum safe altitude� approach course� FAF altitude� DA/MDA altitude� field elevation� VDP� missed approach

– heading

– altitude

– intentions� abnormal implications.


PF PNF

Precision Approach

CALL “Slats.”

CALL “Flaps – 15.”“Approach checklist.”


When slats light green,“Slats indicated.”

CALL “Flaps selected 15.”

When flaps indicate 15°,“Flaps indicate 15.”

ACTION Complete Approachchecklist.

CALL “Approach checklistcomplete.”

CALL “Localizer/course alive.”


CALL “Glideslope alive.”

CALL “Gear – DOWN.”“Before Landingchecklist.”

CALL “Glideslope alive.”

CALL Gear selectedDOWN.

When gear indicates DOWN,“Gear indicatesDOWN.”

ACTION Complete BeforeLanding checklistexcept for full flapsand autopilot.

Prior to Initial Approach Fix

At Initial Convergence of Course Deviation Bar

At Initial Downward Movement of Glideslope Raw Data Indicator



CALL “Localizer captured.” CALL “Localizer captured.”

When Annunciators Indicate Localizer Capture

PF PNF

Precision Approach (cont.)

At One Dot From Glideslope Intercept

CALL “Flaps – 30.”CALL “One dot to go.”



CALL “Glideslope captured.”



CALL “Glideslope captured.”

CALL “Flaps – 52.”

When Annunciator Indicates Glideslope Capture

If the VOR on the PNF’s side is used for crosschecks on the inter-mediate segment, the PNF’s localizer and glideslope status calls areaccomplished at the time the PNF changes to the ILS frequency. Thisshould be no later than at completion of the FAF crosscheck, ifrequired. The PNF should tune and identify his NAV radios to thespecific approach and monitor.


PF PNF


CALL “1,000 ft tominimums.”

CALL “Check.”

At 1,000 Ft Above DA(H)

CALL “Outer marker.” or“Final fix.”

ACTION � Start timing.� Visually crosscheckthat both altimetersagree with crossing altitude.

� Set missedapproach altitudein altitude alerter.

� Check PF and PNF instruments.

� Call FAF inbound.

CALL “Outer marker.” or“Final fix.”“Altitude checks.”

At FAF



PF PNF


CALL “Check.”

CALL “200 ft to minimums.”

CALL “Check.”


At 200 Ft Above DA(H)

CALL “Check.”




Approach Window� Within one dot deflection, both LOC and GS� IVSI less than 1,000 fpm� IAS within VAP ± 10 kts (no less than VREF)� No flight instrument flags with the landing runway or visual references not in sight

� Landing configuration

When within 500 ft above touchdown, the aircraft must be within the“approach window.” If the aircraft is not within this “window,” a missedapproach must be executed.


PF PNF


ACTION Announce intentions.

CALL “Going visual. Land,”or “Missed approach.”

CALL “Minimums.Runway (or visualreference) o’clock.”

ACTION As PF goes visual,PNF transitions toinstruments.

CALL “Going visual. Land,”or “Missed approach.”

CALL “Runway(or visual reference)

o’clock.”

ACTION As PF goes visual,PNF transitions toinstruments.

At Point Where PNF Sights Runway or Visual References

At DA(H)



PF PNF

Precision Missed Approach

At DA(H)

CALL “Missed approach.”

ACTION Apply power firmly andpositively. Activate go-around mode andinitially rotate the noseto the flight director go-around attitude.


CALL “Minimums.Missed approach.”

ACTION Assist PF in settingpower for go-around.


When flaps indicate 30°,“Flaps indicate 30.”“Flaps selected 15.”


CALL “Gear – UP.”

CALL “Positive rate.”

CALL Gear selected UP.


ACTION Announce headingand altitude formissed approach.




CALL “VREF.” or“VREF minus (knots below VREF).”


CALL “Sink (amount)hundred and(increasing, holding,decreasing).”

At or Below V REF

Rate of Descent Exceeds 1,000 FPM

PF PNF

Precision Approach Deviations


CALL “One half dot (high,low) and (increasing,holding, decreasing).”


CALL “One half dot (right,left) and (increasing,holding, decreasing).”


CALL “Speed plus orminus and(increasing, holding,decreasing).”

± One Half Dot – Glideslope

± One Half Dot – Localizer

VAP ±



PF PNF

Non-Precision Approach

CALL “Slats.”










CALL “Localizer/coursecaptured.”

CALL “Localizer/coursecaptured.”

Prior to Initial Approach Fix

At Initial Convergence of Course Deviation Bar

When Annunciators Indicate Course Capture


CALL “Gear – DOWN.”“Before Landing checklist.”


CALL “One mile or 30seconds from FAF.”

CALL “Gear selectedDOWN.”

When gear indicates DOWN, “Gear indicatesDOWN.”

ACTION Complete BeforeLanding checklistexcept for full flapsand autopilot.



CALL “Altimeters check.”

Prior to FAF

PF PNF

Non-Precision Approach (cont.)



PF PNF


CALL “1,000 ft tominimums.”

CALL “Check.”

At 1,000 Ft Above MDA



ACTION � Start timing.� Visually crosscheckthat both altimetersagree.

� Set MDA (ornearest 100 ftabove) in altitudealerter.

� Check PF and PNF instruments.

� Call FAF inbound.

At FAF


PF PNF


CALL “Check.”


CALL “Check.”


At 200 Ft Above MDA

CALL “Check.”



At 100 Ft Above MDA

Approach Window� Within one dot CDI deflection or 5° bearing� IVSI less than 1,000 fpm� IAS within VAP ± 10 kts (no less than VREF)� No flight instrument flags with the landing runway or visual references not in sight

� Landing configuration, except for full flaps

When within 500 ft above touchdown, the aircraft must be within the“approach window.” If the aircraft is not within this “window,” a missedapproach must be executed.

CALL “Check.”

CALL “Minimums. .”(time) to go.” orMinimums. .(distance) to go.”

At MDA



PF PNF


CALL “Going visual. Land.”or “Missed approach.”

CALL “Runway (or visualreference) o’clock.”

At Point Where PNF Sights Runway or Visual References


PF PNF

Non-Precision Missed Approach

At MAP

CALL “Missed approach.”

ACTION Apply power firmly andpositively. Activate go-around mode andinitially rotate the noseto the flight director go-around attitude.


CALL “Missed approachpoint. Missedapproach.”

ACTION Assist PF in settingpower for go-around.



CALL “Gear – UP.”CALL “Positive rate.”

CALL Gear selected UP.


ACTION Announce headingand altitude formissed approach.




PF PNF

Non-Precision Missed Approach (cont.)

At VREF + 35 and 400 Ft Above Airport Surface (Minimum)

CALL “Clean the wing.”

CALL “Clean wing speed.”

ACTION Slats selected.”

When slats indicate 0°,“Slats indicated.”“Clean wing selected.”

When clean wing indicated,“Wing indicates clean.”

CALL “Climb checklist.”

ACTION Complete Climbchecklist.

CALL “Climb checklistcomplete.”

At 1,500 Ft (Minimum) Above Airport Surface andWorkload Permitting



CALL “VREF.” or“VREF minus (knots below VREF).”


CALL “Sink (amount)hundred and(increasing, holding,decreasing).”

At or Below V REF

Rate of Descent Exceeds 1,000 FPM

PF PNF

Non-Precision Approach Deviations


CALL “One dot (right, left)and (increasing,holding, decreasing).”


CALL “ degrees offcourse) (right, left)and (increasing,holding, decreasing).”


CALL “Speed plus orminus and(increasing, holding,decreasing).”

± One Dot – Localizer/VOR

± 5° At or Beyond Midpoint for NDB Approach

VAP ±



CALL “Slats.”








CALL “Gear – DOWN.Before Landingchecklist.”


CALL “Gear selectedDOWN.”

When gear indicates DOWN,“Gear indicatesDOWN.”



ACTION Complete BeforeLanding checklistexcept for full flaps.

Before Pattern Entry/Downwind (1,500 Ft Above Airport Surface)

Downwind

PF PNF

Visual Traffic Patterns


PF PNF

Visual Traffic Patterns (cont.)

CALL “Check.”

CALL “1,000 AGL.”

CALL “Check.”

CALL “500 AGL.”

CALL “Check.”CALL “200 AGL.”

At 1,000 Ft Above Airport Surface

At 500 Ft Above Airport Surface

At 200 Ft Above Airport Surface



PF PNF

Landing

CALL “Going visual. Land.Flaps – 52.”

ACTION Push autopilotdisconnect switch.

CALL “Autopilot off.”



ACTION Continue with:� speed check� vertical speed check� callouts� gear downverification

� flap verification.

CALL “Final gear and flapsrecheck.”“Before Landingchecklist complete.”

At Point on Approach When PF Sights Runway or VisualReference (Landing Assured)

CALL “100 ft.”

At 100 Ft Above Touchdown

CALL “50 ft.”

At 50 Ft Above Touchdown

ACTION Extend airbrakes.

CALL “Airbrakes extended.”

At Touchdown

CALL “Extend airbrakes.”


CALL “Two green lights.”

At Thrust Reverser Deployment

CALL “80 kts.”

At 80 Kts

PF PNF

Landing (cont.)

CALL “60 kts.”

At Thrust Reverser Idle Speed (60 KIAS)

Falcon 10/100 For SimuFlite training only 3D-1April 2000

Maneuvers

Chapter 3D

This chapter includes a written description of various maneuversand techniques during normal operation and one-engineoperation.

3D-2 For SimuFlite training only Falcon 10/100April 2000


Maneuvers

Table ofContents

Normal Operation . . . . . . . . . . . . . . . . . . 3D-7

Taxiing . . . . . . . . . . . . . . . . . . . . . . . . 3D-7

Before Takeoff . . . . . . . . . . . . . . . . . . . . 3D-8

Takeoff (General) . . . . . . . . . . . . . . . . . . . 3D-8

Normal Standing Takeoff . . . . . . . . . . . . . . 3D-8

Rolling Takeoff . . . . . . . . . . . . . . . . . . . 3D-8

Crosswind Takeoff . . . . . . . . . . . . . . . . . 3D-8

Reduced Thrust Takeoff . . . . . . . . . . . . . . 3D-9

Takeoff Rotation . . . . . . . . . . . . . . . . . . . 3D-9

Rejected Takeoff . . . . . . . . . . . . . . . . . . 3D-9

Initial Climbout . . . . . . . . . . . . . . . . . . . . 3D-9

Noise Abatement Climbout . . . . . . . . . . . . 3D-10

Climb . . . . . . . . . . . . . . . . . . . . . . . . 3D-10

Cruise . . . . . . . . . . . . . . . . . . . . . . . . 3D-10

Thrust Setting . . . . . . . . . . . . . . . . . . . 3D-10

Cabin Temperature . . . . . . . . . . . . . . . . 3D-10

Turbulent Air Penetration . . . . . . . . . . . . . . 3D-10

Operation in Icing Conditions . . . . . . . . . . . . 3D-11

Inflight Procedures . . . . . . . . . . . . . . . . . 3D-11

Airbrake Deployment . . . . . . . . . . . . . . . 3D-11

Change of Airspeed . . . . . . . . . . . . . . . . 3D-11

Steep Turns . . . . . . . . . . . . . . . . . . . . . 3D-11

Stall Recognition and Recovery . . . . . . . . . . . 3D-12

Approach to Stall . . . . . . . . . . . . . . . . . 3D-12

Clean Configuration – Flaps and Gear Up . . . . 3D-12

Takeoff Configuration – Slats + 15°Flaps and Gear Down . . . . . . . . . . . . . . . 3D-12

Landing Configuration – Full Flapsand Gear Down . . . . . . . . . . . . . . . . . . 3D-12


Unusual Attitudes . . . . . . . . . . . . . . . . . 3D-13

Recovery from Nose-High Attitude . . . . . . . . 3D-13

Recovery from Nose-Low Attitude . . . . . . . . . 3D-13

Instrument Procedures . . . . . . . . . . . . . . . 3D-13

Holding . . . . . . . . . . . . . . . . . . . . . . 3D-13

Flight Director . . . . . . . . . . . . . . . . . . . 3D-13

Instrument Approach Considerations . . . . . . . 3D-14

Additional Instrument Systems . . . . . . . . . . 3D-15

Normal Descent . . . . . . . . . . . . . . . . . . . 3D-15

Emergency Descent . . . . . . . . . . . . . . . . . 3D-15

VFR Traffic Pattern . . . . . . . . . . . . . . . . . 3D-16

Approaches . . . . . . . . . . . . . . . . . . . . . 3D-16

Checklist and Configuration . . . . . . . . . . . . 3D-16

Power Settings . . . . . . . . . . . . . . . . . . 3D-17

Typical Precision Approach . . . . . . . . . . . . 3D-17

Typical Non-Precision Approach . . . . . . . . . 3D-17

Slats-Only/Clean Wing Approach and Landing . . 3D-17

Go-Around/Missed Approach/Balked Landing . . . 3D-18

Go-Around Procedure . . . . . . . . . . . . . . . 3D-18

After a Missed Approach –Proceeding for Another Approach . . . . . . . . . . 3D-18

After a Missed Approach – Departing Area . . . . . 3D-18

Circling Approach/Circling Pattern . . . . . . . . . 3D-19

Landing . . . . . . . . . . . . . . . . . . . . . . . 3D-19

Thrust Reversers . . . . . . . . . . . . . . . . . 3D-19

Crosswind . . . . . . . . . . . . . . . . . . . . . 3D-19

Touch-and-Go Landings . . . . . . . . . . . . . . 3D-19


Maneuvers

Wet/Contaminated/Very Slippery Runways . . . . . 3D-20

Definitions . . . . . . . . . . . . . . . . . . . . . 3D-20

Recommendations . . . . . . . . . . . . . . . . 3D-20

Wet Runways . . . . . . . . . . . . . . . . . . 3D-20

Contaminated Runways . . . . . . . . . . . . . 3D-20

Compacted Snow or Icy Runways . . . . . . . . 3D-21

After Landing . . . . . . . . . . . . . . . . . . . . 3D-21

One Engine Inoperative Operation . . . . . . . . 3D-23

Engine Failure After V1 – Takeoff Continued . . . . 3D-23

Single Engine Precision Approach and Landing . . 3D-23

Single Engine Go-Around/Missed Approach/Balked Landing . . . . . . . . . . . . . . . . . . . 3D-23

Profiles . . . . . . . . . . . . . . . . . . . . . . . 3D-25

Takeoff . . . . . . . . . . . . . . . . . . . . . . . 3D-27

Rejected Takeoff . . . . . . . . . . . . . . . . . . 3D-29

Engine Failure After V1 . . . . . . . . . . . . . . . 3D-31

Stalls at Altitude . . . . . . . . . . . . . . . . . . . 3D-33

Steep Turns . . . . . . . . . . . . . . . . . . . . . 3D-35

Precision Approach and Landing . . . . . . . . . . 3D-37

Single Engine Precision Approach and Landing . . 3D-39

Non-Precision Approach and Landing . . . . . . . 3D-41

Single Engine Non-Precision Approachand Landing . . . . . . . . . . . . . . . . . . . . . 3D-43

Visual Approach/Balked Landing . . . . . . . . . . 3D-45

Circling Approach . . . . . . . . . . . . . . . . . . 3D-47

Go Around/Missed Approach . . . . . . . . . . . . 3D-49

Clean Wing Approach and Landing . . . . . . . . . 3D-51

Slats-Only Approach and Landing . . . . . . . . . 3D-53


Maneuvers

TaxiingPrior to taxiing the Falcon 10/100,complete all items of the Before Taxichecklist. Obtain clearance from theappropriate control agency and ensureboth pilots understand the taxi routeprior to aircraft movement. Both pilotsshould visually check the area aroundthe aircraft for ground equipment, otherobstructions, and personnel.

Also, a visual check should be madeof the passenger cabin to note that bag-gage and equipment are stowed, emer-gency exit access is clear, galley equip-ment and supplies are secure, andpassengers are seated with seat beltsfastened. If necessary, a verbal or PAannouncement can be made that theaircraft is being taxied.

When ready to taxi, release the parkingbrake. Depress and hold the nose steer-ing wheel while advancing the powerlevers. Remember that the nose steer-ing wheel can be turned 120° in eachdirection, with the first 60° of rotationproducing up to 6° of nosewheel turn-ing, and the remaining 60° of rotationproducing approximately 54° addi-tional nosewheel turning.

Smoothly pressure the nose steeringwheel into and out of each turn to pro-duce a lurch-free ride. Releasing thedownward pressure on the steeringwheel allows the nosewheel to returnto its center position rather abruptly;this can cause lurching while the air-craft is moving, especially in a turn.

When applying power to taxi, use careand good judgment to avoid exhaustblast to other aircraft, personnel, equip-ment, and buildings. Apply sufficientpower to start the aircraft rolling; checkproper operation of the wheel brakesand then reduce power to idle. Atlighter weights and higher elevations,

the aircraft may accelerate easily, evenat idle power making it easy to gener-ate taxi speeds much higher thandesired.

If it is necessary to make a sharp turnafter moving from the parking spot,maintain above idle power until suffi-cient speed is gained to complete theturn with idle thrust. The additionalspeed prevents the aircraft from stop-ping during the turn and then requir-ing excess thrust to get it movingagain. If taxiing in a congested areaand close to other aircraft, hangars, orother obstacles, use ground personnelto ensure adequate clearance.

When clear of other aircraft, checkanti-skid operation by depressing onebrake at a time and pushing and releas-ing the anti-skid test button. Ensurethat the anti-skid lights illuminate andextinguish, and that the brake releasesand engages. Repeat this procedurewith the other brake pedal.

Both pilots should maintain good look-out discipline while taxiing. Avoidtests, checks, and paperwork activitythat compromise necessary visualclearing. Taxi speed should be kept tothe minimum practical for safety andfor passenger comfort.

Items of the Taxi and Before Takeoffchecklists should be accomplished byflow pattern, then verified by check-list reading and response when visualclearing is not compromised.

Whenever it is necessary to stop theaircraft movement with the enginesrunning, hold firm pressure on thebrake pedals until coming to a com-plete stop, then set the Park Brakehandle in the aft detent. Plan ahead –be sure that the aircraft and its pilotsand passengers are ready for flightbefore calling for takeoff clearance.

NormalOperation


Before TakeoffPrior to takeoff, consider the following:� use of flight director� thrust application� brake release� time to 100 kts� runway alignment� proper use of controls� proper rotation� gear retraction� noise abatement procedures andthrust reduction to climb power� adherence to airport area speedlimits� icing conditions.

The PF’s takeoff briefing, in accord-ance with SOP, should be clear, con-cise, and pertinent to the specific take-off. Set airspeed bugs according to theSOP. Navigation aids should be tunedand identified; the specific coursesshould be set. The altitude alertershould be set to the proper altitude.When cleared for takeoff, complete allitems of the Takeoff/ Line Up checklist.

Takeoff (General)The primary instruments for settingtakeoff thrust are the N1 gages. Therequired takeoff power settings areobtained from the manufacturer’sAFM or from the SimuFlite OperatingHandbook. The AFM and the Oper-ating Manual state that for normal take-offs, this power is set statically and thecharted takeoff performance is basedon such a setting.

With the fuel computers operating nor-mally, advancing the power leversquickly full forward may producemomentary overshoots of the targetvalues for N1 and ITT; these shouldreturn to normal ranges within a shorttime. Both pilots should, however,monitor these indicators to ensure lim-iting values do not remain exceeded.

Normal Standing TakeoffHold the brakes firmly and advancethe power levers to approximately 80%N1. Allow the engines to spool up andstabilize at nearly equal N1 indicationsbefore advancing the levers to thedesired takeoff N1. When power is set,check engine instruments and releasethe brakes smoothly.

To optimize coordination, the PNFshould monitor the instruments andassist with the power levers to enablethe PF to concentrate on directionalcontrol.

The PF should control direction ini-tially with the nose steering wheelwhile maintaining power lever controlwith the right hand. The PNF initiallymaintains slight forward pressure onthe control column while holding neu-tral ailerons.

At 80 KIAS, the PF’s left hand releas-es the nose steering wheel and movesto the control yoke; directional con-trol is then accomplished with the rud-der pedals. At V1, the PF’s right handmoves to the yoke in preparation fortakeoff rotation. Refer to the profileon page 3D-27.

Rolling TakeoffA rolling takeoff may be accomplishedwhen actual runway length adequate-ly exceeds balanced field length andobstacle clearance is not a factor. Oncethe aircraft is aligned with the runway,the brakes are applied and power leversadvanced to 80% N1. The brakes arethen released and power adjusted tothe takeoff N1 setting prior to 60 KIAS.Remember that AFM takeoff fieldlength data and takeoff N1 settingsassume a standing start.

Crosswind TakeoffWhen required, a crosswind takeoffmay be combined with any other take-off. Directional and lateral controlthroughout a crosswind takeoff arecritical.


Maneuvers

The PNF holds the yoke forward tokeep the nosewheel firmly on theground until takeoff rotation. The ailer-ons should remain in the neutral posi-tion, and a combination of rudder andnosewheel steering should be useduntil 80 kts when the rudder becomeseffective enough to control direction.In strong/gusty crosswinds, apply fullrudder in the downwind direction priorto starting the takeoff roll. During thetakeoff roll, rudder displacement de-creases as the rudder gains effectiveness.

Nosewheel steering use is restrictedby speed; it should not be used above80 KIAS.

Reduced Thrust TakeoffThe Falcon 10/100 AFM authorizesreduced thrust takeoffs at the pilot’sdiscretion. The practice of using reducethrust takeoffs has proven beneficialin improving engine service lifebecause it sets the minimum requiredthrust to meet certification require-ments for takeoff.

The assumed temperature method isused to calculate the takeoff perfor-mance requirements. When consider-ing reduced thrust takeoffs:� only dry, hard surface runways maybe used� use of clearways or stopways is notpermitted� use in icing conditions is notallowed� the anti-skid must be working� no outstanding mechanical prob-lems affecting the engines or the air-craft’s performance exist� the maximum N1 reduction cannotexceed 5% below the rated N1 speed� a rated (full power) thrust takeoffmust have been made within the last10 to 20 takeoffs.

The AFM has strict rules for reducedthrust takeoffs. Supplement 4 of theAFM provides the approved proceduresfor use of reduced thrust takeoffs.

Takeoff RotationPrecisely at VR, smoothly rotate to atakeoff attitude of 16° (if takeoff ismade with slats + 15° flaps) or to thecharted takeoff attitude (if takeoff ismade with a speed increase or withslats-only, or if obstacle clearance is afactor).

Smooth rotation prevents a decreasein airspeed. Early or late rotationdegrades takeoff performance.

Rejected TakeoffFor abort prior to V1, immediately andsimultaneously apply wheelbrakes,retard power levers to idle, move air-brake handle to EXTEND, and deploythe thrust reversers. When the thrustreversers are deployed, increasereverse thrust to slow the aircraft.

If necessary, use maximum reversethrust to 60 KIAS, then reduce to idle.Use reverse thrust cautiously on wetor slippery runways. Use caution alsoduring strong crosswind conditionsbecause reverse thrust may aggravateany weather vaning tendency. Maintaindirectional control with differentialbraking and nosewheel steering toremain on the runway centerline.Thrust reverser operation is limited to60 seconds in a 30 minute period.

Initial ClimboutOnce a positive rate of climb is indi-cated by the altimeter and verticalspeed indicator, move the landing gearlever to UP. Confirm gear retractionand monitor annunciators and engineinstruments.

When the airspeed increases to a min-imum of V2 + 35 KIAS, clean the wingby retracting the leading edge slats andflaps in two separate motions.

After retraction of the slats and flaps,climb power should be set (860°C, fiveminutes maximum). The initial settingis made by reference to 795°C ITT(maximum 832°C) or the climb N1.


After making the initial setting, con-sult the climb N1 chart for the maxi-mum allowable N1; adjust powerlevers accordingly unless N1 is belowthe recommended value and addition-al power is not required.

Noise AbatementClimboutSection V of the AFM provides theFalcon 10/100 pilot with a noise abate-ment procedure which, if followedexactly, allows the aircraft to meet FAR36 noise levels.

After gear retraction in the initialclimb, maintain takeoff power andwing configuration. The aircraft climbsat V2 + 10 KIAS until reaching 2,600ft AGL in a 18,300 lb or 18,740 lbgross weight aircraft or 2,350 ft in a19,300 lb gross weight aircraft. Reducepower to a charted value until clear ofthe noise-sensitive area, then increasepower to the normal climb N1 values.

ClimbAfter setting the climb power, andwhen clear of the airport traffic area,both pilots complete the climb check-list. Use a flow pattern with the PNFverifying completion and indicationswith the checklist.

Throughout the climb, the PNF com-pares the indicated N1 with the climbN1 chart. N1 RPM increases with alti-tude; several power adjustments maybe necessary during climb to maintainthe specified setting required by theclimb charts. If a temperature inver-sion is encountered during the climb,closely monitor the climb N1 settingto stay within the climb N1 limits.

Observe the differential pressure/ cabinaltitude and cabin vertical speed gagesfor proper programming and comfortrate.

CruiseThrust SettingClimb power is normally maintainedupon level-off until acceleration to thedesired cruise Mach, then adjusted tothe appropriate setting. During theclimb and acceleration to cruise speed,the ITT should not be between 795°and 832°C for more than 30 minutes.

Cabin TemperatureMonitor the temperature control panelto ensure proper comfort level for thepassengers and crew. Normally, thetemperature control selector is inAUTO at the 12 o’clock position.

Turbulent AirPenetrationAlthough the aircraft is not opera-tionally restricted in rough air, do notfly into known severe turbulence.

Carefully plan turbulence avoidancestrategy with an understanding ofmountain wave dynamics, thunder-storm characteristics, and weight ver-sus altitude buffet margins. When tur-bulence is encountered, the followingsteps are recommended:

1. Maintain 250 KIAS (0.75 M).

2. Set thrust to maintain target air-speed. Change thrust only for extremeairspeed variation.

3. With the autopilot not engaged,keep control movements moderate andsmooth. Maintain wings level anddesired pitch attitude. Use the attitudeindicator as the primary instrument. Inextreme drafts, large attitude changesmay occur. Do not make sudden largecontrol movements. After establishingthe trim setting for penetration speed,do not change the trim.


M a n e u v e r s

4 . L a rge altitude changes are possiblein severe turbulence. Reduce altitude toincrease the buffet boundary marg i nif necessary. Do not chase altitude ora i r s p e e d .

5 . Ensure yaw damper is engaged toreduce yaw/roll oscillations.

6 . If turbulence is penetrated with theautopilot on, engage the Tu r b u l e n c e(Soft Ride) mode.

7 . If turbulence is known or suspectedto be moderate or severe, the startselectors should be placed to AIRS TA RT and the FASTEN BELT signi l l u m i n a t e d .

Operation inIcing ConditionsThe engine anti-ice systems and thewing anti-ice system prevent the accu-mulation of icing; they should beturned on prior to encountering suchc o n d i t i o n s .

Engine anti-ice should be used for taxiand takeoff when the ambient temper-ature is 5°C or below, and visible mois-ture, precipitation, or a wet runwayexists. Engine anti-ice is used in flightwhen the total air temperature is 5°C orbelow with visible moisture, precipi-tation, or icing.

Wing anti-ice should be used underthe same conditions, but must not beused on the ground. On takeoff, itshould be turned on after gear retrac-tion. During approach and landing, itshould be turned off prior to touch-d o w n .

If it becomes necessary to de-ice theaircraft in flight, use the following pro-c e d u r e .

1 . Place engine start selectors to AIRS TA RT.

2 . Turn on engine anti-ice to eachengine, one at a time, pausing momen-tarily between each step.

3 . Turn wing anti-ice on.

For proper anti-ice operation, ensureadherence to the minimum engine N1settings for the existing phase of flight(see the AFM or the SimuFliteOperating Handbook).

Inflight ProceduresAirbrake DeploymentAirbrakes may be used to expedite adescent or reduce airspeed. Buff e t i n gis noticeable with airbrakes extended.

Airbrakes may be used at any speedup to the VM O/ MM O and, if necessary,with the landing gear and/or wingslats/flaps extended. During approach,extension of the airbrakes is not per-mitted below 500 ft AGL.

Change of AirspeedAirbrakes may be used in conjunctionwith thrust reduction when reducingairspeed quickly. Reduce thrust to theappropriate setting for the desired air-speed, then extend the airbrakes. Uponreaching the desired airspeed, retractthe airbrakes.

Smoothly coordinate all power andflight control inputs to maintain desiredheading, airspeed, and altitude. Air-brakes may also be used to control air-speed during inflight operation of theengine and airframe anti-icing systemswhen higher-than-normal enginepower settings are required.

Steep TurnsSteep turns (45° bank) confirm theaerodynamic principle that increasingbank requires increased pitch andpower to maintain altitude. Refer tothe profile on page 3D-35.

At intermediate altitudes, approxi-mately 10,000 ft MSL, practice steepturns at 250 KIAS. Start the maneu-ver on a cardinal heading and altitude.

3D-12 For SimuFlite training only Falcon 10/100June 2006

The initial engine power setting isabout 78% N1. When passing through30° bank, increase power setting about2% N1 and pitch attitude approxi-mately 4°. Trim out back pressure asneeded. Lead the roll-out headingapproximately 15° and reduce thrustand pitch to the original settings.

Stall Recognitionand RecoveryApproach to StallThe approach to stall should be con-tinued only to the first warning indi-cation of a stall (audible stall warningor airframe buffet, whichever occursf i r s t ) .

At the first warning indication, initi-ate an immediate recovery. Do notallow the aircraft to go into a full stall.Refer to the profile on page 3D-33.

Perform the approach to stall in clean,t a k e o ff, and landing configurations.Practice altitude should be no higherthan 20,000 ft MSL (AFM limitation),and no lower than 5,000 ft above ter-r a i n .

Before practicing approaches to stall,clear cockpit area of loose articles;visually clear the practice area, com-pute VR E F, and set airspeed indicatorbugs to VR E F.

Clean Configuration –Flaps and Gear UpWhile maintaining altitude and head-ing (wings level), retard power leversto 40% N1. As the aircraft slows, main-tain altitude with back pressure. Usetrim to reduce stick force (but notbelow 160 KIAS).

Observe the angle-of-attack indicatorinformation; stall warning occurswhen the indicator approaches the redband. At the first indication of a stall(i.e., stall warning horn), perform thefollowing.

1 . Advance power levers to maximump o w e r.

2 . Maintain pitch attitude (approxi-mately 11° nose-up) and wings level.

3 . As airspeed increases, reduce pitchattitude to maintain altitude.

4 . Accelerate to 180 KIAS, thenreduce power to approximately 70%N1 to maintain 180 KIAS.

Takeoff Configuration –Slats + 15° Flaps andGear DownEstablish a level turn using 15 to 20°bank; retard power levers to 50% N1.As the aircraft slows, maintain altitudewith back pressure. Use trim to reducestick forces; however, stop trimmingat 130 KIAS.

Observe the angle-of-attack indicatorinformation; stall warning occurs whenthe indicator approaches the red band.At the first indication of a stall (i.e.,stall warning horn), perform the fol-l o w i n g .

1 . Advance power levers to maximumpower while rolling the wings level.

2 . Maintain pitch attitude (approxi-mately 13° nose-up) and wings level.

3 . As the airspeed increases, reducepitch attitude to maintain altitude.

4 . Accelerate to 160 KIAS, thenreduce power to approximately 70%N1 to maintain 160 KIAS. If complet-ing a stall series, maintain gear andflap configuration.

Landing Configuration –Full Flaps and Gear DownWhile maintaining altitude and head-ing (wings level), retard power leversto 60% N1. Trim to reduce stick forces,but do not trim below VR E F. At thefirst indication of a stall (i.e., stallwarning horn), perform the following.

C A U T I O N : The adjacent dis-cussion is presented only in thecontext of recovery training.Stalls in high performance air-craft should not be deliberatelyexecuted unless they are part ofa supervised pilot training pro-gram. Safety of flight consider-ations dictate that the utmostcaution be employed during suchexercises.


Monitor the warning lights for indica-tion of malfunction. If the computer isnot working properly, erroneous infor-mation may be presented.

Instrument ApproachConsiderationsSeveral factors should be consideredprior to commencing an approach ina high performance jet aircraft. Thepilot must have a thorough knowledgeof the destination and alternate weath-er conditions before descending out ofthe high altitude structure. Manyweather and traffic advisory sourcesare available, including:� Flight Service Stations that may beused enroute any time to obtain the lat-est destination and alternate weatherconditions� ARTCC where controllers canobtain information (if requested) per-taining to traffic delays and whetheraircraft are successfully completingapproaches� ATIS� destination tower and/or ApproachControl.

If weather is at or near minimums forthe approaches available, review howmuch time and fuel is needed to go toan alternate.

To continue the approach to a landingafter arrival at minimums, FAR 91.175requires that:

(c) Operation below DH or MDA.Where a DH or MDA is applicable, nopilot may operate an aircraft, excepta military aircraft of the United States,at any airport below the authorizedMDA or continue an approach belowthe authorized DH unless:

(1) The aircraft is continuously in aposition from which a descent to alanding on the intended runway can

be made at a normal rate of descentusing normal maneuvers, and for oper-ations conducted under part 121 orpart 135 unless that descent rate willallow touchdown to occur within thetouchdown zone of the runway ofintended landing;

(2) The flight visibility is not less thanthe visibility prescribed in the stan-dard instrument approach being used;and

(3) Except for a Category II orCategory III approach where any nec-essary visual reference requirementsare specified by the Administrator, atleast one of the following visual refer-ences for the intended runway is dis-tinctly visible and identifiable to thepilot:

(i) The approach light system, exceptthat the pilot may not descend below100 ft above the touchdown zone ele-vation using the approach lights as areference unless the red terminatingbars or the red side row bars are alsodistinctly visible and identifiable.

(ii) The threshold.

(iii) The threshold markings.

(iv) The threshold lights.

(v) The runway end identifier lights.

(vi) The visual approach slope indi-cator.

(vii) The touchdown zone or touch-down zone markings.

(viii) The touchdown zone lights.

(ix) The runway or runway markings.

(x) The runway lights.

(d) Landing. No pilot operating anaircraft, except a military aircraft ofthe United States, may land that air-craft when the flight visibility is lessthan the visibility prescribed in thestandard instrument approach proce-dure being used.


Maneuvers

Additional InstrumentSystemsThe following additional equipment isavailable on most aircraft and shouldbe set according to company SOP:� radio altimeter� terrain advisory voice encodingaltimeter� vertical navigation computer con-troller� long-range navigation equipment.

Normal DescentAs descent is initiated, set the pres-surization control for landing. Thecabin pressure controller may initiallybe set to:� QNH and true landing field altitude� QFE of landing field and altitude 0� 29.92 inches of Hg and QNE.

The manufacturer then permits settingthe pressure controller to 300 ft belowfield elevation. The latter setting resultsin the aircraft landing slightly pres-surized, but it depressurizes within 30seconds after touchdown. Many oper-ators, however, use a final settingslightly above field elevation to landdepressurized.

Continue to monitor the differentialpressure, cabin altitude, and cabin verti-cal speed throughout descent. The mostcomfortable condition occurs whencabin descent is distributed over themajority of the aircraft descent time.

The engine and wing anti-ice systemsshould be on when operating in visi-ble moisture if the total air temperatureis 5°C or below. The radar altimetermay be bugged to the decision height,the height above the airport, or asdesired in VFR operation for terrainproximity warning.

Double-check landing field informa-tion and estimated arrival gross weight;check runway requirements and deter-mine VREF and VAP (VAP equals VREFplus configuration correction, if thereis any, plus wind factor; minimum is10 kts, maximum is 20 kts). Whendescending through the transition alti-tude, set the altimeters to field pres-sure and check for agreement.

EmergencyDescentAn emergency descent moves the air-craft rapidly from a high altitude to alower altitude; it is most often used inconjunction with a loss of pressuriza-tion.

Put on oxygen masks, establish com-munications, disconnect autopilot,retard power levers to idle, extend air-brakes, illuminate the seat belt/nosmoking signs, and roll into a moder-ate bank while lowering the nose (ini-tially to approximately 20°) below thehorizon.

Adjust pitch as necessary to approach,but do not exceed, VMO/MMO (approx-imately 10° below the horizon). If fly-ing in turbulent air, or if structuralintegrity is questionable, make thedescent at a lesser and more prudentspeed. The PNF should set thetransponder to 7700.

When conditions permit, the enginestart selectors may be placed to AIRSTART, the condition of the passen-gers checked, and ATC contacted forassistance and instructions. The PNFshould monitor the descent progress,and complete the appropriate check-lists on command.


VFR Traffic PatternThe traffic pattern altitude is normal-ly at 1,500 ft AGL. At uncontrolledairports, comply with the prescribedtraffic flow for that airport. Refer toprofile on page 3D-45.

The specific power settings stated inthe following paragraphs apply to aflight weight of about 16,000 lbs at asea-level airport with standard dayatmospheric conditions.

Before entering downwind leg, com-plete the Approach checklist. Set slats+ 15° flaps and slow to 160 KIAS.Target power setting is approximate-ly 63 to 65%. Abeam the end of therunway, select gear down and main-tain airspeed. Complete the BeforeLanding checklist.

In the base turn, select flaps 30° andslow to 140 KIAS; power is set atabout 63% N1 and a descent rate of500 to 600 FPM is maintained. Uponintercepting the glidepath, set landingflaps (52°). As airspeed approachesVAP, set power to maintain VAP (about68% N1). Cross the threshold at VREF+ wind factor.

ApproachesChecklist andConfigurationFor instrument approaches where aprocedure turn is flown, the Approachchecklist should be completed andflaps set at 15°. The aircraft is slowedto 160 KIAS with power set to approx-imately 63 to 65% N1 when passingthe IAF outbound.

Table 3D-A; Maneuvering Guidelines

Configuration Airspeed

Clean (Above 10,000 ft) —85%/8803300 kts

Clean (10,000 ft) 85%/1,10078%/7004250 kts

Clean (Below 10,000 ft) 85%/1,10068%/6504250 kts

Clean 80%/1,10063%/6007180 kts

Slats + 15° Flaps 80%/1,10065%/6504160 kts

Slats + 30° Flaps(Gear Down, Level Flight) 93%/1,30072%/7004140 kts

Slats + 30° Flaps(Gear Down, Glideslope) 75%/90065%/6004140 kts

Slats + 52° Flaps(Gear Down, Glideslope) —68%/6504VREF +10

Slats + 30° Flaps(Gear Down,Non-Precision Descent

68%/60058%/4001140 kts

Slats Only(Gear Down, Glideslope) —50%/3004Bug +10

PitchPower/PPH

2 Engines 1 Engine


Maneuvers

If the aircraft is receiving radar vec-tors for an approach, the Approachchecklist and aircraft configurationchanges should be completed whenabeam the FAF, or three to five milesbefore the FAF for a straight-inapproach.

At uncontrolled airports, make allrequired position/intention reports onthe appropriate Traffic Advisory fre-quency.

Power SettingsTable 3D-A depicts typical power set-tings and fuel burn rates duringapproach.

Typical PrecisionApproachAn ILS approach is considered nor-mal when all engines, the appropriateILS facilities, and the airborne equip-ment are operating normally. Refer toprofile on page 3D-37.

1. When established on the localizerinbound to the FAF, ensure flaps areset to 15°.

2. Maintain airspeed at 160 KIAS withpower set at approximately 63 to 65%.

3. When the glideslope indicates alive,lower the landing gear. Complete theBefore Landing checklist.

4. When the glideslope indicates onedot prior to intercept, set flaps to 30°.

5. At glideslope intercept, begindescent and extend flaps to 52°.

6. Maintain airspeed at VAP withpower set at approximately 68% N1.

7. At or before DA/DH, establish visu-al contact with the runway.

8. Reduce power slightly to ensurecrossing the runway threshold at VREFplus wind factor.

Typical Non-PrecisionApproachRefer to the profile on page 3D-41.

When established on the inboundcourse to the FAF, perform the fol-lowing.

1. Set slats + 15° flaps and completethe Approach checklist.

2. Adjust airspeed to 160 KIAS; thepower setting should be about 63 to65% N1.

3. One mile outside of FAF, extendlanding gear, select flaps to 30°, andcomplete the Before Landing check-list (except for final flap setting).Maintain 140 KIAS or VREF + 20 min-imum with the power approximately72% N1.

4. Upon crossing the FAF, start tim-ing, notify ATC, and descend to theMDA while maintaining 140 KIASwith power at about 58% N1. Verticalspeed in descent should normally be1,000 to 1,500 fpm.

5. After leveling off at MDA, increasepower to approximately 72% N1 tohold airspeed at 140 KIAS or VREF +20 minimum while proceeding to theVDP or the MAP.

6. With the runway landing environ-ment in sight, set landing flaps (52°)and slow to VAP while interceptingthe proper visual glidepath for land-ing.

Slats-Only/Clean WingApproach and LandingMaintain a minimum airspeed of 180KIAS while maneuvering with powerset to about 63% N1. Plan for a longfinal approach. Extend slats and main-tain 160 KIAS minimum until glide-slope alive.


If no slats are available, maintain 180KIAS until glideslope alive. Lowerlanding gear early in the approach tohelp control airspeed. Complete theBefore Landing checklist.

Once established on final, reduce toVREF + 15 KIAS + wind factor (withslats extended) or VREF + 35 KIAS +wind factor (with clean wing). The sta-bilized power settings on final shouldbe about 50% N1.

The aircraft has a tendency to floatbecause of increased airspeed and lowdrag configuration; this can be coun-tered by flying the aircraft onto the run-way and using minimal flare to breakthe descent rate. Expect landing fieldlength to be longer than normal. Referto profiles on pages 3D-51 and 3D-53.

Go-Around/Missed Approach/Balked LandingAccomplish the go-around/missedapproach/balked landing at the DA/DH or MDA with time expired (ifapplicable) and runway visual refer-ence either not in sight or not in a posi-tion from which a normal visual land-ing approach can be accomplished.Refer to the appropriate profile on page3D-49.

Go-Around ProcedureAccomplish the following.1. Apply go-around power.

2. Push the go-around button; rotateto the flight director go-around atti-tude (approximately 14° nose-up).Ensure airbrakes are retracted.

3. With airspeed at a minimum ofVREF, set flaps to 15°, one detent at atime. Retract gear at indication of apositive rate of climb.

4. When clear of obstacles (400 ftAGL minimum) and at a minimum air-speed of VREF + 35 KIAS, clean thewing and accelerate. Adjust pitch atti-tude and power as necessary.

5. When clear of obstacles, reducepower to climb N1. At the relativelylight gross weight at which missedapproaches are normally accom-plished, the aircraft accelerates quick-ly. Pitch and power need to be adjustedaccordingly.

6. Set the flight director as required.Use the heading bug and the headingmode to fly a desired heading, and anavigation mode and the course selec-tor to capture a desired radial/track.After the initial fixed (14° nose-up)climb attitude is established, variableclimb attitudes may be commandedwith the pitch synch button on the con-trol yoke. Desired climbs or altitudesmay then be captured and maintainedby using one of the vertical modes.

7. Confirm the level-off altitude andheading/course needed for the missedapproach. Comply with the publishedmissed approach instructions unlessother directions are received fromATC.

After a MissedApproach –Proceeding forAnother ApproachAccomplish the following.

1. After level-off, complete the Climbchecklist and maintain 180 KIASminimum.

2. Complete the Approach checklist.In the slats + 15° flap configuration,maintain 160 KIAS. Refer to the pro-file on page 3D-49.

After a MissedApproach –Departing AreaAccomplish the following.

1. Accelerate to normal climb speed.

2. Complete the Climb checklist.

3. Follow normal climbout procedures.


Maneuvers

Circling Approach/Circling PatternA circling approach is an instrumentapproach requiring a heading changeof 30° or more to align the aircraft withthe landing runway. Once visual con-ditions are reached, the circlingapproach is a modified version of theVFR traffic pattern. Refer to profileon page 3D-47.

Turbulence, strong winds, poor visi-bility, and low maneuvering altitudeare factors that must be consideredwhen planning a circling approach.Plan to use a minimum circling alti-tude and distance appropriate to theairspeed or approach category. Atuncontrolled airports, observe localtraffic direction and restrictions.

It is recommended that the approachbe flown with gear down and flaps 30°until arriving at a position from whicha normal descent for landing can bemade. At that time, begin descent,select flaps 52°, and slow to VAP.

While maneuvering during the circlingapproach, fly a minimum of 140 KIAS.When established on final in the land-ing configuration, fly at VAP untilreducing power slightly to cross therunway threshold at VREF + windfactor.

LandingWith slats + 52° flaps, cross the thresh-old at 50 ft AGL with a speed of VREF+ wind factor. The aircraft pitch atti-tude is approximately 2° nose-up.Reduce thrust slightly. At about 30 ftAGL, gradually increase pitch todecrease the rate of descent (resultingin a pitch attitude of approximately 8°nose-up at touchdown), then reducethe thrust to idle.

For approach and landing with slats +15° flaps, pitch attitudes are approxi-mately 4° nose-up on approach andapproximately 10° nose-up on touch-

down. Thrust may be reduced to idle at50 ft AGL and minimum flare used toreduce float.

Upon touchdown, lower the nosewheelsmoothly to the runway, extend the air-brakes, and apply braking as neces-sary. Use rudder and differential brak-ing to maintain directional control, anddeploy the thrust reversers. Use nose-wheel steering below 80 KIAS.

Thrust ReversersIf necessary, reverse thrust may beused to shorten the landing rollbetween touchdown and full stop. Pullthe reverser levers up and aft; after thethrust reverser doors are fullydeployed (as indicated by the greenREV lights), smoothly pull the leversfurther aft to increase reverse thrustto maximum N1.

CrosswindThe maximum demonstrated cross-wind for the Falcon 10/100 is 25 kts.On the final approach in a crosswind,the crab approach or the wing-downmethod may be used.

Do not allow the aircraft to float withpower off prior to touchdown. Fly totouchdown with little, if any, flare.Deploy airbrakes on touchdown. Atnosewheel touchdown, neutralize theailerons. Use rudder and differentialbraking for directional control.Nosewheel steering may be usedbelow 80 KIAS after nosewheel touch-down.

Touch-and-Go LandingsIf touch-and-go landings are to be prac-ticed, they should be preplanned andbriefed. The thrust reversers and air-brakes should not be used on landing.The PNF should reset the flaps to 15°,set the stabilizer trim in the takeoffrange, and confirm these settings tothe PF before the power levers areadvanced to takeoff power.


Wet/Contaminated/Very SlipperyRunwaysTo assist the Falcon 10/100 pilot inidentifying the many factors involvedin operating on other than dry, hardsurface runways, the manufacturer hasprovided Service Newsletter No. 34(April 1987). A synopsis of the defin-itions and considerations from this doc-ument is provided below.

DefinitionsWet runway – Not covered to anyextent with standing water; water depthis not measurable or less than 1/8 inch.

Contaminated runway – Covered bystanding water, slush, wet snow, orloose dry snow. The depth of such mat-ter is greater than 1/8 inch and coversat least 25% of the required length.

Compacted snow runway – Thesnow has been packed into a solidmass into which the aircraft’s wheelsdo not sink.

Very slippery runway – Coveredwith ice or black ice; some parts of arunway can also be very slippery dueto a mixture of oil, rubber, and water.

Hydroplaning speed (V h) – Thegroundspeed at which the hydro-dynamic pressure build-up betweenthe aircraft’s tires and the water filmon the runway is sufficient to lift thetire surface off the ground.

In this situation, wheel braking andsteering effectiveness are diminished.Manufacturer’s tests have shown thatthe hydroplaning speed for the Falcon10/100 can be approximated by theformula:

Vh = 6.4 √tire pressure (PSI)

At normal tire inflation pressures, Vhfor the nose and main wheels is 62 ktsand 75 kts, respectively.

When operating on runways contam-inated by standing water,in addition tothe aerodynamic drag, the aircraft isalso subject to hydrodynamic drag andplume drag. The total (hydrodynamicplus plume) drag increases propor-tionally up to Vh, and reaches about3,100 lbs for a 3/8 inch water depth,then decreases slowly at speeds aboveVh.

If the runway is contaminated by slush,wet snow, or dry snow, the equivalentwater depth should be determinedbefore estimating the Vh. A chart fordetermining equivalent water depth isfound in the Falcon 10/100 Perform-ance Manual.

When braking on non-dry runways,the braking coefficient of friction maybe reduced by 20% to 50% on wet run-ways, by up to 75% on contaminatedrunways, by 60% to 90% on snow-compacted runways, and by 90% onvery slippery runways.

RecommendationsWet Runways

Throughout the ground phase of oper-ation on such runways, ensure that thestart selectors are in the AIR STARTposition. If a crosswind is present,maintain neutral ailerons and hold for-ward yoke pressure for better nose-wheel steering.

For landing operations, increase land-ing distance and landing field lengthby 15%. For takeoff operations, do notuse reduced thrust and add 15% of thelanding distance to the takeoff distance.

Contaminated Runways

Operation on contaminated runwaysshould be avoided whenever possibleespecially during and immediatelyafter heavy rainfall. If the surface con-taminant is slush or snow, use thePerformance Manual charts to deter-mine the equivalent water depth.


Maneuvers

Maximum recommended equivalentwater depth is 0.5 inch; the anti-skidsystem must be operable and maxi-mum crosswind is limited to 15 kts.

In a crosswind, use the same controltechniques as for a wet runway. Startselectors should be in the AIR STARTposition.

For landing operations, multiply thenormal landing distance by 2.3 forlanding field length. Use thrustreversers as soon as possible; do notapply wheel brakes until belowhydroplaning speed (75 kts on water,slightly more on snow), then applythem fully.

For takeoff operations, adopt V1 = VRand do not use reduced thrust. Use thePerformance Manual charts to deter-mine the takeoff distance with theequivalent water depth, then add dou-ble the landing distance; this gives anaccelerate/stop distance with a safetymargin.

Select a landmark along side the run-way at the expected takeoff distance. Ifnot airborne shortly before this land-mark, abort the takeoff.

In the event the takeoff is rejected, usethe above landing recommended stop-ping procedures. Use neutral aileronsand nosewheel steering until Vh; abovethat speed, use the rudder.

After takeoff, delay gear retraction (toblow off contaminant accumulation)and then cycle the gear several times ifnecessary (to shake off any accumu-lation from the landing gear andbrakes).

Compacted Snow orIcy Runways

When considering operation on com-pacted snow or icy runways, the man-ufacturer says: “It is safer to neithertakeoff nor land.”

Maximum crosswind is limited to fiveknots. If a landing on such a runwayis contemplated, the landing distance isequal to the dry runway distance times2 (compacted snow) to 3.4 (black ice).Add 15% to the result for the newlanding field length.

Upon touchdown, apply brakes withanti-skid fully, and use the thrustreversers as soon as possible.Nosewheel steering is of little help.Takeoffs on such runways should beavoided if at all possible because direc-tional control is poor to nil. Do not usereduced thrust.

The approximate distance is the nor-mal balanced field length plus once(compacted snow) or twice (black ice)the landing distance under the sameconditions. If the takeoff must be abort-ed, use stopping techniques as recom-mended for landing above.

After LandingAfter clearing the runway, completethe After Landing checklist. Theengines should be operated at idle forat least two minutes (taxi time may beincluded) prior to shutdown. After theaircraft is parked, complete theShutdown checklist.


Maneuvers

Engine FailureAfter V 1 –Takeoff ContinuedWith an engine fire or failure indicationafter V1, continue the takeoff. Referto the profile on page 3D-31.

Maintain directional control using therudder and accelerate to VR/V2. At VR,rotate the aircraft to 16° nose-up (orthe pre-computed pitch attitude), andclimb at V2 minimum. If the indica-tion occurs after exceeding V2, main-tain the existing airspeed. Retract thelanding gear when a positive rate ofclimb is established.

When clear of obstacles (minimum400 ft AGL), accelerate to V2 + 35KIAS and clean the wing. Increasespeed to VFS and continue the climbto 1,500 ft (or required altitude) abovethe takeoff field elevation. Limit themaximum takeoff power to five min-utes, then reduce to maximum contin-uous power.

When time and conditions permit,complete After V1 Engine Fire and/orEngine Failure checklist(s).

Single EnginePrecision Approachand LandingA one-engine inoperative approach isflown essentially the same as anapproach with all engines operating.Landings may be made with slats +30° flaps or slats + 52° flaps. On finalapproach, however, flaps are notextended beyond 30° until landing isassured (normally 200 ft AGL). Referto the profile on page 3D-39.

Up to the final descent point, the air-craft is configured normally with thepreviously recommended speeds flownfor each configuration.

Single engine thrust settings are slight-ly higher than comparable normalengine settings. If adequate runway isavailable, a landing in the slats + 30°flap configuration avoids the large trimand power changes required by selec-tion of slats + 52° flaps.

If rudder trim has been used during theapproach to counter the asymmetricthrust, zero the rudder trim prior to orduring the landing power reduction toprevent unwanted yaw. Thrust reduc-tion and flare are similar to a normallanding. Thrust reduction should beslower than normal to counter roll dueto yaw effect. Consequently, slightlyless flare than normal is required toprevent floating.

After touchdown, lower the nose,extend the airbrakes, apply wheel brak-ing, and keep the wings level. Use rud-der as required. Reverse thrust may beused if the runway is dry and nose-wheel steering is available.

Single EngineGo-Around/Missed Approach/Balked LandingRefer to the profiles on page 3D-49.

Apply maximum power on the oper-ating engine, check that the airbrakesare retracted, and push the flight direc-tor go-around button to select the Go-Around mode. Rotate to approximately14° nose-up as commanded by theflight director and retract the flaps to15°. As thrust is increased, apply rud-der pressure as required to counteryaw.

Maintain the go-around pitch attitudeand minimum airspeed of VAC. Retractthe landing gear when a positive rate ofclimb is established.

One EngineInoperativeOperation


Climb to 400 ft AGL (minimum), thenaccelerate to VREF + 35 and clean thewing. Continue accelerating to VFS,set climb power, and continue climbon the published missed approach.

When time permits, the PNF sets thePF’s heading bug on the missed

approach heading and selects therequested modes on the flight direc-tor. At the appropriate time, adviseATC of the missed approach andrequest further clearance (anotherapproach or a diversion to the alter-nate airport).


Maneuvers

The following flight profiles illustrate how selected maneuvers areperformed. Each maneuver is broken down into sequential eventsthat illustrate appropriate configurations.

� Takeoff� Rejected Takeoff� Engine Failure After V1

� Stalls at Altitude� Steep Turns� Precision Approach and Landing� Single Engine Precision Approach and Landing� Non-Precision Approach and Landing� Single-Engine Non-Precision Approach and Landing� Visual Approach/Balked Landing� Circling Approach� Go-Around/Missed Approach� Clean Wing Approach and Landing� Slats-Only Approach and Landing

Profiles


Maneuvers

� PITCH – MAINTAIN AS CALCULATED � AIRSPEED – ALLOW TO INCREASEAT VFR � CONFIGURATION – CLEAN WING � POWER – SET CLIMB POWER N1 (ITT 795°C)

AT 3,000 FT AGL MINIMUM � CLIMB CHECKLIST – COMPLETE � AIRSPEED – 250 KIAS

TAKEOFF � FLIGHT DIRECTOR – SET � BRAKES – HOLD � POWER – SET � BRAKES – RELEASE

5

2

1

ROLLING TAKEOFF � POWER – SET T/O POWER BY 60 KIAS

80 KIAS CROSSCHECK � NOSE STEERING WHEEL – RELEASE � LEFT HAND – MOVE TO CONTROL WHEEL

AT V1 � V1 – CALL � RIGHT HAND – MOVE TO CONTROL WHEEL

AT VR � ROTATE TO 16° (OR APPROPRIATE PITCH ATTITUDE)

AT POSITIVE RATE OF CLIMB � CONFIGURATION - GEAR – UP

7

83

9

6

TIME TO 100 KIAS4

NOTE: ALL POWER SETTING ARE ARE APPROXIMATE.

Takeoff


Maneuvers

3D-27

Takeoff


Maneuvers

2

TAKEOFF� FLIGHT DIRECTOR – SET� BRAKES – HOLD� POWER – SET� BRAKES – RELEASE AT 80 KIAS

� NOSE STEERING WHEEL– RELEASE� LEFT HAND – MOVE TO CONTROL WHEEL

PRIOR TO V1� DECISION TO REJECT TAKEOFF� CALL “ABORT”� BRAKES- AS REQUIRED� POWER- IDLE� AIRBRAKE- EXTEND� REVERSE THRUST- AS REQUIRED

TIME TO 100 KIAS

ROLLING TAKEOFF� POWER – SET T/O POWER BY 60 KIAS

5

3

1

4

Rejected Takeoff


Maneuvers

AIRSPEED – MAINTAIN V2 (MINIMUM)UNTIL 400 FT AGL (MINIMUM)

AT SAFE ALTITUDE (400 FT AGL (MINIMUM) � AIRSPEED – ACCELERATE TO VFR � CONFIGURATION - CLEAN WING � AIRSPEED – ACCELERATE TO VFS

NORMAL TAKEOFF � FLIGHT DIRECTOR – SET � BRAKES – HOLD � POWER – SET T/O POWER � BRAKES – RELEASE � TIMING – START

5

2

1

ROLLING TAKEOFF � POWER – SET T/O POWER BY 60 KIAS

80 KIAS CROSSCHECK � NOSE STEERING WHEEL – RELEASE � LEFT HAND – MOVE TO CONTROL WHEEL

AT V1 � V1 – CALL � RIGHT HAND – MOVE TO CONTROL WHEELENGINE FAILURE RECOGNIZED

AT VR � ROTATE TO 16° (OR APPROPRIATE PITCH ATTITUDE)

AT POSITIVE RATE OF CLIMB � CONFIGURATION - GEAR – UP � HEADING MODE – SELECT

7

83

9

6

TIME TO 100 KIAS4AFTER 5 MINUTES � POWER – SET CLIMB POWER � FAILED ENGINE – IDENTIFY/CHECKLIST

Engine Failure After V 1


Maneuvers

3D-31

Engine FailureAfter V 1


Maneuvers

BEFORE BEGINNING STALL PRACTICE � VREF – COMPUTE � CLEARING TURNS – COMPLETE

TAKEOFF CONFIGURATION STALL – SLATS + FLAPS 15°, GEAR EXTENDED � POWER – 50% � TRIM – TO 130 KIAS � ALTITUDE – MAINTAIN � BANK – 15 TO 30° � EXPECT STALL WARNING – APPROXIMATELY VREF � RECOVERY - POWER – MAXIMUM - ATTITUDE – ROLL WINGS LEVEL, MAINTAIN PITCH - AIRSPEED – 160 KIAS - ALTITUDE – MINIMUM LOSS

2

3

1

CONSTANT ALTITUDE

CONSTANT ALTITUDEMINIMUM ALTITUDE 10,000 AGL

MAXIMUM ALTITUDE 20,000 MSL

CLEAN CONFIGURATION STALL � SET AIRSPEED BUG - VREF � POWER – 40% � TRIM – TO 160 KIAS � ALTITUDE – MAINTAIN � EXPECT STALL WARNING - APPROXIMATELY 140 TO 145 KIAS � RECOVERY - POWER – MAXIMUM - ATTITUDE – MAINTAIN - CONFIGURATION – CLEAN - AIRSPEED – 180 KIAS - ALTITUDE – MINIMUM LOSS

LANDING CONFIGURATION STALL – SLATS + FLAPS 52°, GEAR EXTENDED � POWER – 60% � TRIM TO VREF � ALTITUDE – MAINTAIN � EXPECT STALL WARNING – APPROXIMATELY VREF –15 � RECOVERY - POWER – MAXIMUM - ATTITUDE – MAINTAIN - CONFIGURATION - FLAPS 30° AT VREF - CONFIGURATION – SLATS + FLAPS 15° AT POSITIVE RATE - GEAR – UP AT VREF + 35 - CONFIGURATION – CLEAN WING - ALTITUDE – MINIMUM LOSS


Stalls At Altitude


Maneuvers

3D-33

Stalls AtAltitude


Maneuvers

LEAD ROLL OUT TO ASSIGNED HEADING BY APPROXIMATELY 15° � WINGS – SMOOTHLY ROLL LEVEL � TRIM – AS REQUIRED � PITCH – AS REQUIRED TO MAINTAIN ALTITUDE � POWER – 78% N1

TOLERANCES: SPEED ± 10 KIAS ALTITUDE ± 100 FT BANK ± 5° HEADING ± 10°

CLEAN CONFIGURATION � POWER – 78% N1 � AIRSPEED – 250 KIAS � ATTITUDE – 4°

THIS MANEUVER MAY BE USED FOR A 180° OR 360° TURN,AND MAY BE FOLLOWED BY A REVERSAL IN THE OPPOSITE DIRECTION.

THE PNF MAY ASSIST AS DIRECTED BY THE PF.

1

2 BANK – SMOOTHLY ROLL TO 45° � ALTITUDE – MAINTAIN � TRIM – AS DESIRED � PITCH – INCREASE TO APPROXIMATELY 5° � POWER – 80% N1

4

3 � ALTITUDE – MAINTAIN � AIRSPEED – MAINTAIN � BANK – MAINTAIN


Steep Turns


Maneuvers

3D-35

Steep Turns


Maneuvers

LANDING ASSURED

RADAR VECTORS

WITHIN RANGE � POWER – 60% N1 � CONFIGURATION – SLATS + FLAPS 15° � AIRSPEED – 180 KIAS � APPROACH CHECKLIST – COMPLETE � POWER – 63 TO 65% N1 � AIRSPEED – 160 KIAS

WITHIN 3 MINUTES OF IAF � POWER – 60% N1 � CONFIGURATION – SLATS + FLAPS 15° � AIRSPEED – 180 KIAS � APPROACH CHECKLIST – COMPLETE � POWER – 63 TO 65% N1 � AIRSPEED – 160 KIAS

IAF OUTBOUND � TIMING – START

THRESHOLD � POWER – SLIGHTLY REDUCE (3 TO 4% N1) � AIRSPEED – VREF + WIND FACTOR

6

TOUCHDOWN � BRAKES – AS REQUIRED � POWER – IDLE � AIRBRAKES – EXTEND � REVERSE THRUST – AS REQUIRED

7

3

5

WHEN GLIDESLOPE ALIVE � CONFIGURATION - GEAR – DOWN � BEFORE LANDING CHECKLIST – COMPLETE

AT GLIDESLOPE INTERCEPT � TIMING – START � POWER – INCREASE APPROXIMATELY 5% N1 � CONFIGURATION – SLATS + FLAPS 52° � AIRSPEED – VAP

4

PROCEDURE TURN INBOUND

1A

1

2

ONE DOT BELOW GLIDESLOPE � CONFIGURATION – SLATS + FLAPS 30°

RADAR VECTORS

TERMINAL AREA2A 3A

RADAR VECTORS

WITHIN 5 TO 10 NM OF FAF

NOTE: ALL POWER SETTING ARE APPROXIMATE.VAP = VREF + WIND FACTOR BUT NOT LESSTHAN VREF + 10 KTS.

Precision Approach and Landing


Maneuvers

Precision Approachand Landing

3D-37


Maneuvers

LANDING ASSURED � POWER – 75% N1

RADAR VECTORS

WITHIN RANGE � POWER – 80% N1 � CONFIGURATION – SLATS + FLAPS 15° � AIRSPEED – 180 KIAS � APPROACH CHECKLIST – COMPLETE � POWER – 80% N1 � AIRSPEED – 160 KIAS

WITHIN 3 MINUTES OF IAF � AIRSPEED – 180 KIAS � POWER – 80% N1 � CONFIGURATION – SLATS + FLAPS 15° � APPROACH CHECKLIST – COMPLETE � AIRSPEED – 160 KIASIAF OUTBOUND

� TIMING – START

THRESHOLD � AIRSPEED – VREF + WIND FACTOR � POWER – REDUCE TO IDLE

6

TOUCHDOWN � BRAKES – AS REQUIRED � AIRBRAKES – EXTEND � REVERSE THRUST – AS REQUIRED

RADAR VECTORS

TERMINAL AREA2A 3A

7

3

5

WHEN GLIDESLOPE ALIVE � CONFIGURATION - GEAR – DOWN � BEFORE LANDING CHECKLIST – COMPLETE

AT GLIDESLOPE INTERCEPT � TIMING – START � AIRSPEED – VAP

4

RADAR VECTORS



1A

1

2

ONE DOT BELOW GLIDESLOPE � POWER – DECREASE APPROXIMATELY 5% N1 � CONFIGURATION – SLATS + FLAPS 15°/30°

NOTE: ALL POWER SETTING ARE APPROXIMATE.VAP = VREF + CONFIGURATION CORRECTION(+5 KTS FOR S+30°; +10 KTS FOR S+15°) + WINDFACTOR BUT LESS THAN VREF + CONFIGURATIONCORRECTION + 10 KTS.

Single Engine Precision Approach and Landing


Maneuvers

3D-39

Single EnginePrecision Approachand Landing


Maneuvers

RADAR VECTORS

TERMINAL AREA

TOUCHDOWN � BRAKES – AS REQUIRED � POWER – IDLE � AIRBRAKE – EXTEND � REVERSE THRUST – AS REQUIRED

WITHIN 3 MINUTES OF IAF � POWER – 60% N1 � AIRSPEED – 180 KIAS � CONFIGURATION – SLATS + FLAPS 15° � APPROACH CHECKLIST – COMPLETE � POWER – 63 TO 65% N1 � AIRSPEED – 160 KIAS

THRESHOLD � AIRSPEED – VREF + WIND FACTOR � POWER – SLIGHTLY REDUCE (3 TO 4% N1)

LANDING ASSURED � CONFIGURATION – SLATS + FLAPS 52° � AIRSPEED – VAP

AT FAF � TIMING – START � DESCENT – 1,500 FPM MAXIMUM � POWER – 58% N1


RADAR VECTORS

WITHIN RANGE � POWER – 60% N1 � AIRSPEED – 180 KIAS � CONFIGURATION – SLATS + FLAPS 15° � APPROACH CHECKLIST – COMPLETE � POWER – 63 TO 65% N1 � AIRSPEED – 160 KIAS

RADAR VECTORS



6

7

4

2A 3A

8

5

3

1A

1

PRIOR TO FAF � POWER – 72% N1 � CONFIGURATION - SLATS + FLAPS 30° - LANDING GEAR – DOWN � AIRSPEED – 140 KIAS � BEFORE LANDING CHECKLIST – COMPLETE

AT MDA � ALTITUDE – MAINTAIN � POWER – 72% N1 � AIRSPEED – 140 KIAS (VREF + 20 [MINIMUM])

2

NOTE: ALL POWER SETTING ARE APPROXIMATE.VAP = VREF + CONFIGURATION CORRECTION(IF ANY) + WIND FACTOR, BUT NOT LESS THANVREF + CONFIGURATION CORRECTION + 10 KTS.

Non-Precision Approach and Landing


Maneuvers

3D-41

Non-PrecisionApproach and Landing


Maneuvers

RADAR VECTORS

TERMINAL AREA

TOUCHDOWN � BRAKES – AS REQUIRED � AIRBRAKES – EXTEND � REVERSE THRUST – AS REQUIRED

WITHIN 3 MINUTES OF IAF � POWER – 80% N1 � AIRSPEED – 180 KIAS � CONFIGURATION – SLATS + FLAPS 15° � APPROACH CHECKLIST – COMPLETE � AIRSPEED – 160 KIAS

THRESHOLD � AIRSPEED – VREF + WIND FACTOR � POWER – REDUCE TO IDLE

LANDING ASSURED � POWER – REDUCE AS REQUIRED � CONFIGURATION – SLATS + FLAPS 15°/30° � AIRSPEED – VAP

AT FAF � TIMING – START � DESCENT – 1,500 FPM MAXIMUM � POWER – 78% N1 � CONFIGURATION – SLATS + FLAPS 30°


RADAR VECTORS

WITHIN RANGE � POWER – 80% N1 � AIRSPEED – 180 KIAS � CONFIGURATION – SLATS + FLAPS 15° � APPROACH CHECKLIST – COMPLETE � AIRSPEED – 160 KIAS

RADAR VECTORS



6

7

4

2A 3A

8

5

3

1A

1

PRIOR TO FAF � POWER – 93% N1 � CONFIGURATION – SLATS + FLAPS 15°/30° – LANDING GEAR – DOWN � AIRSPEED – 140 KIAS (VREF + 20 KTS [MINIMUM]) � BEFORE LANDING CHECKLIST – COMPLETE

AT MDA � ALTITUDE – MAINTAIN � AIRSPEED – 140 KIAS (VREF + 20 [MINIMUM]) � POWER – 93% N1

2


Single Engine Non-Precision Approach and Landing


Maneuvers

3D-43

Single EngineNon-PrecisionApproach and Landing


Maneuvers

BASE LEG � CONFIGURATION – SLATS + FLAPS 30° � RATE OF DESCENT – ESTABLISH AT 500 TO 600 FPM � AIRSPEED – 140 KIAS

THRESHOLD – BALKED LANDING � FLIGHT DIRECTOR – GO-AROUND MODE � PITCH – 14° � POWER – MAXIMUM � CONFIGURATION – SLATS + FLAPS 15° (BY STEP) � AIRSPEED – VREF MINIMUM

LANDING ASSURED � POWER – AS REQUIRED � CONFIGURATION – FLAPS 52° � AIRSPEED – VAP

AT SAFE ALTITUDE (400 FT AGL MINIMUM) � PITCH – 14° � POWER – CLIMB N1 � CONFIGURATION – CLEAN � AIRSPEED – VREF + 35

POSITIVE RATE OF CLIMB � CONFIGURATION - GEAR – UP � PITCH – 14° � CONFIGURATION – SLATS + FLAPS 15°

THRESHOLD – LANDING � AIRSPEED – VREF + WIND FACTOR � POWER – SLIGHTLY REDUCE (3 TO 4%)

TOUCHDOWN � BRAKES – AS REQUIRED � POWER – IDLE � AIRBRAKES – EXTEND � REVERSE THRUST – AS REQUIRED

DESCENT � DESCENT CHECKLIST – COMPLETE � AIRSPEED BUGS – SET TO VREF

5

4

1

6A

6B

8B

7A

BEFORE PATTERN ENTRY (DOWNWIND 1,500 FT) � POWER – 60% N1 � CONFIGURATION – SLATS + FLAPS 15° � AIRSPEED – 180 KIAS � APPROACH CHECKLIST – COMPLETE � POWER – 63 TO 65% N1 � AIRSPEED – 160 KIAS

2ABEAM LANDING THRESHOLD � CONFIGURATION - GEAR – DOWN � BEFORE LANDING CHECKLIST – COMPLETE

3

NOTE: ALL POWER SETTING ARE APPROXIMATE.VAP = VREF + WIND FACTOR BUT NOT LESSTHAN VREF + 10 KTS.

7B

Visual Approach/Balked Landing


Maneuvers

3D-45

Visual Approach/Balked Landing


Maneuvers

Circling Approach

RECOMMENDATIONS � CONFIGURATION – SLATS + FLAPS 30° � GEAR – DOWN � AIRSPEED – 140 KTS (VREF + 20 MINIMUM) (MAINTAIN CONSTANT SPEED FOR TIMING) � POWER – 72% N1 � F/D ALTITUDE HOLD – SELECT � F/D HEADING – SELECT � USE OF AUTOPILOT IS OPTIONAL � SLIGHT ADJUSTMENTS TO TIME OR HEADING MAY BE USED TO ADJUST FOR WIND

ENTER BASIC PATTERN AS APPROPRIATEFOR AIRCRAFT POSITION

START TIMING ABEAM APPROACH END OF RUNWAY

START TURN TO FINAL, MAXIMUM 30° BANK

WITH RUNWAY IN SIGHT AND IN POSITION TO MAKEA NORMAL DESCENT TO LANDING� BEFORE LANDING CHECKLIST – COMPLETE� DESCENT FROM MDA – BEGIN� AIRSPEED – VAP� IF NOT IN A POSITION TO MAKE A NORMAL LANDING - GO-AROUND – EXECUTE� AT THRESHOLD - AIRSPEED – VREF + WIND FACTOR

NOTES � BASED ON 30° BANK TURNS � USE CATEGORY D MINIMUMS � A MINIMUM OF 300 FT OBSTACLE CLEARANCE PROVIDED AT CATEGORY D CIRCLING MINIMUMS (MDA) TO 2.3 NM FROM ANY RUNWAY.

CAUTION: FAR 91.175 requiresimmediate execution of themissed approach procedurewhen an identifiable part ofthe airport is not distinctlyvisible to the pilot during thecircling maneuver, unless theinability to see results from anormal bank of the aircraftduring the approach.

3

15 SEC

30 SEC

4

2

1

30° BANK

15 SEC� RUNWAY IN SIGHT AND WITHIN CIRCLING APPROACH AREA� TURN 45° FROM RUNWAY CENTERLINE� TIMING – START� AFTER 30 SECONDS – TURN TO DOWNWIND

� FLY 90° TO RUNWAY� START TIMING CROSSING RUNWAY CENTERLINE� AFTER 15 SECONDS TURN TO DOWNWIND

� FLY OVER RUNWAY� WHEN ESTABLISHED ON CENTERLINE - 30° BANK TURN TO DOWNWIND

� TURN OVER RUNWAY� AT RUNWAY END, 30° BANKED TURN TO DOWNWIND

ABEAM POINT

1

2

3

4

BASIC CIRCLING PATTERN

1

1

1

45°

KEY POINT



Maneuvers

3D-47

Circling Approach


Maneuvers

MISSED APPROACH � FLIGHT DIRECTOR – GO-AROUND MODE � PITCH – 14° � POWER – MAXIMUM � CONFIGURATION – SLATS + FLAPS 15° (BY STEPS) � AIRSPEED – VREF MINIMUM

AT SAFE ALTITUDE (400 FT AGL MINIMUM) � POWER – CLIMB N1 � CONFIGURATION - CLEAN � AIRSPEED – VREF + 35

POSITIVE RATE OF CLIMB � CONFIGURATION - GEAR – UP

ADVISE ATC1

2

4

3

Go-Around/Missed Approach


Maneuvers

3D-49

Go-Around/Missed Approach


Maneuvers

WITHIN 3 MINUTES OF IAF � POWER – 60% N1 � SLATS/FLAPS – UP � APPROACH CHECKLIST – COMPLETE � AIRSPEED BUGS – SET TO VREF, VAP � SPEED – REDUCE TO 180 KTS � POWER – 63% N1

IAF OUTBOUND � TIMING – START � SPEED – 180 KTS

THRESHOLD � SPEED – VREF + 35 + WIND FACTOR

5

TOUCHDOWN � BRAKES – DEPLOY � BRAKES – AS REQUIRED � REVERSE THRUST – APPLY

6

3

GLIDESLOPE ALIVE � GEAR – DOWN � LANDING CHECKLIST – COMPLETE � SPEED – SLOW TO VAP

AT GLIDESLOPE INTERCEPT � TIMING – START � DESCENT – BEGIN � POWER – REDUCE TO MAINTAIN VAP (APPROXIMATELY 50% N1)

4


1

2


Clean Wing Approach and Landing


Maneuvers

3D-51

Clean WingApproachand Landing


Maneuvers

MISSED APPROACH � POWER – APPLY T/O POWER � PITCH – ROTATE TO 14°

WITHIN 3 MINUTES OF IAF � POWER – 60% N1 � SPEED – REDUCE TO 180 KTS � SLATS – SELECT � APPROACH CHECKLIST – COMPLETE � SPEED – REDUCE TO 160 KTS � POWER – 63% N1

IAF OUTBOUND � TIMING – START � SPEED – 160 KTS

THRESHOLD � SPEED – VREF + 15 + WIND FACTOR

6

TOUCHDOWN � AIRBRAKES – DEPLOY � BRAKES – AS REQUIRED � REVERSE THRUST – APPLY

7

3

5

WHEN GLIDESLOPE ALIVE � GEAR – DOWN � LANDING CHECKLIST – COMPLETE � SPEED – SLOW TO VAP

AT GLIDESLOPE INTERCEPT � TIMING – START � SPEED – VAP

4

PROCEDURE TURN INBOUND � SPEED – 160 KTS

1

2


Slats-Only Approach and Landing


Maneuvers

3D-53

Slats-OnlyApproachand Landing


FlightPlanning

Chapter 4

Flight planning is critical to flight safety.

This section parallels groundschool instruction and providesinstruction in and examples of flight planning procedures. Asample problem is used throughout the chapter with the appro-priate chart opposite each lesson’s procedural steps.

Problem examples and data drawn from the charts are italicized.


Flight Planning

Table ofContents

General Information . . . . . . . . . . . . . . . . . 4-7

Limiting Factors . . . . . . . . . . . . . . . . . . . . 4-7

Structural Limits . . . . . . . . . . . . . . . . . . . 4-7

Runway and Climb Performance Limits . . . . . . . 4-7

Trip Fuel Loads . . . . . . . . . . . . . . . . . . . . 4-8

Fuel Requirements . . . . . . . . . . . . . . . . . . . 4-8

Reserve Fuel . . . . . . . . . . . . . . . . . . . . . 4-8

Fuel Burn Rates . . . . . . . . . . . . . . . . . . . 4-8

Takeoff and Landing Performance . . . . . . . . . . . 4-9

Trip Planning Data . . . . . . . . . . . . . . . . . . 4-11

Definitions . . . . . . . . . . . . . . . . . . . . . . 4-13

Weight and Balance . . . . . . . . . . . . . . . . . 4-19

Basic Empty Weight and Moment . . . . . . . . . . 4-22

Passenger and Crew Weights . . . . . . . . . . . . 4-24

Miscellaneous Supplies and Baggage . . . . . . . . 4-26

Baggage Weights and Adjustments . . . . . . . . . 4-28

Zero Fuel Weight Computations . . . . . . . . . . . 4-30

Zero Fuel Weight Limits . . . . . . . . . . . . . . . . 4-32

Takeoff Weight and Moment . . . . . . . . . . . . . 4-34

Center of Gravity . . . . . . . . . . . . . . . . . . . 4-36

Trip Planning Data . . . . . . . . . . . . . . . . . . 4-39

Estimated Trip Plan . . . . . . . . . . . . . . . . . . 4-41

Initial Cruise Altitude . . . . . . . . . . . . . . . . . 4-42

Alternate Reserve Fuel Requirement(Landing Weight) . . . . . . . . . . . . . . . . . . . 4-44

Trip Fuel Requirement and Takeoff Weight . . . . . . 4-46

Final Fuel Calculation . . . . . . . . . . . . . . . . . 4-46

Takeoff Performance . . . . . . . . . . . . . . . . 4-49

Runway and Climb Weight Limits . . . . . . . . . . . 4-50


TOLD Card . . . . . . . . . . . . . . . . . . . . . . 4-51

Takeoff Gross Weight . . . . . . . . . . . . . . . . . 4-52

Takeoff Weight Limited by Climb Requirements . . . 4-53

Balanced Field Length . . . . . . . . . . . . . . . . 4-54

VMBE Speed Computations . . . . . . . . . . . . . . 4-56

Tire Speed . . . . . . . . . . . . . . . . . . . . . . 4-58

V1 Speed Computations . . . . . . . . . . . . . . . 4-60

Takeoff Speed and Nose-Up Attitude(VR, ATT, V2, Slats + Flaps 15°) . . . . . . . . . . . 4-62

Takeoff Clean Wing Speed (VFR) . . . . . . . . . . . 4-62

Final Takeoff Climb Speed (VFS) . . . . . . . . . . . 4-64

Enroute Climb Speed (VENR) . . . . . . . . . . . . . 4-64

N1 Setting Computations – Takeoff Thrust . . . . . . 4-66

N1 Setting Computations – Climb Thrust . . . . . . . 4-66

Time to 100 Knots . . . . . . . . . . . . . . . . . . 4-68

Cruise Climb N1 Setting . . . . . . . . . . . . . . . . 4-68

Takeoff – Slats-Only . . . . . . . . . . . . . . . . . 4-70

Maximum Takeoff Weight Limited byClimb Requirements – Slats-Only . . . . . . . . . . 4-70

Takeoff Climb: First Segment – Slats-Only . . . . . . 4-72

Balanced Field Length – Slats-Only . . . . . . . . . 4-72

Second Segment Climb – Slats-Only . . . . . . . . . 4-74

V1 Safety Speed – Slats-Only . . . . . . . . . . . . 4-76

Takeoff Speeds and Nose-Up Attitude,VR, Attitude, V2, Slats-Only . . . . . . . . . . . . . . 4-78

Maximum Tire Speed – Slats-Only . . . . . . . . . . 4-80

Clean Wing Speed – Slats-Only . . . . . . . . . . . 4-82

Slats-Only TOLD Card Calculation . . . . . . . . . . 4-82

Cruise and Landing Performance . . . . . . . . . 4-83

Climb Data . . . . . . . . . . . . . . . . . . . . . . 4-84

Cruise Data . . . . . . . . . . . . . . . . . . . . . . 4-86

Descent Data . . . . . . . . . . . . . . . . . . . . . 4-88


Flight Planning

VREF/VAP Speed Computations . . . . . . . . . . . . 4-88

Approach Climb Gradient/Speed . . . . . . . . . . . 4-90

Missed Approach/Go Around Clean Wing Speed . . 4-92

Landing Distance . . . . . . . . . . . . . . . . . . . 4-92

N1 Setting – Missed Approach or Go Around . . . . . 4-94

Supplemental Information . . . . . . . . . . . . . 4-97

Adjustment of Weight and Balance . . . . . . . . . . 4-98

Obstacle Clearance . . . . . . . . . . . . . . . . . 4-100

Maximum Takeoff Weight Limited byObstacle Clearance Requirement . . . . . . . . . . 4-104

Takeoff Noise Reduction Procedure – Maximum Takeoff Weight (18,300 Lbs – Basic Without SB F10-0052),or (18,740 Lbs – Basic With SB F10-0052) . . . . . 4-106

Takeoff Noise Reduction Procedure – Maximum Takeoff Weight (19,300 Lbs – Basic With SB F10-0238) . . . . . . . . . . . . . . 4-106

Contaminated Runway Operations . . . . . . . . . 4-108

Hydrodynamic Pressures . . . . . . . . . . . . . 4-108

Hydroplaning . . . . . . . . . . . . . . . . . . . . 4-108

Drag . . . . . . . . . . . . . . . . . . . . . . . . 4-108

Braking . . . . . . . . . . . . . . . . . . . . . . . 4-110

Summary: Contaminated Runway Operations . . . 4-110

Conclusions . . . . . . . . . . . . . . . . . . . 4-110

Recommendations . . . . . . . . . . . . . . . . 4-111

Landing – Contaminated Runway . . . . . . . . . 4-112

Takeoff – Contaminated Runway . . . . . . . . . 4-112

Landing – Slippery Runway . . . . . . . . . . . . 4-112

Takeoff – Slippery Runway . . . . . . . . . . . . . 4-113

Contaminated Runway Calculations . . . . . . . . 4-113

Equivalent Water Depth . . . . . . . . . . . . . . 4-114

Effect of Precipitation on Takeoff Distance . . . . . 4-116

Effect of Precipitation on Landing Distance . . . . 4-118


Flight Planning

Flight planning begins with data gathered on:� payload� airport data for departure, arrival, and alternate� navigation information for departure, enroute, and arrival� weather for departure, enroute, destination, and alternate� NOTAMs.

The operator provides payload information. Airport data, such aselevation, available runways, and runway length, width, and light-ing, is in Jeppesen or NOS charts. If runway bearing strength isquestioned, other sources of airport information may be required.Navigation information is in Jeppesen or NOS charts. In some cases,additional information is needed (e.g., NAT tracks if crossing theNorth Atlantic).

Obtain weather conditions for departure, arrival, and alternate airportsas well as enroute weather, winds, and temperatures aloft from theFAA Flight Service System by telephone, computer, or a flight plan-ning vendor. The same sources provide NOTAMs.

Unless otherwise noted, the charts in this section are from the Falcon10/100 Airplane Flight and Operational Instructions Manuals.

Limiting FactorsTo initiate flight planning, first consider possible limiting factorssuch as structural, runway, and climb performance limits and tripfuel loads.

Structural LimitsStructural limits restrict heavy payloads. A dense load (e.g., goldbars) means weight is the primary planning concern; begin flightplanning with weight and balance to determine payload restrictionsrequiring significant plan changes (e.g., two aircraft, load left atdeparture).

Runway and Climb Performance LimitsIf takeoff, runway, or climb performance limits affect the maximumtakeoff weight, fuel or payload may need adjustment. Determine themaximum takeoff weight, then adjust factors to accommodate it(e.g., an enroute refueling stop).

GeneralInformation


Trip Fuel LoadsThe total capacity or the trip requirements limit the trip fuel load; carryno more than the required amount plus desired reserves. Determineweight and balance to find the zero fuel weight, then calculate the tripfuel, and add it to the zero fuel weight to determine the takeoff grossweight. Finally, examine takeoff performance.

Fuel RequirementsIn practical operation, estimate fuel by rules of thumb relating toexperience with aircraft operation. Large variances in fuel burn arebased on factors such as gross weight, selected cruise altitude, andspeed.

Reserve FuelReserve fuel requirements vary with the destination location andtraffic density as well as the weather. While a 1,200 lb reserve isadequate for a low traffic density and good VMC destination, antici-pated traffic delay or IMC weather increases fuel reserves to 2,000lbs or more. According to the FARs, 1,200 lbs of fuel at the destinationis adequate to proceed to an alternate approximately 120 NM (stillair) from the primary destination and arrive with the required reserves.

Fuel Burn RatesHourly fuel burn rates, Table 4-A , vary with takeoff weight, climbschedule, cruise altitude, and cruise Mach. For takeoffs near maxi-mum takeoff gross weight, the following fuel burns occur if thecruise altitude is maintained at or near the highest possible.

Table 4-A; Fuel Burn Rates

For cruise at low altitudes, such as 15,000 feet MSL, a fuel flow of600 lbs/hr per engine produces approximately 255 KIAS, decreas-ing to 220 KIAS at 1,000 ft MSL.

Hour

Climb260 kts/0.72 M

Cruise0.75 M

(Lbs Fuel Burned)

Climb300 kts/0.80 M

Cruise0.80 M

(Lbs Fuel Burned)

1 1,628 1,701

2 1,320 1,250

3 1,240 1,130

4 1,200 1,130


Flight Planning

Takeoff and Landing PerformanceIn general, runway lengths are shorter for slats + flaps 15° than forslats-only takeoffs.

Slats + flaps with speed increase and slats-only takeoffs producebetter climb performance than slats + flaps 15° takeoffs.

Slats + flaps 52° landings produce the shortest landing distances.


Flight Planning

Use the charts throughout this chapter to plan a long range trip on ahot day from Redbird Airport (RDB), Dallas, Texas, to Teterboro,New Jersey, with an alternate of Harrisburg, Pennsylvania.

The charts are appropriate for an aircraft with SB F10-0052. Assumea basic empty weight of 11,460 lbs and a moment of 400, whichresults in a typical operating weight of 11,800 lbs. The payload is threepassengers, baggage, and supplies weighing 920 lbs.

This example uses the trip fuel load method of determining limitingfactors.

Trip PlanningData

NOTE: Most aircraftincorporate SB F10-0052, whichincreases the maximum takeoffweight (MTOW) to 18,740 lbs.The charts for an 18,740 lbaircraft are the same as for an18,300 lb or 19,400 lb aircraft,except the values are commensu-rate with takeoff weights.

Although moment is recorded inpound-inches, the notationpounds-inches (lbs/in) is notused hereafter because it isunderstood. All numbers arepositive unless otherwise noted.


Flight Planning

To understand flight planning, it is necessary to be thoroughly familiarwith the terms involved. This section reviews the definitions forterms used throughout this chapter.

Accelerate-stop distance– Distance necessary to accelerate the air-craft to a given speed (V1), and to come to a full stop, assuming oneengine fails at V1.

AGL – Above ground level.

Approach climb – The steady gradient of climb with one engineinoperative may not be less than 2.1%. Engine rating is takeoff thrust.The stabilized airspeed is VAC for a slats + flaps 15° or slats + flaps30° configuration as found in the AFM or the SimuFlite OperatingHandbook.

Balanced field length– Distance obtained by choosing the enginefailure speed V1 so that takeoff distance and accelerate-stop distanceare the same.

Basic empty weight– Weight of airframe, powerplant, interioraccommodation, systems, and equipment that are an integral part ofa given version (i.e., the weight without usable fuel, including allfluids contained in closed systems, the unusable and undrainablefuel and the engine oil).

Basic operating weight– Basic empty weight plus operational itemsand crew.Clearway – Area beyond the runway:� not less than 500 ft (152 m) wide� centrally located on the extended centerline of the runway� under the control of the airport authorities� upward slope not exceeding 1.25%� above which no object nor terrain protrudes (threshold lights

should be at each side of the runway and not more than 26 inches(65 cm) high.

Configurations – The configurations referred to by name in theAFM charts correspond to the settings in Table 4-B on the followingpage (Figure 4-1 , page 4-15).

Engine failure speed associated with balanced field length–Distance obtained by choosing the engine failure speed V1 so thattakeoff distance and accelerate-stop distance are the same. V1 mustbe greater than V1 MIN and less than VMBE and VR. If the deter-mination of V1 gives a value outside one of these limits, V1 mustbe selected equal to the limit value. The field length found in thechart is the higher of either the takeoff distance or the accelerate-stop distance associated with this limit value of V1.

Definitions

High Lift


Final segment– Segment extending from the end of the transitionsegment to a height no less than 1,500 ft AGL. The gradient of climbmay not be less than 1.2%. Engine power is reduced from takeoff tomaximum continuous.First segment– Segment extended from the point at which the air-craft becomes airborne to the point at which gear retraction isachieved. The climb gradient without ground effect may not be lessthan 0.0% (positive). The speed increases from VLOF to V2, to beattained at a height not greater than 35 ft (10.7 m).Gross climb gradient– Demonstrated ratio expressed in percentof change in height to horizontal distance travelled.Height – Vertical distance from the lower point of the aircraft to theairport surface.IAS – Indicated airspeed. Airspeed indicator reading, as installedin the aircraft.Landing climb – The steady gradient of climb in landing configu-ration with all engines operative may not be less than 3.2%. Enginerating is takeoff thrust. The stabilized airspeed is 1.3 VSO (VREF)for slats + flaps 52° landing configuration.

FlightCondition

Engines

Configuration

Gear Gross ClimbEffectiveSpeed

No.

Speed

FAR Conditions

Power Setting

S + Flaps 15°or Slats

Takeoff Down 0.0 0 to VLOF2 0 to VLOFTakeoff


First Segment Down 0.0 VLOF to V21 VLOF to V2Takeoff


Second Segment Up 2.4 V21 ≥ V2Minimum

Takeoff

In RetractionTransient Segment Up Available 1.2 In Acceleration1 0.0Takeoff

CleanFinal Takeoff Up 1.2 1.4 VS1 ≥ 1.25 VSMaximumContinuous

CleanEnroute Climb Up 0.0 1.5 VS1 —MaximumContinuous

S + Flaps 30°Approach Climb

Up 2.1 1.3 VS1 ≤ 1.5 VSTakeoff

S + Flaps 15° Up 2.1 1.35 VS1 ≤ 1.5 VSTakeoff

S + Flaps 52°Landing Climb Down 3.2 1.3 VS2 ≤ 1.3 VSTakeoff

S + Flaps 52°Landing Down 0.0 1.3 VS2 ≥ 1.3 VSIdle

Table 4-B; Configuration Settings

FINALSEGMENT

SECONDSEGMENT

FIRSTSEGMENT

V2VLOF

400 FTMINIMUM

1,500 FTMINIMUM

ENGINES

THRUST

AIRSPEED

LANDINGGEAR

FLAPS ANDSLATS

MIN. T.O. FLIGHTPATH CLIMB

GRADIENT

ONE INOPERATIVE

MAX CONT THRUST

VFS

RETRACTED

RETRACTED

1.2%2.4%POSITIVE

TAKEOFF SETTING

V2ACCELERATING

DOWN

TAKEOFF THRUST 5 MINUTES MAXIMUM

RETRACTING

* GRADIENT AT ALL POINTS 0.8% LESS THAN TAKEOFF FLIGHT PATH

GROUNDROLL

GEARUP

35 FT

TAKEOFFDISTANCE

REFERENCEZERO

TRANSITION(ACCELERATION)

TAKEOFFFLIGHT PATH

*NET TAKEOFFFLIGHT PATH

BOTH

ACCELERATING

RETRACTION

BRAKERELEASE

TOTAL TAKE OFF PATH HORIZONTAL DISTANCE

V1 VR


Flight Planning

Minimum Climb/Obstacle ClearanceOne Engine Inoperative

4-1


Landing distance– Horizontal distance required to land and cometo a complete stop from a point at a height of 50 ft (15.2 m) abovethe landing surface. The stabilized airspeed is VREF for slats + flaps52° landing configuration.

Landing field length – The demonstrated landing distance multipliedby 1.67.

Landing weight – Maximum weight permissible at landing basedon landing field limitations and other associated limitations. Thelanding weight must not exceed the maximum landing weight (MLW)defined by the structural weight limitations.

M – True Mach number. Indicated Machmeter, corrected for staticand total pressure ports position error.

Maximum landing weight (MLW) – Maximum weight at landing,based on structural weight limitations.

MI – Indicated Mach number. Machmeter reading, as installed inthe aircraft.

MSL – Mean sea level.

Net climb gradient – Gross climb gradient reduced by:� 0.8% for takeoff flight path� 1.1% for enroute flight path with one engine inoperative.

Operational takeoff weight – Maximum weight permissible fortakeoff based on takeoff field or flight limitations, or other asso-ciated limitations. The operational takeoff weight must not exceedthe maximum takeoff weight (MTOW) defined by structural weightlimitations.

Payload– Weight of passengers, cargo, and baggage.

QFE – Field pressure. Actual atmospheric pressure at the elevationof the airport.

Second segment– Segment extending from the end of the firstsegment to a height of at least 400 ft. The gradient of climb may notbe less than 2.4%. Aircraft speed is stabilized at V2.

Stopway– Area beyond the runway:� no less wide than the runway� centrally located upon the extended centerline of the runway

� designated by the airport authorities for use in decelerating theaircraft during an aborted takeoff

� able to support the aircraft without causing structural damage.


Flight Planning

Takeoff distance– Greater horizontal distance along the takeoffpath from start of takeoff roll to the point at which the aircraft is 35ft (10.7 m) high with either:� one engine failure at V1

� all engines operating (factored by 115%).

Takeoff run (takeoff with clearway) – Greater horizontal distancealong the takeoff path from start of takeoff roll to a point equidistantbetween the point at which VLOF is reached and the point at whichthe aircraft is 35 ft (10.7 m) high with either:� one engine failure at V1

� all engines operating (factored by 115%).

TAT – Total air temperature. Outside air temperature, includingadiabatic compression rise; assumed recovery factor is 99%.

VLOF – Liftoff speed. Speed at which the aircraft first becomes air-borne.

VMBE – Maximum brake energy speed. Maximum engine failurespeed V1 at which the maximum demonstrated brake energy is notexceeded. V1 must not exceed VMBE.

VR – Rotation speed. Speed at which rotation is initiated.

VREF – Reference speed. Minimum speed at the height of 50 ft (15.2m) during a normal landing. VREF should not be less than 1.3 VSfor landing configuration (slats + flaps 52°).

VS– Stalling speed. Minimum speed obtained during the stall man-euver in the specific configuration.

VSO – Stalling speed. Minimum speed obtained during the stallmaneuver with gear and flaps fully extended.

V1 – Engine failure speed. Speed at which one engine is assumedto become suddenly inoperative during takeoff, after which thetakeoff must be continued. This speed is always greater than VMCG(100 kts).

V2 – Takeoff safety speed. Initial climb speed reached by the air-craft before it is 35 ft (10.7 m) above the takeoff surface with oneengine inoperative.

Wind components– velocity and direction recorded at the heightof 50 ft (15.25 m):� headwind or tailwind – component parallel to the flight path� crosswind – component perpendicular to the flight path.

Zero fuel weight (ZFW) – Certified empty weight plus payload.The zero fuel weight must not exceed the maximum zero fuel weightdefined by structural weight limitations (i.e., ZFW MZFW).


Flight Planning

Weight andBalance

Precise weight computations are required to operate the aircraft with-in limitations and for performance calculations. Balance computa-tions are required to operate the aircraft within center of gravitylimitations.

The Weight and Loading subsection of AFM Section 1, Limitations,states that maximum weight limits are determined by structurallimitations that may in turn be reduced by performance limitations.The section states the center of gravity limits must stay within limitsgraphically indicated. This section then refers to Performance ManualSection 2, Loading, for weight and balance determinations.

Structural weight limitations are shown below in Tables 4-C , 4-D,and 4-E.

Table 4-C; Aircraft with Minimum Takeoff Weight of 18,300 lbs(Basic without SB F10-0052)

Weight Pounds Kilograms

Maximum Ramp Weight 18,300 8,300

Maximum Takeoff Weight 18,300 8,300

Maximum Landing Weight 17,200 7,800

Maximum Zero Fuel Weight 12,460 5,650

Minimum Flight Weight 9,920 4,500

Table 4-D; Aircraft with Maximum Takeoff Weight of 18,740 lbs(Basic with SB F10-0052)








Loading Manual Section 2, Balance Control, states that the center ofgravity position is determined and checked by means of a weightand balance diagram. There are two steps to the procedure:

1. Determine that the zero fuel weight and moment are within thezero fuel weight limits.

2. Determine that the takeoff weight (ZFW + fuel weight) andmoment are within the Weight and Balance Diagram limits, thendetermine the center of gravity position in percent MAC.

The Loading Manual provides the following to determine weightand center of gravity:

� definitions� fuel moment graphs

� loading examples for various airplane configurations andblank loading forms for those configurations

� change in CG position for all configurations� equipment lists� weight and balance diagrams� numerical weight and balance data tables.

The information is in both metric and U.S. units. The proceduresused are exactly the same, whether in metric or U.S. units.

The pilot must use a configuration variation form that correspondsto a particular Falcon 10/100.

Table 4-E; Aircraft with Maximum Takeoff Weight of 19,300 lbs(Basic with SB F10-0238)








Flight Planning

For this example, use Version 1 Weight and Balance form, U.S units,Three Windows on RH Side (Figure 4-2 ).

The form contains an illustration of the aircraft in the center; on theleft are tables of moments for standard weights for specific loca-tions; on the right is a loading schedule.

The loading schedule lists the location in the first column. The sec-ond column contains the weight for the area or item. The third andfourth columns contain the moments for the weights and expressnegative and positive, respectively. When a certain moment cannothave a negative or positive sign, that column for that row is shadedto prevent entry.

In this example, some of the specific locations are BAGGAGECOMP. A, COAT RACK (COMP. B), and GALLEY.

Weight and Balance(Version No. 1, U.S. Units)

4-2

Flight Planning

Weight andBalance

4-21 4-22 For SimuFlite training only Falcon 10/100March 2001

Basic Empty Weight and MomentObtain the basic empty weight and moment from the latest aircraftweighing form or from the latest weight and balance computationform after an aircraft alteration. Empty weight changes are made onthe weight and balance form if added or removed equipment affectsthe basic empty weight.

1. Enter the basic weight and moment on the Determination of CGLocation/Load Distribution form (Figure 4-3 ).

For this example, the basic aircraft weight is 11,460 lbs with acorresponding moment of 400. Moments are entered in thousands.The figures 11,460 and 0.4 are entered in the BASIC EMPTYWEIGHT row in the WEIGHT column and the + MOMENTScolumn, respectively.

2. Record changes in empty weight and moment in the EMPTYWEIGHT CHANGES row, then add them to or subtract themfrom the BASIC EMPTY WEIGHT and MOMENT row. Recordthe results in the EMPTY WEIGHT row.

For this example, there are no changes to the basic empty weight.The entry in the EMPTY WEIGHT row is the same as in theBASIC EMPTY WEIGHT row.

NOTE: Maintenance usuallyrecords changes to the basicempty weight and moment withan alteration form.

1111,,446600..44

1111,,446600 ..44


Flight Planning


4-3


Passenger and Crew WeightsThe manufacturer assumes standard passenger and crew weights.The assumption is that variations in these weights from the actualweights are insufficient to significantly affect weight and balance.

Assume crew and passenger weights of 170 and 165 lbs,respectively.

1. On the weight and balance configuration form (Figure 4-4 ), linethrough the weight and moment areas of seats not occupied.

For this example, the two crewmembers occupy the pilot andcopilot seats; the jump seat is not occupied. Three passengersare in seats 1, 3, and 4. Corresponding weight and moment rowsfor unoccupied seats 2, 5, 6, and 7 are lined through.

NOTE: In rare cases, it ispossible to exceed forward CGlimits on some Falcon 10/100s ifa very heavy individual sits on aforward seat.

1111,,446600 ..44

1111,,446600 ..44

4-4


Flight Planning



Miscellaneous Supplies and BaggageThe moments for loads in various areas are determined from theknown weight of items placed in the areas.

Assume the following weights:

Galley (full stocks and aircraft documentation) 70 lbsBar (full stocks) 20 lbsBaggage A 275 lbsBaggage B 30 lbsBaggage D 30 lbs

1. Locate the corresponding rows on the loading schedule on theright side of the weight and balance form and record the weights(Figure 4-5 ).

2. Find the corresponding weight/moment table on the left of theform. Interpolate if necessary.

3. Record the moment for the applicable weights in the rowsprovided on the loading schedule.

The moments are:Galley (full stocks and aircraft documentation) -8.0Bar (full stocks) -2.4Baggage A +10.6Baggage B -3.0Baggage D +3.3

4. Record the values.

1111,,446600 ..44

1111,,446600 ..44


Flight Planning

Configuration Form (Baggage)(Version No. 1, U.S. Units)

4-5

227755

3300

33..00

1100..66

33..33

3300

7700

2200

88..00

22..44


Baggage Weights and AdjustmentsBaggage is loaded in the baggage compartment or in the interior.

If baggage loaded in the interior is stowed in one of the locationsalready calculated, the new weight is calculated and the new momentdetermined from the moment table. The revised values then replacethe existing values in the corresponding row.

The following scenario is for illustration only, and the figures arenot used in the example problem.

A passenger requests additional baggage weighing 60 lbs bestored in the coat rack (compartment B). Add that 60 lbs to the30 lbs presently stored in compartment B for a total of 90 lbs,and determine a new moment from the chart on the left of theform. The revised baggage compartment B weight, 90 lbs, and therevised moment, -9.2, are recorded in the space (**) on the weightand balance form on the opposite page (Figure 4-6 ).

1111,,446600 ..44

1111,,446600 ..44

227755

3300

33..00

1100..66

33..33

3300

7700

2200

88..00

22..44


Flight Planning


4-6

�1

�1 REVISED WEIGHT IS 90 LBS.,AND REVISED MOMENT IS -9.2


Zero Fuel Weight ComputationsComplete the zero fuel weight computations. Record the results onthe loading schedule (Figure 4-7 ).

1. Total the columns in the loading schedule, then enter the resultsin the TOTAL row.

For this example, the weight column total is 12,720 lbs. Thenegative and positive moment columns total 86.6 and 14.3,respectively.

2. Subtract the total negative moment from the total positivemoment. If the value is negative, enter the number without a signin the left box below the TOTAL row; if the number is positive,enter the number without a sign in the right box below the TOTALrow.

When -86.6 is subtracted from 14.3, the result is -72.3; enter thisnegative value in the left box below the TOTAL row.

3. Enter the weight from the TOTAL row into the ZERO FUELWEIGHT row, then enter the combined moment result in theZERO FUEL WEIGHT row in the corresponding column.

Repeat the zero fuel weight (12,720 lbs) in the WEIGHT columnof the ZERO FUEL WEIGHT row. Enter the combined momentof -72.3 in the left column of the ZERO FUEL WEIGHT row.

1111,,446600 ..44

1111,,446600 ..44

227755

3300

33..00

1100..66

33..33

3300

4-7

7700

2200

88..00

22..44

1122,,772200 8866..66 1144..33

--7722..33

1122,,772200 --7722..33


Flight Planning



Zero Fuel Weight LimitsThe Zero Fuel Weight and Balance diagram depicts an envelope forzero fuel weight (Figure 4-8 ). The zero fuel weight and momentmust fall within the envelope to be within zero fuel weight limits.

The envelope is subdivided into zones; the zone into which the zerofuel weight and moment falls determines possible fuel loadinglimitations. Zone limitations are defined below the diagram.

Zone 1 is the most desirable because there are no fuel loading limit-ations. If the zero fuel weight and moment is in Zone 1 and normalfuel management procedures occur, the aircraft remains within limitsfor aircraft attitudes between 0° and 11° nose-up pitch. If the zero fuelweight and moment fall within one of the other zones, there arelimits on fuel loading.

1. Enter the chart from the left with the aircraft’s weight (12,720 lbs).Move right to the general area of the applicable moment.

2. Enter the chart from the top at the aircraft moment (-72.3). Movedown to intersect the aircraft weight.

3. Identify the limitations associated with the zone in which theintersection occurs.

In this case, the intersection of the zero fuel weight is withinZone 1. There are no limitations.


Flight Planning

Zero Fuel Weight Envelope

4-8

ZONE 1 No limitation of filling of feeder tanks if wing tanks are full.

ZONE 2 Filling of the feeder tanks is limited to regulation level 2 x 600 lb(2 x 272 kg), if the wing tanks are not full, or to 2 x 1,070 lb(2 x 485 kg), if the wing tanks are full.

ZONE 3 Emptying of the feeder tanks is limited to no less than 2 x 300 lb/2 x 136 kg (illumination of LO FUEL light).

ZONE 4 Filling of wing tanks is limited to 2 x 800 lb (2 x 362 kg).

ZONE 5 Limitation of filling of feeder tanks, and/or wing tanks.

ZONE 6 Filling of wing tanks to full and limitation of filling of feeder tanks.

-72.3

17.8% MAC

12,720 LBS�

�

�


Takeoff Weight and MomentWhen the zero fuel weight and moment and the fuel weight andmoment are combined, the result is the takeoff weight and moment.Use the FUEL table at the lower left of the weight and balance form(Figure 4-9 ) to obtain the moment for the fuel load.

The table is organized with moments for each 1,000 lbs and 200 lbsof fuel. Fuel moments are always positive. Use the fuel moment tothe nearest 200 lb weight. The fuel moment chart provides momentsfor weights to the nearest 200 lbs from the left side of the weightand balance form.

1. Record the weight of the fuel loaded in the WEIGHT column ofthe FUEL row. Enter the fuel moment from the FUEL momenttable in the positive column.

Normally, the fuel load is calculated at this time, then that figureis used to complete the takeoff weight and CG exercise. For con-tinuity, the fuel load for this example is 5,380 lbs. Using theFUEL moment table, determine the moment for 5,400 lbs of fuel(i.e., 84.2). Enter the values 5,380 and 84.2 in the WEIGHT andpositive MOMENT columns of the FUEL row, respectively.

2. Add the zero fuel weight and the fuel weight, and record theresult in the WEIGHT column of the TAKEOFF WEIGHT row.

In this example, the total weight is 12,720 lbs plus 5,380 lbs, or18,100 lbs. Enter this figure in the WEIGHT column of the TAKE-OFF WEIGHT row.

3. Add the zero fuel weight moment and the fuel moment, thenrecord the result (11.9) in the positive MOMENTS column inthe TAKEOFF WEIGHT row.

The zero fuel weight moment may be negative or positive, but thefuel moment is positive. Be sure to observe the positives andnegatives when combining the zero fuel weight and fuel momentsso that the final moment is entered into the correct column.

The combined zero fuel weight moment, -72.3, and the fuelmoment, 84.2, is 11.9; enter this value in the positive momentcolumn of the TAKEOFF WEIGHT row.


Flight Planning

1111,,446600 ..44

1111,,446600 ..44

227755

3300

33..00

1100..66

33..33

3300

4-9

7700

2200

88..00

22..44

1122,,772200 8866..66 1144..33

--7722..33

1122,,772200 --7722..33


55,,338800

1188,,110000

8844..22

1111..99


Center of GravityThe computed takeoff weight and moment are plotted on the BalanceDiagram (Figure 4-10 ) to determine whether the intersection is inthe maximum takeoff weight envelope.

1. Enter the chart from the left with the aircraft’s weight (18,100 lbs).Move right to the general area of the applicable moment.

2. Enter the chart from the top at the aircraft moment (11.9). Movedown to intersect the aircraft weight.

3. From the intersection, parallel the guidelines down to the bot-tom of the chart. Read the % MAC CG for the plotted takeoffweight and moment (25.9%).

4. The CG derived from the Weight and Balance Diagram (25.9) isentered on the loading schedule on the CG POSITION row nextto the “%” sign.


Flight Planning

Balance Diagram

4-10

18,100 LBS

25.9%

11.9

�

�

�

13126°

17174°

35354° 31

309°

ELEV 659 '

ELEV 646'

ELEV 651'

ELEV 657'

3801

'

5452'


Flight Planning

4-11

Redbird Airport, Dallas

The Flight Planning section of the Performance Manual containscharts used to determine the initial cruising altitude, the highestcruise altitude as limited by maximum continuous thrust, the variouscruise modes, and the reserve fuel requirement.

To use the Flight Planning section, the following information isrequired:

� distance from destination to alternate� average head- or tailwind from departure to destination� average head- or tailwind from destination to alternate� expected or desired reserve time for holding or cruise� Zero Fuel Weight (ZFW)� selected cruise mode� selected final cruise altitude� selected climb profile� distance from departure point to destination� average temperature enroute.

This example plans a trip from Redbird Airport, Dallas, Texas,(Figure 4-11 ) to Teterboro, New Jersey (Figure 4-12 , followingpage). The distance is 1,190 NM. The alternate, Harrisburg,Pennsylvania (Figure 4-13 , following page), is 143 NM from thedestination.

Trip PlanningData


17

ELEV 9'

6013'

7000

'

24241°

O19186°

1016°

6061°

ELEV 9'ELEV 7'

ELEV 6'

4-12

Teterboro, New Jersey

ELEV 310'

13127 O

31307°

9501'

4-13

Harrisburg, Pennsylvania

The weight information is the same derived during the precedingweight and balance discussion.

Assume the wind from departure to destination averages a 40 ktheadwind. From the destination to the alternate, the windaverages a 30 kt headwind. The temperature enroute is ISA.

Using the 260 kt (0.72M) climb profile, calculate the initial cruisealtitude by estimating a flight plan and takeoff weight.

Apply standard U.S. reserve fuel/time regulations, and calculateexact trip time and fuel by using the appropriate charts from theAFM and OIM.


Flight Planning

Estimated Trip PlanBefore accomplishing actual flight planning, calculate a trip estimateso that the maximum initial cruise altitude can be computed.

1. Carry over the zero fuel weight from the weight and balancechart. Then calculate the estimated time enroute by assumingthe distance covered in the first hour of flight to be 370 NM, andsubsequent distances at 460 NM per hour, plus estimated head-wind or tailwind.

2. Calculate the hours or parts of hours needed to complete the trip.

For example:Planned Trip 1,190 NMHour 1 370 NM - 40 kt wind = 330 NMHour 2 460 NM - 40 kt wind = 420 NMHour 3 460 NM - 40 kt wind = 420 NMHour 4 (3 min or 0.05 hr) 460 NM - 40 kt wind = 20 NM

The result is three hours and 3 minutes enroute.

3. Calculate the fuel needed by allowing 1,800 lbs for the first hour,1,400 lbs for the second hour, 1,300 lbs for the third hour, and1,200 lbs for every hour thereafter.

Hour 1 1,800 lbsHour 2 1,400lbsHour 3 1,300lbsHour 4 (1/20 hr X 1,200 lbs/hr) 60 lbs

Total trip fuel 4,560 lbs

4. Calculate fuel required for alternate and reserves.

The alternate airport is 143 NM from the destination and thereis an average 30 kt headwind enroute. At an approximate 300kt average speed, the trip should take approximately 25 minutes.A first-hour fuel burn of 1,800 lbs per hour should use about800 lbs of fuel. Add to that a holding time of 45 minutes at approx-imately 1,050 lbs per hour, or 790 lbs, to calculate an estimated1,590 lbs required for alternate and reserves.

Zero Fuel Weight 12,720 lbsEstimated Trip Fuel 4,560lbsEstimated Alternate and Reserve Fuel 1,590lbs

Estimated T/O Gross Weight 18,870 lbs


5. Add the zero fuel weight, the estimated trip, alternate, and reservefuel to derive the estimated takeoff weight (18,870 lbs).

Because the estimated weight is more than the maximum takeoffweight allowed, assume a gross weight of 18,740 lbs, and continuethe calculations.

Initial Cruise AltitudeTo determine initial cruise altitude, use the Two Engine FlightPlanning Maximum Initial Cruise Altitude chart for the type of climbperformed. The chart is divided into groups by the climb mode con-sidered and subdivided by climb rate values. Charts exist for each ofthe following conditions.

Climb Mode 260 kt/0.72M Climb Mode 300 kt/0.25M

Rate of Climb 100 fpm Rate of Climb 100 fpm



1. Use the Initial Cruise Altitude chart (Figure 4-14 ) to calculatethe highest initial cruise altitude at the calculated gross takeoffweight at a climb speed (260 kts/0.72M).

In the example, a weather service advises the temperature athigh altitudes is ISA.

2. Enter the chart from the bottom with the takeoff weight(estimated 18,740 lbs).

3. Move up to intersect the cruise temperature (ISA).

4. Move left to the edge of the chart. Read the maximum initialcruise altitude (38,900). Select the initial cruise altitude at or lessthan the maximum initial cruise altitude, as appropriate for direc-tion of flight (37,000 ft).

4-14

18,740 LBS

38,900 FT

�

�


Flight Planning

Two Engine Flight PlanningMax Initial Cruise Altitude/Climb 260 kt/0.72M/Limited by 300 ft/min Rate of Climb


Alternate and Reserve Fuel Requirement(Landing Weight)To determine the required fuel for the trip, calculate the aircraftweight at the destination by adding the reserve fuel to the zero fuelweight.

Use the Two Engine Operating Reserve Fuel chart (Figure 4-15 )(referred to hereafter as Reserve Fuel chart) to determine the reservefuel weight and landing weight at the destination.

U.S. FARs require the reserve fuel to the alternate be based on fuelconsumption at normal cruising speed for both VFR and IFRoperations; long range cruise is considered one of the four normalcruise modes. The FARs do not specify any altitude for the reservecruise time.

Compute fuel to the alternate based on a missed approach at thedestination, a climb to 25,000 ft, and long range cruise.

Time adjustments for reserve fuel are based on long range cruise at25,000 ft to the alternate, holding at 5,000 ft for 45 minutes, anddescent to sea level. The holding fuel burn rate at 5,000 ft is the sameas the long range cruise fuel burn rate at approximately 15,000 ft.

1. Enter the Reserve Fuel chart at the Tailwind-kt-Headwind scaleat the average wind value from destination to alternate (30 ktsheadwind).

2. Move up to the distance from the destination to the alternate(143 NM).

3. Move left crossing the Time to Alternate-min scale. Read thetime to the alternate (30 min).

4. Move right to the appropriate Holding Time at Alternate line(45 min).

5. Move down from the reference line, following the guidelines,to intersect the planned zero fuel weight (12,720 lbs).

6. Move down to the bottom of the chart to derive the minimumreserve fuel (1,280 lbs).

7. Return to the zero fuel weight point (12,720 lbs). Move left to theedge of the chart and read the landing weight at destination(14,000).

With the landing weight at destination computed, eitherre-address landing performance or continue flight planning.

NOTE: For more convenientuse, rotate the Reserve Fuel chart90° clockwise, as shown on theopposite page.


Flight Planning

4-15

Two Engine Flight PlanningReserve Fuel Chart

30 min 143 N

M

30 KTS

14,000 LBS

45 min

12,720 LBS

1,280 LBS

�

�

�

�

�

�

Flight Planning

Two EngineFlight PlanningReserve FuelChart

4-45 4-46 For SimuFlite training only Falcon 10/100April 2000

Trip Fuel Requirement and Takeoff WeightWith cruise mode chosen and the final cruise altitude determined,select the Two Engine Flight Planning: Mi=.80 chart (Figure 4-16 ),hereafter referred to as the Mi chart.

1. Enter the chart from the upper left with the value of the averageheadwind or tailwind (40 kts headwind).

2. Move right to intersect the trip distance (1,200 NM).

3. Move down and read the trip time where the temperature (ISA)line crosses the trip line (2:59).

4. Continue down to intersect the landing weight at the destination(14,000 lbs).

5. Move left to the edge of the chart and read the takeoff weight(18,100 lbs).

6. Find trip fuel by subtracting the landing weight from the takeoffweight (4,100 lbs).

7. Add trip fuel to the alternate and reserve fuel to determine totalfuel required. This total fuel does not include taxi fuel.

8. Record total fuel on the weight and balance form.

Final Fuel CalculationThe total fuel on board at takeoff is the takeoff weight (18,100 lbs)minus the zero fuel weight (12,720 lbs). Use this formula in theexample to calculate fuel weight (5,380 lbs).


Flight Planning

4-16

Mi Chart

1,200 NM40 KTS

14,000 LBS18,100 LBS

�

�

�


Flight Planning

With the trip planned and the desired takeoff weight determined,compute takeoff and takeoff flight path performance. Review FAR91.605(b) at this time.

FAR 91.605(b): No one may operate a turbine-engine-poweredtransport category airplane certificated after September 30,1958, contrary to the AFM, or take off that airplane unless:

� the takeoff weight does not exceed the takeoff weight specifiedin the AFM for the elevation of the airport and for the ambienttemperature existing at the time of takeoff

� normal consumption of fuel and oil in flight to the airport ofintended landing and to the alternate airports leaves anarrival weight not in excess of the landing weight specified inthe AFM for the elevation of each of the airports involvedand for the ambient temperatures expected at the time oflanding

� the takeoff weight does not exceed the weight shown in theAFM to correspond with the minimum distances required fortakeoff, considering the elevation of the airport, the runwayto be used, the effective runway gradient, and the ambientwind and temperature component existing at the time oftakeoff.

The Falcon 10/100 is certified under FAR 25, which prescribestakeoff flight path limits and directs the manufacturer to presenttakeoff limits by weight or distance and takeoff climb performancelimits by weight.

TakeoffPerformance

NOTE: See Quick Reference forweight limitations.


Runway and Climb Weight LimitsThe aircraft takeoff weight is limited by certified maximum grosstakeoff weight, runway length, or climb requirements. Of these, onlythe maximum gross takeoff weight is fixed: 18,300 lbs (basic), 18,740lbs on aircraft with SB F10-0052, or 19,300 lbs with SB F10-0238.Use charts in the AFM to determine if aircraft weight is limited byrunway or climb requirements.

Acquire the following information to determine the maximum take-off weight limited by runway and climb requirements and to deter-mine runway requirements:

� departure airport elevation in pressure altitude

� length of runway in use

� runway slope

� field temperature

� runway wind

� desired takeoff weight if other than maximum

� any SID or obstacle climb requirement.

Dallas RBD is 660 ft MSL. Assume the conditions are 90°F, wind210° at 5 kts, and altimeter 29.77 in/Hg. The pressure altitude forRBD is calculated as 510 ft. Runway 13, which is 5,452 ft long,is in use. The slope is not indicated; assume it is less than 2%.There is an 11 ft difference in elevation between runway ends,but that does not take runway rise or dip into consideration.

NOTE: SB F10-0052; MaximumTakeoff Weight of 18,740 Lbs(8,500 Kg) and Maximum ZeroFuel Weight of 13,560 Lbs(6,150 Kg).

SB F10-0238; Maximum TakeoffWeight of 19,300 Lbs (8,755 Kg)and Maximum Zero Fuel Weightof 14,400 Lbs (6,540 Kg).


Flight Planning

TOLD CardA Takeoff and Landing Data (TOLD) card displays takeoff and land-ing data as a convenient reference aid in the cockpit.

The takeoff side of the card provides spaces for the following infor-mation:� ATIS� V1 – Takeoff Decision Speed� VR – Rotation Speed� V2/VREF – Safety Speed/Emergency Return Landing Speed� VFR – Slat + Flap Retraction Speed� VFS – Final Segment Climb Speed� VENR – Enroute Climb Speed� Weight – Takeoff Weight� Flaps – Configuration Flap Setting (15° or 0°)� Pitch – Rotation Pitch Attitude� Takeoff N1

� Climb N1

� Runway Required – Computed Takeoff Field Length� Time to 100 kts.� Clearance.

The approach side of the TOLD card provides spaces for thefollowing information:� ATIS� VREF – Landing Configuration Speed – 50 ft point.� VAP – Approach Target Speed� VAC – Single Engine Minimum Approach Speed for Flaps 15°

or 30°� VFR – Slat + Flap Retraction Speed (Go Around)� Weight – Computed Landing Weight� S + Flaps – Required Landing Configuration� Max N1 – Maximum Go Around N1

� RWY RQD – Computed Landing Distance or Landing FieldLength

� Notes.

ATIS

WEIGHT

FLAPS

VREF

VAP

APPROACH FALCON 10/100

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY

TAKEOFF FALCON 10/100

VAC

VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°


Takeoff Gross WeightEnter the takeoff weight determined by weight and balance in the WTblock of the Takeoff side of the TOLD card (18,100 lbs).

Takeoff gross weight may be limited by any of the following:� brake energy� tire speed� field length� climb performance� obstacle clearance� approach climb gradients� landing distance� structural limits (maximum takeoff and maximum landing

weight).

Evaluate the gross weight determined by weight and balance (18,100lbs) against the manufacturer’s guidelines for each of the limitationslisted above.

Use performance charts in the AFM to determine the maximumtakeoff gross weight permitted by FAR 25 and the associated speedsand flight paths. The aircraft may be limited in takeoff gross weightby field length, climb gradient, tire speed, obstacle clearance, orbrake energy.

NOTE: In addition to the infor-mation on the right, the aircraft isalways limited as specified in the AFM,Limitations section (Structural).


Flight Planning

Takeoff Weight Limited byClimb RequirementsThe takeoff weight may be limited by climb requirements, espe-cially at high runway elevations or hot temperatures. The MaximumTakeoff Weight Limited by Climb Requirements chart (Figure 4-17 )encompasses all the minimum climb requirements of certificationfor the takeoff flight path but may not include the first segment. Ifa takeoff weight is not limited by the chart, the takeoff flight path withan engine failure at that weight meets all FAR 25 minimum climbgradients.

The desired weight is 18,100 lbs. At Dallas RBD the conditionsare:

� pressure altitude of 500 ft

� temperature of 90°F

� runway slope negligible

� 0 headwind/tailwind components.

1. Enter the chart from the left with the ambient temperature (90°F).Move right to intersect the airport pressure altitude (500 ft).

2. Move down to intersect the bottom of the chart and read the max-imum weight allowable at takeoff to meet all minimum climbrequirements. (The result, 17,550 lbs, is not adequate for theplanned trip.)

3. If a slats + flaps 15° configuration does not meet climb require-ments, a slats-only configuration may, although it is unlikelybecause the runway available is less than 6,000 ft. An alternativeis to use up to a 10% speed increase, which uses more runway thanthe first configuration, but less runway than that needed for slats-only; speed can be adjusted to meet the need.

4. Move down to the speed increase reference line, then parallelthe slant lines to the appropriate speed increase (5%). Movedown to read the maximum weight allowed (18,100 lbs).

Ensure the runway required does not exceed the runway available.Address several charts if necessary to refine the calculations to fit therequirement. Work the chart in the same manner for all configurations.

4-17

Maximum Takeoff Weight Limited by Climb Requirements

NOTE: Use the Balanced FieldLength chart to determine if therequired speed increase is limitedby runway length.

500 FT90°F

17,550 LBS

18,100 LBS

�

�

�

�

Flight Planning

MaximumTakeoff WeightLimited by ClimbRequirements


Balanced Field LengthThe procedure specified by the manufacturer in the AFM utilizesthe Balanced Field Length chart (Figure 4-18 ) to determine runwayrequirements and weight limits.

Assume the desired takeoff configuration is slats + flaps 15° witha 5% speed increase.

1. Enter the balanced field length chart from the upper left with thefield temperature (90°F). Move right to intersect the departurefield pressure altitude (500 ft).

2. Move down to the reference line and parallel the weight guide-lines to intersect a line drawn from the takeoff weight at the left(18,100 lbs).

3. To determine values for slope and wind:

a. Move straight down to the slope reference line. Enter from theleft at the slope (0%) and move right across the slope guide-lines.

b. Move up or down from the reference line and parallel theslope guidelines to intersect the slope value line (0), thendown the wind reference line.

Because the values are zero, there is no correction for slope.

c. Enter from the left at the wind (0) and move right across thewind guidelines.

d. Continue from the reference line and parallel the windguidelines to intersect the wind value line and then downto read the value (0).

Because the wind value is zero, there is no correction forwind.

4. Parallel the speed increase guidelines to the appropriate point(5%) and move straight down to read the balanced field lengthrequired (5,300 ft).


Flight Planning

4-18

Balanced Field Length

5,300 FT

0 KTS

0%

18,100 LBS

500 FT90°F �

�

�

�

�

�


VMBE Speed ComputationsGenerally, maximum brake energy speed is not a limiting factor,particularly in the slats + flaps 15° configuration. A limiting speedmay require a decrease in takeoff weight or a change in runwayconditions (wind, slope, etc.).

Determine VMBE from the Maximum Brake Energy Speed (VMBE)chart (Figure 4-19 ).

1. Enter the chart from the left at the ambient temperature (90°F).Move right to intersect the pressure altitude (500 ft). Move downto the reference line.

2. Enter left at the takeoff weight and move right, across the weightguidelines.

3. Return to the temperature-altitude reference line intersection andparallel the guidelines to intersect the takeoff weight line.

4. Move down to the runway slope reference line.

Assume no slope.

5. Move down to the wind reference line.

Assume no wind.

6. Move down to read the maximum brake energy speed(146 KIAS).

NOTE: V1 should never begreater than VMBE.


Flight Planning

4-19

Maximum Brake Energy Speed (V MBE)

90°F500 FT

18,100 LBS

146 KIAS

�

�

�

�


Tire SpeedUse the Maximum Allowable Takeoff Gross Weight Permitted by TireSpeed chart (Figure 4-20 ) to determine the maximum allowable take-off gross weight based on tire speed.

Generally the maximum takeoff weight is not limited by tire speed.It happens only at high gross weights, tailwind, high elevation, andhot temperatures; it most likely occurs in a slats + flaps 15° plusspeed increase or in a slats-only configuration.

1. Enter the chart from the left at the ambient temperature (90°F).Move right to intersect the Field Pressure Altitude-Ft line(500 ft).


Assume no wind.

3. Move down to intersect the tire speed limit/speed increase lines,then across to the maximum takeoff weight allowable.

The vertical line did not intersect a tire speed line; therefore, tirespeed is not limiting.


Flight Planning

4-20

Max Allowable T/O Gross Wt Permitted by Tire Speed

90°F500 FT

�

�


V1 Speed ComputationsOnce the runway and weight requirements are met, determineairspeeds for takeoff and climb. The first speed required is V1. Usethe Engine Failure Speed Associated with Balanced Field Lengthchart – Slats + Flaps 15° (Figure 4-21 ); the takeoff configurationdetermines which chart to use.

1. Enter the chart from the left at the ambient temperature (90°F).Move right to intersect the pressure altitude (500 ft). Move downto the weight reference line and parallel the weight guidelines.

2. Enter the chart from the left at the desired takeoff weight (18,100lbs).

3. Move right across the weight guidelines to the intersection ofthe ambient temperature/pressure altitude line and the weightreference line. From this intersection, move down to the slopereference line.

4. Adjust V1 for slope.

Assume no slope. Continue to the wind reference line.

5. Adjust V1 for wind.

Assume no wind. Continue to the speed increase reference line.

6. Adjust V1 for speed increase.

a. Enter the chart from the left at the speed increase (5%).Move to the right through the speed increase guidelines.

b. Return to the temperature, altitude, and weight line, andfollow the speed increase guidelines to the speed projectionintersection.

c. Move down to the bottom and read V1 for a 5% speedincrease (118 kts).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 550000,

TTEEMMPP -- 9900∂∂FF

111188

55,,330000

1188,,110000

1155


Flight Planning

4-21

Engine Failure Speed Associated with Balanced Field Length

90°F 500 FT

18,100 LBS

118 KTS

�

�


Takeoff Speed and Nose-Up Attitude(VR, ATT, V2, Slats + Flaps 15° )To compute VR, takeoff attitude and V2, refer to the Takeoff Speedsand Nose-Up Attitude/ S + Flaps 15° chart for the takeoff configu-ration (Figure 4-22 ).

Compute VR (top chart).

1. Enter the chart from the takeoff weight at the bottom of the chart(18,100 lbs), move up to the speed reference line, move left, andread the speed (117 kts).

2. When, as in this case, a speed increase is used, move right fromthe speed reference line to the speed increase reference line, thenparallel the speed increase guidelines to the 5% point and moveright to read the rotation speed (123 kts).

Compute attitude.

1. Enter the chart from the takeoff weight at the bottom and moveup to the pitch line. Read the pitch attitude left (16°).

2. With a speed increase, move right to the speed increase refer-ence line and parallel the speed increase guidelines to the 5%point, then right to the takeoff pitch attitude (14.5°).

Compute V2 takeoff safety speed.

1. Enter the chart at the takeoff weight and move up to intersectthe V2 speed line. Read the speed at the left (120 kts).

2. Calculate the speed increase adjustment by moving from the V2speed line to the right to the speed increase reference line; parallelthe speed increase guidelines to the 5% point, then move right toread V2 (125 kts).

Takeoff Clean Wing Speed (V FR)Clean wing speed is a weight-controlled number, regardless of con-figuration. To determine VFR for any configuration, use the TakeoffSafety Speed (V2) chart, Slats + Flaps 15° (refer to Figure 4-22 ).

To calculate clean wing speed, take V2 previously calculated, speed increase not considered (120 kts), and add 35 kts (120 kts +35 kts = 155 kts VFR).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. --550000,


111188

112233

112255

115555

55,,330000

1188,,110000

1155 1144..55

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. --550000,


111188

112233

112255

55,,330000

1188,,110000

1155 1144..55


Flight Planning

4-22

Takeoff Speeds and Nose-Up Attitude

117 KTS123 KTS

14.5°

125 KTS

120 KTS

16°

18,100 LBS

18,100 LBS

18,100 LBS

��

�

��

�

�

�

�


Final Takeoff Climb Speed (V FS)Determine VFS speed using the Final Takeoff Climb Gradient chart(Figure 4-23 ).

1. Find the 1.4 VS speed chart in the lower right side of the abovechart.

2. Enter gross weight at the bottom and move up to intersect the1.4 VS guideline.

3. Move left from that point to read the final segment climb speed(168 kts).

Enroute Climb Speed (V ENR)Determine VENR using the Enroute Climb Gradient One-EngineInop/Clean chart (Figure 4-24 ).

1. Find the 1.5 VS speed chart on the lower right side of the abovechart.

2. Enter gross weight at the bottom and move up to intersect the1.5 VS guideline.

3. Move left from that intersection to read the enroute climb speed(180 kts).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 550000,


111188

112233

112255

115555

116688

118800

55,,330000

1188,,110000

1155 1144..55


Flight Planning

4-23

Final Takeoff Climb Gradient

4-24

Enroute Climb Gradient One Engine InopClean

168 KTS

18,100 LBS

180 KTS

18,100 LBS

�

�

�

�

Flight Planning

Enroute ClimbGradient OneEngine Inop


N1 Setting Computations – Takeoff ThrustUse the Takeoff Thrust Setting chart (Figure 4-25 ) to calculate N1 set-ting based on static air temperature and pressure altitude. This thrustbegins the takeoff roll and achieves all takeoff performance values.Full advancement of the thrust levers should provide, as a minimum,the charted N1 value for takeoff.

Assume ice protection is not required.

1. Enter the chart from the bottom with the OAT (90°F). Move upto intersect the pressure altitude (500 ft), or the limit line, which-ever is first.

2. Move left from the intersection to the edge of the chart. Readthe N1 setting (95.3%).

Use a similar chart in the AFM in the same manner to calculatetakeoffs with wing and/or engine anti-ice on. Wing anti-ice isprohibited during takeoff roll (on the ground).

N1 Setting Computations – Climb ThrustUse the Maximum Continuous Thrust – Final Takeoff and EnrouteClimb chart (Figure 4-26 ) to determine climb N1 at the maximumcontinuous thrust.

(The Maximum Continuous Thrust – Final Takeoff chart is used inthe same manner for takeoff performance.)

Assume ice protection is not required.

1. Enter the chart from the bottom with the RAT (90°F). Move upto intersect the pressure altitude (500 ft), or limit line, whicheveris first.

2. Move left from the intersection to the edge of the chart. Readthe N1 setting (93.4%).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 550000,


111188

112233

112255

115555

116688

118800

55,,330000

1188,,110000

1155 1144..55

9955..33

9933..44


Flight Planning

4-25

Thrust Setting/Takeoff

4-26

Thrust Setting/Maximum Continuous

95.3°F

90°F 90°F

93.4°F

�

�

�

�

Flight Planning

Thrust Setting/Takeoff

Thrust Setting/Maximum Continuous


Time to 100 KnotsUse the Time To 100 Kts chart (Figure 4-27 ) to ensure the aircraft’soperational performance is normal (e.g., that the thrust is appropri-ate, the brakes are not partially engaged, or the flaps are not extendedmore than required). The chart is valid for slats + flaps 15° or slats-only takeoffs.

1. Enter from the left at ambient temperature (90°F), move rightto field pressure altitude (500 ft), then down to the gross weightreference line.

2. Enter from the left at the gross weight (18,100 lbs), and move rightto the center of the chart.

3. Continue the temperature/altitude line at the weight referenceline and parallel the weight guidelines to intersect the horizontalweight line drawn previously.

4. Continue down to the slope reference line.

Assume no slope.

5. Continue down to the wind reference line.

Assume no wind.

6. Continue down to read the time to 100 kts (21 seconds).

Cruise Climb N 1 SettingUse the Thrust Setting Maximum Climb chart (Figure 4-28 ) to setclimb power during normal climb to cruise altitude. Climb speedcould be 260 kts/0.72M, or 300 kts/0.75M; ice protection is not afactor in calculating cruise climb N1.

1. Enter from the top at pressure altitude (10,000 ft), and movedown the vertical column. Intersect the horizontal row from theram air temperature (15°C). The square at the intersection con-tains the appropriate maximum N1 setting (97.4%).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. --550000,


111188

112233

112255

115555

116688

118800

55,,330000

1188,,110000

1155 1144..55

9955..33

9933..44

2211


Flight Planning

4-27

Time to 100 Kts (IAS)Valid for Takeoff with S + Flaps 15° or Slats Only

4-28

Thrust Setting/Maximum Climb

90°F 500 FT

18,100 LBS

21 SECONDS

�

�

�

Flight Planning

Time to 100 Kts (IAS)

Thrust Setting/Maximum Climb


Takeoff – Slats-OnlySometimes temperature and elevation severely limit the Falcon10/100’s load-carrying capabilities. In high density altitudes wherethe runway length is not a factor, loads may be increased by usingspeed increases discussed earlier or slats-only takeoffs.

Slats-only requires more runway than any other takeoff. If the run-way length is less than 7,000 ft, slats-only may not improve loadmore than the speed increase technique.

The following example is for illustration only and is not related toprevious problems:

Airport conditions:

� Elevation: 4,500 ft

� Temperature: 100°F

� Runway: 11,000 ft

Find the maximum gross weight allowable under these conditions andcalculate a TOLD card.

Maximum Takeoff Weight Limited ByClimb Requirements – Slats-OnlyUse the Maximum Takeoff Weight Limited by Climb Requirementschart (Figure 4-29 ).

1. Enter from the left at the temperature (100°F).

2. Move right to the pressure altitude (4,500 ft).

3. Move down to read the maximum takeoff weight (15,600 lb).

NOTE: The following slats-onlyexercises are for illustration onlyand do not relate to the mainexample for this chapter.

NOTE: SB F10-0077: Takeoffwith Slats Extended and FlapsRetracted.

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 44,,550000,


1155,,660000


Flight Planning

4-29

Maximum Takeoff Weight Limited by Climb Requirements/Slats

100°F 4,500 FT

15,600 LBS

�

�


Takeoff Climb: First Segment – Slats-OnlySlats-only takeoffs are limited by first segment, rather than secondsegment as slats + flaps 15° takeoffs. Use the Takeoff Climb-FirstSegment chart (Figure 4-30 ).


2. Move right to the pressure altitude (4,500 ft), then down theweight reference line.

3. Parallel the weight guidelines.

4. Enter from the left at the takeoff weight (15,600 lbs), then moveacross the weight guidelines to intersect the temperature/altitudeline.

5. From the intersection, move down to read the gradient (-.1%).

The minimum gradient is 0% gross climb gradient. If the weight ismore than what produces a positive gradient, then weight must bereduced (15,500 lbs), so that the gradient is positive (0%).

Balanced Field Length – Slats-OnlyUse the Balanced Field Length chart (Figure 4-31 ).

New Gross Takeoff Weight: 15,500 lbs


2. Move right to the pressure altitude (4,500 ft), then down to theweight reference line.

3. Enter from the left at the takeoff weight (15,500 lbs) and moveright across the weight guidelines.

4. Continue the temperature/altitude line from the reference lineand parallel the weight guidelines to intersect the weight line.

5. From the intersection move down to the slope reference line.

Assume no slope.


Assume no wind.

7. Move down to read the balanced field length (6,750 ft).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 44,,550000,


66,,775500

1155,,550000


Flight Planning

4-30

Takeoff ClimbFirst Segment/Slats

4-31


100°F 4,500 FT

15,500 LBS

15,600 LBS

100°F4,500 FT

15,500 LBS

6,750 FT

-.1% 0%

�

�

�

�

�

�

Flight Planning

Takeoff Climb

Balanced FieldLength


Second Segment Climb – Slats-OnlyUse the Takeoff Climb – 2nd Segment chart (Figure 4-32 ) if obstacleclearance or minimum climb is a consideration. This gradient is alsonecessary in calculating the slats-only takeoff pitch attitude.

1. Enter from the left at the ambient temperature (100°F).

2. Move right to the pressue altitude (4,500 ft).

3. Move down to the weight reference line.

4. Enter from the left at the takeoff weight (15,500 lbs).

5. Move right across the weight guidelines.

6. Continue the temperature/altitude line from the reference line,paralleling the weight guidelines.

7. Intersect the weight line and move down to read the gradient(1.8%). A minimum of 1.6% is required.

With a net climb gradient of 1.8%, a slats-only takeoff is notlimited by second segment climb.


Flight Planning

4-32

Takeoff ClimbSecond Segment

100°F 4,500 FT

15,500 LBS

1.8%

�

�

�

�


V1 Safety Speed – Slats-OnlyUse the Engine Failure Speed Associated with Balanced Field Lengthchart (Figure 4-33 ).



3. Move down to the takeoff weight reference line.

4. Enter from the left at the takeoff weight (15,500 lbs).

5. Move right across the weight guidelines.

6. Continue from the weight reference line, paralleling the weightguidelines.

7. Intersect the weight line and move down to the slope referenceline.

Assume no slope.


Assume no wind.

9. Move down to read the V1 speed (124 kts).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 44,,550000,


112244

66,,775500

1155,,550000


Flight Planning

4-33

Engine Failure Speed Associated with Balanced Field Length

100°F4,500 FT

15,500 LBS

124 KTS

�

�

�

�


Takeoff Speeds and Nose-Up Attitude,VR, Attitude, V 2 – Slats-OnlyUse the Takeoff Speeds and Nose-Up Attitude chart (Figure 4-34 ) tocalculate VR, attitude, and V2.

To determine VR, perform the following.

1. Enter from the bottom of the Rotation speed VR chart at the take-off weight (15,500 lbs).

2. Move up to the reference line.

3. Move left to read the rotation speed (126 kts).

To determine attitude, perform the following.

1. Enter from the bottom of the Attitude chart at the second segmentnet climb gradient (1.8%).


3. Move left to read the Takeoff Pitch Attitude (13.4°).

To determine V2, perform the following.

1. Enter from the bottom of the Take-off safety speed V2 min chartat the takeoff weight (15,500 lbs).


3. Move left to read the takeoff safety speed (127 kts).

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 44,,550000,


112244

112266

112277

66,,775500

1155,,550000

1133..44∂∂


Flight Planning

4-33

Takeoff Speeds and Nose-Up Attitude

126 KTS

15,500 LBS

13.4°F

1.8%

127 KTS

15,500 LBS

�

�

�

�

�

�


Maximum Tire Speed – Slats-OnlyUnder certain conditions, tire speed can be very high when using aslats-only takeoff. Use the Maximum Takeoff Weight Limited byTire Speeds chart (Figure 4-35 ) to ensure these tire speed limits arenot exceeded.


2. Move right to the pressure attitude (4,500 ft).


Assume no wind.

4. Move down to the tire speed (180 mph) then right to read themaximum weight (17,900 lbs).

NOTE: Two types of tires areavailable, 190 kt and 180 kt.


Flight Planning

4-35

Maximum Takeoff Weight Limited by Maximum Tire Speed

100°F4,500 FT

17,900 LBS

�

�


Clean Wing Speed – Slats-OnlyThe clean wing speed is a speed at which it is safe to fly withouthigh lift devices extended.

To calculate clean wing speed, take V2 previously calculated(127 kts), and add 20 kts (127 kts + 20 kts = 147 kts VFR).

Slats-Only TOLD Card CalculationAfter calculating clean wing speed (VFR), all following calculationsare the same as for a slats + flaps 15° takeoff. Follow the same stepsfor:

� final takeoff climb gradient and speed

� enroute climb gradient and speed

� N1 setting – takeoff thrust

� N1 setting – climb thrust

� time to 100 kts

� maximum climb N1 setting

� maximum brake energy speed.

ATIS

WEIGHT

FLAPS TRIM

V1

VR

VFR

VFSTIMETO100 KTS

CLEARANCE

RWY


VENR

RQD FT.

V2/REF

S+TAKEOFF

N1:CLIMB

N1:

PP.. AA.. -- 44,,550000,


112244

112266

112277

114477

66,,775500

1155,,550000

1133..44∂∂


Flight Planning

Normal trip planning encompasses climb, cruise, and descent fortotal time and fuel used. It does not give specific climb, cruise,and descent information. Such information for all-engine climb,cruise, and descent is presented only in the Operating Manual.

Cruise andLandingPerformance


Climb DataTwo Engine Flight Planning tables (Figure 4-36 ), one for climb at 260kts/0.72M and one for climb at 300 kts/0.75M, are organized byaltitude, weight, and temperature deviations from ISA.

1. Enter the appropriate chart (260 kt/0.72M) for the pressurealtitude (37,000 ft) from the top with the aircraft weight(18,000 lbs).

2. Move down the column to the row of blocks directly across fromthe temperature (ISA) and read the values.

For this example, the values are:TAT -56.5°CRAT -37°CTime 16 minutesDistance 97 NMFuel Used 594 lbs


Flight Planning

4-36

Two Engine Flight PlanningClimb 260 kt/.72


Cruise DataUse data tables for detailed flight planning or enroute performanceverification.

Data tables for all four normal cruise modes (i.e., long range, indicated0.75M, indicated 0.80M, and maximum cruise thrust) are in theOIM.

The engine RPM, TAS, and RAT, and fuel flow are verified againstindicated and trip planning values. Use the specific fuel consump-tion or fuel flow to estimate required fuel for the remaining cruise;specific fuel consumption provides more accurate information, butis more difficult to use.

Use the Two Engine Operating Chart: Mi = .80 (Figure 4-37 ) fromthe Operational Instructions Manual. Note that, as the trip progresses,the aircraft gross weight decreases to 15,000 lbs.

1. Enter the chart from the pressure altitude (37,000 ft) at the top withthe aircraft weight (15,000 lbs).

2. Move down the column to the row of blocks directly across fromthe temperature and read the values.

For this example, the values are:TAT -56.5°CRAT -33°CTAS 454 ktsN1 RPM 94.1%Fuel Flow 620 PPHNM per lb 0.365


Flight Planning

Two Engine OperatingIndicated Mach .80

4-37


Descent DataUse the Two-Engine Operating Normal Descent Mi .80/320 kt chart(Figure 4-38 ) for detailed flight planning and to verify the descentpoint in flight. The chart is based on descent at 85% N1 and powersettings that maintain a 500 lb/hr fuel flow for the remainder ofdescent.

Assume the flight weight is about 14,500 lbs near the descentpoint.

1. Enter the chart at aircraft weight (14,500 lbs).

Use the columns for 14,000 and 15,000 lbs and interpolate.

2. Move down to the appropriate pressure altitude (37,000 ft).Interpolate between columns if necessary. Read the values at theintersection.

The values for this example are:

15,000 lbs 14,000 lbs Interpolation

Time (Min) 17 16 17

Distance (NM) 106 102 104

Fuel Used (Lbs) 259 250 255

3. Compute the landing weight by subtracting the descentfuel from the weight at the beginning of descent(14,500 - 255 = 14,245).

VREF/VAP Speed ComputationsDetermine VREF from the applicable Landing Speeds chart (Figure4-39). The VAP is equal to VREF plus wind correction, with a mini-mum value of 10 kts.

1. Enter the chart from the bottom with the landing weight (14,000lbs). Move up to intersect the diagonal reference line.

2. Move left to the edge of the chart and read the value for VREF flaps52° (107 kts).

The wind is calm, thus VAP is VREF+ 10 kts, or 117 kts.

ATIS

WEIGHT

FLAPS

VREF

VAP


VAC

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°

110077

111177

1144,,000000


Flight Planning

4-38

Two Engine OperatingNormal Descent Mi .80/320 Kt

4-39

Landing Speeds

107 KTS

14,000 LBS

�

�

Flight Planning

Two EngineOperating

LandingSpeeds


Approach Climb Gradient/SpeedThe following example is for illustration only and is not a part ofthe ongoing example.

Assume an engine fails after takeoff from Dallas Redbird Airport. Thefuel used in returning to the airport is approximately 400 lbs, andthe resulting landing weight is 17,700 lbs.

Using the One Engine Inoperative Approach Climb Gradient chart(Figure 4-40 ), determine the approach climb gradient slats + flaps 30°.

1. Enter the chart from the left at the ambient temperature (90°F).Move right to pressure altitude (500 ft), then down to the referenceline.

2. Enter from the left at the aircraft weight (17,700) and move rightacross the weight guidelines.

3. Continue the temperature/altitude line from the reference line,paralleling the weight guidelines to intersect the weight line.

4. Move down to read the gross climb gradient (1.2%).

The minimum gross climb gradient is 2.1%. The gradient at thisweight and condition is insufficient to meet the minimum required;however, changing the configuration to slats + flaps 15° mayincrease performance.

Use the One Engine Inoperative Approach Climb Gradient/Slats +Flaps 15° chart (Figure 4-41 ) to determine the gradient and speed.

1. Enter the chart from the left at the ambient temperature (90°F)and move right to the pressure altitude (500 ft). Then move downto the reference line.

2. Enter from the left at the gross weight (17,700 lbs), and move rightacross the weight guidelines.

3. Continue the temperature/altitude line from the reference line,paralleling the weight guidelines to intersect the gross weightline; move down to read the gross climb gradient (3.6%).

By retracting the flaps from 30° to 15°, the climb gradient increases2.4%, three times the climb gradient at 30° flaps or 1.2%.

Calculate the approach climb speed by using the small chart on thelower right side.

1. Enter from the bottom at the gross weight (17,700 lbs)and moveup to the 1.35 VS line, then left to read VAC, slats + flaps 15°(131 kts).

ATIS

WEIGHT

FLAPS

VREF

VAP


VAC

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°

111199

112299

1144,,000000

113311 112233

TTOOLLDD CCAARRDD FFOORR

EEMMEERRGGEENNCCYY RREETTUURRNN

TTOO RREEDDBBIIRRDD


Flight Planning

4-40

One Engine Inoperative Approach Climb GradientS + Flaps 30°

4-41

One Engine Inoperative Approach Climb GradientS + Flaps 15°

90°F500 FT

131 KTS

17,700 LBS17,700 LBS

3,6%

90°F 500 FT

17,700 LBS

1.2%

�

�

�

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�

�

�

�

�

�

Flight Planning

One EngineInoperativeApproachClimb Gradient


Missed Approach/Go Around Clean Wing SpeedCalculate, as previously discussed, the minimum speed at which theslats and flaps can be retracted after a missed approach or go-aroundmaneuver.

V2 slats + flaps 15° + 35 kts = VFR. Generally, VREF is within oneor two knots of V2 for the same condition; in this case, add 35 kts toVREF to find VFR.

V2 - S+F 15 at 14,000 lbs = 108VREFat 14,000 lbs = 107107 + 35 = 142 VFR

Landing DistanceTo determine landing distance, use the Landing Distance S + Flaps52° chart (Figure 4-42 ).

1. Enter the chart from the left with the landing weight (14,000lbs). Move right to intersect the pressure altitude (0 ft).

2. Move down to intersect the wind (0 kts). From the intersection,move right to the landing distance scale. Read the landing distance(2,200 ft).

3. To convert the landing distance to landing field length, multiplylanding distance by 1.67 (2,200 ft X 1.67 = 3,674 ft), or contin-ue to move the line right to the reference line. Parallel the guide-lines to the edge of the chart, and read the landing field lengthvalue (3,675 ft).

ATIS

WEIGHT

FLAPS

VREF

VAP


VAC

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°

110077

111177

1144,,000000

111177 111100

114422

5522

ATIS

WEIGHT

FLAPS

VREF

VAP


VAC

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°

110077

111177

1144,,000000

111177 111100

114422

5522

33,,667744

4-42

Landing DistanceS + Flaps 52°


Flight Planning

14,000 LBS

2,200 FT 3,675 FT

�

�

�


N1 Setting – Missed Approach orGo AroundFind missed approach/go around thrust settings by using the TakeoffThrust Setting chart (Figure 4-43 ) or Takeoff Thrust Setting withWing and Engine Anti-Ice chart (Figure 4-44 ), depending on thecondition.

1. For this example, anti-ice is not needed. Enter the chart(Figure 4-43 ) from the bottom at the OAT (90°F).

2. Move up to the pressure altitude or limit line, whichever is reachedfirst.

3. Move left to read N1 (95.4%).

For illustration only, refer to Figure 4-44 . With 0°C and 2,000 ftpressure altitude given on the chart for thrust with wing andengine anti-ice, N1 is 91.8%.

ATIS

WEIGHT

FLAPS

VREF

VAP


VAC

VFR

RWY

RQD FT.

S+MAX

N1:

NOTES:

15° 30°

110077

111177

1144,,000000

111177 111100

114422

5522

9955..44

33,,667744

4-43

Thrust SettingTakeoff


Flight Planning

4-44

Thrust SettingTakeoff/Engine and Wing Anti-Ice

95.4%

90°F

0°C

91.8%

�

�

�

�

Flight Planning

Thrust Setting



Flight Planning

The supplemental information section addresses specific situationsother than normal routines presented elsewhere in this chapter. Theseprocedures apply to unusual situations and should not be confusedwith normal performance procedures.

There is no relationship between the information computed in theprevious example and the information exhibited on the followingpages. Consider each subject independently, based on the data given.

SupplementalInformation


Adjustment of Weight and BalanceSometimes it is necessary to adjust balance calculation when a pas-senger or cargo moves from the planned position to an alternateposition.

In the example, a passenger weighing 165 lbs moves from seat6 to the jump seat. This forward movement causes the moment tobecome negative. Use the Moment Change chart (U.S. Units)(Figure 4-45 ):

1. From the previous position (seat 6), move up to read themoment change (+10 inches aft of datum).

2. Follow, or parallel, the guidelines down and left to thevertical line extending from the new position (jump seat).

3. From that intersection move left and read the change inmoment (21,400 in/lb).

4. Add the change (-21,400) to the initial moment and calculate thenew CG.


Flight Planning

4-45

Moment of Change

�

�

��


Obstacle ClearanceObstacle clearance calculations depend on the distance from thepoint of liftoff to the top of the obstacle. Two charts, one for close-in obstacle and one for distant obstacle clearance, calculate in per-centage the net climb gradient required to clear the obstacle by 35 ft.This percentage is used on a climb gradient chart to find the maxi-mum weight for this climb under current airport conditions.

Use the Close-In Obstacle Clearance chart (Figure 4-46 ) to calculatethe climb gradient necessary to clear the obstacle in the flight path.

1. Enter from the left at the obstacle’s height (140 ft) andparallel the clearance guideline to the reference line, thencontinue right across the chart.

2. Enter the chart from the bottom at the obstacle’s distance fromreference zero (4,500 ft), and move up to the wind reference line.

Assume no wind.

3. Continue up to intersect the height line and read the net climbgradient required (3.8%).

4. Go to the Takeoff Climb – 2nd Segment chart in the AFM tofind the maximum gross weight.


Flight Planning

4-46

Close-In Obstacle Clearance

140 FT

3.8%

4,500 FT

�

�

�

Flight Planning

Close-InObstacleClearance


The example on the Distant Obstacle Clearance chart (Figure 4-47 )is for illustration only. Use this chart if the obstacle is in the secondsegment or beyond.

The obstacle is 1,400 ft high, 10 NM (60,000 ft) from the lift-offpoint; the minimum gradient to clear it is 2.4% net.


Flight Planning

4-47

Distant Obstacle Clearance

1,400 FT2.4%

60,000 FT

�

�

Flight Planning

DistantObstacleClearance


Maximum Takeoff Weight Limited byObstacle Clearance RequirementTo continue the obstacle clearance problem on the previous page,use the Takeoff Climb – 2nd Segment chart (Figure 4-48 ) because theobstacle height (140 ft) falls into the second segment of the takeoffprofile.

1. Enter the chart from the left at the ambient temperature (90°F)and move right to the pressure altitude (500 ft), then down to thereference line.

2. Enter from the bottom at the required climb gradient.

In this case, use a 5% speed increase; a 3.8% net is equivalentto a 4.6% gross climb gradient, or 3.8% + .8% = 4.6%.

3. Move up to a line extending across from the speed increase (5%),then parallel the speed increase guidelines to the reference line.Continue straight up through the weight guidelines.

4. Resume the temperature/altitude line at the reference line andparallel the weight guidelines to intersect the gradient line.

5. From the intersection, move left and read the maximum weightallowable to meet climb requirements (15,800 lbs).

Use the Final Climb Gradient – One Engine Inoperative chart (Figure4-49) to find the maximum gross takeoff weight limited by climbfor obstacles in the final segment (illustration only).

The maximum weight allowed is 17,900 lbs, using the minimumclimb gradient (2.4%).

If the obstacle in the enroute segment is higher than 1,500 ft, workthe Final Segment Climb Gradient chart in the same manner.

NOTE: If a speed increase is notused, work the chart in the samemanner; however, begin thegradient line at the net climbgradient (3.8%) reference abovethe speed increase box, and moveup to intersect the temperature/altitude line, paralleling theweight guideline. From theintersection, move left to themaximum weight. Notice that themaximum weight is less thanwhen a 5% speed increase is used(about 15,500 lbs).


Flight Planning

4-49

Final Climb GradientOne Engine Inoperative

90°F500 FT

17,900 LBS

2.4%

�

�

�

�4-48

Takeoff Climb2nd Segment

90°F500 FT

15,800 LBS

4.6%

�

�

�

�

�

�

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Flight Planning

Takeoff Climb

Final ClimbGradient


Takeoff Noise Reduction Procedure –Maximum Takeoff Weight(18,300 Lbs – Basic Without SB F10-0052), or(18,740 Lbs – Basic With SB F10-0052)The following procedure reduces takeoff noise level at point A to aminimum after takeoff.

1. Maintain takeoff thrust up to 2,600 ft above the takeoff surface,with high lift devices in the takeoff configuration and maintain-ing a steady speed equal to V2 + 10 kts.

2. At 2,600 ft, set the N1 speed as defined on Takeoff NoiseReduction Procedure/N1 Setting After Cutback chart (Figure4-50) on the next page. This N1 speed setting permits level flightwith one engine inoperative.

3. As soon as possible, resume climb and increase thrust to maxi-mum continuous power setting.

Takeoff Noise Reduction Procedure –Maximum Takeoff Weight(19,300 Lbs – Basic With SB F10-0238)The following procedure reduces takeoff noise level at point A to aminimum after takeoff.

1. With high lift devices in the takeoff configuration, maintain take-off thrust to 2,350 ft above the takeoff surface and maintain asteady speed equal to V2 + 10 kts.

2. At 2,350 ft, set the N1 speed as defined on the Takeoff NoiseReduction Procedure/N1 Setting After Cut Back chart (Figure4-51). This N1 speed setting permits level flight with one engineinoperative.

3. As soon as possible, resume climb and increase thrust to maxi-mum continuous power setting.

NOTE: See Quick Reference fornoise reduction takeoff andapproach configurations andslope diagrams.

NOTE: See Quick Reference fornoise reduction takeoff andapproach configurations andslope diagrams.


Flight Planning

4-50

N1 Setting After Cut-BackMaximum Takeoff Weight 18,300 Lbs (Basic, Without SB F10-0052)or 18,740 Lbs (Basic With SB F10-0052)

4-51

N1 Setting After Cut-BackMaximum Takeoff Weight 19,300 Lbs (With SB F10-0238)

Flight Planning

Takeoff NoiseReduction Procedure


Contaminated Runway OperationsThe manufacturer, in its Service Newsletter 34, defines a wet runwayas one with less than 1/8 inch standing water and a contaminatedrunway as one with more than 1/8 inch standing water, or equiva-lent water depth, on more than 25% of its length. Other runway-contaminating fluids include slush, wet snow, loose dry snow, com-pacted snow, ice, or black ice (a mixture of water, oil, and rubber).

Hydrodynamic PressuresAs a tire travels over a fluid-covered surface, hydrodynamic pressuresincrease with speed, creating both drag and lift under the tire. Theforce of the tire against the fluid produces drag, and as speed increas-es, the tire hydroplanes, or lifts clear of the runway surface. Testsshow that at the moment the wheel lifts clear of the surface, wheelrotation diminishes rapidly and steering becomes less effective.When the main wheels slow to below 20 kts tire speed, the anti-skidsystem deactivates.

HydroplaningThe manufacturer and other authorities support a recently devel-oped formula to calculate hydroplaning speed: 6.4 times the squareroot of tire pressure (VH = 6.4√P). Prior to Service Newsletter 34, theaccepted formula was VH = 9√P, which recent tests suggest was anoverestimation. Calculated by the new formula, Falcon 10/100 nose-wheel hydroplaning speed is 62 kts rather than 87 kts with the for-mer formula, and main wheel hydroplaning speed is 75 kts ratherthan 106 kts. These speeds are, respectively, 25 and 27 kts less thanpreviously believed.

DragPlume drag occurs when each wheel (especially the nose wheel)throws a water plume that strikes the aircraft. Plume drag, which isproportional to the speed squared, is greater for small aircraft thanfor large aircraft. Both hydrodynamic and plume drag can be exten-sive. For example, 0.4 inch water produces drag near the maximumpower of one engine; above 0.5 inch water exceeds the force of oneengine. The Speed-Drag-Thrust Chart (Figure 4-52 ) illustrates dragincrease as speed increases in 0.79 inch (20 mm) water.

NOTE: See Maneuvers chapterfor further discussion oncontaminated runways.


Flight Planning

4-52

Speed-Drag Thrust


BrakingAlthough braking measurement is difficult, authorities believe the fol-lowing braking coefficients illustrate the use of anti-skid is indis-pensible.

Runway Braking Friction Decreased BrakingCondition Coefficient Factor Relative to

Dry Condition

Dry 0.5 – –Wet 0.4 to 0.25 Up to 2 timesContaminated 0.125 Up to 4 timesVery Slippery 0.05 Up to 10 times(e.g, Black Ice)

Summary: Contaminated Runway OperationsConclusions� Nosewheel hydroplaning speed is approximately 62 kts1.

� Main wheel hydroplaning speed is approximately 75 kts1.

� The tire lifts clear of runway surface.

� Steering is less effective and unpredictable.

� The wheels spin down rapidly.

� The anti-skid system deactivates below 20 kts wheel speed.

� Total drag can exceed power of one engine.

� Braking effectiveness is half or less.1 Speed increases as fluid density decreases (e.g., slush = 12%increase).


Flight Planning

Recommendations

Landing on Wet Runway

� Increase landing distance by 15%.

� Increase factored landing distance by 67% for the landingfield length.

Takeoff on Wet Runway

� Increase the balanced field length by adding 15% of thelanding distance for the same conditions.

Contaminated Runway

If takeoff cannot be delayed, observe the following.

� Limit takeoff operations to less than:

– 1/2 inch standing water

– 3/4 inch slush

– 1 inch wet snow

– 2 inches dry snow.

� Adopt V1 = VR.

– For new accelerate-stop distance, determine takeoffdistance factored for equivalent water depth (Refer to OIM 2-60(1) and OIM 2-60(2)).

– Add two times the landing distance for the same conditions.

� Limit crosswind to 15 kts for takeoff or landing.

� Multiply the landing distance required by 2.3.

� Ensure functional anti-skid availability.

� Delay braking action until below hydroplaning speed(75 kts).

If runways are very slippery (e.g., with compacted snow, ice, blackice), observe the following.

� Avoid if possible.

� Anticipate up to 1/10 braking effectiveness.

� Multiply the landing distance by at least 2.0 for compactedsnow or at least 3.4 for ice.

� Use maximum reverse thrust.

� Limit crosswind to 5 kts for takeoff or landing.


Landing – Contaminated Runway1. Increase landing distance by at least 2.0.

2. Delay braking until below VH.

3. Use maximum thrust reverse.

4. Ensure anti-skid operation.

5. LFL is landing distance (as above) plus 15%.

Takeoff – Contaminated Runway1. Avoid. If the runway is covered with water, wait until the water

flows away or is absorbed.

2. Calculate as previously discussed.

3. Pick a landmark alongside the runway as a go/stop point.

4. Above VH, only the rudder is effective for steering.

5. Ailerons neutral.

6. If takeoff is aborted, proceed as for landing on contaminated run-way.

7. After takeoff, cycle the gear several times to clear anyaccumulated snow or ice.

Landing – Slippery Runway1. Landing distance is equal to landing distance times 2 for

compacted snow, or times 3.4 for black ice.

2. LFL is landing distance as above plus 15%.

3. Apply maximum brakes at touchdown.

4. Use maximum thrust reverse.

5. Nose wheel may be of little use in steering.

NOTE: Although landing fieldlength is normally calculated bydividing the landing distance by0.6, the calculations used herehave safety factors built in andonly an additional 15% is neces-sary.


Flight Planning

Takeoff – Slippery Runway1. Avoid if possible. Control is uncertain.

2. Required takeoff distance (same conditions):

� for compacted snow: BFL plus the landing distance

� for ice: BFL plus twice the landing distance.

3. Leave a safety margin.

4. If takeoff is aborted, proceed as for landing.

Contaminated Runway CalculationsThe following example is for illustration only and is notrelated to previous problems.

Plan takeoff at an airport with:

� 0.75 inch slush on more than 25% of the runway

� 3,000 ft pressure altitude

� temperature 33°F

� headwind 20 kts.

� takeoff weight 18,000 lbs

� balanced field length(Figure 4-53 ) 4,000 ft.

4-53


4,000 FT

20 KTS

18,000 LBS

500 FT

33°F �

�

�

�

Flight Planning

BalancedField Length


Equivalent Water DepthFirst find the equivalent water depth for slush. All other charts arecalculated by the amount of standing water; go to the EquivalentWater Depth Chart (Figure 4-54 ) to determine how much water depthcauses the same resistance as 0.75 inch slush.

1. Enter the chart from the left at relative density (slush).

2. Move right to equivalent water depth (0.75 inch).

3. Move down to read equivalent water depth (0.5 inch water).


Flight Planning

4-54

Equivalent Water Depth

0.75

IN

0.5 IN

�

�


Effect of Precipitation on Takeoff DistanceUse the Effect of Precipitation On Takeoff Distance For Slats/Flaps15° chart (Figure 4-55 ) to determine runway length required.

1. Enter the chart from the left at the ambient temperature (33°F).


3. Move down to the weight reference line and parallel the weightguidelines.

4. Enter left at the takeoff weight (18,000 lbs) and move right tointersect the line paralleling the weight guidelines.

5. From the intersection move down to the equivalent water depth(EWD) reference line and parallel the EWD guidelines.

6. Enter left from the EWD (0.5 inch) and move right to intersectthe line paralleling the guidelines.

7. Move down to the slope reference line.

Assume no slope.

8. Move down to the wind reference line and parallel the windguidelines.

9. Enter from the left at the wind (20 kts). Move right to intersectthe line paralleling the guidelines.

10. From the intersection, move down to read the corrected takeofffield length (5,700 ft).

To complete the calculations, add two times the landing distance tothe corrected takeoff distance for safety margin. Use the LandingDistance S + Flaps 52° chart (Figure 4-56 ).

1. Enter from the left at the landing weight (18,000 lbs)

2. Move right to the pressure altitude (3,000 ft) and down to theheadwind (20 kts).

3. Move right to read landing distance (2,600 ft).

Takeoff Distance 5,700 ftLanding Distance (2,600 X 2) 5,200 ftTotal runway required 10,900 ft

The total runway required (10,900 ft) is a considerable change froman uncorrected balanced field length (4,000 ft).


Flight Planning

4-55

Effect of Precipitation on Takeoff DistanceFlaps 15°/Two Engines Operating

4-56


33°F

18,000 LBS

0.5 IN

20 KTS

5,700 FT

2,600 FT

3,000 FT18,000 LBS

��

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Flight Planning

Effect of Precipitationon Takeoff Distance

Landing Distance


Effect of Precipitation on Landing DistanceTo determine landing field length on wet or contaminated runway,use the Landing Distance/S + Flaps 52° chart (Figure 4-57 ).

Assume a landing is to be made into an airport with:

� 3,000 ft elevation

� 33°F temperature

� 20 kt headwind

� 14,000 lbs landing weight.

1. Enter the chart from the left at the landing weight (14,000 lbs).

2. Move right to the pressure altitude (3,000 ft) and down to theheadwind (20 kts).

3. Move right to read the landing distance (2,000 ft).

4. To determine landing distance, multiply the distance (2,000 ft) bythe factors shown below.

a. by 1.5 for wet runways (2,300 ft)

b. by 2.0 for more than 1/8 inch equivalent water depth(4,000 ft)

c. by 2 (compacted snow) to 3.4 (black ice) (4,000 ft to 6,800 ft).

5. To determine landing field length, perform the operations shownbelow.

a. For wet runways, multiply the dry landing field length by1.15.

b. For contaminated runways, multiply the landing distance by2, then multiply the result by 1.15.

c. For very slippery runways (i.e., compacted snow or blackice), multiply the landing distance by 2 (compacted snow)to 3.4 (black ice) depending on conditions, then multiply theresult by 1.15.


Flight Planning

4-57


14,000 LBS

2,000 FT

�

�

3,00

0 FT

�

Falcon 10/100 Developed for Training Purposes 5-1September 2007

Several chapters contain multiple systems to facilitate a morecoherent presentation of information. The systems coveredare listed below in alphabetical order opposite the chapter inwhich they are located. ATA codes are noted in parentheses.

Air Conditioning (21). . . . . . . . . . . . . . . . . . . . . . PNEUMATICAircraft Structure (51) . . . . . . . . . . . . AIRCRAFT OVERVIEWAutopilot (22). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSBrakes (32) . . . . . . . . . . . . . . . . . . . . . . . . . . . LANDING GEARCommunications (23). . . . . . . . . . . . . . . . . . . . . . . . AVIONICSDimensions and Areas (6) . . . . . . . . . AIRCRAFT OVERVIEWDoors (52) . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWElectrical (24). . . . . . . . . . . . . . . . . . . . . . . . . . . . ELECTRICALEngine (71). . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTEngine Controls (76) . . . . . . . . . . . . . . . . . . . . . POWERPLANTEngine Fuel and Control (73). . . . . . . . . . . . . . . POWERPLANTEngine Indicating (77) . . . . . . . . . . . . . . . . . . . . POWERPLANTEquipment/Furnishings (25). . . . . . . . AIRCRAFT OVERVIEWFire Protection (26) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FIREFlight Controls (27) . . . . . . . . . . . . . . . . . FLIGHT CONTROLSFuel (28) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUELFuselage (53) . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWHydraulics (29). . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULICSIce and Rain Protection (30). . . . . . . . . . . . . . . . ICE AND RAINIgnition (74) . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTLanding Gear (32) . . . . . . . . LANDING GEAR AND BRAKESLighting (33) . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELECTRICALNavigation (34) . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSOil (79) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTOxygen (35) . . . . . . . . . . . . . . . . . . . . . . . . MISCELLANEOUSPitot/Static (34) . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSPneumatic (36) . . . . . . . . . . . . . . . . . . . . . . . . . . . PNEUMATICPressurization (21) . . . . . . . . . . . . . . . . . . . . . . . . PNEUMATICStabilizers (55) . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWStall Warning (27) . . . . . . . . . . . . . . . . . . FLIGHT CONTROLS

Systems

SYSTEM (ATA Code) CHAPTER

CAE SimuFlite

5-2 Developed for Training Purposes Falcon 10/100September 2007

Thrust Reversers (78) . . . . . . . . . . . . . . . . . MISCELLANEOUSWarning Lights (33) . . . . . . . . . . . . . . . . . . MISCELLANEOUSWindows (56) . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWWings (57) . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEW


AircraftOverview

Chapter 5A

This section presents an overview of the Falcon 10 aircraft. Itincludes major features, aircraft structures, dimensions, andhazard areas as well as a summary of service bulletins (SBs).

This section references the manufacturer’s serial number, andwhere system differences warrant, publishes separate data andschematics.


Aircraft Overview

Table ofContents

Aircraft Features Diagram . . . . . . . . . . . . . 5A-4

Airframe Description . . . . . . . . . . . . . . . . 5A-5

Fuselage . . . . . . . . . . . . . . . . . . . . . . . 5A-5

Nose Gear Box . . . . . . . . . . . . . . . . . . . 5A-5

Pressurized Section . . . . . . . . . . . . . . . . . 5A-5

Cockpit . . . . . . . . . . . . . . . . . . . . . . 5A-5

Passenger Cabin . . . . . . . . . . . . . . . . . 5A-7

Center Section Wing Box . . . . . . . . . . . . . . 5A-8

Main Landing Gear Box . . . . . . . . . . . . . . . 5A-9

Collector Tank Box . . . . . . . . . . . . . . . . . 5A-9

Fin Stub . . . . . . . . . . . . . . . . . . . . . . . 5A-9

Auxiliary Structures . . . . . . . . . . . . . . . . . 5A-9

Tail Unit . . . . . . . . . . . . . . . . . . . . . . . 5A-10

Vertical Stabilizer . . . . . . . . . . . . . . . . . 5A-10

Horizontal Stabilizer . . . . . . . . . . . . . . . . 5A-10

Engines . . . . . . . . . . . . . . . . . . . . . . . 5A-10

Wing . . . . . . . . . . . . . . . . . . . . . . . . . 5A-11

Landing Gear . . . . . . . . . . . . . . . . . . . . 5A-11

Aircraft Dimensions . . . . . . . . . . . . . . . . 5A-13

Hazard Areas . . . . . . . . . . . . . . . . . . . . 5A-15

Radar . . . . . . . . . . . . . . . . . . . . . . . . 5A-15

Engine Inlet Draw . . . . . . . . . . . . . . . . . . 5A-15

Engine Exhaust Plume . . . . . . . . . . . . . . . 5A-15

Service Bulletins . . . . . . . . . . . . . . . . . . 5A-17

VERTICALSTABILIZER

RUDDER

FLAPS

AILERON

NACELLE

PYLON

HORIZONTALSTABILIZER

ELEVATOR

AIRBRAKES

LEADINGEDGE SLATS

ENTRANCE DOOR


Aircraft Features


Aircraft Overview

The two engine Falcon 10 aircraft is amedium range, swept wing, high speedaircraft. It is a transport category air-craft certified under FAR parts 25 and36 and operates at altitudes up to45,000 ft.

Failure design and construction guar-antees aircraft life for 30,000 hours or20,000 cycles (takeoffs).

FuselageFuselage construction consists ofaluminum frames and folded stringerswith a riveted aluminum skin. Sevenmain frames carry the loads and stress-es generated by the nose gear, cabinpressurization, wings, and the empen-nage. Forty-eight secondary framesbrace the fuselage. The stringerstogether with the frames form a latticestructure to which the skin is attached.Additional fuselage structures includeribs, spars, and beams. Chemicallymilled skin panels utilize milledreinforcing structures for addedstrength around window, door, andaccess openings.

The fuselage consists of a main struc-ture, auxiliary structure, and fairings.

The main structure consists of the nosegear box, pressurized section, centersection wing box, main landing gearbox, collector tank box, and fin stub(Figure 5A-1 ).

Nose Gear BoxAn auxiliary frame and a single mainframe form the front and rear of thenose gear box. Beams on each side ofthe box and fittings on the main frameconnect the nose gear to the aircraftfuselage. The main frame of the nosegear box also forms the forward pres-sure bulkhead.

Pressurized SectionThe pressurized section contains thecockpit, entrance area, passenger cabin,and rear baggage area.

Cockpit

The two-crew cockpit is insulated fromheat, cold, and noise (Figure 5A-2 ,following page). Each pilot has a flightstation with a conventional controlcolumn and an instrument panel. Acenter instrument panel contains theengine instruments, fuel quantity andflow indicators, landing gear controls,

AirframeDescription

PRESSURIZED ZONEENGINEPYLON

FIN STUBSTEM

FIN STUBLEADING EDGE

FIN STUBROOT FAIRING

FIN STUB TAIL CONE

COLLECTOR TANKBOX STRUCTURE

MAIN L/G BOXSTRUCTURE

CENTERSECTIONWING BOX

NOSE GEARWELL BOX

NOSE COWL

5A-1


weather radar/flight management dis-plays, and communication and navi-gation control units. A pedestalbetween the pilot stations contains thepower levers, secondary flight controlselectors, and autopilot and flight direc-tor equipment controls.

Side consoles contain oxygen masksand microphone jacks. The pilot’s con-sole also has a nosewheel steeringwheel. An overhead console containsengine starting, fuel, and electrical andlighting controls. Panels behind thecrewmembers’ heads contain most ofthe circuit breakers accessible in flight.

Six windshields/windows provide vis-ibility from the cockpit. These consistof two curved windshields, two open-ing windows, and two fixed rear win-dows (Figure 5A-3 ).

Each windshield consists of four glasspanes separated by three layers ofbutyryl plastic with a heating elementfor anti-icing.

Each triangular opening window con-sists of two panes of oroglass bondedto a metal frame. Air circulatingbetween the two panes from two holesdemists the window. An opening andlatching mechanism allows each win-dow to be opened inward at an anglefor ventilation and direct viewing.

Depending on service bulletin status,the rear windows consist of either twopanes of oroglass bonded to an orlonshim or three panes of oroglass bondedto a dacron shim. Air circulating throughan airspace between the panes demiststhe windows. An electrical heating ele-ment demists the three-pane windows.

5A-2 5A-3

5A-55A-4


Aircraft Overview

Passenger Cabin

The pressurized, air conditioned pas-senger cabin is certified for eight pas-sengers, though typical configurationscomfortably seat seven (Figure 5A-4 ).Accommodations in the aircraftinclude a galley, lavatory, work tables,and a coat closet. The entrance dooris on the left side of the fuselage. Asingle overwing emergency exit is onthe right side.

The two piece (upper and lower)entrance door (Figure 5A-5 ) can beopened from either inside or outside;three steps are built into the lowerpanel. Hooks on the lower panelengage catches in the upper panel tosecure the door halves. A counter-balance system assists door openingand closing (Figure 5A-6 ).

To open the door from the outside,unlock the door and remove the key.Push on the locktab to release the outerdoor control handle. Lift the handle.If the two halves catch, press on thelower half while lifting the top. Thetwo door halves are interconnectedthrough a counterbalance system.Restrain either half to control open-ing. To lock the doors in the open posi-tion, lift on the upper half.

To close the door from the outside, liftthe lower door to drive the upper doordown. Once the lower half is closed,push the upper half against the doorframe. Push the door opening handledown until it is flush with the door.Lock the door with the key if necessary.

To open the door from the inside,depress the pushbutton to release the

CABLE PULLEY

BUFFER

PULLEYASSEMBLYRAIL

5A-6


door locks and open the depressuriza-tion door. Lift up on the internal con-trol handle to retract the side studs andrelease the door hooks. Push down onthe handle to open the door halves.Once the doors are fully open, pushthe control handle up to engage thesafety lock.

To close the door from the inside,depress the latch to release the door.Move the operating handle upwardsto drive the upper and lower halvestoward the closed position. Once thehalves close, pull the door handle outand rotate it downwards to lock thedoor.

An emergency locking mechanismlocks the lower door half for ditchingpurposes. The upper half can beopened for emergency egress in theevent of ditching.

There are three passenger windows oneach side (Figure 5A-8 ); the secondwindow on the right is part of theemergency exit. Some aircraft have aseventh window opposite the entrancedoor (three left, four right); the thirdwindow on the right is part of theemergency exit. Each window consists

of two panes of plastic with an airspacebetween each pane. Cabin air circu-lating between the panes demists thewindows.

The emergency exit consists of a 20by 36 inch panel with a window on thelower part and a locking mechanismon the upper part (Figure 5A-7 ). Theexit is hinged on the bottom to balljoints. The exit opens from either theinside or outside; once unlatched, theexit opens inward.

Center Section Wing BoxThe center section wing box, con-structed of spars, ribs, skin panels, andplates, carries wing bending movementand transmits twist, lift, and drag fromthe wings to the fuselage. The centersection wing box also forms the fuse-lage fuel tanks. The aft spar of thisstructure also contains fittings for themain landing gear side stays. Twomain frames connect the structure tothe fuselage. Dual skin constructionisolates the wing box from the pres-surized fuselage section. The airspacebetween the dual skins is drained andventilated.

5A-7

5A-8


Aircraft Overview

Main Landing Gear BoxThe main landing gear box enclosesthe landing gear and protects the air-craft structure against tire burst.

Collector Tank BoxThe collector tank box is aft of the cen-ter section wing box. The tank boxcontains two separate feeder tanks forthe engines.

Fin StubThe fin stub secures the empennage tothe fuselage frame. It consists of a frontspar, rear spar, oblique rib, ribs, andstiffeners. The horizontal stabilizer ishinged to the rear of the fin stub. Thevertical stabilizer attaches directly tothe stub.

Auxiliary StructuresThe auxiliary structure consists of thefloors and the engine pylons. Thenosecone cowl, tailcone, fin stub root,wing-to-fuselage fillet, fin stub lead-ing edge, and stem fairings provideaerodynamic smoothing of structuresto each other.

Fixed and removable panels form thefloor of the pressurized section.Removable panels allow access to con-trol cables and equipment. Each panelis secured to beams that are supportedby stringers.

The engine pylons are box structuresconstructed of skin panels, ribs, stiff-eners, and firewalls (Figure 5A-9 ). Thestructures transmit the forces producedby the engines to the fuselage. Eachpylon also includes pylon-to-enginefittings and pylon-to-fuselage fittings.

Each pylon consists of a front sectionthat forms the leading edge, a centerbox structure for engine electrical, fuel,hydraulic, and bleed air lines, and arear section.

A hinged nosecone cowl (Figure 5A-10) allows easy access to avionicsequipment forward of the cockpit. Thecowl is constructed of frames bracedby stringers, plates, and skin panels.A honeycomb structure covered byfiberglass forms the radome.

The tailcone is constructed of frames,stringers, skin panels, doors, and anend cap. The end cap contains the

5A-9 5A-10


Freon line condenser, air-conditioninglines, heat exchanger, and tail naviga-tion light. Doors in the tailcone allowaccess to various equipment for main-tenance purposes. An optional com-partment provides 28.6 cubic feet forbaggage storage.

The wing-to-fuselage fillet, construct-ed of beams, plates, stiffeners, and skinpanels, aerodynamically fairs the wingsto the fuselage structure. Removablefillets and doors allow access to vari-ous equipment and the center sectionfuel tank.

Tail UnitThe tail unit (Figure 5A-11 ) consists ofthe vertical stabilizer (fin), moveablehorizontal stabilizer, elevators, and arudder (see Flight Controls).

Vertical StabilizerThe vertical stabilizer is a torsion-typebox structure composed of a forwardand aft spar connected with ribs andangles covered with aluminum skinpanels (Figure 5A-12 ). The structureencloses the hinge and control systemsfor the rudder and the VOR antennamounts. It is attached to the fin stub attwo points. The stabilizer consists of aleading edge, box structure, upper fair-ing, and rudder. The rudder is hingedat three points. A spring-attachedshroud seals the gap between the ver-tical stabilizer and the rudder.

The vertical stabilizer transmits theaerodynamic loads created by the sta-bilizer and rudder to the fuselage.

Horizontal StabilizerThe horizontal stabilizer consists of abox structure that forms the main struc-ture, two detachable leading edges,two detachable tips, and two elevators(Figure 5A-13 ). The stabilizer and ele-vators are constructed of spars, ribs,stiffeners, doublers, and skin panels.The moveable horizontal stabilizer pro-vides elevator trim. Two spring-loadedribs seal the gap between the move-able horizontal stabilizer and the ver-tical stabilizer.

A fixed hinge at the rear of the stabi-lizer connects the unit to the fin stub.A scissors attachment driven by anelectric motor allows the horizontalstabilizer to be trimmed along the hor-izontal axis.

An elevator on each side of the verti-cal stabilizer is hinged to the horizon-tal stabilizer at three points. A cou-pling unit joins the left and rightelevators.

EnginesTwo Garrett Airesearch TFE-731-2-1C front fan engines power the air-craft. Each engine produces 3,230 lbsof thrust up to 86°F at sea level (seePowerplant chapter).

5A-12 5A-135A-11


Aircraft Overview

The TFE-731 engine is of modulardesign that consists of a fan, planetarygearbox, compressor and high pressureturbine section, low pressure turbinesection, and accessory gearbox module.

WingEach swept-back wing (Figure 5A-14 )consists of a main frame wing centerbox structure with an auxiliary struc-ture and moveable components. Theauxiliary structure consists of inboardand outboard fixed leading edge boxes,flap and aileron baffle boxes, mainlanding gear wheel well ceiling, andwing tip fairing. The moveable com-ponents include aileron, inboard andoutboard slats, inboard and outboardflaps, and inboard and outboardairbrakes.

The wing center box structure carriesand transmits bending and torsion loadsand stresses to the center box structure.The wing center box structure containsa front spar, rear spar No. 1, rear sparNo. 2, 18 ribs, and skin panels.

The rear spars contain fittings for themain landing gear front hinge bearingand the wing flaps and ailerons. Thefront spar carries hinges for the inboardand outboard leading edge slats.

On the trailing edge of each wing aretwo flaps and a single aileron; each ishinged at three points (Figure 5A-15 ).

These control surfaces are of conven-tional construction with aluminumskins, spars, and ribs. On the top of thewing, there are inboard and outboardairbrakes constructed of aluminumhoneycomb with aluminum skins. Theinboard and outboard slats are on thewing leading edge (Figure 5A-14 ).The slats consist of a fixed inboard boxwith a moveable outboard box. Theinboard box provides a passagewayfor aileron control linkages, hydrauliclines, and electrical wires. All compo-nents are constructed of aluminum.Each slat rides on a roller and is drivenby irreversible jacks.

Landing GearThe aircraft is supported by a conven-tional tricycle landing gear that con-sists of a single wheel nose gear anddual wheel main landing gear (Figure5A-16). An oleo-pneumatic (air-oil)shock absorber on each gear assemblyabsorbs landing forces (see LandingGear chapter).

When retracted, the nose landing gearis completely covered by a pair ofdoors. Each main landing gear has asingle strut-connected door that onlycovers the strut assembly. The out-board wheel on each main landinggear assembly has an embellisher(hub cap).

5A-15 5A-165A-14

19' 1"

42' 11"

45' 5.5"

40' 11"

17' 4.7"

15' 1

"

9' 5"


Aircraft Dimensions


Aircraft Overview

Length . . . . . . . . . . . . . . . . . . . . . . . . . . 45´ 5.5˝Height . . . . . . . . . . . . . . . . . . . . . . . . . . 15´ 1.0˝Fuselage Length . . . . . . . . . . . . . . . . . . . . . 40´ 11.0˝Wing Span . . . . . . . . . . . . . . . . . . . . . . . 42´ 11.0˝Horizontal Stabilizer Span . . . . . . . . . . . . . . . . 19´ 1.0˝Wheel Base . . . . . . . . . . . . . . . . . . . . . . . . 17´ 4.7˝Wheel Track . . . . . . . . . . . . . . . . . . . . . . . . 9´ 5.0˝

Passenger Cabin:Length . . . . . . . . . . . . . . . . . . . . . . . . . 16´ 3.0˝Width (maximum) . . . . . . . . . . . . . . . . . . . 4´ 9.0˝Height (minimum) . . . . . . . . . . . . . . . . . . . 4´ 7.0˝Height (maximum) . . . . . . . . . . . . . . . . . . 4.´ 10.0˝

Cabin Windows:Height . . . . . . . . . . . . . . . . . . . . . . . . . . 1´ 3.7˝Width . . . . . . . . . . . . . . . . . . . . . . . . . . 1´ 1.3˝

Emergency Exit:Height . . . . . . . . . . . . . . . . . . . . . . . . . . 3´ 0.0˝Width . . . . . . . . . . . . . . . . . . . . . . . . . . 1´ 8.0˝

Entrance Door:Height . . . . . . . . . . . . . . . . . . . . . . . . . . 4´ 2.2˝Width . . . . . . . . . . . . . . . . . . . . . . . . . . 2´ 5.5˝

AircraftDimensions


7' 0" RADIATION

7' 0" RADIATION

12' 0"

INLET DRAW

12' 0"

ENGINE EXHAUST MAXIMUM POWER130'

ENGINE EXHAUST MAXIMUM POWER130'

Aircraft Hazard Areas


Aircraft Overview

Radiation emissions from the weatherradar, engine inlet draw, and engineexhaust are primary dangers duringground operations of the aircraft.

RadarHazards exist to personnel, equipment,and other aircraft when operating theweather radar on the ground. A poten-tial fire ignition hazard exists within300 ft of refueling aircraft. Radiationhazards to personnel exist within 7 ftforward of the aircraft in a 270° sector.Pointing the aircraft toward largeobstructions, hangars, buildings, andother metallic objects within 300 ft canresult in radar equipment damage.

Engine Inlet DrawEngine inlet draw at full power is haz-ardous within 12 ft of the engine inlets.The draw increases as the distance tothe inlet decreases.

Inlet draw at high power settings cancause the engine to ingest small objects(i.e., bolts, gravel) that can result inserious engine damage. Personnel haz-ards exist immediately forward of theinlet.

Engine ExhaustPlumeHazards are present due to the hightemperature, high velocity exhaustfrom the engines. Engine exhaust haz-ards exist within 130 ft of the nozzle.

Advise ground personnel of imminentengine starts. Do not start an enginewithout first verifying that the imme-diate area behind the aircraft is clearof personnel, small articles, or otheraircraft and vehicles. Excessive thrustduring taxi can result in serious dam-age to other aircraft, vehicles, struc-tures, and personnel.

Hazard Areas


Aircraft Overview

The Service Bulletins (SBs) addressed by the manual are listedbelow in numeric order.SB 24-006 (0154);Electrical Power – DC Power – 22G RelaySupplied by the Main Bus Bar (S/Ns 001 to 123, except 108, 113,118, and 121).SB 27-013;Flight Controls – Dual Horizontal Stabilizer TrimControl Microswitch Equipped Control Wheels (Optional – All ).SB 29-007;Hydraulic Power – Intake Filter in the Reservoirs –Installation of a Paper Filter (All ).SB 29-009;Hydraulic Power – Engine-Driven Pumps – Replacementof ABEX Pumps by Vickers Pumps (S/Ns 001 to 175 and 185).SB 30-001;Ice and Rain Protection – Cockpit Electrical De-misting of Rear Glass Panels (S/Ns 002 to 019 except 005, 007,014, and 017).SB 30-010 (0121);Ice and Rain Protection – Nacelle – EngineAnti-Icing (S/Ns 001 to 085, except 071, 076, and 082).SB 30-012;Ice and Rain Protection – Passenger Cabin Windows –Installation of Dessicators (S/Ns 001 to 119).SB 30-017;Ice and Rain Protection – Wing Leading Edge SlatAnti-Icing – Hot Air Leak Warning (S/Ns 001 to 191).SB 72-001 (0180);Engines With Cone-Shaped Spinners and NoAnti-Icing (All ).SB 77-003 (0074);N1 and N2 Tachometers – Doubled VoltageSupply With Battery Start on Low Temperature Start (S/Ns 001 to052).SB 78-001;Exhaust Grumman Tailpipes – Installation with a BackupRing (Aircraft with Tailpipes; F10A5B10-03-T).

ServiceBulletins


Avionics

Chapter 5B

Specific avionics systems vary with customer preference; manyoptions are available. Therefore, this section provides only abrief introduction to some of the basic types of avionics equip-ment available on the Falcon 10/100.

In addition, this chapter includes the pitot/static system andinstrumentation not addressed in other chapters. The pitot/static system supplies dynamic and static air pressure for theair data computer and some flight instruments.

Cockpit panel art is provided to locate instruments.


Avionics

Table ofContents

Cockpit Instrument Panel Schematic . . . . . . . 5B-6

Pedestal Instrumentation . . . . . . . . . . . . . . 5B-7

Pitot/Static System Schematics . . . . . . . . . . 5B-8

Air Data Systems . . . . . . . . . . . . . . . . . . 5B-9

Pitot/Static System . . . . . . . . . . . . . . . . . . 5B-9

Air Data Computer (ADC) . . . . . . . . . . . . . 5B-10

Total Temperature System . . . . . . . . . . . . . 5B-10

Altimeters . . . . . . . . . . . . . . . . . . . . . 5B-10

Pilot’s Altimeter . . . . . . . . . . . . . . . . . 5B-10

Copilot’s Altimeter . . . . . . . . . . . . . . . . 5B-11

Standby Altimeter . . . . . . . . . . . . . . . . 5B-11

Mach/Airspeed Indicators . . . . . . . . . . . . . 5B-11

Vertical Speed Indicators . . . . . . . . . . . . . 5B-12

Speed, Mach, and Altitude Warning . . . . . . . . 5B-12

Additional Instrumentation . . . . . . . . . . . . . . 5B-13

Digital Clock . . . . . . . . . . . . . . . . . . . . 5B-13

Magnetic Compass . . . . . . . . . . . . . . . . 5B-13

Standby Gyro-Horizon . . . . . . . . . . . . . . . 5B-13

Avionics Equipment . . . . . . . . . . . . . . . . 5B-15

Communications . . . . . . . . . . . . . . . . . . 5B-15

VHF Communications . . . . . . . . . . . . . . . 5B-15

HF Communications . . . . . . . . . . . . . . . . 5B-15

Interphone . . . . . . . . . . . . . . . . . . . . . 5B-16

Audio Control Units . . . . . . . . . . . . . . . . 5B-16

Static Discharging . . . . . . . . . . . . . . . . . 5B-17

Navigation . . . . . . . . . . . . . . . . . . . . . . 5B-17

VHF Navigation . . . . . . . . . . . . . . . . . . 5B-17

Instrument Landing System . . . . . . . . . . . . 5B-18

Automatic Direction Finding . . . . . . . . . . . . 5B-18


Distance Measuring . . . . . . . . . . . . . . . . 5B-19

Transponder . . . . . . . . . . . . . . . . . . . . 5B-19

Radio Magnetic Indicators . . . . . . . . . . . . . 5B-19

Radio Altimeter . . . . . . . . . . . . . . . . . . 5B-19

Vertical Gyros . . . . . . . . . . . . . . . . . . . 5B-20

Directional Gyro . . . . . . . . . . . . . . . . . . 5B-20

Avionics Switching Panel . . . . . . . . . . . . . 5B-20

Weather Radar System . . . . . . . . . . . . . . . 5B-22

Autopilot/Flight Director System . . . . . . . . . 5B-23

Autopilot . . . . . . . . . . . . . . . . . . . . . . . 5B-23

Autopilot Computer . . . . . . . . . . . . . . . . 5B-23

Autopilot Amplifier . . . . . . . . . . . . . . . . . 5B-23

Servos . . . . . . . . . . . . . . . . . . . . . . . 5B-23

Autopilot Controller . . . . . . . . . . . . . . . . 5B-23

Controls and Indicators . . . . . . . . . . . . . . 5B-24

Flight Director . . . . . . . . . . . . . . . . . . . . 5B-24

Flight Director Computer . . . . . . . . . . . . . 5B-24

Mode Selector . . . . . . . . . . . . . . . . . . . 5B-25

Instrumentation . . . . . . . . . . . . . . . . . . 5B-25

Attitude Director Indicator . . . . . . . . . . . . 5B-25

Horizontal Situation Indicator . . . . . . . . . . 5B-26

Electronic Flight Instrument System . . . . . . . 5B-27

Display Processor Unit . . . . . . . . . . . . . . . 5B-27

Multifunction Processor Unit . . . . . . . . . . . . 5B-27

EADI and EHSI . . . . . . . . . . . . . . . . . . . 5B-27

Display Control Panels . . . . . . . . . . . . . . . 5B-28

Multifunction Display Unit . . . . . . . . . . . . . . 5B-28

Reversionary Switching Panels . . . . . . . . . . . 5B-29

Course Heading Panel . . . . . . . . . . . . . . . 5B-29


5B-6

CockpitInstrument

Panel

Avionics

Preflight and Procedures . . . . . . . . . . . . . 5B-31

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5B-31

Abnormal Procedures . . . . . . . . . . . . . . . . 5B-31

Pilot’s Pitot/Static Pressure . . . . . . . . . . . . 5B-31

Copilot’s Pitot/Static Pressure . . . . . . . . . . . 5B-31

Landing without Speed Information . . . . . . . . 5B-31

Abnormal Airspeed Indication at High Altitude . . 5B-31

Air Data Computer Inoperative . . . . . . . . . . 5B-31

Compass System Failure . . . . . . . . . . . . . 5B-32

Pitch Trim Failure . . . . . . . . . . . . . . . . . 5B-32

Out-of-Trim Condition . . . . . . . . . . . . . . . 5B-32

Emergency Procedure . . . . . . . . . . . . . . . 5B-32

Pitch Servo-Actuator Runaway . . . . . . . . . . 5B-32

1.51.41.31.2

.0.2

.4

.6

.81.0V/Vs

ANGLE

OF

ATTACK

RE

F

SE

T

200

350

300

250180

160140

120

100

MACH

KNOTS

PUSH

TEST

AIRSPEED

DIMNM MIN KTS TEST

TAT

C

PUSHTAS

EMG

VHF 1

VHF 1

SPK

VHF 2

VHF 2

S.T.

VHF 3

VHF 3

VOL

HF

HF

NAV 1

ADF 1

MASK

NAV 2

ADF 2

PAGE

DME 1

FILT

CABIN

DME 2

MKR

C PT

AUDIO

MICROPHONE

AUDIO

SLAVE HDG

+

+PANELONLYNORMAL

PITOT/STATICSELECTOR

DG MAG

20 20

10 10

10 10

20 20

FAST

SLOW

DH0 0 0

DH

DHSETTEST

ATTTEST

N

W

S

E

30

24 21

1512

63

33

MAG

NM ET

ANG

GS

CRS0 4 4

NAV 1

NAV 2

COMP 1

COMP 2

VG 1

VG 2ADCFAIL

EM PWRON

F/S A/A

F/S A/S

VHF NAV

NAVEM PWROFF

ANG DEV

LIN DEV

HSI/ADI

VG

FAST ERECT

BRTADC TESTDIM

TRIM

TRIMFAILTRIMFAIL

VLFDRVD

FAILAP

DISC

VLFWPT

HDG

APXFER

VLFBATTLATARMGS

ARM

ALTSFILATCAPGS

CAP

ALTHOLDBACKLOG

GA

VERT

FOGDR

SMOKEDET

C

M

A

COMPWARN

MAX SPEED SLATSAND FLAPS DOWN

SLATSSLATS + 15¡SLATS + 30¡SLATS + 50¡

200Kts190Kts165Kts165Kts

190Kts220Kts

MAN G VLOEXTENDED VLV

MAN G SPEED. VA 220Kts

MAX SPEED L G

DI VE

CLI MB

PULL

TO

CAGE

ALTALERT

PUSH TO

CANCEL

, 002 0ALTITUDE SET

PUSH100

20

15

10 54

3

2

1

RAD ALT

DH

1 2

4.5

.5

12

4

UP

DN60

VS

8

90

2

3

45

6

7

192

001

901

901

2 9 9 2

EMERSTOW

NORM

EMERSTOW

NORM

T/R

TRANS

REV

TRANS

REV

200250

300

350400

3

3

6

5

4

6080

100

120

160

140

180

MACH

KNOTS

0

3 69

1215

18

212427

3033

ADF

ADF

VOR VOR

20 20

10 10

10 10

20 20

FAST

SLOW

DH0 0 0

DH

DHSETTEST

ATTTEST

N

W

S

E

30

24 21

1512

63

33

MAG

NM ET

ANG

GS

CRS0 4 1

DIMNM MIN KTS TEST

SLAVE HDG

+

+ PANELONLYNORMAL

PITOT/STATICSELECTOR

DG

MAG

EMG

VHF 1

VHF 1

SPK

VHF 2

VHF 2

S.T.

VHF 3

VHF 3

VOL

HF

HF

NAV 1

ADF 1

MASK

NAV 2

ADF 2

PAGE

DME 1

FILT

CABIN

DME 2

MKR

C PIT

AUDIO

MICROPHONE

AUDIO

ANG DEV

LIN DEV

HSI/ADIVD

FAST

ERECTDIM

1 24

6

1 24

VERTICALSPEED

1000 FT PER MIN.

IVSI.5UP

DOWN

.5

0

9

8

7

65

4

3

2

1

ALT

109

2 0 0MILLIBARS

IN HG

29.92

1013

0

TIME

ET

FT

ZERO

RUN

STOP

SET DIM

UP D 1 hrup

B

COMPWARN

10

8642 DC AMPS

TRIM

TRIMFAILTRIMFAIL

VDFAILAP

DISC

FLITEFONE

HDG

APXFER

AFTSMOKE

LATARMGS

ARM

ALTHOLDLATCAPGS

CAP

AUXHEATBACKLOG

GA

FOGDRALT

WARN

C

M

A

OXYGEN PSI X 100

5 10 15 20

PASSENGER

OXYGEN

OFF MAN

AUTO+

ONSURGE

USE NO OIL

MAX PRESSURE 160 PSI PU

SH

TOT

UR

N

30

29

31

IN.HS

1

2

3

4

56

7

8

9

0

23 25

30

3540

45

50

FLIG

HT

ALTITUDE

CABIN

1000

FT.

IDC

PUL

L

FOR

BA

R

OPULL

+60 +40+20

0

-20-40-60

RAM AIRC

FUEL USED

POUNDS1 0 6 4

012

34 5 6

789

10CABIN

PSI

4

0

12

8

16CABINFT x 1000

1.5

1.5

2

VERTICALSPEED

X1000 FT.PER MIN.

1

1UP

DOWN

0

.5

.5

0

5

1015

20

25

30

FUEL QTY

LBS x 100

0

5

1015

20

25

30

FUEL QTY

LBS x 100

8

7

6

54 3 2

10

9 10EXH

TEMP

C x 100

1

23

4 5 67

89

0

N1%RPHx10

024

6

89

10N2

%RPHx10

PPHx 100

20

15

10

5

0

OIL¡C PSI

150100

50 20

8060

40

8

7

6

54 3 2

10

9 10EXH

TEMP

C x 100

1

23

4 5 67

89

0

N1%RPHx10

024

6

89

10N2

%RPHx10

PPHx 100

20

15

10

5

0

OIL¡C PSI

150100

50 20

8060

40

¡F

CABIN0

40 80

120AIR

VALVE

COLD HOT180160140120¡F

OFF

180160140120¡F

OFFWARM HOT

1 2BAT. TEMP

COLD HOT

HOTCOLDMANUAL

HYDR

QTY PSI x 1000

4/4

1/2

0

4

3

2

1

0

HYDR

QTY PSI x 1000

4/4

1/2

0

4

3

2

1

0

FLAPSUP

DN52 30

15

0

AIRBRAKE

FLAPASYM

BRIGHT

DIM

TEST

LDG GEAR

MOVING

DOWN

FIRE2

PULL

FIRE1

PULL

XTING

TEST

210

TEST

210

T/R HORNTEST

TOT

REAR TANK

TEST

GENE. 1

PITOT 1

HYD. 1

HYD. TK1

AIRBRAKE

GENE. 2

PITOT 2

HYD. 2

HYD. TK2

STAB TRIM

BATT

Q. UNIT

ST. BYPUMP

P. BRAKE

BRAKE

ENG. C1

FUEL P1

OIL 1

GUST

ENG . C2

FUEL P2

OIL 2

LO. FUEL

HOT BAT

CABIN

CONDO. HEAT

TGT 25

20

5

1015

TEST1 2 3

RDR

NAV

RMT

PGE

ENG

PWR INT

RCL SKP CLR

TCT HLD STB

DATA

MODE

MAP

WXNORMTEST

STBY

OFF

GAIN

MAXMIN

TILT

-10-5

0

+5+10

RANGE

10

2550 100

200

300

VOLOFF

109.50NAV-1

MKRHI

NAV TEST

HOLD

DME

VOLOFF

VOLOFF

VOLOFF

119.10 127.40COMM-1

135.40COMM-2

HOLD

DME

MKRHI

NAV TEST

115.90NAV-2

TEST

TFR

TEST

TFR

373.0ADF

OFFANT

ADFIDNT

OFF

TEST

FREQUENCY

7325ATC

ALTRPTG

OFF

TFR

1 2IDNT TESTREPLY

STBYON

LO

RANGEEXTENDER

TESTBATTERY

FLTTIME

RESET

FLTTIME

RESET

FPAFAIL

SATKNOTS 0

3 69

1215

18

212427

3033

ADF

ADF

VOR VOR

DE ICING AIR DATA

BRT


Cockpit Instrument Panel

AIRBRAKE

WS

HL

D

FL

OO

R

OPEN

CLOSE

BO

OM

PH

ON

ES

BO

OM

PH

ON

ES

30TO

30L RAILERON

RUDDERL R

40 40TO

NOSEDN

NOSEUP

STAB

0

2

46

810

12

ANTI SKID HYDR.ST. BY. PUMP

TESTON

OFF

ON

OFF

BRAKEL

BRAKER

FREON

AUTO

HPBLEED

OPEN

COND GVALVE

AUTO

OFF CLOSED OFF

PRESSURIZATION

LIFT

AUTO

DUMP

UP

DOWN

S + FLAPS 15¡

S + FLAPS 30¡

S + FLAPS 52¡

SLATS

CLEAN

%

ON

EXT

DOWN

UP

STABEMG

AILERON

TRIM

RUDDER

EMERGSLATS

CAB

VMO

STALL 1

STALL 2

CRS CRS

HDGCRS 1 CRS 2

HDG TTGSPD

ET

TTGSPDET

VOR LOC APPRHDG1/2 BANK

MACH IASVSALT SEL ALT

VOR LOC APPRHDG1/2 BANK

AP XFR

TURB

DN

UP

TURN YD

DISENGAGED

APENGAGED

HORN SILENCE

PARK BRAKENORMAL

OMEGA NAV

ALT NAV

NORMAL

ON OFF

VOICEADVISORY

Z *

GLOBAL

D

HOLD

321

4 5 6

987

# 0 –

AFIS

BACK

NXT

PRV

MSG

I J K L M N O

A B C D E F G H

P Q R

S T U YXWV S

ENTER

NAV YNAV FPL ON BRTN PLAN HDG TUNE


5B-8

Pitot/StaticSystems

Avionics

Pedestal Instrumentation

PILOT'S STATICPRESSURE

SELECTOR VALVE

PILOT'S PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

STANDBY STATIC

AIR DATAINDICATOR

ALTITUDE/ALERTPRESELECTOR

ALTIMETER

VERTICAL SPEEDINDICATOR

MACH/AIRSPEEDINDICATOR

ADC AP IAS ANDQ SENSE

ALTIMETER

CABIN DIFFEREN-TIAL PRESSURE

INDICATOR

APZ CAPSULE

PITCHARTHUR-Q UNIT

RUDDER ARTHUR-Q UNIT

VMO/MMOWARNING

STANDBYALTIMETER

DUAL SHUTOFFVALVES

PILOT'S STATIC PRESSURE

COPILOT'S STATIC PRESSURE

COPILOT'S PITOTPORT (HEATED)

PILOT'S PITOTPORT (HEATED)

TEMPERATUREPROBE

STANDBYSTATIC PORT

MMODETECTOR

VERTICAL SPEEDINDICATOR

MACH/AIRSPEEDINDICATOR

ELECTRICAL CONNECTION

MECHANICAL CONNECTION

PRESSURECONTROL BOX

FLIGHTRECORDER

AILERON/ELEVATOR

ARTHUR JACK

PILOT'S PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

STANDBY STATIC

LHSTATIC PORT

CONNECTINGDRAIN UNIT

IASMI

VSIALTIASMI

VSI ALT

DIFFERENTIALPRESSUREINDICATOR

DRAINUNIT

RH STATICPORT

DUALCUT-OUTVALVE

MDETECTOR

MOSTATICSELECTORVALVE

PILOT'SDISTRIBUTOR

AND DRAIN UNIT

COPILOT'SDISTRIBUTOR

AP AND ARTHURIAS CAPSULE

DRAINUNIT

DRAINUNIT

PILOT'S PITOTPROBE

COPILOT'S PITOTPROBE

(3) STANDBYSTATIC PORT

MAIN CONNECTING UNIT

AP ALTCAPSULE

RUDDERARTHUR JACK


Pitot/Static SystemWithout Air Data Computer

Pitot/Static SystemWith Air Data Computer


Avionics

Air Data Systems include pitot probes,static ports, air temperature sensors,air data computers, and related instru-mentation.

Pitot/Static SystemThe pitot/static system consists of twoindependent systems for the pilot andcopilot. Each system consists of:� two heated pitot probes� two heated dual static ports� one unheated standby static port.

A pitot probe (Figure 5B-1 ) on thelower fuselage on each side of the noselanding gear supplies pitot pressure tothe pilot’s and copilot’s pitot/static sys-tems.

A heated dual static port (Figure 5B-2) is on each side of the fuselage. Oneport on each side supplies the pilot’sstatic system; one port on each sidesupplies the copilot’s system. A selec-tor valve on the pilot’s instrumentpanel supplies static pressure from thestandby static valve in the noseconein the event of pilot’s system mal-function.

The pilot’s pitot/static system direct-ly supplies the:� Mach/airspeed indicator� barometric altimeter� vertical speed indicator.

Some aircraft use an air data computer(ADC) to supply pitot, static, and ramair temperature information to thepilot’s instruments. The ADC suppliesthe:� electric Mach/airspeed indicator� electric vertical airspeed indicator� electric altimeter� altitude alerter� temperature indicator.

The copilot’s pitot/static systemdirectly supplies the:� Mach/airspeed indicator� barometric altimeter� vertical speed indicator� Mach/airspeed warning switch� Arthur and autopilot airspeedswitches.

The copilot’s system indirectly suppliesthrough a shutoff valve the:� cabin differential pressure indicator� Arthur unit actuator for pitch androll� Arthur unit for rudder control� autopilot altitude sensitive switch� flight recorder (if installed)� air data computer (if installed).

The copilot’s pitot/static selector valveis a dual shutoff valve. In NORMAL,

5B-1 5B-2

Air DataSystems


both the instruments and auxiliaryequipment are supplied with pitot/static pressure. In PANEL ONLY, thevalve supplies the copilot’s instrumentsand isolates the auxiliary equipment.

Drain traps within the system allowresidual water removal from the pitotand static pressure lines. Plastic bowlswithin the drains indicate the presenceof water at the traps.

Air Data Computer (ADC)The ADC receives pitot and static pres-sure data from the pilot’s pitot/staticsystem and ram air temperature datafrom a probe. The ADC then computesaltitude, corrected altitude, verticalspeed component, indicated airspeed,Mach, maximum speed, and air tem-perature. This information is then fedto the pilot’s instruments, autopilot,and flight director systems.

The ADC is in the nosecone and issupplied with 28V DC from the A busthrough a circuit breaker.

A test button on the front of the ADCallows for system testing on theground. The AIR DATA light illumi-nates for one-half second to indicateproper system operation. If the AIRDATA light illuminates steadily inflight, the system is malfunctioning.

Total Temperature SystemAn external total temperature probeon the forward right side of the fuse-lage provides ram air temperature datato the ADC. The ADC provides totalair temperature (TAT) information tothe total air temperature indicator onthe pilot’s instrument panel (Figure5B-3). The indicator reads from -60°Cto +60°C.

A heating element powered by 28VDC from the A bus provides anti-icingprotection.

AltimetersTwo altimeters are on the Falcon10/100 aircraft: pilot’s and copilot’s.A third standby altimeter is optional.

Pilot’s Altimeter

On aircraft without an ADC, the pilot’saltimeter is a barometric unit thatreceives data from the static ports. Theunit displays altitude from -999 to+50,000 ft with a pointer that makesone revolution per 1,000 ft and a drumdisplay. An adjustable pressure displayallows the setting of local barometricpressure on the indicator.

On aircraft with an ADC, the pilot’selectric altimeter receives data fromthe ADC. The altimeter (Figure 5B-4) displays altitude with a single point-

5B-3 5B-4


Avionics

er that makes one revolution per 1,000ft and a drum display. The drum dis-play indicates aircraft altitude from -1,000 to +50,000 ft in 20-ft increments.A window indicates the barometric ref-erence pressure in inches of mercuryand millibars. A BARO knob on thelower right corner of the instrumentallows for the adjustment of baromet-ric pressure. Pushing the knob inadjusts for inches of mercury (In Hg);pulling it out adjusts for millibars.

The pilot’s altimeter provides encodedaltitude information to the transponder.

Pressing the TEST button on thealtimeter displays a red flag and thealtimeter indicates 750 ft. Absence ofeither indicates altimeter or servo-sys-tem failure.

Copilot’s Altimeter

The copilot’s altimeter (Figure 5B-5 )is a barometric unit that displays alti-tude from -999 to +50,000 ft with apointer that makes one revolution per1,000 ft and a drum display. The unitreceives static pressure data from thecopilot’s pitot/static system.

Two small windows display baromet-ric pressure in inches of mercury (InHg) and millibars. A knob on the lowerright corner allows for the adjustmentof barometric pressure.

On aircraft with a second ADC, thecopilot’s altimeter is similar to thepilot’s electric altimeter.

Standby Altimeter

Some aircraft have a standby baro-metric altimeter on the copilot’s instru-ment panel. The unit displays altitudefrom -1,000 to +60,000 ft with threeconcentric pointers. Barometric pres-sure is indicated in inches of mercury(In Hg) with a drum type display.

The unit receives static pressure datafrom the copilot’s pitot/static system.

Mach/Airspeed IndicatorsOn aircraft without an ADC, the pilotand copilot Mach/airspeed indicatorsare identical (Figure 5B-6 ). Each unitreceives pitot/static data from itsrespective pitot/static system. The unitsdisplay accurate airspeed informationfrom 60 to 400 kts and Mach numbersimultaneously.

A moving pointer indicates maximumairspeed (VMO/MMO). A moveable ref-erence pointer (bug), controlled by aknob on the lower left corner of theinstrument, allows the setting of aspeed reference marker.

On aircraft with an ADC, the pilot’sMach/airspeed indicator is an electricunit that receives data from the ADC.

5B-65B-5


The unit displays airspeed from 60 to420 kts with a pointer and Mach num-ber from 0.30 to 0.99 on a drum dis-play. If the power supply or indicatormalfunctions, a red warning flagappears on the display.

Vertical Speed IndicatorsOn aircraft without an ADC, twoidentical vertical speed indicators(VSIs) indicate aircraft vertical speedfrom 0 to 6,000 feet per minute (FPM)up or down (Figure 5B-7 ). The grad-uations on the instrument indicate 100FPM between 0 and 1,000 FPM and500 FPM above 1,000 FPM.

On aircraft with an ADC, the pilot’svertical speed indicator is an electricunit that receives data from the ADC.The unit displays vertical speed from0 to 6,000 FPM up or down. A red flagappears in the event of power failure ora system malfunction detected by the

ADC monitoring circuits. A moveablereference marker, controlled by knobon the lower right corner of the indi-cator, allows the setting of a climb ordescent speed marker.

Speed, Mach, andAltitude WarningA Mach limit and overspeed warningsystem provides an aural warningwhen airspeed reaches 350 to 370 kts(depending on altitude) or Mach num-ber reaches 0.87 above 25,000 ft. Thesystem receives 28V DC power fromthe B bus.

An altitude alerting system provides avisual and aural signal whenever theaircraft is outside a preselected alti-tude. The ALT ALERT light illumi-nates and a 800 Hz tone sounds fortwo seconds when the aircraft is with-in 1,000 ft of the preselected altitude.

5B-7


Avionics

The light extinguishes and remainsextinguished as long as the aircraft flieswithin 300 ft of the selected altitude. Ifaltitude varies more than 300 ft, thetone sounds and the ALT ALERT lightilluminates.

The system has a built-in test featurethat illuminates the light and displaysa warning flag when the ALT ALERTbutton is pressed.

AdditionalInstrumentationAdditional flight instrumentation notrelated to the air data system includesdigital clocks, a magnetic compass,and a standby gyro-horizon.

Digital ClockA single digital clock provides timeand elapsed time information in a 24-hour format (Figure 5B-8 ). The clock

is capable of displaying elapsed time,flight time, and actual time.

Magnetic CompassA standard liquid-filled compass(Figure 5B-9 ) is on the center wind-shield post. The unit has provisions formaintenance personnel to adjust thecompass to compensate for magneticfields generated by aircraft equipment.A compass correction card on the unitprovides a written record of compassadjustments.

Standby Gyro-HorizonThe standby gyro-horizon (Figure 5B-10) provides roll and pitch attitude datain the event of a system failure. Theunit operates on AC power suppliedby an internal inverter from the A bus.The emergency battery supplies oper-ating current if the aircraft electricalsystem fails.

5B-8 5B-9

5B-10


Avionics

Avionics equipment on the Falcon10/100 aircraft includes communica-tions and navigation equipment.

Depending upon customer preferenceand aircraft modifications, avionicssystems can vary widely. This sectionprovides a brief overview of the mostcommon systems.

CommunicationsCommunications equipment includesvery high frequency (VHF), high fre-quency (HF), interphone and radiotelephone systems, and static dis-charging systems.

VHF CommunicationsVery high frequency (VHF) trans-ceivers provide air-to-air, air-to-ground,and ground-to-ground communicationsin the 116 to 151.975 MHz frequencyrange. Typical systems (118 to 135.975MHz) provide 25 kHz frequency spac-ing with a possible 720 distinct fre-quencies. Extended frequency rangesystems provide 1,440 distinct fre-quencies in the 116 to 151.975 MHzrange.

Each VHF communications systemconsists of a control unit, transceiver,and antenna. The No. 1 and No. 2 VHFtransceivers are in the nosecone. DC

power from the A bus powers the No.1 transceiver; the D bus powers theNo. 2 transceiver. Control units(Figure 5B-11 ) on the center instru-ment panel provide frequency selec-tion and display, volume control, trans-fer between two frequencies, powercontrol (on/off), and system testing.

Each VHF transceiver has its ownantenna. The No. 1 system antenna ison the lower fuselage; the No. 2 systemantenna is on the vertical stabilizer.

HF CommunicationsVery long range, high frequency (HF)communication is provided by a sin-gle transceiver and control system.This system operates in the 2 to 30MHz range with two transmissionmodes: amplitude modulation (AM)and single sideband (SSB). Frequencyspacing of 100 Hz provides a possible288,000 distinct frequencies.

The system consists of:� transceiver� control unit� coupler-amplifier� antenna.

The transceiver in the nosecone oper-ates on 28V DC from the D bus. TheHF control unit (Figure 5B-12 ) on the

5B-11 5B-12

AvionicsEquipment


right console contains controls for sys-tem mode, frequency selection and dis-play, squelching, and volume.

A coupler-amplifier in the baggagecompartment boosts transceiver out-put and balances the antenna load tothe selected frequency. The wire HFantenna runs from a mast on the leftwing to the left engine nacelle.

InterphoneA cabin address and external commu-nications system allows the crew tocommunicate with the passengersthrough speakers and with ground per-sonnel through jacks in the nose gearwheel well and in the baggage com-partment.

When the crew communicates with thepassengers, an attention tone precedessounds in the cabin. The passengeraddress system also includes an inputfor a tape player or other audio source.

Audio Control UnitsAn audio control unit (Figure 5B-13 )for each crewmember on the instru-ment panel allows audio source out-put and microphone input selection.

Each unit has inputs for a handset, oxy-gen mask, and headset microphone andoutputs for a cockpit loudspeaker andheadset. The units use either toggleswitches (Figure 5B-14 ) or push-buttons for input and output selection.

A single potentiometer controls thevolume of the selected audio channels.The volume control knobs on equip-ment control heads control individualreceiver volume levels. Volume levelsof the interphone, sidetone, markerreceiver, and radio compass audio out-puts are controlled by the volume con-trol when they are filtered.

Multiple pushbuttons or toggle switch-es select audio sources from the com-munication and navigation receiversand microphone output to the com-munication transmitters, the publicaddress system, and the interphone sys-tem.

On the pushbutton unit, an interlockingsystem for the VHF 1, VHF 2, VHF3, and HF 1 microphone input switch-es prevents the selection of more thanone input. Depressing the CABIN but-ton overrides the VHF and HF micro-phone inputs.

5B-13


Avionics

Selecting EMG or EMER (emergency)sends all selected audio outputs direct-ly to the headset amplified by thereceivers only. Depressing the SPK(speaker) button sends all selected out-puts to the loudspeaker in the cockpit.

The system automatically mutes allaudio outputs except cabin, aural warn-ings, and sidetone during VHF, HF,and cabin interphone transmissions.

Static DischargingAs an aircraft moves through the air,friction between the surface of the air-craft and air molecules creates staticelectricity. A static charge build up onthe skin of the aircraft has an adverseeffect on communication equipment;communication becomes unintelligible.Likewise, lightning strikes are destruc-tive to the airframe, antennas, andavionics systems.

Static dischargers (Figure 5B-15 ) onthe wings, stabilizers, and pylon trailingedges minimize the effect of static elec-tricity and lightning strikes. The dis-chargers bleed off accumulated staticcharges to the atmosphere. They min-

imize static-induced noise in the com-munication systems.

During the preflight inspection, thesecurity and presence of the discharg-ers should be checked.

The number of static dischargers varieswith owner preference, type of dis-charger used, and modification status.Typical installations employ static dis-chargers on the wing tips, ailerons,pylon trailing edges, rudder, elevators,horizontal stabilizer trailing edge, andtailcone.

NavigationNavigation equipment on the Falcon10/100 includes VHF receivers, auto-matic direction finding (ADF), dis-tance measuring (DME), radio alti-tude, attitude and heading gyros, andlong range navigation equipment.

VHF NavigationVHF navigation receivers provide veryhigh frequency omni-range (VOR),localizer (LOC), glideslope (GS), andmarker beacon navigation informationto the flight crew through various indi-cating equipment.

5B-14

5B-15


Each receiver has 200 VHF frequen-cies from 108.00 to 117.95 MHzspaced at 50 kHz, 40 paired glide-slopefrequencies from 329.15 to 335.00MHz spaced at 150 kHz, and 40 LOCfrequencies from 108.10 to 111.95MHz. Automatic DME channeling isthrough the system control head.Multiple outputs from the receiversdrive the flight director, radio mag-netic indicators (RMIs), autopilot,course deviation indicators, and areanavigation equipment (RNAV). Audiooutput from the receiver is supplied tothe audio control units.

Receiver control, frequency selection,and display are controlled throughheads on the center instrument panel(Figure 5B-16 ).

As part of the VHF navigation receiv-er, a marker beacon receiver providesvisual and aural indications of beaconpassage. The system receives on asingle 75 MHz frequency and provideselectrical outputs to two sets of threeindicating lights on the instrumentpanel. The receiver also provides audiooutput to the audio control units forbeacon passage notification.

The A bus and the pilot’s inverter sup-ply 28V DC and 26V AC, respective-ly, to the No. 1 VHF system. The Cand D buses and pilot’s inverter supply28V DC and 26V AC to the No. 2system.

Instrument LandingSystemAn instrument landing system (ILS)combines outputs from the VHF nav-igation receiver, a UHF glideslopereceiver, and the marker beacon receiv-er to display ILS information on theattitude director indicator and the hor-izontal situation indicator.

The system consists of a glideslopereceiver operating in the 329.15 to335.00 MHz frequency range, the VHFreceiver in LOC mode operating in the111.95 to 118.10 MHz frequencyrange, a glideslope antenna in the nose,and a LOC antenna on each side of thevertical stabilizer.

Automatic DirectionFindingAutomatic direction finding (ADF)systems consist of dual receivers, dualcontrol heads (Figure 5B-17 ), and loopand sense antennas. They operate inthe 190.0 to 1749.5 kHz frequencyrange with 0.5 kHz spacing to provide3,120 distinct frequencies.

Typical ADF systems provide threebasic modes of operation: antenna(ANT), automatic direction finding(ADF), and tone (TONE/BFO). Inantenna mode, the RMI pointer parksand the system provides audio only.ADF mode provides RMI bearing to

5B-16 5B-17


Avionics

the station and station audio. Tone/BFO mode provides a 1000 Hz tonefor identification of morse code stationidentifiers.

The No. 1 ADF system receives 28VDC from the B bus and 26V AC fromthe Y bus. The No. 2 system receivespower from the C bus and the Z bus.

The No. 1 and No. 2 system loopantennas are in the wing-to-fuselagefillets below the wing center sectiontank. The No. 1 sense antenna is in thefin stub root; the No. 2 sense antennais the fuselage belly.

Distance MeasuringDistance measuring equipment (DME)provides slant range distance betweenthe aircraft and a VORTAC facility.The system transmits in the 1025 to1150 MHz range, and receives in the962 to 1213 MHz range. Each DMEchannel is paired with a VHF naviga-tion frequency between 108.00 and117.95 MHz. DME channel selectionis automatic through the VHF naviga-tion receiver.

The DME system provides distance,speed, and time information to the hor-izontal situation indicators (HSIs), andDME displays.

The No. 1 DME receives 28V DCfrom the B bus, and the No. 2 systemreceives power from the C bus.

TransponderTypical transponder systems with4,096 individual Mode A/C code capa-bility provide identification and alti-tude reporting to surveillance radarinstallations.

The system consists of dual trans-ceivers that transmit on 1090 MHz andreceive on 1030 MHz with a singlesource selectable control head on thecenter instrument panel (Figure 5B-18). Altitude information is from thepilot’s encoded altimeter. A trans-mit/receive antenna for the transceiv-er is on the lower front fuselage.

Radio Magnetic IndicatorsTwo radio magnetic indicators (Figure5B-19) display aircraft heading infor-mation on a calibrated servo-drivencompass card. Bearing to either VORor ADF stations is provided by point-ers that are read against the card.

The radio magnetic indicators (RMIs)receive compass information from theHSIs, VOR bearing information fromthe VHF navigation receivers, andADF bearing information from theADF system.

Radio AltimeterA single radio altimeter (Figure 5B-20) system provides precise altitudereadings for approaches and landing.

5B-18 5B-19 5B-20


The system consists of a transceiver,indicator, transmit antenna, and receiveantenna; the antennas are on the aftfuselage belly. The system transmits a4250 to 4350 MHz (4.3 GHz ± 5MHz) signal toward the ground,receives the bounced signal, and com-putes distance from the delay betweentransmission and reception. The sys-tem provides altitude information from-20 to +2,500 ft on an indicator on thepilot’s instrument panel or on theADI/EADI.

Vertical GyrosTwo vertical gyros provide pitch androll data to the attitude director indi-cators, automatic flight control sys-tem, radar antenna pedestal, and yawdamper system. The No. 1 verticalgyro supplies the pilot’s systems, andthe No. 2 gyro supplies the copilot’ssystems and yaw damper.

The system consists of two verticalgyros in the nosecone. The W bus sup-plies 115V AC to the No. 1 gyro, andthe X bus supplies 115V AC to the No.2 gyro.

Directional GyroTwo independent directional gyro sys-tems provide 360° of magnetic head-ing information to the HSIs, RMIs,autopilot, and inertial navigation sys-tems. Each system consists of a gyro,flux valve, compensator, and controller.

The flux valves measure the earth’smagnetic field to act as a synchro trans-mitter for the directional gyros. A com-pensator for each flux valve counter-acts magnetic disturbances affectingthe flux valve.

The compass controllers on the pilot’sand copilot’s consoles allow the set-ting of a heading or course on the HSIsand RMIs.

The No. 1 system receives 28V DCfrom the A bus and 115V AC from theW bus. The No. 2 system receives 28VDC from the C bus and 115V AC fromthe X bus. The directional gyros sup-ply their flux valves with 26V AC.

Avionics Switching PanelA custom manufactured panel on thelower left and right sides of the instru-ment panel provides switching, trans-ferring, and monitoring of instrumentsand navigation equipment. Panel onfig-uration and operation varies on equip-ment installed and customer preference.

A typical panel (Figure 5B-21 ) for thepilot includes:� EMERG ON-OFF. Arms and dis-arms the emergency battery. This sys-tem powers the emergency attitudeindicator, altimeter, VSI, No. 1 COMand NAV radios, No. 1 transponder,and the lights for these instruments.With the switch off and the main bat-teries on, the light is yellow. When theswitch is on, the light is green.� NAV 1-NAV 2. Directs No. 1 NAVor No. 2 NAV information to the pilot’sHSI; selected mode lights green.� VHF NAV-Omega NAV. DirectsVHF NAV or Omega NAV informa-tion to the pilot’s HSI; selected modelights green.� ANG DEV-LIN DEV. Selectsangular or linear deviation mode onthe pilot’s HSI. In angular deviationmode, each dot of course line dis-placement is measured in degrees. In


Avionics

linear deviation mode, each dot ofcourse line displacement is measuredin actual distance.� ADC FAIL. Illuminates if the ADCfails.� COMP 1-COMP 2. Directs direc-tional gyro (DG) information from DGNo. 1 or DG No. 2 to the pilot’s HSI;selected mode lights green.� FS/AA-FS/AS. FS/AA directs fast/slow information to the AOA index-ers and FS/AS directs airspeed infor-mation to the AOA indexers; selectedmode lights green.� VG 1-VG 2. Directs vertical gyro(VG) information from VG No. 1 or

VG No. 2 to the pilot’s ADI; selectedmode lights green.

The copilot’s panel (Figure 5B-21 )contains switching for the angular andlinear deviation displays. It includes:� FPA FAIL. Illuminates if the flightpath advisory system fails (FPA).� ANG DEV-LIN DEV. Selects angu-lar or linear deviation mode on thecopilot’s HSI. In angular deviationmode, each dot of course line dis-placement is measured in degrees. Inlinear deviation mode, each dot ofcourse line displacement is measuredin actual distance.

5B-21


Weather RadarSystemWeather radar systems consist of anantenna, a receiver-transmitter, a dis-play, and system controls (Figure 5B-22). The vertical gyro system providesattitude information to the radar sys-tem to stabilize the antenna. The sys-tem operates by transmitting a highfrequency radio signal (X-Ray band),receiving the reflected signal, and dis-playing the reflected signals as echoeson the display. Controls on and belowthe indicator select system mode, scan-ning range, antenna tilt, and receivergain (sensitivity).

Typical systems provide:� selectable scanning range� ground mapping� weather cell contouring� adjustable antenna tilt and scan� target alerting.

Some radar systems provide a movingmap navigation display with the radarindicator. The navigation displayincludes VORTAC stations, waypoints,magnetic compass heading, stationbearing, course lines, and other help-ful navigation information with weath-er information.

When operating radar on the ground,precautions should be taken to avoidpersonnel injury, equipment damage,or fuel ignition. Avoid operating theradar during refueling or within 300 ftof refueling aircraft. Caution personnelto remain outside of an area within270° and 7 ft forward of the radome.Direct the nose of the aircraft so that a240° sector forward of the aircraft isfree of large obstructions, hangars, andother buildings for a distance of 300ft. Tilt the antenna up 15°.

For more information concerning radarhazards, refer to FAA AdvisoryCircular AC 2068B.

5B-22


Avionics

Autopilot/flight director systems com-bine an autopilot, flight director, yawdamper, indicators, and controls to pro-vide automatic flight control of highperformance aircraft.

AutopilotThe autopilot provides stabilization ofthe aircraft around the pitch and rollaxes. The system also provides:� heading hold and commanded turns� altitude capture and hold� Mach/airspeed hold� vertical speed hold� VOR/LOC course capture and track� Loran/LRN course capture andtrack� ILS beam capture and track.A typical autopilot system consists of:� autopilot computer or shared auto-pilot/flight direcor computer� autopilot amplifier� servos� autopilot controller� controls and indicators.

Autopilot ComputerEach autopilot computer interpretscommands from the vertical gyro,flight director computers, autopilot

controller, magnetic heading referencesystem, and vertical accelerometers tocommand the autopilot amplifier todrive the aileron and elevator servos.

Autopilot AmplifierThe autopilot amplifier provides ampli-fied steering commands from the flightdirector systems, vertical gyros, and di-rectional gyro to the aileron and eleva-tor servos. The servos in turn drive thecontrol surfaces to the desired position.

ServosServos for the elevator and aileronsconsist of a DC torque motor, a tach-ometer-generator, and an electro-magnetic clutch. Signals from the auto-pilot amplifier drive the servo motorto position the control surface throughcables. A feedback signal produced bythe tachometer-generator provides con-trol surface position information to theamplifier. Once the control surfacereaches the commanded position, theamplifier signals the servo to stop. Theelectromagnetic clutch connects theservo motor to its output shaft duringautopilot engagement.

Autopilot ControllerA single autopilot controller on thepedestal (Figure 5B-23 ) contains an

5B-23

Autopilot/Flight DirectorSystem


autopilot engage lever or pushbutton,yaw damper engage lever, pitch controlwheel, turn control knob, and transferand turbulence buttons. On some sys-tems, the pitch and turn controls arespring-loaded to the center position.Engaging the autopilot also engagesthe yaw damper.

Manual control of the autopilot throughthe controller automatically disengagessome functions of the system.Manually changing aircraft pitchthrough the pitch control knob auto-matically disengages all flight direc-tor vertical modes; using the turn con-trol knob also disengages all flightdirector lateral modes.

The AP XFR button selects the flightdirector system supplying the auto-pilot. A mode coupler controls auto-pilot to flight director coupling; push-ing the AP XFR button transfers theautopilot between flight director No.1 or No. 2. In turbulent flight condi-tions, pressing the TURB buttonreduces autopilot commands to the ser-vos to prevent the system from over-reacting to rapid aircraft attitudechanges.

Controls and IndicatorsControls and indicators for the auto-pilot are distributed throughout thecockpit. Pushbuttons on the controlyoke disengage the autopilot and ini-tiate a go-around command (Figure5B-25). Pushing the disengage buttondisengages the autopilot, releases theautopilot engage lever, and illuminatesthe AUTOPILOT light on the instru-ment panel. Pushing the go-around but-ton disengages the autopilot and initi-ates a fixed, nose-up pitch commandfor a missed approach maneuver (14°).

The aileron and elevator trim indicatorson the pedestal allow the crew to ver-ify autopilot operation. Signals fromthe autopilot amplifier and servos drivethe indicators.

Flight DirectorA typical flight director system con-sists of:� flight director computer� mode selector� mode coupler� attitude director indicator� horizontal situation indicator.

SYNC

RADIO

AP DISC

ICS

P-T-T INTERPHONE(OPTIONAL)

GO-AROUND

P-T-T RADIO

PITCH SYNCHRO

AUTOPILOTDISENGAGE

5B-25

5B-24


Avionics

Flight Director ComputerThe flight director computers in thenosecone provide pitch and roll com-mands to the autopilot and the indi-cators. The navigation receivers, airdata computer (if installed), gyros,and accelerometers supply data to theflight director computer. The flightdirector computer in turn providespitch and roll commands to the auto-pilot and indicator command bars.

The mode coupler passes pitch androll commands through the flightdirector to the autopilot, and from theautopilot to the aileron and elevatorservos.

Mode SelectorA single mode selector on the pedestalallows the selection of flight directormodes for the pilot’s and copilot’sflight directors. Pushbuttons allow:� selection of altitude and headinghold� capturing and tracking of a radiobeam� approach path (ILS)� selection and coupling of the auto-pilot or autopilot/flight director com-puters to the autopilot amplifier.

In addition, pushbuttons allow theselection of vertical modes. Theseinclude:

� indicated airspeed

� Mach number hold

� vertical speed hold

� altitude selection.

Pressing a pushbutton selects the modeand illuminates the button to indicatemode activation.

InstrumentationDepending on the system installed, theFalcon 10/100 uses mechanical or elec-tronic flight instrumentation (EFIS).

Attitude Director Indicator

Depending on aircraft equipment,modification status, and customer pref-erence, the attitude director indicator(ADI) is either a mechanical unit oran electronic unit (EADI). The ADI(Figure 5B-24 ) provides a three-dimensional representation of aircraftattitude and flight director generatedroll and pitch commands.

The indicator displays:� attitude and steering commands� localizer and glideslope deviation� rate-of-turn� radio altitude� decision height� speed deviation.

A rate-of-turn sensor detects aircraftlateral turn rate and drives the ADIdisplay.

Aircraft attitude information is dis-played by the relationship of a sta-tionary aircraft symbol with respect toa moving horizon line. The horizonline is servo-driven through the pitchand roll axes. Degree lines above andbelow the horizon line indicate aircraftpitch angles in climb and descent.Aircraft roll angle is displayed by ascale at the top of the indicator.

Two bars flanking the aircraft symbolindicate flight director steering com-mands. The bars are servo-driven forcombined roll and pitch commands.Numerous warning flags within theindicator alert the crew to failures.

The EADI provides the same infor-mation as its mechanical counterpartusing a cathode ray tube (CRT).


Horizontal Situation Indicator

The horizontal situation indicator (HSI)(Figure 5B-26 ) can be either a mech-anical unit or an electronic display. Itdisplays:� aircraft position and heading withrespect to magnetic or true north� selected heading and selected course� DME slant range� deviation from selected VOR, LOC,or other navigation aid� vertical deviation from GS

� TO/FROM indication� bearing/track information.

An aircraft symbol shows aircraft posi-tion and heading in relation to an azi-muth card, lateral deviation bar, andselected heading. The azimuth card dis-plays heading information from a gyro-stabilized magnetic compass. Headingis read on the card beneath the lubberline at the top center of the indicator.

The EHSI provides the same infor-mation as its mechanical counterpartusing a CRT.

5B-26


Avionics

The Electronic Flight InstrumentSystem (EFIS) reproduces informa-tion presented on the ADI, HSI, andweather radar on color CRTs. Itreplaces the standard mechanical ADIand HSI with an electronic attitudedirector indicator (EADI) and an elec-tronic horizontal situation indicator(EHSI) that are high-resolution, black-matrix, shadow-mask color cathoderay tubes (CRT).

Additional information such as weath-er radar, navigation waypoint locations,and moving map displays are alsoavailable for display on the EADI andEHSI.

The weather radar indicator and con-trol panel are replaced by a multi-function display that is capable of dis-playing radar and navigationinformation simultaneously.

The complete EFIS system includes:� two display processor units� one multifunction processor unit� four identical CRTs for the EADIsand EHSIs� two display control panels� one multifunction display unit� one course heading panel� two reversionary switching panels.

Display ProcessorUnitThe display processor unit (DPU)receives data from the navigationequipment, flight director computer,vertical gyros, and directional gyros.The DPU processes this informationfor presentation on the EADI andEHSI. Each DPU controls a set of atti-tude and heading displays.

If a DPU fails, the multifunctionprocessor unit (MPU) replaces thefailed unit. If both DPUs fail, the MPUreplaces both units to display identi-cal data on both pair of EADIs andEHSIs.

MultifunctionProcessor UnitThe multifunction processor unit(MPU) receives data from the weath-er radar system and both pilot’s navi-gation and attitude systems. The MPUprocesses radar information and/ornavigation information for display onthe multifunction display unit.

In the event of DPU failure, the MPUreplaces the failed DPU. If both DPUsfail, the MPU replaces both DPUs.Each pair of displays (EADIs andEHSIs) displays the same information.

EADI and EHSIThe EADI (Figure 5B-27 , followingpage) and EHSI (Figure 5B-28 , fol-lowing page) present displays similarto their mechanical counterparts. Eachpair of units (EADI and EHSI) isdriven by a separate DPU. In the eventof EADI or EHSI failure, EADI is pre-sented on the functional display andthe EHSI is presented on the multi-function display unit. If both displaysfail, a combined image of attitude andheading is displayed on the remainingdisplay.

The EHSI can present information inthree different modes: full compassrose, 60° compass rose, map, and com-binations thereof. Full compass rosemode is similar to a mechanical HSI’sdisplay.

ElectronicFlightInstrumentSystem


Display ControlPanelsA display control panel (DPU) for eachcrewmember contains controls for eachEADI and EHSI pair. Each DPU (Fig-ure 5B-29 ) allows:� selection of display modes on theEHSI� display of navigation informationon the EADI and EHSI� brightness adjustment.

If one of the display control panelsfails, the remaining panel operates bothpairs of displays.

MultifunctionDisplay UnitThe multifunction display unit (MFD)(Figure 5B-30 ) presents informationin either radar, radar and navigation,or navigation only modes. In radar andnavigation mode, navigation informa-tion (waypoints, course, etc.) is pre-

5B-295B-285B-27

5B-315B-30


Avionics

sented with radar information. Modeselection is through pushbuttons onthe unit’s face. In the event of EADIor EHSI failure, the MFD presents thefailed display’s information.

MFD control is through a radar con-trol panel below the display. This panelcontains the normal controls found ona standard weather radar unit.

ReversionarySwitching PanelsTwo reversionary switching panels(RSP) control the operation of eachEFIS display pair (Figure 5B-31 ).Each panel contains switches for:� display transfer

� image transfer� display of attitude and headinginformation from the INS or attitudeand heading system� replacement of a failed DPU by theMPU� system testing.

Course HeadingPanelA single course heading panel (Figure5B-29) for the entire system allowseach crewmember to select a desiredcourse through two knobs. A singleknob allows a heading to be selected;it also allows the display of navigationdata on the EHSIs.


Avionics

PreflightDuring the preflight inspection, removethe covers and verify that the pitottubes and static ports are clear and freeof obstructions. Check the conditionand presence of the 17 or more staticdischargers.

AbnormalProceduresAbnormal procedures for the Falcon10/100 concern the pitot/static andautopilot systems. Refer to Section 3 ofthe AFM for abnormal procedures con-cerning the electronic flight informa-tion system (EFIS).

Pilot’s Pitot/StaticPressureInaccurate airspeed, altitude, or verti-cal airspeed indications on the pilot’sinstruments indicate pitot/static sys-tem malfunction, probe and static portclogging, or air data computer systemfailure. Set the pilot’s pitot/static selec-tor to EMERG to select the nose- conestatic port. If indications are still inac-curate, continue the flight using thecopilot’s instruments.

Copilot’s Pitot/StaticPressureInaccurate indications on the copilot’spitot/static instruments indicate pitot/static system failure. Set the copilot’spitot/static selector to PANEL ONLY.The copilot’s pitot/static selector is adual shutoff valve that acts on the pitotand static ports. PANEL ONLY iso-lates:� cabin pressure differential indicator� Arthur unit actuators for pitch, roll,and rudder control� autopilot altitude sensitive switch� flight recorder (if installed).

As the PANEL ONLY position isolatesthe Arthur units from the pitot/static

system, the Q UNIT light illuminatesand the artificial feel system no longeroperates. Disengage the autopilot andreduce airspeed to 260 kt/Mach 0.76.

Landing WithoutSpeed InformationIf all airspeed indicators fail due topitot/static system or instrument fail-ure, maintain a normal glideslope anda N1 speed 2% above normal approachspeeds.

Abnormal AirspeedIndication at High AltitudeAbnormal or jammed Mach/airspeedindications at high altitudes due tocomplete pitot/static system failure orpitot port/static probe freezing is usu-ally accompanied by the illuminationof the AIR DATA and Q UNIT lights,sounding of the VMO/MMO warninghorn, disengagement of the autopilot,and disagreement between the airspeedindicators and the standby airspeedindicator. If it is certain that the over-speed warning is false, do not changeaircraft airspeed or attitude. Use thestandby altimeter to maintain aircraftaltitude.

Disengage the autopilot and avoid sud-den or large control movements. Pullthe VMO/MMO warning horn circuitbreaker to silence it. Reset the circuitbreaker at intervals and leave it reset ifthe horn stops. Refer to the AFM forclimb, level flight, and descent proce-dures.

Air Data ComputerInoperativeIllumination of the AIR DATA lightindicates failure of the air data com-puter. The pilot’s altimeter, verticalspeed indicator, and Mach/airspeedindicator is inoperative. Use the co-pilot’s instruments and continue theflight. Disengage the autopilot becausethe ADC supplies the autopilot and theflight director system with data.

Preflight andProcedures


Compass System FailureFailure of the compass system is indi-cated by the HDG flag on the RMI andcourse indicator or a disagreementbetween the pilot’s and copilot’s sys-tems. The pilot’s RMI receives com-pass information from the copilot’ssystem and power from the pilot’sinverter. The copilot’s RMI receivesinformation from the pilot’s systemand power from his own inverter.Failure of an inverter is correctable byusing the standby inverter. Failure of agyro renders the system inoperable.

Once a system fails, it is possible torestore heading information by switch-ing the MAG-DG switch to DG. Thisallows the directional gyro to provideheading information to the malfunc-tioning system. Using the SLAVEHDG switch resets the secondary head-ing information.

If switching to the standby inverter orusing the DG switch fails to correctthe problem, the operating gyro andthe standby compass can provide head-ing information.

Pitch Trim FailureDepending on the autopilot installedin the aircraft, procedures for failureof the pitch trim system are similar;the only difference is in the failure indi-cation. Illumination of the ELEV TRor AP TRIM light indicates failure ofthe pitch trim system.

If the pitch trim system fails, proce-dures require holding the control yokefirmly to maintain level flight, disen-

gaging the autopilot, and retrimmingthe aircraft using normal stabilizer con-trol. Do not re-engage the autopilot toprevent a reoccurrence of the problem.

Out-of-Trim ConditionUnusual readings on the trim indica-tors and illumination of the MISTRIMlights indicate an out-of-trim condi-tion. Hold the control yoke firmly tomaintain level flight and disengage theautopilot. Trim the airplane with nor-mal stabilizer control and attempt toreconnect the autopilot. If the problemreoccurs, disengage the autopilot.

EmergencyProcedureA single emergency procedure con-cerns the autopilot system. Proceduresfor the Collins AP-105 and APS-80and Sperry SPZ-500 autopilot systemsare similar with the exception of alti-tude loss between failure recognitionand recovery.

Pitch Servo-ActuatorRunawayA rapid nose-down movement of theaircraft with the autopilot engaged indi-cates a nose-down trim runaway.Procedures require the immediate dis-engagement of the autopilot and con-trol yoke movement to bring the air-craft out of a dive. At cruise altitudesand speeds, demonstrated reaction timebetween recognition and recovery canresult in an approximate altitude loss of400 to 550 ft.


ElectricalSystem

Chapter 5C

This section details the Electrical and Lighting Systems of theFalcon 10/100 aircraft.

The Falcon 10/100 Electrical System includes direct and alter-nating current.

Two engine-driven starter-generators provide direct current(DC) to their respective distribution buses for primary electri-cal power. Two nickel-cadmium (ni-cad) batteries providesupplementary DC for power buffering and engine starting.

Two primary inverters and one standby inverter provide alter-nating current (AC) by converting DC power to 115V, 400 Hzand 26V, 400 Hz AC power.

The Falcon 10/100; lighting system is divided into five majorsubsystems: flight deck, passenger compartment, emergency,rear commpartment, anad exterior lighting.


Electrical Systems

Table ofContents

DC System Schematic . . . . . . . . . . . . . . . . 5C-6

DC System . . . . . . . . . . . . . . . . . . . . . . 5C-7

Components . . . . . . . . . . . . . . . . . . . . . . 5C-7

Batteries . . . . . . . . . . . . . . . . . . . . . . . 5C-7

Cooling and Temperature . . . . . . . . . . . . . 5C-8

Protection . . . . . . . . . . . . . . . . . . . . . 5C-8

Battery Controls . . . . . . . . . . . . . . . . . . 5C-9

Battery Charging . . . . . . . . . . . . . . . . . . 5C-9

Starter/Generators . . . . . . . . . . . . . . . . . 5C-10

Protection . . . . . . . . . . . . . . . . . . . . . 5C-10

Engine Starting . . . . . . . . . . . . . . . . . . 5C-11

Generator Controls . . . . . . . . . . . . . . . . 5C-12

External DC Power . . . . . . . . . . . . . . . . . 5C-13

Distribution and Protection . . . . . . . . . . . . . . 5C-13

Distribution . . . . . . . . . . . . . . . . . . . . . 5C-13

Distribution Buses . . . . . . . . . . . . . . . . . 5C-14

Protection . . . . . . . . . . . . . . . . . . . . . 5C-16

AC System Schematic . . . . . . . . . . . . . . . 5C-22

AC System . . . . . . . . . . . . . . . . . . . . . 5C-23

Inverters . . . . . . . . . . . . . . . . . . . . . . . 5C-23

Distribution . . . . . . . . . . . . . . . . . . . . . 5C-23

AC Controls . . . . . . . . . . . . . . . . . . . . 5C-23

Indication . . . . . . . . . . . . . . . . . . . . . . 5C-24

Protection . . . . . . . . . . . . . . . . . . . . . 5C-24

Preflight and Procedures . . . . . . . . . . . . . 5C-25

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5C-25

Abnormal/Emergency Procedures . . . . . . . . . . 5C-25

Abnormal Procedures . . . . . . . . . . . . . . . 5C-25

Failure of One Generator . . . . . . . . . . . . . 5C-25


Battery Light On . . . . . . . . . . . . . . . . . 5C-26

Battery Overheat . . . . . . . . . . . . . . . . . 5C-26

Failure of One Inverter . . . . . . . . . . . . . . 5C-27

Bus Failure . . . . . . . . . . . . . . . . . . . . 5C-27

Emergency Procedures . . . . . . . . . . . . . . 5C-27

Electrical Fire or Smoke . . . . . . . . . . . . . 5C-27

Complete Electrical Distribution Failure . . . . . 5C-28

Failure of Two Generators . . . . . . . . . . . . 5C-28

Electrical Data Summary . . . . . . . . . . . . . . 5C-31

Electrical System Distribution . . . . . . . . . . . 5C-33

DC Powered Electrical Systems . . . . . . . . . . . 5C-33

AC Powered Electrical Systems . . . . . . . . . . . 5C-35

Lighting Systems . . . . . . . . . . . . . . . . . . 5C-37

Interior Lighting . . . . . . . . . . . . . . . . . . . 5C-37

Cockpit Lighting . . . . . . . . . . . . . . . . . . 5C-37

Ambient . . . . . . . . . . . . . . . . . . . . . . 5C-37

Instrument Integral . . . . . . . . . . . . . . . . 5C-37

Glareshield Lighting . . . . . . . . . . . . . . . 5C-38

Passenger Cabin . . . . . . . . . . . . . . . . . . 5C-38

Entrance and Bar Lighting . . . . . . . . . . . . 5C-38

Cabin Ambient Lighting . . . . . . . . . . . . . . 5C-38

Passenger Call Signs . . . . . . . . . . . . . . 5C-39

Rear Compartment . . . . . . . . . . . . . . . . . 5C-39

Emergency Lighting . . . . . . . . . . . . . . . . 5C-39

Exterior Lighting . . . . . . . . . . . . . . . . . . . 5C-39

Navigation and Strobe . . . . . . . . . . . . . . . 5C-40

Anticollision . . . . . . . . . . . . . . . . . . . . 5C-40


5C-6

Electrical Systems

DC SystemLanding Lights . . . . . . . . . . . . . . . . . . . 5C-40

Taxi and Identification . . . . . . . . . . . . . . . 5C-40

Ice Detection . . . . . . . . . . . . . . . . . . . . 5C-41

Lighting Data Summary . . . . . . . . . . . . . . 5C-43


STANDBYINVERTER

NO. 2COPILOT'SINVERTER

BATT

BUS

START BUS

BATT

BATT

GCU2

BATTERY RELAYSREAR BAY AND PILOT'S DOME LIGHTSEXTINGUISHER EMERGENCY POS. 2IGNITION

EXT. PWR.BATTERYCHARGE

BATTERYCHARGE

LOW TEMPSTARTRELAY

NO. 1START

EXT POWERBATTERY START

RELAY

BATTERY 1RELAY

LOWTEMPSTARTRELAY

NO. 1BATTERYNORMALGROUNDSUPPLY

GEN2

ALT #2 ENGINECOMPUTER

POWER SOURCE

ALT #1 ENGINECOMPUTER

POWER SOURCE

GEN1

LOW TEMP

GEN 1RCR

DIFF PROTRELAY

GCU1

BATTERY 2RELAY

GEN 2RCR DIFF PROT

RELAY

SENSE PINPOWER PIN

A

BUS

B

BUS

D

BUS

C

BUS

TOAUXBUS

SWITCH

STBYHYDPUMP

FREON

EMERA BUSRELAY

LH PITCH SENSOR (STALL 1)NORMAL HORIZONTAL STABILIZERARTHUR QYAW DAMPER CONTROLAIRBRAKESTHRUST REVERSERS

STANDBY HORIZONPILOT'S FD INSTRUMENTS

RH BOOST PUMP CONTROLRH LEVEL SENSORSRH FUEL TRANSFER

LH STARTINGLH IGNITIONLH ENGINE INDICATIONLH ENGINE COMPUTER

PILOT'S STATIC AND PITOT HEADSLEFT STALL VANEWING ANTI-ICING VALVEWINDSHIELD WIPERSPILOT'S WINDSHIELD HEATERCONTROL OF PILOT'S WINDSHIELD HEATERLH ENGINE ANTI-ICING

RAPID DEPRESSURIZING (DUMP)

L/G CONTROLNOSEWHEEL STEERING

LH ENGINE FIRE DETECTIONNORMAL EXTINGUISHINGCONTROL POS. 1

MACH NUMBER LIMITAUDIO WARNINGS

BELLYANTI-COLLISION LIGHTPASSENGER CABIN LIGHTINGPOSITION LIGHTINGSTROBE LIGHTS

ATCVHF 1ICS 1VOR 1GLIDE 1

FLIGHT CONTROLS

FLIGHT INSTRUMENTS

ENGINES

FUEL

ANTI-ICING AIR CONDITIONING

HYDRAULIC CIRCUIT

FIRE DETECTION

WARNING

COMMUNICATION/NAVIGATION

RH PITCH SENSOR (STALL2)EMERGENCY SLAT CONTROL

FLIGHT CONTROLS

RH STARTINGRH IGNITIONRH ENGINE INDICATIONRH ENGINE COMPUTER

ENGINES

TR ARMING REQUIRES #1 BATT SWITCH ON, OR IF SB154 COMPLIED WITH, MAIN BUS POWERED

TOTAL TEMP GAUGEFREON EVAPORATOR BLOWERCONDITIONING VALVE

AIR CONDITIONING

HYDRAULIC SYSTEM INDICATINGSTANDBY ELECTRO PUMP CONTROLL/G POSITION INDICATORS

RH ENGINE ANTI-ICINGSIDE WINDOW PANEL ANTI-ICING

HYDRAULIC CIRCUIT

ANTI-ICING

ICS 2DME 1ADF 1MARKER


LIGHTING

FIN ANTI-COLLISION LIGHTAUDIO INTEGRATING UNIT LIGHTPASSENGER SIGNSCB PANEL(S)RADIO CONTROL BOXES

LIGHTING

MALFUNCTION TESTSMALFUNCTION LIGHTS

ANNUNCIATOR WARNINGLH BOOST PUMP CONTROLLH LEVEL SENSORSLH FUEL TRANSFER

FUEL

RH ENGINE FIRE DETECTION

FIRE DETECTION

FLAP AND AUTO SLAT INDICATIONAILERON TRIM

COPILOT'S ALTIMETERANNUNCIATORS, COPILOTCOPILOT'S FD INSTRUMENTS

CROSSFEED AND INTERCOMFUEL FLOW AND TEMPERATURE

CABIN TEMPERATUREFREON CONTROL

BRAKES (ANTI-SKID)

HP BLEED VALVE CONTROL

RH LANDING LIGHTPILOT AND COPILOT'S SEAT ADJUSTMENT HEATED CONTAINERTOILET CIGARETTE LIGHTER

RADAR 28V DCVOR 2DME 2ADF 2TAPE RECORDERVHF 3

FLIGHT CONTROL FUEL

AIR CONDITIONING

HYDRAULIC CIRCUIT

MISCELLANEOUS


FLIGHT INSTRUMENTS

COMPRESSED AIR

EMER STABILIZER TRIMTRIM INDICATNGFLAP MOTORFLAP CONTROLRUDDER TRIM

AP SERVO CONTROL FOR AILERON AND ELEVATOR

COPILOT'S PITOT AND STATIC HEADSCOPILOT'S WINDSHIELD HEATINGCONTROL OF COPILOT'S WINDSHIELD HEATINGSIDE WINDOW HEAT

TAXIING LIGHTLH LANDING LIGHT

VHF 2GLIDE 2AOA INDICATOR (OPTION)ATC 2HF

FLIGHT CONTROLS ANTI-ICING

LIGHTING


MAIN

BUS

115V AC

26V AC

BATTERY

GEN 1

GEN 2

EXTERNAL POWER

NO.1PILOT'S

INVERTER

AUTOMATIC PILOT

BATT2

BATT1

DC System


Electrical Systems

The Falcon 10/100 DC electrical sys-tem provides and distributes 28.5V DCpower from various sources to busesfor aircraft electrical systems.

ComponentsDirect current is provided by:� two nickel-cadmium (ni-cad) bat-teries� two engine-driven starter/gener-ators� external DC power� various emergency batteries for sub-systems, including emergency light-ing, standby gyro, optional navigationequipment, and optional cockpit power.

The starter/generators are the primarysource of electrical power when theengines are operating. Batteries pro-vide engine starting power, buffer DCpower for the electrical system, andsupply emergency DC power if thegenerators are inoperative. An exter-nal power unit allows a ground powerunit (GPU) to power the entire DCelectrical system when the Main busis connected.

The batteries, starter/generators, andexternal power unit supply power tothe Main bus. The Main bus in turnsupplies two primary (A and B) andtwo auxiliary (C and D) buses. A relay

system allows the left starter/generatorto directly supply the A bus if the MainDC bus fails.

In addition to the Main, A, B, C, andD buses, there is a Start bus suppliedeither by the Main bus or by a GPU aswell as a Battery bus that is fully inde-pendent of the other buses. The bat-teries supply the Battery bus.

A load-shedding switch for the C andD buses allows both buses to be dis-connected from the Main bus.

BatteriesTwo 20 cell, 26V, 23 amp/hour, or two19 cell, 24V, 23 amp/hour nickel-cad-mium batteries are in the aft compart-ment. They are normally connected inparallel to the Main and Battery busesfor power buffering, battery charging,and power supplying during emer-gencies. The batteries can be connectedin series to the Start bus to provideadditional power for cold weatherengine starts.

During normal external power use, thebatteries are disconnected from the DCsystem. Battery charging is also pos-sible with external power.

If the generators fail, the batteries canprovide essential power for up to 20minutes.

DC System

5C-1

Battery LimitationMinimum battery voltage forengine starting:

19 cell . . . . . . . . . . . 22V

20 cell . . . . . . . . . . 23 V


Cooling and Temperature

Air circulating in the rear compartmentnormally cools the batteries. A suctionvent on the fuselage bottom connectsto the battery vent to provide additionalcooling air to the batteries. When theaircraft is on the ground, a blowerdirects air to cool the batteries. A micro-switch on the landing gear controlspower from the batteries or externalpower for the blower.

A temperature probe in each batteryconnects to a dual-reading battery tem-perature indicator on the instrumentpanel (Figure 5C-1 , previous page)that provides battery temperature andoverheat indications.

An amber WARM light on the indica-tor illuminates when either battery tem-perature exceeds 120°F. The red HOTlight on the indicator illuminates whenbattery temperature exceeds 150°F.

Because the indicator scale is limited,a button on the indicator provides ameans to shift the display up when theindicator is off-scale. Pressing theRANGE EXTENDER button shiftsthe display up 50° from the actual tem-perature. Subtracting 50 from the indi-cated reading gives accurate batterytemperature.

An amber HOT BAT light on theFailure Warning Panel illuminateswhen battery temperatures reach 160°F(71°C) for SAFT batteries or 135°F(57°C) for G.E. batteries. If the HOTBAT light illuminates, the batteriesmust be checked for damage.

Protection

Make-and-break relays in the DC sup-ply box protect each battery. Themake-and-break relays normally areclosed when the battery switches areplaced to ON, and thus connect thebatteries to the Main DC bus. Theyalso protect the batteries from exces-sive reverse current and the electricalsystem from low battery voltage (lessthan 16V). If reverse current or lowbattery voltage is detected, the make-and-break relay trips to isolate the bat-tery from the bus.

A reverse current of approximately250 amps causes the make-and-breakrelays to disconnect the affected bat-tery. When the relay trips, the BATlight on the Failure Warning Panel illu-minates, and the battery is discon-nected from the Main bus.

5C-45C-35C-2


Electrical Systems

Battery Controls

The Battery bus always is poweredwhen the batteries are connected.Individual battery switches on theoverhead panel (Figure 5C-2 ) controlthe batteries.

If battery voltage is 18V or greater,placing the battery control switch oncloses the make-and-break relays andconnects the batteries to the Main bus;the BATT light on the Failure WarningPanel then extinguishes.

Pressing the RESET button above andbetween the BAT 1 and BAT 2 switch-es resets the battery make-and-breakrelays if an electrical fault causes oneto open.

Setting the Auxiliary Bus switch toON supplies battery power to the Cand D buses when the Main bus ispowered.

A source selectable voltmeter andammeter on the overhead panel

(Figure 5C-3 ) provide DC source volt-age and current. Placing the voltmeter/ammeter selector switch to BAT dis-plays Main bus voltage and current onthe indicators.

The power selector switch (Figure 5C-4) below the voltmeter/ammeter selec-tor switch controls the connection ofthe batteries for engine starting pur-poses. The three position switch (EXTPOWER/LOW TEMP START/NOR-MAL) selects the DC power sourceused for engine starting. NORMALconnects the batteries in parallel fornormal engine starts, LOW TEMPSTART connects the batteries in series(Figure 5C-5 ) for additional power forengine starts in cold weather, and EXTPOWER selects external power forground operations or engine starts.

Battery Charging

An external power unit can charge thebatteries with the electrical system

A

BUS

B

BUS

C

BUS

D

BUS

BATT1

BATT2

EXTERNAL POWERRECEPTACLE

INVNO. 1

INVNO. 2

STBYINV

AUXBUSSWITCH

EMERA BUSRELAY

24V 48V24V

LOW TEMP START-RIGHT ENGINE

START BUS

MAIN

BUS

BATTBUS

GEN1

GEN2

5C-5


powered or unpowered (Figures 5C-6 and 5C-7, respectively). A chargingunit in the rear compartment controlledby a magnetically held Battery Chargeswitch on the main DC power supplybox allows the batteries to be chargedat voltages between 24V and 29V. Ifthe charging voltage falls below 24Vor exceeds 29V or if external power isdisconnected, the unit trips off-line.

A BATTERY CHARGE light on themain DC power supply box illumi-nates when the batteries are beingcharged by an external power unit.

Starter/GeneratorsTwo engine-driven starter/generatorsprovide DC power for aircraft electri-cal needs. Each engine has a starter/generator on the accessory gearbox.

At normal operating speeds, each gen-erator provides approximately 10.5kilowatts of electrical power regulatedat 28.5V DC. The maximum load ofeach starter/generator should notexceed 350 amps. Generator control

units and reverse current relays regu-late and protect the generators.

Protection

Generator control units (GCUs), make-and-break relays, and relays provideprotection for the generators. A gen-erator control unit (GCU) for each gen-erator provides voltage regulation andprotection. The GCUs in the cabinbehind the right coat compartment:� regulate generator voltage to 28.5VDC� provide generator equalizing� protect against overvoltage� limit maximum generator currentto 350 amps� provide field weakening duringengine start� provide anti-runaway protection� ensure voltage reduction to limitbattery charging current to approx-imately 27V for three minutes afterengine start.

MAKE-AND-BREAKRELAY

20P1

MAKE-AND-BREAKRELAY

20P2

MA

IN B

US12P

GPUCONTACTOR

BATT 1

BATT 2

GPUCONNECTOR

BATTERY CHARGE CONTACTORS

18P2

18P1

5C-6

Generator LimitationMaximum . . . . . . 350 amps


Electrical Systems

Overvoltage protection varies withvoltage; as voltage increases, the cut-off time decreases. At 31V, overvoltagetriggers a generator cut-off in sevenseconds, 45V triggers a cut-off in threeseconds, and a 60V condition causesan instantaneous cut-off. If an over-voltage condition is detected, the fieldshunt is cut, the make-and-break relayopens, the generator is cut off from theelectrical system, a generator warninglight illuminates, and the generatorcontrol switch automatically moves toOFF.

The make-and-break relays connectthe generators to the electrical systemwhen generator voltage exceeds theMain bus voltage by 0.6V. The relaysalso trip a generator off-line if a reversecurrent between the generators exists.If one generator produces a reversecurrent to the other generator of 0.6Vor 25 amps or more, the make-and-break relay trips the generator pro-ducing the lower output off-line. If thisoccurs, the respective GENE light illu-

minates and the isolated generatorswitch remains on.

A differential protection relay for eachgenerator protects against shortsbetween a generator and the Main bus.If a 150 amp short occurs, the relayenergizes and isolates the generator,the respective generator warning lightilluminates, and the generator controlswitch automatically moves to OFF.This failure requires system main-tenance.

Engine Starting

During a battery start, the battery relaysdirect power from the batteries to theMain bus. Pushing a start button closesthe external power/battery start relay todirect power to the Start bus. The startrelay closes and the Start bus powersthe respective starter.

With a GPU connected and providingpower to the Start bus, selecting EXTPOWER on the power selector switchcloses the external power/battery start

MAKE-AND-BREAKRELAY

20P1

MAKE-AND-BREAKRELAY

20P2

MA

IN B

US

12P

GPUCONTACTOR

BATT 1

BATT 2

GPUCONNECTOR

BATTERY CHARGE CONTACTORS

18P2

18P1

5C-7


relay closes and provides power to theMain bus; the battery and generatorrelays open. Pushing a start buttoncloses the start relay and directs powerto the respective starter.

Selecting LOW TEMP START con-nects the batteries in series to provideapproximately 48V DC to the Start busand the selected starter. The No. 2battery provides 24V DC to the Mainbus and the No. 1 battery disconnectsfrom the Main bus; the No. 1 batterysupplies only the Start bus. Duringlow temperature engine starts, startonly one engine and perform a gen-erator assisted start for the secondengine.

Generator Controls

Controls and volt/amp meters for thegenerators are on the overhead panel.Two generator warning lights are onthe Failure Warning Panel.

Separate generator control switches(Figure 5C-8 ) operate the No. 1 andNo. 2 generators. Placing a controlswitch on with the engines operatingenergizes the generator make-and-break relay, connects the generator to

the Main bus, and extinguishes therespective GENE light.

The generator control switches arenormally in the ON position. Over-voltage and differential current (short)energize the magnet and allow theswitch to move to the OFF position.Normally the switches are kept in theON position for engine starting andgenerator excitation. The OFF posi-tion is only used during electricalemergencies, fire, or complete sys-tem failure.

A guarded Emergency Generatorswitch (Figure 5C-9 ) allows the left(No. 1) generator to bypass the Mainbus and supply the A bus directly.

A source selectable voltmeter andammeter on the overhead panel(Figure 5C-10 ) provides generator out-put indications. Placing the switch ineither GEN 1 or GEN 2 displays therespective generator’s output voltageand amperage.

The GENE 1/GENE 2 lights illu-minate when the respective genera-tor is not connected. The lights alsoilluminate when external power is inuse.

5C-95C-8


Electrical Systems

External DC PowerAn external power receptacle on theright side of the aft fuselage allows theconnection of a GPU. External powerprovides power for:� aircraft electrical system� battery charging� engine starting.

The power selector switch on the over-head panel (Figure 5C-11 ) is the solecontrol for external power. Placing theswitch in EXT POWER suppliespower from the receptacle through theStart bus to the Main bus, de-energizesthe battery make-and-break relays, andilluminates the BATT light.

With the engines operating and exter-nal power connected, the generatormake-and-break switches do not close;the GENE 1/GENE 2 lights illuminate,and it is impossible to connect any ofthe generators regardless of powerselector switch position.

External power voltage and amper-age is read on the voltmeter andammeter on the overhead panel byselecting the BAT position on theselector switch.

Distribution andProtectionDirect current from the batteries, gen-erators, or the external power systemprovide power to operate the majorityof the aircraft systems. Power fromthese sources is supplied throughoutthe aircraft through distribution buses.

DC distribution buses in the Falcon10/100 aircraft are either non-loadshedding or load-shedding.

DistributionThe batteries and generators supply28.5V DC power through circuitbreakers and protection devices direct-ly to the Main bus. External powerfrom a GPU supplies the Start bus thatin turn supplies the Main bus. TheMain bus in turn supplies the A and Bbuses directly, and the C and D busesindirectly through the Auxiliary Buscontrol switch.

The Main, A, B, C, and D buses arein the main DC power supply box inthe rear compartment. The powersupply box is an insulated fiberglass-reinforced plastic box that also con-tains protective devices for the buses.

5C-115C-10

External Power LimitationGPU . . . . 1,000 AMPS, 28 V


A Battery bus in the battery box, inde-pendent of the other buses, connectsthe batteries together. The batteriesconnect to the Battery buses throughdiodes, and the batteries connect to theMain bus through separate make-and-break relays for each battery.

Distribution BusesNormally, electrical power from theMain bus powers the A, B, C, and Dbuses through four 130 amp currentlimiters. The A and B buses supplypower to the essential aircraft electri-cal loads (Figure 5C-12 ). The C and Dbuses carry non-essential loads andcan be disconnected (load-shed) fromthe Main bus (Figure 5C-13 ). In theevent of Main bus failure, the left gen-erator can directly power the A busthrough the emergency generatorswitch (Figure 5C-14 , page 5C-17).

The Main bus powers:� standby hydraulic pump� Freon air conditioning compressor,right blower, and condenser blower� A, B, C, and D buses.

The A bus supplies:� left engine computer, starting, firedetection, anti-icing, and indication� normal fire extinguishing systemposition No. 1� left boost pump control, fuel level,and fuel transfer� pilot’s essential communication andnavigation radios� Failure Warning Panel lights

� No. 1 inverter� landing gear control� steering.

The A bus also supplies parts of theflight control, flight instrument, anti-icing, air conditioning, and lightingsystems.

The B bus supplies:� right engine computer, starting, firedetection, anti-icing, and indication� Mach number and audio warningsystems� right boost pump, fuel level, andfuel transfer� hydraulic system indicating andstandby electric pump control� standby inverter� landing gear position indicators.

The B bus also supplies parts of theflight control, air conditioning, lighting,communications equipment, and thestandby inverter systems.

The C and D buses supply:� No. 2 inverter� brake anti-skid system� flaps� Freon air conditioning control� HP bleed air.

The C and D buses, controlled by theAuxiliary Bus switch, also power thenon-essential electrical loads on theaircraft. This includes No. 2 commu-nication and navigation equipment andthe copilot’s instruments.

NOTE: SB 77-003 (0074);N1 and N2 Tachometers –Double Voltage Supply withBattery Start on Low Temp Start(S/N 001 to 052).

NORMAL OPERATION

LOAD SHED OPERATION

START BUS

MAIN

BUS

B

BUS

C

BUS

D

BUS

BATTBUS

GEN1

GEN2

BATT1

BATT2


INVNO. 1

INVNO. 2

STBYINV

AUXBUSSWITCH

EMERA BUSRELAY

START BUS

MAIN

BUS

A

BUS

B

BUS

D

BUS

C

BUS

GEN2

BATT1

BATT2


INVNO. 1

INVNO. 2

STBYINV

AUXBUSSWITCH

EMER A BUSRELAY

A

BUS

BATTBUS

GEN1


Electrical Systems

5C-13

5C-12


The Start bus supplies power neces-sary for engine starting. The bus ispowered by:� batteries in parallel during a normalbattery start� batteries in series during a lowtemperature start� ground power receptacle during aGPU start� batteries in parallel with one gen-erator during a generator-assisted start.

The Battery bus, independent of theother buses, receives power directlyfrom the batteries, the generators, or aGPU (Figure 5C-15 ). Normally, bothbatteries are connected in parallel tothe bus. The bus powers:� battery relays� rear compartment and pilot’s domelights� fire extinguisher backup power con-trol position No. 2

� engine ignition alternate source� N1 and N2 indicators (SB 77-003).

ProtectionCircuit breakers, relays, and fuses pro-tect the aircraft electrical system fromoverloads, shorts, and other electricalmalfunctions. The circuit breaker (CB)panels are on either side of the cockpit.

The CB lettering is color coded: whitelabeled CBs protect circuits on the Abus, yellow labeled CBs protect cir-cuits on the B bus, and red labeled CBsprotect circuits on the C and D buses.

Additional CBs and fuses in the bat-tery boxes (Figure 5C-16 , page 5C-18, and Table 5C-A , page 5C-19) andmain DC supply (Figure 5C-17 , andTable 5C-B , page 5C-20) providebasic protection for the DC electricalsystem at the source.

GPU OPERATION

START BUS

MAIN

BUS

B

BUS

D

BUS

C

BUS

BATTBUS

GEN1

GEN2

BATT1

BATT2


INVNO. 1

INVNO. 2

STBYINV

AUXBUSSWITCH

EMERA BUSRELAY

A

BUS

B

BUS

D

BUS

BATT1

BATT2


INVNO. 1

INVNO. 2

STBYINV

AUXBUSSWITCH

EMERA BUS RELAY

A BUS ONLY OPERATIONA

BUS

MAIN

BUS

C

BUS

BATT

BUS

START BUS

GEN2

GEN1


Electrical Systems

5C-15

5C-14


14J114J2

59P54P62P

66P67P81P

68P73P74P

78P

64P18K

3J1

3J2

5K1

5K2

4M5J

56P

35P

2M

18P2

18P1

7K 28P

4J 24K 13P

30P

12K

22J1 22J2 42P 45P45P2 53P

46P1 46P2 61P 63P 65P

41P1 41P2

43P

23K1

22K1

23K2

22K2

69P

77P

HB

85E

79P

36P

31P

32P

55P1

55P2

21J1

21J2

22W

101L

37P

9P1

38P

25P

11K 6K1 6K2

5C-16

NOTE: See Table 5C-A(opposite page) for battery boxcontents.


Electrical Systems

Table 5C-A; Battery Box Contents

Breakers 9P1 BATTERY 1 CONTROL 37P BATTERIES COUPLING9P2 BATTERY 2 CONTROL 38P BATTERY 1 CONTROL21J1 NO. 1 ENGINE IGNITION 55P1 NO. 1 GENERATOR VOLTMETER21J2 NO. 2 ENGINE IGNITION 55P2 NO. 2 GENERATOR VOLTMETER22W EXTINGUISHERS CONTROL 79P BATTERY (2 P2) STARTING25P BATTERIES VOLTMETER 85E TACHOMETERS SUPPLY36P BATTERIES CHARGING 101L LIGHTING

Diodes

Capacitors

4M 42P 54P 67P14J1 43P 59P 68P14J2 45P1 61P 73P22J1 45P2 62P 74P22J2 46P1 63P 81P41P1 46P2 65P41P2 53P 66P

23K1 STARTING RELAY SUPPLY23K2 STARTING RELAY SUPPLY

Fuses 35P BATTERIES CHARGING56P BATTERY 1 CHARGING

Relays 2M SUPPLY – EMERGENCY 11K BATTERIES IN SERIES STARTINGHYDRAULIC PUMP 12K BATTERIES IN SERIES STARTING

3J1 IGNITION LEFT ENGINE 13P GROUND POWER UNIT CONNECTION3J2 IGNITION RIGHT ENGINE 18K CROSS START RELAY4J AUTOMATIC IGNITION 18P1 NO. 1 BATTERY CHARGING5J TIME RELAY FOR 18P2 NO. 2 BATTERY CHARGING

AUTOMATIC IGNITION 24K 10 K RELAY CONTROL5K1 STARTING LEFT ENGINE 28P BATTERY 1 RELAY5K2 STARTING RIGHT ENGINE 30P BATTERIES CHARGING6K1 STARTING LEFT ENGINE 64P GROUND POWER UNIT CONNECTION6K2 STARTING RIGHT ENGINE 69P GROUND POWER UNIT CUT OUT7K BATTERIES IN SERIES COUPLING 77P BATTERY (2 P2) STARTING

GENERATORS VOLTAGE10K REDUCTION TIME RELAY

Resistors 22K1 STARTING RELAY SUPPLY22K2 STARTING RELAY SUPPLY78P BATTERY BUS SUPPLY


8P1 8P2

12P

51P

16P

20P2 20P1 34P1

BATBAT

BAT

BUS52P1

BAT

BUS52P2

34P2

GEN GEN

GEN GEN

1M

47P

50P

49P

48P

131H

DISP

DISP

MAIN BUS5C-17

Table 5C-B; Battery Box Contents

Relays 1M 100 AMP HYDRAULIC STANDBY PUMP RELAY12P GROUND POWER UNIT RELAY16P C BUS SUPPLY RELAY20P1 BATTERY RELAYS20P2 BATTERY RELAYS8P1 NO. 1 GENERATOR EXCITATION8P2 NO. 2 GENERATOR EXCITATION51P D BUS SUPPLY RELAY131H 200 AMP FREON PUMP RELAY

Switches

CurrentLimiters

34P1 GEN TO MAIN BUS SWITCH34P2 GEN TO MAIN BUS SWITCH52P1 DIFFERENTIAL PROTECTION SWITCH52P2 DIFFERENTIAL PROTECTION SWITCH

47P 135 AMP CURRENT LIMITERS (A BUS)48P 135 AMP CURRENT LIMITERS (B BUS)49P 135 AMP CURRENT LIMITERS (C BUS)50P 135 AMP CURRENT LIMITERS (D BUS)DISP 135 AMP CURRENT LIMITERSDISP 135 AMP CURRENT LIMITERS


Electrical Systems

A BUS

A BUS

W

BUS

X

BUS

NO.1PILOT'S

INVERTER

STANDBYINVERTER

NO. 2COPILOT'SINVERTER

LIGHTING

FLIGHTINSTRUMENTS

COPILOT'S INSTRUMENT LIGHTING

COPILOT'S HEADING CHANNELRADIO ALTIMETERRADARCOPILOT'S ATTITUDE FLIGHT DIRECTORCOPILOT'S FLIGHT DIRECTOR INSTRUMENTSDME COUNTERPUBLIC ADDRESS AMPLIFIER

PILOT'S ENGINE CONTROLS ANDINTEGRAL INSTRUMENTLIGHTINGCOCKPIT GLARESHIELD LIGHTINGPEDESTAL LIGHTING

PILOT'S HEADING CHANNELNORMAL CABIN PRESSURE CONTROLPILOT'S FLIGHT DIRECTION INSTRUMENTSPILOT'S ATTITUDE FLIGHT DIRECTORTURN AND BANKYAW COMPUTER

FLIGHTINSTRUMENTS

LIGHTING

Y

BUS

POWER SUPPLYPILOT'S NAVIGATION INSTRUMENTS

#1

#2

STANDBY INVERTER SWITCH

Z

BUS

POWER SUPPLYCOPILOT'S NAVIGATION INSTRUMENTS

NO. 2 ACFAIL LT

NO. 1 ACFAIL LT

A

BUS

B

BUS

C

BUS

AUX BUS SWITCH


AC System


Electrical Systems

The Falcon 10/100 AC electrical sys-tem provides and distributes 115V and26V AC to buses to power various air-craft systems.

Alternating current is provided by:� two primary inverters� one standby inverter.

Two inverters supply the pilot’s andcopilot’s separate AC electrical sys-tems. A third standby inverter providesan alternate power source should oneof the inverters fail. All three invert-ers are in the nose compartment.

InvertersThe single-phase 750 VA static invert-ers convert 28V DC into 115V and26V, 400 Hz AC power for the pilot’s(No. 1) standby, and copilot’s (No. 2)systems. The batteries, external power,or the starter/generators provide 28VDC to the DC electrical system topower the three inverters. The pilot’sinverter receives power from the Abus, the copilot’s inverter receivespower from the C bus, and the standbyinverter receives power from the B bus.

DistributionThe inverters supply 115V, 400 Hz,and 26V AC power to their respectivedistribution buses. Table 5C-C liststhe inverters, their power output, andthe buses they power.

The standby inverter supplies eitherthe pilot’s system or the copilot’s sys-tem through a three-position transferstandby inverter transfer switch.

AC ControlsThree switches on the overhead panelcontrol the inverters (Figure 5C-18 ,following page). The left inverter switchcontrols the No. 1 inverter that receivespower from the A bus. The right invert-er switch controls the No. 2 inverterthat receives power from the load-shedC bus. The three-position center inverterswitch (Figure 5C-18 , following page)controls the buses that the standbyinverter powers. The standby inverterreceives power from the B bus.

In the center position, the standbyinverter control switch isolates theinverter from the AC electrical sys-tem. In NO. 1 AC BUS, the switch iso-lates the pilot’s inverter from the ACelectrical system and the standbyinverter powers the W and Y buses. InNO. 2 AC BUS, the switch isolates thecopilot’s inverter and powers the Xand Z buses.

Because the copilot’s inverter receivesDC power from the C bus, load shed-ding of the C bus results in the loss ofthe copilot’s inverter and prevents theuse of the standby inverter. The stand-by inverter cannot provide AC powerto the copilot’s AC buses if the auxil-iary bus switch is off.

AC System

Table 5C-C; Inverter Power Supply Distribution

Inverter Power Bus or System

Pilot’s 115V, 400 Hz AC26V, 400 Hz AC

W busY bus

Copilot’s 115V, 400 Hz AC26V, 400 Hz AC

X busZ bus

Standby 115V, 400 Hz AC26V, 400 Hz AC

Pilot’s or copilot’s system


IndicationA single, source-selectable voltmeter(Figure 5C-18 ) on the overhead paneldisplays inverter voltage from 0 to 125volts. Source selection is through thevoltmeter supply switch below thevoltmeter. The three position switch(NO. 1/STD BY/NO. 2) connects thevoltmeter to a 115V AC output of therespective inverter.

Two warning lights on the overheadpanel monitor the pilot’s and copilot’sinverters. Illumination of an AC FAILlight indicates the failure of the pilot’s

or copilot’s inverter to power the 115VAC bus. Because DC power suppliesthe lights, the inverter failure relayscontrol the illumination of the lights.The inverter failure relays controlledby the inverters are de-energized when-ever the 115V AC output of an invert-er fails or fails to reach the relay.

ProtectionCircuit breakers protect the AC elec-trical system. The pilot’s, copilot’s,and standby inverters employ circuitbreakers for DC input and AC output.

5C-18


Electrical Systems

PreflightDuring the preflight inspection, verifythat the No. 1 and No. 2 inverter con-trol switches are in the off position.Verify that the standby inverter switchis in the center (off) position. Set theDC voltmeter/ammeter selector toBAT, and the inverter voltmeter selec-tor to STD BY to check the output ofthe batteries and standby inverter.Verify that the generator and auxiliarybus switches are on because an enginestart cannot be made with the gener-ators off. Perform a battery check toverify that battery voltages areadequate for a normal battery start(23V minimum). When the initialcockpit inspection is finished, turn thebatteries off.

Abnormal/EmergencyProceduresThe following section provides a briefdiscussion of what happens within theelectrical system during abnormal andemergency procedures.

For a complete list of specific proce-dural steps for emergency and abnor-mal procedures for the Falcon 10/100electrical system, please refer to theAFM or the SimuFlite OperatingHandbook.

Illumination of the GENE 1/GENE 2lights on the Failure Warning Panelindicate that the generator is not con-nected to the Main bus. A reverse cur-rent, overvoltage, short, or mechani-cal failure can cause a generator to gooff-line. Depending on the condition,the generator control switch may ormay not move to the off position.

A reverse current flowing from onegenerator to the other de-energizes themake-and-break relay of the lower out-

put generator and removes the gener-ator from the electrical system; thegenerator switch remains on.

An overvoltage condition de-energizesthe make-and-break relay, removes thegenerator from the electrical system,and moves the generator switch to off.The voltmeter and ammeter indicatezero for the affected generator. Resetthe generator by moving the switch toon.

A short between the generator and theMain bus opens the differential pro-tection relay and removes the genera-tor from the electrical system; the gen-erator switch trips off. The voltmeterand ammeter indicate zero for theaffected generator. The generator can-not be reset as maintenance is requiredto correct the short.

A mechanical failure, either internalor external, directly affects the outputof the generator. The generator switchremains on and no voltage or amperageindicates on the meters. Maintenanceis required to return the aircraft toservice.

Abnormal ProceduresAbnormal procedures provide stepsfor dealing with single battery, gener-ator, or inverter failures. They alsoinclude procedures for overheated bat-teries.

Failure of One Generator

Illumination of either GENE light indi-cates generator failure due to mechan-ical failure, overvoltage, reverse cur-rent, or electrical short between thegenerator and the Main bus. A shortor overvoltage automatically movesthe generator control switch to off; amechanical failure or reverse currentrequires moving the switch to off. Theassociated voltmeter and ammeter indi-cations are zero.



Attempt a maximum of two resets onthe affected generator; if the genera-tor fails to reset, remove non-essentialequipment from the electrical systemto reduce operating generator load to300 amps.

A list of reasons why the GENE lightilluminates and procedures to correctthem is provided in Table 5C-D .

Battery Light On

Illumination of the BATT light on theFailure Warning Panel indicates that abattery reverse current relay is open.On the ground, verify that the powerselector switch is not in EXT POWERor LOW TEMP START; either posi-tion illuminates the BATT light.

In the air, attempt a maximum of tworesets of the battery system using the

RESET button. If the light remainsilluminated, it is difficult to identifythe failed battery.

Battery Overheat

Illumination of the HOT BAT lightindicates battery temperatures are160°F (71°C) for SAFT batteries or135°F (57°C) for G.E. batteries. Athermal runaway of the batteries thatcan cause aircraft structural damage ispossible.

Turn both battery switches off andallow the batteries to cool. If the air-craft has a battery temperature moni-tor it is possible to identify the over-heating battery; turn the unaffectedbattery on.

If the HOT BAT light remains illumi-nated, land as soon as possible.

Table 5C-D; Generator Light ON, Check Generator Switch Position

Generator Switch ONCheck Volts and Amps

Generator Switch OFF

Differential Trip Overvoltage Trip Reverse Current Mechanical Failure

Attempt reset(2 maximum).

Non-resetable.

If required, reduce load onoperating generator to 300amps maximum.

Attempt reset(2 maximum).

Check volts and amps.

Overvoltage generatorswitch OFF.


Check opposite generatorvolts and amps.

Abnormal generatorswitch OFF.

Normal generator AUTOreset.


Generator switch OFF.

Cause Cause Cause Cause

� 150 amp differentialbetween the generatoroutput and Main bus.

� 31 to 33V for 7 seconds

� 65V spike

� 0.6V difference in generator output andMain bus voltage

� 25 amp reverse currentflow

� sheared generator shaft

� G.C.U.

� broken wire

� other mechanical failure


Electrical Systems

Failure of One Inverter

Illumination of either AC FAIL light,zero voltage readings, or the appear-ance of warning flags on flight instru-ments indicates inverter failure. Thestandby inverter can power either thepilot’s or copilot’s AC electrical buses.Move the standby inverter switchtoward the failed inverter side. If theAC FAIL light extinguishes, the prob-lem is a failed inverter.

If the light remains illuminated, theinverter is not at fault, and the prob-lem exists elsewhere. Move the stand-by inverter switch to the center posi-tion, and reset the controlling CB ifnecessary. Turn the failed inverter off.Continue the flight using instrumentspowered by the other system.

Bus Failure

Failure of the A, B, C, or D bus canresult in the loss of flight critical sys-tems. Monitoring of certain systemsassists in the identification of a busfailure. Please refer to Table 5C-E ,page 5C-29, for lists of the key itemsto verify a bus failure and the affectedsystems critical to flight.

If only the A bus fails due to a short, orother electrical malfunction, key flightsystems are rendered inoperative.Procedures require that the copilot flythe aircraft. Because the thrust revers-er system is inoperative, airport fieldlength and braking conditions are crit-ical landing considerations. Loss ofthe left engine and wing anti-icing sys-tem prohibits flight into known icingconditions. Land as soon as possible.

Please refer to the AFM or the Simu-Flite Operating Handbook for a com-plete list of systems affected and land-ing considerations with bus failures.

Emergency ProceduresEmergency procedures for the Falcon10/100 electrical system concern elec-trical fire or smoke, complete failureof the DC and AC electrical distribu-tion systems, and failure of both gen-erators.

Electrical Fire or Smoke

If the source of smoke or electrical fireis known, isolate the fault by pullingcircuit breakers or removing electricalsources.

If the source is unknown, proceduresrequire the immediate donning of crewoxygen masks and verifying commu-nication between the pilot and copilot.Procedures then require the isolationof the C and D buses through the aux-iliary bus control switch and smokeevacuation by setting the WSHLD andFLOOR valves to full HOT. Use of thewindshield and floor heating controlsprovides increased air flow to clear thecockpit of smoke.

Isolating the C and D buses renders allred labeled (load-shed) CB’s unusable.If smoke decreases or stops, try tolocate the source of the fire and iso-late the circuit. If time permits, pull allred labeled CBs and reset one at a time.

Land the aircraft as soon as possible,refer to the Aux bus off landing con-siderations and proceed to a suitableairport. Aux bus off landing consider-ations include:� no flap landing� no anti-skid� 8,000 ft minimum runway.

If the fire persists, perform an emer-gency descent and determine that oxy-gen use by the passengers does notincrease the fire hazard.


Override the passenger oxygen con-trol to provide oxygen to the passen-gers. Once the passengers are usingoxygen, begin a complete shutdownof the electrical system by turning thegenerators and batteries off. Provideelectrical power to the essential air-craft systems by turning the Emer-gency Generator switch ON. Thisbypasses the generator make-and-breakrelay and provides left generator out-put directly to the A bus.

If the fire decreases or stops, locate thesource of the by pulling the affectedsystem’s CB; reset the other CBs. Landas soon as possible. Refer to the AFMor the SimuFlite Operating Handbookfor landing considerations.

If all efforts fail to reduce the smoke orfire, pull and reset the white labeled(A bus) circuit breakers to try to isolatethe electrical fault. Fight the fire andland as soon as possible. Refer to theAFM or the SimuFlite Operating Hand-book for landing considerations.

Complete ElectricalDistribution Failure

Indications of total electrical failureare:� total loss of lighting� appearance of warning flags oninstruments� activation of emergency lightingsystem.

Procedures for a complete distributionsystem failure require turning off the

generators and batteries and turningon the Emergency Generator switch.Turning on the Emergency Generatorswitch bypasses the generator make-and-break relay and directly powersthe A bus from the left generator. TheA bus powers the minimum equipmentrequired for IFR flight; this includesthe pilot’s flight instruments, the leftengine indicators, the No. 1 VHF andVOR radios and all equipment pro-tected by white labeled circuit break-ers (A bus).

Check the condition of the batteries byturning the cockpit dome lights on. Ifthe dome lights increase in brightness,the batteries are available. Land assoon as possible. Refer to the AFM orthe SimuFlite Operating Handbook forA bus only landing considerations.

Failure of Two Generators

Illumination of the GENE 1 andGENE 2 lights indicate failure of bothgenerators. Remove non-essentialloads by setting the auxiliary busswitch off. Try resetting each genera-tor switch; a maximum of two attemptsis permissible for each generator.

If both lights remain illuminated,remove as many electrical loads fromthe system by pulling circuit breakersor shutting off equipment. The batteriesare capable of providing power for upto 20 minutes; land as soon as possible.Refer to the AFM or the SimuFliteOperating Handbook for landing con-siderations with the Aux bus off.


Electrical Systems

Bus Failure Indication � Flight Critical Systems Affected �

A Bus

� No. 1 inverter fails with no red light� Annunciation (warning) lights fail� Overhead lights fail

Left booster pump . . . . . . . . . . . . . . . . . . . . . . . . . INOPLeft fuel gage . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPLeft fuel transfer . . . . . . . . . . . . . . . . . . . . JP FAILS OPENFuel level sensor . . . . . . . . . . . . . . . . . . . . . . . . . . INOPLeft fire detection . . . . . . . . . . . . . . . . . . . . . . . . . . INOPNo. 1 fire bottle . . . . . . . . . . . . . . . . . INOP, No. 2 AVAILABLELeft engine indication . . . . . . . . . . . . . . . . . . . . . . . . INOPWing anti-ice . . . . . . . . . . . . . . . . . . . . . . . . . . FAIL OFFPilot pitot/static heat . . . . . . . . . . . . . . . . . . . . . . . . INOPPilot windshield heat . . . . . . . . . . . . . . . . . . . . . . . . INOP

B Bus

� Right fuel P-2 fail� Right fuel gage fail� Landing gear annunciator fail� Right engine anti-ice – light on

Copilot audio panel . . . . . . . . . . . . . . . . . . . . . . . . . INOPNo. 1 ADF/DME/Markers . . . . . . . . . . . . . . . . . . . . . . INOPRight engine indications . . . . . . . . . . . . . . . . . . . . . . INOPRight fuel transfer . . . . . . . . . . . . . . . . . . . . JP FAILS OPENRight fuel level sense . . . . . . . . . . . . . . . . . . . . . . . . INOPRight engine fire detection . . . . . . . . . . . . . . . . . . . . . INOPA/C temperature control . . . . . . . . . . . . . . . . LAST POSITIONHydraulic STBY PUMP control . . . . . . . . . . . . . . . . . . . INOP

C Bus

� No. 2 inverter fails� Fuel flow gages fail� No. 2 F/D fails� No. 2 VOR fails

VOR 2/DME2/VHF 2 . . . . . . . . . . . . . . . . . . . . . . . . INOPRadar/FD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPCrossfeed valve . . . . . . . . . . . . . . . . . . . . LAST POSITIONIntercom valve . . . . . . . . . . . . . . . . . . . . . . FAILS CLOSEDFreon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPAnti-skid . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPHP valve . . . . . . . . . . . . . . . . . . . . . . . . . FAIL CLOSEDSeat adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . INOPAileron trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOP

D Bus

� No. 2 pitot heat fails� Trim indicators full deflection (3)

COM 2/GS 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPEmergency stabilizer trim . . . . . . . . . . . . . . . . . . . . . . INOPFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPRudder trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INOPCopilot pitot/static heat . . . . . . . . . . . . . . . . . . . . . . . INOPCopilot windshield heat . . . . . . . . . . . . . . . . . . . . . . . INOP

Table 5C-E; Bus Failures and Indications


Electrical Systems

DataSummary

Electrical SystemDC and AC Electrical Systems

Power Source Batteries (2)24/26V DC, 23 amp hour (20-cell)22/24V DC, 23 amp hour (19-cell)

Engine generators (2)28.5V DC, 350A, 10.5 kW

Inverters (3) – 750VA115V AC, 400Hz26V AC, 400 Hz

Ground power28.5V, 1,000A maximum

Distribution DC busesBatteryStartMainA,B,C,D

AC busesW bus (pilot) – 115V AC, 400 HzY bus (pilot) – 26V AC, 400 HzX bus (copilot) – 115V AC, 400 HzZ bus (copilot) – 26V AC, 400 Hz

Control DC system switchesPower selectorGenerator controlBattery controlEmergency powerAuxiliary bus (load shed)Battery reset

AC system switchesInverter controlStandby inverter

Monitor DC systemVoltmeter/ammeterBattery temperature

AnnunciatorsGENE. 1/2BATTHOT BAT

AC SYSTEMVoltmeterAC FAIL NO. 1/2 annunciators


Current limiters

Relays

Generator control units


Electrical Systems

ElectricalSystemDistribution

DC Powered Electrical SystemsThe following are partial listings of systems powered by the DC andAC electrical systems.

A Bus 28V DC Aircraft configuration/failure warning/etc

(non-load shedding) Aileron/elevator and rudder Arthur jacks

Airbrake after SB 27-015

Anti-collision fin light

Audio selector/CB panel/radio control It

Cabin pressurization dump valve

Engine synchronization

Flight recorder

Landing gear control

Left angle-of-attack sensor

Left booster pump control

Left engine anti-icing

Left engine computer

Left engine fire detection

Left engine indication

Left engine starting

Left fuel tank indication and transfer

Left pitot,static, and AOA anti-icing

Normal fire extinguisher control

Normal stabilizer control circuit

Passenger warning lights

Pilot’s interphone supply circuit

Pilot’s inverter power supply

Pilot’s windshield anti-ice and control

Standby horizon

Steering control

No. 1 VHF, VOR, GS, and DME

Windshield wipers

Wing anti-icing

Yaw damper servo-control


B Bus 28V DC No. 1 ADF and DME

(non-loading) Air conditioning valve control/total temperature

Anti-collision (belly) and position light

Audio warning device and Mach limit

Copilot’s interphone supply circuit

Engine ignition

Landing gear indication

Left freon heat exchanger blower

Marker beacon receiver

Passenger cabin lighting

Reservoir level/hydraulic pressure indication

Right angle-of-attack sensor

Right booster pump control

Right engine anti-icing

Right engine computer

Right engine fire detection

Right engine indication

Right engine starting

Right fuel tank indicaiton and transfer

Standby electro-pump control

Standby inverter power supply

Standby inverter control

Strobe lights

C Bus 28V DC No. 2 ADF, VOR, DME/VHF No. 3

(load-shedding) Aileron trim

Annunciator panels

Anti-skid

Cabin temperature regulation

Cigarette lighter/toilet

Crossfeed and intercom valves

Copilot’s altimeter

Copilot’s flight director instruments

Copilot’s inverter power supply

Freon system control

Fuel flow and temperature indicationHeating container

HP air bleed control

Pilot’s/copilot’s seat adjustment

Radar

Right landing light

Slat and flap indicator

Tape recorder

DC Powered Electrical Systems (cont.)


Electrical Systems

D Bus 28V DC Airbrake prior to SB 27-015

(load-shedding) Angle-of-attack indicator

Autopilot aileron/elevator servo-motors

Copilot’s windshield heating

Copilot’s windshield heating control

Flap control

Flap gear reduction motor

Glideslope No. 2

HF Communication receiver

Lateral rear glass panel anti-icing

Left landing light

Razor light

Right pitot/static/AOA anti-icing

Rudder trim

Standby stabilizer control circuit

Taxi light

Transponder (ATC) No. 2

Trim display panel

VHF No. 2

DC Powered Electrical Systems (cont.)

W Bus 115 V AC, 400 Hz Automatic cabin pressurization indication

Cockpit glareshield lighting

Pilot’s flight director instruments

Pilot’s heading channel

Pilot’s instrument/pedestal integral

Rate-of-turn

Yaw damper computer

X Bus 115V AC, 400 Hz Copilot’s flight director instruments

Copilot’s heading channel

Copilot’s instrument integral lighting

DME No. 1 and No. 2 indicator supply

Radar

Radio altimeter

Y Bus 26V AC Pilot’s navigation instruments

Z Bus 26V AC Copilot’s navigation instruments

AC Powered Electrical Systems


Electrical Systems

The Falcon 10/100 lighting systemconsists of interior, emergency, andexterior lighting. Interior lightingincludes systems for the cockpit, pas-senger cabin, and the rear compart-ment. Exterior lighting includes thenavigation, landing, and taxi lights.Optional exterior lighting includes anidentification light in the nosecone,wing leading edge scanning lights (icedetection), anticollision, and strobelights.

An emergency lighting system pro-vides lighting in the event of completepower failure.

Interior LightingInterior lighting consists of:� cockpit� passenger cabin� rear compartment.

Cockpit LightingCockpit lighting consists of ambient,integral, and instrument panel lightingsystems. The majority of the lightingcontrols are on either side of the over-head panel. The pilot’s lighting panel(Figure 5C-19 ) controls the pilot’sinstrument integral, pedestal integral,and glareshield lighting, and a readinglight. The copilot’s lighting panel(Figure 5C-20 ) controls the overheadpanel, radio control box, and copilot’s

integral instrument lighting, and a read-ing light.

Ambient

Ambient lights include a single cock-pit dome light, circuit breaker floodlight, and adjustable reading lamps forthe pilot and copilot. The circuit break-er panel and reading lamps receive28V DC power from the A bus. Thedome light receives 28V DC powerfrom the Battery bus with the auxil-iary bus control switch on. The DOMEswitch on the side of the circuit break-er panel controls the dome light. TheCIRCUIT BREAKER PANEL switchon the top of the circuit breaker panelcontrols the panel flood light. SeparateREADING LIGHT controls on eitherside of the overhead panel vary theintensity of the pilot’s and copilot’sreading lamps.

Instrument Integral

Instrument integral lighting consists offour separate systems for pilot, copilot,and pedestal instruments and radiocontrol lighting. The pilot, copilot, andpedestal integral lighting systems use115V AC from the inverters reduced to5.5V AC by separate transformers. Theradio control lighting system uses 28VDC from the A bus. Separate controlsemploy rheostats to vary the intensityof each of the lighting systems fromoff to full intensity (0 to 5.5V).

LightingSystems

5C-205C-19


The pilot and pedestal integral instru-ment lights receive power from the115V, 400 Hz W bus through separatetransformers. The PILOT INSTRU-MENTS rheostat controls the pilot’slights, and the PEDESTAL rheostatcontrols the pedestal instruments. Thecopilot’s lighting system receivespower from the X bus through a trans-former controlled by the COPILOTINSTRUMENTS rheostat. Radio con-trol lighting employs a rheostat to varythe voltage to the lights. The RADIOCONTROL BOX rheostat varies Abus power from 0V to 28V to adjustthe lights from off to full intensity.

Glareshield Lighting

Glareshield lighting consists of a fluor-escent tube under each side of the glare-shield. The tubes shine down on theinstruments to provide lighting if theintegral instrument lighting systemfails.

The 115V W bus supplies AC powerto a power supply unit for glareshieldlighting. The power supply togetherwith the SHIELD rheostat varies thevoltage from 380V to 450V to adjustthe intensity of the lights from dim tofull intensity.

Passenger CabinPassenger cabin lighting systemsinclude entrance, bar, ambient, and

passenger call signs (warning). Asinterior configurations vary from air-craft to aircraft, only a brief discussionof the cabin lighting systems is pre-sented here.

Typical cabin lighting systems employ28V DC from the DC electrical sys-tem to power the lights. The primarycontrols for the cabin lighting systemsare on the overhead panel below theinverter controls. Additional controlsfor the cabin and entrance lights areon the partition to the right of theentrance door.

Entrance and Bar Lighting

Entrance and bar lighting consists offluorescent tubes on a partition to theright of the entrance door and fluores-cent tubes in the ceiling above the lug-gage compartment. Luminous switcheson the partition to the right of theentrance door control the lights. TheENTRANCE and CABIN switches(Figure 5C-21 ) illuminate once theentrance door is unlocked. Switchesin the bar door and galley curtain con-trol the lights for these areas. Openingthe door or the curtain illuminates thelights.

Cabin Ambient Lighting

Cabin ambient lighting includes fluo-rescent tubes above each passengerwindow and above each table. The

5C-22

5C-21


Electrical Systems

CABIN switch (Figure 5C-22 ) in thecockpit controls the lights above eachwindow; switches in the cabin may beavailable. Individual switches controlthe lights over the tables.

Passenger Call Signs

The passenger call sign consists of asingle No Smoking/Fasten Seat Beltssign (Figure 5C-23 ). The NO SMOK-ING and SEAT BELTS luminousswitches on the overhead panel con-trol the operation of the sign (Figure5C-22).

Rear CompartmentA single dome light provides illumi-nation of the rear compartment. Thelight receives power from one of the28V DC electrical buses. Light con-trol is automatic through a microswitchon the rear compartment door and theAux bus control switch.

Emergency LightingTwo independent batteries supply DCto power emergency lighting systemsin the event of complete electrical fail-ure or intentional electrical system shutdown. A three-position (OFF/ON/ARMED) switch on the overheadpanel (Figure 5C-22 ) controls the arm-ing and operation of the lights. Thesystem consists of emergency exit

handle lights, emergency exit spot-lights, entrance door spotlights, exitsign lights, and cockpit dome lights.

During normal operations 115V, 400Hz AC power from the W bus chargesthe batteries.

In ARMED, the emergency lightingsystem automatically provides powerto the lights if the electrical systemfails or is shut down. In ON, the lightsilluminate for aircraft lighting duringengine starting. The main batteriesmust be on to change switch positions.

Exterior LightingExterior lighting consists of:� navigation� anticollision� landing� taxi.

Additional exterior lighting systemsinclude optional nose cone identifica-tion (recognition), leading edge scan-ning (ice detection), anticollision, andstrobe lights. The exterior lighting con-trols are on the lower right side of theoverhead panel (Figure 5C-24 ).

Direct current from the A, B, C, andD buses provides power for the ex-terior lights.

5C-245C-23


Navigation and StrobeThree navigation lights are on the air-craft: a green light on the right wingtip, a red light on the left wing tip, anda white light in the tailcone. Optionalconfigurations combine the naviga-tion lights with a strobe light (Figure5C-25). The two-position (OFF/NAV)or three-position (OFF/NAV/NAVSTROBE) on the overhead panel pro-vides the navigation or combinationnavigation/strobe lights with 28V DCfrom the B bus.

AnticollisionThe high intensity anticollision lightsconsist of red rotating beacons on theaircraft belly (Figure 5C-26 ) and fintip fairing. The lights receive powerfrom separate sources: the belly lightfrom the B bus, and the fin tip lightfrom the A bus.

Landing LightsA single 450 watt landing light (Figure5C-27) is on each main landing gearstrut. The Landing light switch in com-bination with the landing gear handlecontrols the operation of both lights.With the landing gear handle in thedown position, the contacts of the land-ing light switch energize to providepower to the lights. Placing the switchON provides 28V DC from the C andD buses. The right light receives powerfrom the C bus, and the left lightreceives power from the D bus.

Taxi and IdentificationA single, 150 watt taxi light (Figure5C-28) is on the nose landing gearstrut; the light rotates with the nosewheel. The two-position TAXI switch(OFF/TAXI) on the overhead panelcontrols the light with 28V DC fromthe D bus.

5C-27

5C-29

5C-26

5C-28

5C-25


Electrical Systems

The optional identification (recogni-tion) lights in the nosecone consist oftwo 250 watt lights controlled by theTAXI/IDENT switch; they receive28V DC power from the D bus.

Ice DetectionThe ice detection (wing leading edgescanning) lights consist of 40 watt

lights (Figure 5C-29 ) forward of thewing root that shine back on the wingleading edges. The A bus provides 28VDC for the lights through the WINGICE LIGHTS switch on the overheadpanel.


Electrical Systems

DataSummary

Lighting Systems

Power Source 28V DCA and B buses (non-shedding)C and D buses (load-shedding)

115V AC, 400 HzW bus (non-shedding)X bus (load-shedding)

Control Switches and rheostatsFlight deckPassenger individual lightsEntrance and cabinExterior lightsGear handle (landing light)

Monitor Warning and advisory lights



FireProtection

Chapter 5D

The Falcon 10/100 Fire Protection System consists of heat-sensitive, expanding gas detection systems and electricallypowered extinguishers for the two engine fire zones.

Both aural and visual warnings in the cockpit activate in caseof a fire or overheat condition in these areas.

The cabin and cockpit each have a portable fire extinguisherbottle.


Fire Protection

Table ofContents

Fire Protection System Schematic . . . . . . . . . 5D-4

Fire Detection . . . . . . . . . . . . . . . . . . . . 5D-5

Engine Fire Detection . . . . . . . . . . . . . . . . 5D-5

Fire Warning Annunciators and Horn . . . . . . . . . 5D-5

Fire Extinguishers . . . . . . . . . . . . . . . . . . 5D-7

Extinguishers . . . . . . . . . . . . . . . . . . . . . 5D-7

Extinguisher Power Supplies . . . . . . . . . . . . 5D-7

Portable Fire Extinguishers . . . . . . . . . . . . . 5D-7

FIRE PULL Handles/Fuel Shutoff Valves . . . . . . 5D-8

Extinguisher System Protection . . . . . . . . . . 5D-8

Engine Fire Extinguishing . . . . . . . . . . . . . . 5D-8

Preflight and Procedures . . . . . . . . . . . . . . 5D-9

Preflight . . . . . . . . . . . . . . . . . . . . . . . . 5D-9

Fire Protection System Test . . . . . . . . . . . . 5D-9

Emergency Procedure . . . . . . . . . . . . . . . . 5D-9

Engine Fire . . . . . . . . . . . . . . . . . . . . . 5D-9

Data Summary . . . . . . . . . . . . . . . . . . . 5D-11


ELECTRICAL CONNECTIONMECHANICAL CONNECTION

FIRE

1

PULLB BUS

NO. 2 FIREWARNING LIGHT

HORN

HORN SILENCETEST

LOOP FILLEDWITH GAS

DIAPHRAGM

PORTABLEEXTINGUISHERCOCKPIT

RH ENGINEEXTINGUISHINGCONTROL

LH ENGINEEXTINGUISHING

CONTROL

XTING XTING

0TEST TESTRED DISK BOTTLE PRESSURE GAGE

FUEL F/WSHUTOFFVALVES

FROMFEEDER TANKS

1 1

B BUS

FIRE

1

PULL

FIRE

2

PULL

1

PYROTECHNICCARTRIDGECIRCUIT

PORTABLEEXTINGUISHER

CABIN

A BUSBATTERY BUS

0

21

21

Fire Protection System


Fire Protection

Fire detectors in both of the enginenacelles are sensitive to ambient tem-perature rise or localized hot spots.

A stainless steel sensor tube (LindbergLoop detector) in each of the high riskareas (engine nacelles) consists of asensor element attached to a steel hous-ing. The sensor element is protectedby a stainless steel tube helicallyformed and attached to the steel hous-ing. The tube contains an inert gas thatexpands when a relatively high temp-erature is sensed over the entire lengthof the tube. The expanding gas closesa normally open detection contact thattriggers the appropriate fire warningcircuits.

A discrete element in the detecting tubeemits gas when a very hot temperaturein a localized area occurs. This causesthe pressure within the loop to increaseand close the detection switch that trig-gers the appropriate fire warning cir-cuits. Detection circuits are poweredby the A and B buses.

When the temperature lowers, the ele-ment reabsorbs its contracting gas, andthe detection contact reopens to rearmthe detection system.

At one end of the tube is a box thathouses:� the alarm switch for the detectionsystem� the integrity switch for the testsystem.

Engine FireDetectionA 12 ft stainless steel detection tubeis around each engine. Each 1.6 mmdiameter tube contains the inert gasthat triggers the fire warning circuitsat 400°F (204°C). When a core ele-ment senses a local hot spot, it triggersthe warning circuits at 800°F (427°C).The alarm switch housed in the detec-tion box illuminates the appropriateFIRE PULL annunciator in the cock-pit. A warning horn also sounds.

WarningAnnunciators andHornIn the event of fire or overheat condi-tion in a nacelle, the correspondingwarning annunciator on the fire panelilluminates and the fire warning hornsounds. The continuous two-tonesound silences when the pilot pressesthe HORN SILENCE pushbutton onthe pedestal (Figure 5D-1 ).

Each of the fire panel warning annun-ciators contain two parallel-wired bulbsthat test when the pilot presses theTEST pushbutton just below and out-board of each fire handle (Figure 5D-2). A self test of the fire detection loopsand the squibs in the fire extinguisherbottles is also initiated. All compo-nents of the detection system must beintact to achieve a valid warning test.

Fire Detection

5D-1 5D-2


Fire Protection

ExtinguishersTwo fire extinguisher bottles (Figure5D-3) in the aft compartment, arecross-connected to permit both extin-guishers to discharge into either enginenacelle through diffuser nozzles. Thehigh pressure (600 PSI) bottles arefilled with R13B1 Freon (European)or Halon 1301 (US) for firesuppression.

Each bottle incorporates two dischargeoutlets. One outlet is the first shot toan engine; the other is the second shotto the other engine. The extinguishingagent moves to the respective enginethrough check valves that are openedby the extinguishing agent pressure.

Because the discharge is not metered,each bottle empties its contents through

the discharge nozzles in less than onesecond.

Extinguisher PowerSuppliesEach engine extinguisher is controlledby a three-position (0/1/2) safety wiredswitch under its associated FIREPULL handle. See Table 5D-A .

Portable Fire ExtinguishersThe cockpit portable fire extinguishermust be of either the Halon 1211 orCO2 variety; it can be used on ClassA, B, or C fires. The passenger com-partment portable extinguisher mustbe either water (suitable for Class Afires) or Halon (suitable for Class A,B, or C fires).

WARNING: Extinguishers aredirectly supplied by the Batterybus and may be fired with bat-tery switches in the OFF posi-tion with batteries connected.

FireExtinguishers

5D-3

Table 5D-A; Extinguisher Power Supplies1 Observe warning at right.

Switch A Bus Battery Bus 1

ENG 1 Position 1 Position 2

ENG 2 Position 1 Position 2


FIRE PULL Handles/Fuel Shutoff ValvesAn illuminated FIRE PULL annunci-ator in the FIRE PULL HANDLEindicates engine fire or overheat. FIREPULL handles attached to teleforcecables control the fuel shutoff valves ineach pylon (Figure 5D-5 ). Whenpulled, the handle shuts off fuel to itsrespective engine.

Extinguisher SystemProtectionEach fire bottle has an overpressuredischarge device. Excessive internalpressure buildup (from ambient temp-erature of 205°F or more) melts a ther-mal fuse to initiate a discharge throughblowout discs on the aft fuselage undereach engine pylon (Figure 5D-4 ). Thered disc is knocked out by escapingpressure and uncovers the overpressure

relief outlet. When an overboard dis-charge occurs, the bottle empties.

Engine FireExtinguishingEach engine fire extinguisher bottleincorporates two discharge frangiblediscs that are ruptured by indepen-dently controlled pyrotechnic squibs.The volume of each extinguisher isthree pounds of Halon 1301 with anoperating pressure of 600 PSI at 72°F(22°C).

The bottles are cross-connected so thatboth may be discharged into oneengine. The A bus powers the first shotto engines 1 or 2. As a backup for pos-sible primary bus failure, the batterybus powers the second shot to theengines. Refer to the Data Summarytable in this chapter.

TELEFORCE CONTROL CABLE

ENGINEPYLON FIREWALL

2

1FUEL

SHUTOFFVALVE

1

2

5D-5

5D-4


Fire Protection

PreflightDuring preflight, the crew checks inthe cockpit that the extinguisherswitches are set to the normal (0) posi-tion and safetied, and on the exteriorinspection that the red extinguisherdischarge indicators on the fuselageare intact.

A placard on each extinguisher bodylists corresponding pressures and temp-eratures, which should be checkedprior to each flight.

Fire ProtectionSystem TestWhen pressed, the TEST button, justbelow the fire pull handle, activates awarning horn and illuminates theengine 1 and engine 2 FIRE PULLannunciators and checks:� the integrity of the sensing elements(detector head and tube)� the integrity of the bottle pyrotech-nic cartridges (squibs)� the horn and warning annunciators’operating condition� that associated circuit breakers areengaged.

A system does not test if the detectiontube or squibs are not working in oneof the following situations:� if the annunciator is out and no hornsounds, the corresponding system hasfailed (i.e., a detector head or tube isprobably damaged or cut).� if a annunciator is out and the horndoes sound, there may be a warningannunciator malfunction or defectivewires.

� if the horn fails to sound while thewarning annunciator illuminates, thisindicates a horn malfunction or defec-tive wires.

EmergencyProcedureEngine FireA warning horn and an illuminatedFIRE PULL handle annunciator indi-cate fire in the associated engine. Thehorn is silenced by pressing the HORNSILENCE button on the pedestal; donot silence the horn until the problemis identified.

When the problem is identified, silencethe warning horn, and pull the illumin-ated FIRE PULL HANDLE, and retardthe power lever to IDLE to reduce fuelto the engine at the fuel control. Thefuel shutoff valve from the respectivefeeder tank is closed when the FIREPULL handle is pulled and along withswitching off the electric boost pumpensures fuel is no longer supplied tothe engine. Closing the HP bleed airvalves serves to decrease the chanceof contaminated air entering the cock-pit or cabin; turning off the Freon airsystem reduces the demand on the oneremaining generator.

Reducing airspeed reduces airflow inthe nacelle and helps in facilitating fireextinguishing.

Move the appropriate extinguisherswitch to position 1 to discharge thecontents of one fire bottle into theengine nacelle. If the condition per-sists, move the extinguisher switch toposition 2 to discharge the contents ofa second fire bottle into the nacelle.


WARNING: Do not attempt anengine relight after an engine fireor if the integrity of the engineis questioned.


Fire Protection

DataSummary

Fire Protection SystemEngine Fire Protection System

Power Source A busEngine 1 detection circuit (FIRE 1 PULL light)

B busEngine 2 detection circuit (FIRE 2 PULL light)

A busExtinguisher shot 1 to either engine

Battery busExtinguisher shot 2 to either engine

Teleforce cables on FIRE PULL handlesFuel shutoff valves

Distribution Left and right engines

Control TEST button

Horn Silence button

FIRE 1/2 PULL handles

FIRE XTING 1/2 switches

Monitor FIRE 1/2 PULL handles (red lights)

Protection Fire warning test system

Fuel firewall valves

Fire bottle discharge device

Falcon 10/100 For SimuFlite training only 5E-1April 2000

FlightControls

Chapter 5E

The Falcon 10/100 primary flight controls transmit pilot inputsthrough a mechanical linkage system to dual-barrelled, hydraulicservo actuators that move the ailerons, elevators, and rudder.

The secondary flight controls consist of a moveable horizontalstabilizer that provides pitch trim, trailing edge flaps, inboard andoutboard leading edge slats, and airbrakes.

In addition to the primary and secondary flight controls, stallwarning devices complete the flight controls system.

5E-2 For SimuFlite training only Falcon 10/100April 2000


Flight Controls

Table ofContents

Elevator Control Linkage Schematic . . . . . . . . 5E-6

Arthur Q Unit Schematic . . . . . . . . . . . . . . 5E-8

Primary Flight Controls . . . . . . . . . . . . . . . 5E-9

Servo Control Actuators . . . . . . . . . . . . . . . 5E-9

Torque Bar (Secondary AFU) . . . . . . . . . . . 5E-10

Bypass Valve . . . . . . . . . . . . . . . . . . . 5E-10

Mechanical Stops . . . . . . . . . . . . . . . . . 5E-10

Irreversibility Feature . . . . . . . . . . . . . . . 5E-10

Internal Leakage . . . . . . . . . . . . . . . . . . 5E-10

Artificial Feel Unit (AFU) . . . . . . . . . . . . . . . 5E-10

Arthur Q Unit . . . . . . . . . . . . . . . . . . . . 5E-10

Q UNIT Light . . . . . . . . . . . . . . . . . . . . 5E-12

Ailerons . . . . . . . . . . . . . . . . . . . . . . . 5E-12

Amedee Unit . . . . . . . . . . . . . . . . . . . . 5E-13

Roll Trim . . . . . . . . . . . . . . . . . . . . . . 5E-13

Aileron Trim Switch . . . . . . . . . . . . . . . . 5E-13

Elevators . . . . . . . . . . . . . . . . . . . . . . 5E-13

Horizontal Stabilizer/Pitch Trim . . . . . . . . . . 5E-13

Horizontal Stabilizer Actuator . . . . . . . . . . . 5E-14

Autopilot Stabilizer Control . . . . . . . . . . . . 5E-14

Horizontal Stabilizer Emergency Operation . . . . 5E-15

Clacker . . . . . . . . . . . . . . . . . . . . . . 5E-15

Rudder . . . . . . . . . . . . . . . . . . . . . . 5E-15

Rudder Pedals . . . . . . . . . . . . . . . . . . . 5E-15

Rudder Trim . . . . . . . . . . . . . . . . . . . . 5E-15

Yaw Damper . . . . . . . . . . . . . . . . . . . . 5E-16

Gust Damper . . . . . . . . . . . . . . . . . . . 5E-16

Triple Trim Indicator . . . . . . . . . . . . . . . . 5E-16


Normal Slat Control Schematic . . . . . . . . . . 5E-18

Slat Control – Automatic (Stall)and Emergency Schematic . . . . . . . . . . . . 5E-20

Flap Control System Schematic . . . . . . . . . 5E-22

Airbrake Control System Schematic . . . . . . . 5E-24

Secondary Flight Controls . . . . . . . . . . . . 5E-25

Slats . . . . . . . . . . . . . . . . . . . . . . . . . 5E-25

Emergency Slats . . . . . . . . . . . . . . . . . 5E-25

Slats Indication . . . . . . . . . . . . . . . . . . 5E-25

Slat/Flap Control Handle . . . . . . . . . . . . . 5E-26

Stall Warning System . . . . . . . . . . . . . . . 5E-26

Flaps . . . . . . . . . . . . . . . . . . . . . . . . 5E-27

FLAP ASYM Warning . . . . . . . . . . . . . . . . 5E-28

Flap Position Indicator . . . . . . . . . . . . . . . 5E-28

Airbrakes . . . . . . . . . . . . . . . . . . . . . . 5E-28

Airbrake Handle . . . . . . . . . . . . . . . . . . 5E-29

Airbrake Indications . . . . . . . . . . . . . . . . 5E-29

Preflight and Procedures . . . . . . . . . . . . . 5E-31

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5E-31

Abnormal Procedures . . . . . . . . . . . . . . . . 5E-31

Pitch Trim Malfunction . . . . . . . . . . . . . . . 5E-31

Aileron or Rudder Trim Malfunction . . . . . . . . . 5E-31

Yaw Damper Malfunction . . . . . . . . . . . . . 5E-31

Slats Only Approach and Landing . . . . . . . . . 5E-31

Q Unit Failure . . . . . . . . . . . . . . . . . . . 5E-31

Electric Q High-Speed (Cruise) Failure . . . . . . 5E-31

Slat Malfunction . . . . . . . . . . . . . . . . . . 5E-32

Stabilizer Trim Light on Takeoff . . . . . . . . . . 5E-32

Flap Asymmetry . . . . . . . . . . . . . . . . . . 5E-32

Data Summaries . . . . . . . . . . . . . . . . . . 5E-33


Flight Controls

RUDDER SERVOCONTROL

ELEVATORSERVO CONTROL

TRIM JACK

RUDDER LINKAGEMAIN AFUARTHUR JACK

SERVO CONTROL

AILERONTRIM JACK

AILERON/ELEVATORLINKAGE AFU

AILERON/ELEVATORLINEARARTHUR JACK AMEDEE

PUSHRODS


Elevator Control Linkage


5E-8

Flight Controls

Arthur Q Unit


PITOT

STATIC

A BUS

COPILOT

P

R

DIFFERENCE40 KT ACCEL30 KT DECEL

ARTHUR QCOMPARATOR

MONITOR

ARTHUR QSYSTEM

ACTUATOR

PITOT

STATIC

28 VOLTS DC

COPILOT

LOWSPEEDSWITCH(SB 165)

HIGH SPEED

LOWSPEED

DOWNSPRING

UPSPRING

ARTIFICIAL FEEL

UP

UP

FLOORBOARDCOCKPIT

YOKE

250 KT SWITCH(4° CRUISE STOP ANDYAW DAMP 1.5°)

190 KT SWITCH(LANDING GEAR WARNING CIRCUITFLAP OVERSPEED WARNINGAUTO SLAT INHIBIT > 190 KTS)

AIRSPEED FROM Q UNIT

28V DC

MECHANICAL STOP FORMANUAL REVERSIONS

VALVECONTROL

ARM

DOWN UP

AIRSPEEDFROM

REFERENCECAPSULE

Q. UNIT

AIL RUD

1

2

2

1

AIRSPEEDREFERENCE

CAPSULE

A

BUS

Arthur Q Unit

Falcon 10/100 For SimuFlite training only 5E-9March 2001

Flight Controls

The Falcon 10/100 hydraulically pow-ered primary flight controls (Figure5E-1) allow control of the aircraftthrough the roll, pitch, and yaw axes.Control inputs are transmitted throughmechanical linkage systems to dual-barrelled, hydraulic servo actuatorsthat move the following:� ailerons� elevators� rudder.

Servo ControlActuatorsEach control system’s servo actuatorcomprises two barrels in one commonblock. One barrel of the servo actuatorreceives pressurized hydraulic fluidfrom Hydraulic System 1, which ispowered by the left engine hydraulicpump. Hydraulic System 2, poweredby the right engine hydraulic pump,supplies the second barrel. The yokeand rudder pedals move control valvesto port hydraulic fluid to the appro-priate side of each piston.

Once the surface moves to the select-ed position, the appropriate control

valve closes, and trapped hydraulicpressure holds the surface in positionwith no pilot input pressure required.

In the event one hydraulic system fails,maximum allowable airspeed is 260kts/0.76 M.

If both hydraulic systems fail, the flightcontrols revert to a conventional sys-tem. Because no pressure is available,manual control of the yoke moves thecontrol valve arms to a mechanicalstop; this moves the entire actuator anddisplaces the control surface. In thiscase, the maximum allowable airspeedis 260 kts/0.76 M, and the maximumrecommended landing weight is15,400 lbs. Because of control forcesrequired, crosswind landings are moredifficult without hydraulic pressure.

The primary control servo actuatorsincorporate the following protectivedevices:� torque bar (secondary AFU)� bypass valve� mechanical stops� irreversibility feature� internal leakage.

PrimaryFlightControls

HORIZONTALSTABILIZER

RUDDER ELEVATOR

FLAPS

AIRBRAKES

AILERON

OUTBOARD SLAT

INBOARD SLAT

5E-1

CAUTION: In the event ofcomplete hydraulic failure, ex-pect extreme control forces be-cause the pilot must overcomeboth the airloads and the artifi-cial feel system.

NOTE: Because there is no feedback, or feel, to the pilot,the control system is called “irreversible” and requires theaddition of an artificial feeldevice. See Artificial Feel Unit(AFU), this chapter.


Torque Bar(Secondary AFU)Each primary control linkage systemincorporates a torque bar (secondaryAFU) that acts as a centering devicein the event of broken linkage or totalhydraulic failure.

In the event of broken aileron linkage,the torque bar returns the servo actua-tor levers to neutral and fairs the con-trol surfaces. The failed aileronhydraulically locks in neutral position,and the other aileron allows accept-able control of the aircraft.

If total hydraulic failure occurs, thissecondary AFU causes control forcesto become very heavy.

Bypass ValveIn case of hydraulic power failure toone actuator barrel, a bypass valve pre-vents the failed barrel from interfer-ing with the normal function of theoperating barrel.

There is no locking system for eitherthe control surfaces or linkages otherthan active or stored hydraulic pres-sure. When the aircraft is parked andengines are shut down, the hydraulicaccumulators progressively discharge,and the actuator barrels no longerreceive pressure. Both bypass valvesopen when hydraulic pressure drops;friction resistance in the linkage holdscontrol surfaces in a neutral position.

Mechanical StopsThe servo actuator actuating levermoves between two mechanical stopsto provide the pilot with a mechanicalmeans of flying the aircraft if allhydraulic systems fail completely. Inthis case, the forces of friction, sec-ondary AFU, torque tube, and, espe-cially, aerodynamic loads on the con-trol surfaces reduce aircraft responseand could overstress the airframe.Limit airspeed to 260 kts/0.76 M.

Irreversibility FeatureThe check valve in the hydraulic sup-ply line prevents displacement of theservo actuator in the reverse directionif the aerodynamic loads on the controlsurface exceed the force developed bythe servo actuator.

Internal LeakageEach servo actuator slide valve has acalibrated leakage that allows a flowof 600 cm3 per minute to warm thepower servo unit during prolongedflight at high altitude.

Artificial Feel Unit(AFU)Because the primary controls are pow-ered hydraulically, the only forces toovercome are the linkage componentfriction and resistance. Therefore, eachcontrol system incorporates an artificialfeel unit (AFU) (Figure 5E-2 ) thatgives the pilot a control force thatincreases with control deflection; theforce varies as a piston moves againstthree preloaded springs within theAFU tube. An artificial feel unit is oneach of the three primary flight con-trol linkages:� aileron� elevator� rudder.

Arthur Q UnitThe faster the aircraft flies, the small-er the deflection required of controlsurface to get the same response(Figure 5E-3 ). To provide the neces-sary feel and response at higher air-speeds and to further refine the artifi-cial feel system, each primary controllinkage system incorporates an ArthurQ unit (Figure 5E-2 ).

Arthur means motion, and Q refers todynamic pressure, or airspeed. Eitherof the two Falcon 10/100 Arthur Q


Flight Controls

RODAILERON ARTHURSLIDING BLOCK

TRIMACTUATOR

AILERONAFU

ELEVATORARTHURSLIDING BLOCK

ELEVATOR AFU

ARM BELLCRANK

ARTHUR Q UNITACTUATOR

5E-2

260/0.76 M

250

190

140

0 140 190 250 260/0.76 M

Q UNITSLOW SPEEDSTOP

FLAP OVERSPEED WARNINGGEAR WARNINGAUTO SLAT EXTENSION <190 KIASRIGHT AOA >19°

YAW D

AMP L

IMIT

ING AUTO SLAT

EXT <190KIAS

3.5°

1.5°

LIMITSNORMALSTAB TRIMTO 4°NOSE-UPAT 250KIAS

BACKS UPYAW DAMPLIMIT AT250 KIAS

CONTROL FORCE IN

CREASES

LINEARLY

TO 4

00 K

IAS

ARTHUR Q HIGH MODE

ACTION FOR AIL Q FAILURE

MAX AIRSPEED 260 KT/0.76 MVREF + 10 MIN. APPROACH SPEED

1.2.

CAUSE FOR INDICATION

RUD OR AIL ARTHUR Q FAILURECOPILOT PITOT STATIC FAILURESYSTEM 1 HYD FAILURE FOR RUD

1.2.3.

ACTION FOR RUD Q FAILURE

MAX AIRSPEED 260 KT/0.76 MVREF + 10 MIN. APPROACH SPEEDHIGH Q –140 KIAS MIN. APPROACHSPEEDDO NOT TAKE OFF

1.2.3.

4.

Q. UNIT

RUD AIL

ELECTRIC AILERON/ELEVATOR Q (KIAS)

RU

DD

ER

LEFTAOA >17°GEARWARNINGLEFT

AOA >17°

AUTO SLATEXT <190KIAS

5E-3


units (separate from the AFU) mechan-ically responds to changes in airspeedand varies the mechanical advantageof the yoke over the artificial feel sys-tem.� The Primary A bus powers andactuates the aileron and elevator ArthurQ units.� Hydraulic System 1 actuates therudder Arthur Q unit.

Each Q unit, which is a variable lengthbell crank, drives its respective flightcontrol system’s main AFU. The unit’sarm length varies with the aircraft’sindicated airspeed input to control theAFU travel and the loads commensu-rate with indicated airspeed (i.e., lightcontrols at low speed, heavy controlsat high speed).

The rudder’s hydraulic Q system uti-lizes a large spring to move towardlow speed feel as the aircraft deceler-ates and hydraulic pressure to movetoward high speed feel on accelera-tion. Therefore, if Hydraulic System1 fails, the rudder Q system automat-ically goes to the low speed (140 kts)position. This is the reason for the 260kts/0.76 M restriction during hydraulicfailure; the resulting light feel forcescould cause overcontrol and overstressof the empennage.

The electric Q system’s motor posi-tions the mechanical linkage through ajackscrew so that the linkage remainsin its last position if the aileron/elevatorQ system fails.

Q UNIT LightIn flight, an illuminated Q UNIT light(refer to Chapter 5J, MiscellaneousAnnunciator Cross Reference) indi-cates there is a difference betweenArthur Q position and aircraft speed. Itdoes not mean system failure; it meansa difference exists. Under rapid accel-eration or deceleration, it is not uncom-mon for the Q UNIT light to flickeron and off to indicate the Q unit is notkeeping up with the airspeed change.

AileronsThe hydraulically actuated ailerons(Figure 5E-5 ) on the outboard trailingedge of each wing provide the primaryroll control of the aircraft. The aileronlinkage system routes through the fuse-lage and the wing leading edge to theaileron servo actuator (Figure 5E-4 ).See Servo Control Actuators, thischapter.

Control yoke rotation moves eachaileron inversely to the other througha bell crank and torque bar system; the

TORQUE BARATTACHMENT

ATTACHMENTTO AIRFRAME

MECHANICALSTOPS

ATTACHMENT TOCONTROL SURFACE

5E-4


Flight Controls

aileron on one wing moves up at thesame time the aileron on the oppositewing moves down. Full travel rangeis 25°20´ up and 24°50´ down. Bothaileron linkages are routed through adevice called an Amedee.

Amedee UnitThe Amedee unit is a mechanical, vari-able ratio linkage between the yokeand the aileron servo. It reduces ailerontravel at small control yoke displace-ments (in normal smooth flying con-ditions), but it gives proportionatelygreater displacement as control yoketravel increases. The Amedee’s func-tion is comparable to an automobile’svariable rate power steering.

Roll TrimThe AILERON trim switch on thecockpit pedestal provides electricalaileron trim control with aileron deflec-tion up to 30% of normal travel. TheAuxiliary C bus powers the ailerontrim actuator, which is just forward ofthe AFU.

Aileron Trim SwitchThe aileron TRIM switch (Figure 5E-6) incorporates a cancelling device thatprevents trim runaway. On S/Ns 1 to109, the aileron trim is controlled witha single, dual action switch on thepedestal trim switch panel. A double,dual action switch is on S/N 110 andsubsequent.

ElevatorsThe elevators (Figure 5E-7 ), on thetrailing edge of a moveable horizon-tal stabilizer, provide pitch controlthrough fore and aft movement of thecontrol column.

A dual-barrelled, hydraulically pow-ered servo actuator drives the eleva-tors. The pilot’s column controls theactuator through linkage routedbeneath the fuselage floor, aft, andupward to the hinged elevators. SeeServo Control Actuators, this chapter.

Horizontal Stabilizer/Pitch TrimThe electrically driven horizontal sta-bilizer is an airfoil that provides lon-gitudinal stability and replaces the needfor trim tabs. The stabilizer is hingedat two points, fore and aft, to the ver-tical fin stub.

To provide pitch trim, the horizontalstabilizer pivots around the aft mountpoint. A scissors-type hinge anchoredto the fin stub at the forward attachpoint provides lateral stability whilethe dual-motor electrical actuator drivesa vertical shaft to provide trim motion.One motor is used for normal opera-tion; the other is used for emergencyoperation. The Primary A bus powersthe normal trim motor through the 10ANORMAL circuit breaker on the cen-ter pedestal, while the Load Shed Dbus powers the emergency motor.

5E-5 5E-6 5E-7


Horizontal StabilizerActuatorThe stabilizer actuator is a jackscrewmechanism that is driven through areduction gearbox by either of twoelectric motors. Trim buttons on thecontrol yokes (Figure 5E-8 ) controlthe motors. A switch on the pedestalcontrols the emergency trim motor.

The actuator incorporates severalmicroswitches.� The limit microswitches stop thetrim motion as it nears the end of itsmechanical travel. These microswitch-es are available for both normal andemergency trim control systems.Should these microswitches fail, twomechanical stops limit travel prior tostructural damage.� Switches at the 4° and 8° trim posi-tions actuate the STAB TRIM light ontakeoff if trim is less than 4° or greaterthan 8°. A second 4° switch disablestrim beyond 4° nose-up at airspeedsgreater than 250 kts. This switch worksin series with the 250 kts switch on theaileron/elevator electric Q unit to

prevent a runaway nose-up trim fromoverstressing the aircraft. The secondswitch creates a small problem in theevent of Q unit failure in the highspeed position, because normal trimworks only to the 4° microswitch.Emergency trim is required to com-plete trimming for landing.

On S/N 99 and subsequent, splitswitches on each control yoke (Figure5E-8) activate the pitch trim; bothswitch halves must be pressed simul-taneously in the same direction. OnS/Ns 1-98, the switches are not splitunless SB-27-013is installed. The auto-pilot activates the pitch trim as well.

Autopilot StabilizerControlWhen engaged, the autopilot issuespitch trim control signals if the 10Atrim control circuit breaker is engaged.Deflection limits are the same as inmanual control. Actuating manualpitch trim control automatically dis-engages the autopilot in most aircraftwhile leaving the yaw damperengaged.

SYNC

RADIO

AP DIS

C

ICS

PUSH TOTALK RADIO

PITCH SYNCHRO

AUTOPILOTDISENGAGE

DUALSWITCH

P-T-T INTERPHONE(OPTIONAL)

GO-AROUND

5E-8


Flight Controls

Horizontal StabilizerEmergency OperationThe Auxiliary D bus powers theemergency horizontal stabilizer sys-tem, which controls the emergencymotor of the electrical actuator. Whenactivated, the STAB EMGY switch(Figure 5E-9 ) on the center pedestalactuates the emergency motor andmechanically trips the NORMALTRIM circuit breaker.

Clacker

A motion detector inside the stabilizertrim actuator activates an audible cock-pit clacker, which sounds to warn thecrew of horizontal stabilizer motion.

RudderThe rudder provides directional con-trol of the aircraft about its verticalaxis. Full range of motion is 35°. Likethe aileron and elevator actuators, adual-barrelled servo actuator drivesthe rudder. Both hydraulic systemssupply the rudder actuator.

The rudder linkage routes beneath thefuselage floor, past the rear compart-ment, and upward through the verti-cal stabilizer. The rudder is mountedon three bearings with the center bear-ing connected to the servo actuator. Aspring rod under the actuator returnsthe rudder to neutral.

Rudder PedalsThe pilot’s and copilot’s rudder pedalpositions are separately adjustable. Toadjust the pedals, turn the crank handleon the forward face of the controlcolumn until they reach the desiredposition.

Rudder TrimThe electrically actuated rudder trimactivates when the pilot presses:� a single rudder trim switch (Figure5E-10) on center pedestal (S/Ns 1-109)� both sides of a split switch (Figure5E-9) on the center pedestal (S/N 110and subsequent).

The Load Shed D bus powers ruddertrim, which deflects the rudder 40%

5E-9

5E-10


of its maximum travel. The triple trimindicator reflects the amount of trimin percentage of travel left or right.Rudder trim deflection reaches fromzero to full left or right within 15seconds, or 30 seconds for full left tofull right.

Yaw DamperThe electrically controlled andhydraulically actuated yaw dampersystem adjusts a mechanical yawdamper jack to control rapid oscilla-tions about the vertical axis (Dutchroll) in all flight conditions. Controlledby the magnetically held YD switch(Figure 5E-11 ) on the center pedestal,the system also provides coordinationin turns. The Primary A bus powersthe yaw damper system, which doesnot affect pilot control of the rudder.

Gust DamperIn addition to standard servo actuatorprotection features discussed in thischapter, a gust damper with a cali-brated bypass valve is in the rudderservo actuator. In the absence of

System 1 hydraulic pressure, the by-pass valve connects the two actuatorchambers through a restrictor, whichdamps wind effects on the rudder. WithSystem 2 pressure available, the pilot’srudder control is not affected, and theGUST light (see Miscellaneous Sys-tems chapter, Miscellaneous Annun-ciator Cross Reference) is advisoryonly.

When both hydraulic systems are inop-erative, the rudder controls are diffi-cult to move; the pilot utilizes appro-priate abnormal procedures.

Triple TrimIndicatorRoll, pitch, and yaw trim systems areelectrically controlled. The triple trimindicator (Figure 5E-12 ) on the upperleft center pedestal indicates percentof flight control deflection and cali-brates stabilizer position in degreesfrom -2° NOSE DN to +12° NOSEUP. A green range between +5°30´ and+6°30´ shows the deflection rangerequired for takeoff.

5E-11 5E-12


5E-18

Flight Controls

NormalSlat Control


FLAPS

52 30

0

DN

UP

15

RETRACTION PRESSURE

EXTENSION PRESSURE

SUPPLY PRESSURE (EXTENSION/RETRACTION)



SLAT/FLAP HANDLE(MECHANICAL VALVE)

INBOARD SLAT

EXTEND MICROSWITCHES

RETRACT MICROSWITCHES

OUTBOARD SLAT

TELEFORCE CABLE

NO.1ENGINE

PUMP

SYS. 2HYD

STANDBYPUMP

NO.2ENGINEPUMP

CLEAN

SLATS

S + FLAPS 15°

S + FLAPS 52°S + FLAPS 30°

D BUS

C BUS

SLAT CONTROLVALVE SYSTEM

SYS. 1HYD

NORMAL RETRACT

NORMAL EXTEND

Normal Slat Control


5E-20

Flight Controls

Slat ControlAutomatic (Stall)and Emergency


RETRACTION PRESSURE

EXTENSION PRESSURE


EMERGENCY

SLAT CONTROLVALVE SYSTEM

SLAT/FLAP HANDLE(MECHANICAL VALVE)

OUTBOARD SLAT

NO.1ENGINE

PUMP

STANDBYPUMP

NO.2ENGINEPUMP

SYSTEM 1AOA OUTBOARDS

NO. 2SYSTEMRETURN

AOA VANE 19°

AOAVANE17°

EXTENDMICROSWITCHES

RETRACT MICROSWITCHES

B BUSEMER SLAT

SWITCH

SYS. 1HYD

SYS. 2HYD

A BUS

AUTOEXTEND

AUTORETRACT

EMER OUTBOARDSLATS

Slat ControlAutomatic (Stall) and Emergency


5E-22

Flight Controls

Flap ControlSystem


FLAPS

52 30

0

DN

UP

15

FLAPMOTOR

TO SLATCONTROL

VALVE

CLEAN

SLATS

S + FLAPS 15°

S + FLAPS 52°S + FLAPS 30°

D BUS

FLAPASYM

AIRBRAKE

C BUS

ELECTRICAL CONNECTIONMECHANICAL CONNECTION

Flap Control System


5E-24

Flight Controls

AirbrakeControlSystem


RETRACTION PRESSURE

EXTENSION PRESSURE


FLAPS

52 30

0

DN

UP

15

SYS. 1HYD

STANDBYPUMP

AIRBRAKE

ON

EXT

NO.1ENGINEPUMP

AIRBRAKECONTROLVALVE

LH SQUATSWITCH

RETRACTMICROSWITCH

RETRACTPRESSURE

EXTENDPRESSURE

FLAPASYM

AIRBRAKE

A BUS

AIRBRAKE

Airbrake Control System


Flight Controls

The secondary flight controls consistof:� trailing edge flaps� inboard and outboard leading edgeslats� airbrakes.

SlatsEach wing has two teleforce cable-controlled, hydraulically operated lead-ing edge slats (Figure 5E-13 ): oneinboard cuff-type slat and one outboardslotted-type. The slats change thewing’s camber to allow very slowapproach speeds relative to the cruisespeed of the aircraft. Used alone, theslats provide a 20 kt reduction in theclean-wing approach speed and oper-ate in three modes.� In normal mode, Hydraulic System1 powers the inboard and outboardslats; the outboard slats extend firstwhile the inboard slats extend moreslowly because of a restriction in theinboard slat hydraulic line.� In emergency slats mode (standby),Hydraulic System 2 powers extensionof the outboard slats only. This modeprovides no retraction.� In stall protection mode (automatic),either hydraulic system automaticallypowers extension of the outboard slats.

Emergency SlatsThe slats are normally mechanicallycontrolled by the slat/flap handle, butin an abnormal condition or emer-gency, the outboard slats are electri-cally controlled by the EMERG SLATswitch (Figure 5E-14 ) on the pedestal.The Primary B bus powers the emer-gency slats switch.

If Hydraulic System 1 fails, System 2powers the emergency slats extension,but provides no retraction.

Slats IndicationA red/green slat light on the configu-ration panel indicates disagreement/agreement.

The red light illuminates:� following slat selection by the slathandle, emergency slats switch, or thestall warning system, and slats are intransit (when the requested slat con-figuration is complete, the green lightreplaces the red)� when both emergency slats andflaps are selected (the slat/flap lightremains red).

The green light illuminates:� when all four units are fully ex-tended (the slat/flap handle calls forboth inboard and outboard slats)

SecondaryFlightControls

CAUTION: The emergencyslats system does not provideretraction. Once selected, emer-gency slats should never beretracted.

5E-13 5E-14

NOTE: Angle-of-attack sensorsautomatically initiate slat exten-sion. Stall sensors control bothportions of the dual auto-matic (outboard only) slat exten-sion system.


� if the slat/flap handle is in CLEANposition and outboard slats are fullyextended as a result of either the selec-tion of emergency slats or automaticactivation by the stall warning system.

Either emergency or automatic slatextension calls for outboard slats only.

Slat/Flap Control HandleThe slat/flap control handle (Figure5E-16) on the center pedestal mechan-ically controls the slats (Figure 5E-15) through a teleforce cable and elec-trically powers the flap motor. Thehandle stops at each of five detents:� CLEAN (flaps and slats fullyretracted)� SLATS (extension of all the slats)� S + FLAPS 15°� S + FLAPS 30°� S + FLAPS 52°.

Microswitches at each flap positionsignal the flap motor, which runs untilthe flaps reach the corresponding posi-tion. The Auxiliary D bus powers theflap motor.

To move the handle from one detentto the other, pull it slightly upward torelease the latch. If the slat/flap han-dle is moved from any flap positiondirectly to CLEAN, the inboard slatsretract immediately, but the outboardslats remain extended until the flapsare within 2° of fully retracted.

Stall Warning SystemTwo angle-of-attack sensing vanes(Figure 5E-17 ) one on each side of theaircraft nose, activate the stall warn-ing system.

The A bus powers the left stall sensingvane, which controls Hydraulic System1 pressure through solenoid valve A

SLAT RETRACTION MICROSWITCH

SLAT EXTENSION MICROSWITCH

INBOARD SLAT

OUTBOARD SLAT

HYDRAULICDISTRIBUTOR

SLAT/FLAPCONTROL

ANGLE-OF-ATTACKSENSOR AND TRANSDUCER

JACK

JACK

STANDBYJACK

JACK

5E-15


Flight Controls

on the slats control valve to move theoutboard slats through two primaryactuators on each slat.

The B bus powers the right stall vane,which extends the outboard slatsthrough either solenoid valve B on thenormal slat control valve or, ifHydraulic System 1 is inoperative,through the emergency slat controlvalve. If System 1 is inoperative, emer-gency slats do not retract.

The left sensor switch is set at a 17°airflow angle and the right at 19°. Theright vane serves as backup to the left.The switches provide stall protectionfor a clean (flaps and slats retracted)aircraft by:� extending the outboard slats� turning on the ignition system� sounding a stall warning horn.

If the slats are already extended, the17° and 19° switches are bypassed,and the stall warning horn sounds at a27° airflow angle.

The horn continues to sound as longas either contact is maintained. To pre-

vent engine flameout during a stall,the igniter plugs are automaticallyexcited by the ignition boxes; bothgreen IGNITER ON lights illuminate.The ignition remains on for 10 sec-onds after the two contacts arereleased. If IAS is below 190 kts, thesame contacts cause the slats to extendautomatically through electrical cir-cuits independent of the stall warningcircuits.

The stall system is tested on the groundusing the STALL 1 and STALL 2pushbuttons (Figure 5E-18 ), which donot function in flight. Refer to the AFMor the SimuFlite Operating Handbookfor the stall system test checklist.

FlapsThe mechanically actuated and elec-trically driven, double-slotted Fowlerflaps comprise four panels, two oneach wing trailing edge. The panels,which are guided with rollers on rails,reduce approach speeds as much as 15kts and produce high degrees of dragfor short landing distances.

CAUTION: Do not use flapsunless outboard slats (at least)are extended.

5E-16 5E-17 5E-18

5E-28 For SimuFlite training only Falcon 10/100March 2001

When the slat/flap handle is positioned,the twisting forces of mechanical driveshafts actuate irreversible, mechanicaljackscrews that move the panels. Atwo-speed reversible electric motor inthe right wheel well drives thejackscrews. The Auxiliary D bus pow-ers the motor. A special tab on theslat/flap handle stops the handle at theS+FLAPS 15° position, regardless ofwhich direction the handle is moved.

If a drive shaft breaks, the jackscrews’non-reversing devices prevent flapretraction from airload. Each inboardjackscrew has an adjustable retractionstop, and each outboard jackscrew hasa fixed extension stop.

Four microswitches in a limit switchbox near the motor control the flap stopsat 0°, 15°, 30°, and 52°. Another micro-switch in the wing root prevents theretraction of the outboard slats as longas the flaps are extended more than 2°.

A reduction gear box in the motorassembly transmits torque to the flaplinkage and acts as a torque limiter.

The 120° angle gear box also incor-porates torque limit protection in caseof linkage jamming downstream.

FLAP ASYM WarningThe left and right outboard flaps areeach connected to a position poten-tiometer on the aft wing spar. In case offlap asymmetry, an electronic circuitpowered by the Auxiliary C bus tripsthe FLAP CONTROL circuit breakerand causes the flaps to stop; this illu-minates the FLAP ASYM light on thewarning panel. The threshold is equalto or greater than 5° of asymmetry.

Flap Position IndicatorThe left outboard flap potentiometerprovides flap position (0°, 15°, 30°,52°) on the slat/flap indicator (Figure5E-19). The auxiliary D bus powersthe indicator system.

AirbrakesThe airbrakes (Figure 5E-20 ), whichare used for altitude loss and/or speed

CAUTION: At altitudes aboveFL200, do not establish or main-tain a configuration with the flapsor slats extended.

5E-19 5E-20

Slat Extension LimitationUnder all conditions, slat (at leastoutboard) extension must precedeany flap extension. Upon retrac-tion, if the handle is directly setto CLEAN, a flap positionmicroswitch prevents theoutboard slats from retractingbefore the flaps.


Flight Controls

reductions, comprise four panels: twohinged on the upper surface of eachwing. Hydraulic System 1 powers theactuators that operate each panel. Themaximum deflection of each airbrakepanel is 50°. The airbrake panels haveno synchronization system to operatesimultaneously.

While there is no speed limit for oper-ating the airbrakes, they must beretracted below 500 ft AGL.

In the event of System 1 failure, theairbrakes are held stowed by a pres-sure-holding valve in the airbrake sys-tem. Should the failure occur with theairbrakes extended, the airbrakes blowdown, but do not stow.

Airbrake system electrical failure caus-es the airbrakes to stow (fail-safe) ifSystem 1 is pressurized.

Airbrake HandleThe airbrake handle (Figure 5E-21 )moves to two detents:� IN (all four panels retracted)

� EXT (all panels extended).

When selecting a change in the air-brake position, grip the handle to allowa quick response if airbrake asymmetrydevelops.

Airbrake IndicationsThe amber AIRBRAKE light on thegear/slat/flap configuration panel illu-minates when any panel is not in theretract position. Pressing the gear panelTEST pushbutton tests the amber AIR-BRAKE light.

The red AIRBRAKE light illuminates(refer to the Miscellaneous chapter, 5J,Annunciator Cross Reference):� on the ground if either throttle isadvanced for takeoff with airbrakesnot retracted� in the air if airbrakes and slats areextended at the same time.� on the ground or in the air when themain annunciator panel TEST push-button (Figure 5E-22 ) is pressed.

CAUTION: Airbrakes must befully retracted by 500 ft AGL.

5E-21 5E-22

NOTE: In early aircraft, the redAIRBRAKE light is on theupper portion of the pilot’sinstrument panel; on later mod-els, it is on the bottom row of themain annunciator panel.


Flight Controls

PreflightDuring the external preflight inspec-tion, check all control surfaces forsecurity and general condition.

During the cockpit inspection, ensurethat the slat/flap lever agrees with thecorresponding position indicator andthat the stabilizer trim is within take-off range. Refer to the Preflight chapter.

AbnormalProceduresThis section provides a brief discus-sion of flight controls abnormal pro-cedures. For a list of specific proce-dural steps, refer to the AFM or theSimuFlite Operating Handbook.

Pitch Trim MalfunctionIf a pitch trim malfunction occurs inflight, trim in the opposite direction,ensure the autopilot is disengaged, anduse emergency trim as required. If anemergency trim runaway occurs, usethe Auxiliary bus switch to stop thetrim. If both trim systems are affected,consult the AFM or the SimuFliteOperating Handbook for jammed sta-bilizer procedures.

Aileron or RudderTrim MalfunctionIf aileron or rudder trim runawayoccurs, trim the affected switch in theopposite direction to cancel the run-away. Maintain aircraft control andrestore attitude with aileron or rudder.Pull the circuit breaker of the faultysystem and ensure the autopilot is dis-engaged.

Yaw Damper MalfunctionThe illuminated amber YD FAILannunciator indicates a failure in the

yaw damper function. If a malfunctionoccurs in flight, disengage the yawdamper.

Slats Only Approachand LandingDuring a slats-only approach and land-ing, adjust landing distance and air-speed according to the AFM or Simu-Flite Operating Handbook procedures.

Q Unit FailureAbnormal feel force and an illumin-ated Q UNIT annunciator indicate fail-ure of the rudder Q unit or aileron/elevator Q unit. To determine whichsystem malfunctioned, check the RUDand AIL annunciators on the circuitbreaker panel behind the copilot’shead. If a failure occurs, disengage theautopilot/yaw damper; in the case ofelectric Q failure, also select emer-gency trim to trim nose-up past 4°.Using slow, smooth movements on theflight controls and cautious controlforce, adjust speed to no more than260 kts/0.76 M.

Electric Q High-Speed(Cruise) FailureThe electric Arthur Q unit’s most com-mon malfunction is that of high-speed(cruise) failure due to cold soaking.Cold can cause the system lubricantto gel and stop the unit’s motion. Acomparator-monitor illuminates theARTHUR Q annunciator, and the pilotlands the aircraft using the Arthur QJammed High-Speed procedure. Afterlanding, actuation of the LOW SPEED/NORMAL switch to LOW SPEEDdrives the Q actuator to LOW to allowa ferry flight to the maintenance base.The LOW SPEED/NORMAL switchis normally in the forward right sideof the passenger cabin.



Slat MalfunctionIllumination of the red slat light indi-cates the slats are not in agreementwith the slat/flap handle. With the slat/flap handle out of the CLEAN detent,select emergency slats. If the light turnsgreen, the problem was with the out-board slats and has been resolved.

If the light remains red, select CLEANwith the slat/flap handle. If the lightturns green, the problem is with theinboard slats; add 10 kts to VREF anduse flaps as necessary.

The light turns red and remains redwhen flaps are selected, which is a nor-mal indication in this condition. Followabnormal procedures in the AFM orthe SimuFlite Operating Handbook.

If the slat light remains red after select-ing CLEAN, follow the abnormal pro-cedures for a clean wing landing in the

AFM or the SimuFlite OperatingHandbook.

Stabilizer Trim Lighton TakeoffThe advisory STAB TRIM annuncia-tor illuminates on the ground if thestabilizer is out of the 4° to 8° aircraftnose-up range and either power leveris in the takeoff range. Do not take offuntil the condition is corrected.

Flap AsymmetryIllumination of the FLAP ASYMannunciator indicates a difference inflap positions, which causes asymet-ric lift. If flap asymmetry occurs, dis-engage the autopilot, control the air-craft with ailerons and aileron trim,add 10 kts to the approach speed, andland as soon as practical.

CAUTION: Once emergencyslats are selected, do not deselectbecause the emergency slat sys-tem does not provide retraction.


Flight Controls

DataSummaries

Flight ControlsPrimary Flight Controls

Power Source Power Source, Hydraulic System 1Flight controlsRudder Arthur Q

Hydraulic System 2Flight controlsYaw damper

Primary A busAileron and elevator Arthur Q unitsNormal stabilizer trimYaw damper

C busAileron trim

D busEmergency stabilizer trimRudder trim

Control Control yoke

Rudder pedals

Trim switches

Arthur Q low speed switch

Monitor AnnunciatorsQ UNITRUDAILSTAB TRIM (takeoff warning)GUST

Hydraulic pressure gages

Trim gages

Protection Arthur Q system overstress protection

Electric Arthur Q – automatic shutdown uponelevator/aileron Q failure

Torque bar (secondary AFU) centers flight controlsupon primary control linkage failure

Yaw damp limiting to 1.5° at high speed

Pitch trim limiting to 4° nose-up above 250 kts

Rudder actuator pressure limiting (1,850 PSIpermitted)

5E-34 For SimuFlite training only Falcon 10/100March 2001

CAUTION: The emergencyslats system does not provideretraction. Once selected, emer-gency slats should not beretracted.

CAUTION: Do not use flapsunless outboard slats (at least)are extended.

Secondary Flight Controls

Power Source Primary A bus – normal trim

Auxiliary D bus – emergency trim

Load-Shed D bus – rudder trim

Control Control yoke trim switches

AP DISC switch on control yoke

STAB EMG switch on pedestal

Monitor Trim gage

STAB TRIM annunciator

Audio warning (clacker)

Protection 4° cruise stop

Trim limit switches

Slats/Flaps System

Power Source Hydraulic System 1 (normal mode)Inboard/outboard slats

Hydraulic System 2 (emergency slats mode [standby])Outboard slats extension only

Hydraulic Systems 1/2 (stall protection mode [automatic]Outboard slats automatic extension

Primary B busEmergency slats

Auxiliary C busFlap position indicator system

Auxiliary D busFLAP CONTROL CBFlaps motor

Control SLATS/FLAPS handle

EMERG SLATS switch

Monitor Slat/flap position indicators

FLAP ASYM annunciator

Flap overspeed aural warning

Stall aural warning

Slat configuration agreement lights (red/green)

Airbrakes

Power Source Hydraulic System 1Primary A bus

Control Airbrake handle

Monitor AIRBRAKE configuration light (amber)AIRBRAKE annunciator (red)

NOTE: Angle-of-attack sensorsautomatically initiate slat exten-sion. Stall sensors control bothportions of the dual automatic(outboard only) slat extensionsystem.

Falcon 10/100 For SimuFlite training only 5F-1April 2000

Fuel System

Chapter 5F

The Falcon 10/100 is equipped with two independent fuelsystems; each supplies its respective engine. The Fuel Systemsconsist of the following major subsystems.

� A Fuel Storage System consisting of a tank in each wing aswell as left and right feeder tanks in the fuselage. The fuelvent system is considered a part of the storage system.

� A Fuel Distribution System that includes the pumps, valvesand plumbing required to transfer fuel through the aircraft tothe engines. Fuel filttration is considered part of the distri-bution system.

� A Fuel Quantity Indication System that consists of fuel indi-cators and annunciators, and the probes and sensors thatprovide them with information.

Refueling of the aircraft is accomplished through single-pointrefueling or through overwing gravity refueling.

5F-2 For SimuFlite training only Falcon 10/100April 2000


Fuel System

Table ofContents

Fuel System Schematic . . . . . . . . . . . . . . . 5F-6

Fuel Storage . . . . . . . . . . . . . . . . . . . . . 5F-7

Wing Tanks . . . . . . . . . . . . . . . . . . . . . . 5F-7

Feeder Tanks . . . . . . . . . . . . . . . . . . . . . 5F-7

Fuel Capacities . . . . . . . . . . . . . . . . . . . . 5F-7

Vent System . . . . . . . . . . . . . . . . . . . . . 5F-8

Wing Tank Vent Valves . . . . . . . . . . . . . . . 5F-8

Feeder Tank Vent Valve . . . . . . . . . . . . . . . 5F-8

Internal Venting . . . . . . . . . . . . . . . . . . . 5F-8

Fuel Distribution . . . . . . . . . . . . . . . . . . . 5F-9

Fuel Pumps . . . . . . . . . . . . . . . . . . . . . . 5F-9

Jet Pumps . . . . . . . . . . . . . . . . . . . . . . 5F-9

Electric Booster Pump . . . . . . . . . . . . . . . 5F-9

Fuel Filters . . . . . . . . . . . . . . . . . . . . . . 5F-10

Fuel Valves . . . . . . . . . . . . . . . . . . . . . 5F-10

Fuel Shutoff Valve . . . . . . . . . . . . . . . . . 5F-10

Fuel Manifold . . . . . . . . . . . . . . . . . . . 5F-10

Pressure Switch . . . . . . . . . . . . . . . . . . 5F-10

Fuel Control/Indicating System . . . . . . . . . . 5F-11

Fuel Quantity System . . . . . . . . . . . . . . . . 5F-11

Control . . . . . . . . . . . . . . . . . . . . . . . 5F-11

Feeder Tank Low-Level Warning . . . . . . . . . 5F-11

Booster Pumps . . . . . . . . . . . . . . . . . . 5F-11

Low Fuel Pressure . . . . . . . . . . . . . . . . . 5F-11

Fuel Temperature Indicating System . . . . . . . . 5F-12

Fuel Flowmeters . . . . . . . . . . . . . . . . . . . 5F-12

Fire Pull T-Handles . . . . . . . . . . . . . . . . . 5F-12


Fuel Transfer System . . . . . . . . . . . . . . . . 5F-12

Rear Tank Intercom System . . . . . . . . . . . . 5F-13

Crossfeed System . . . . . . . . . . . . . . . . . 5F-13

Normal Feed . . . . . . . . . . . . . . . . . . . . 5F-13

Pressure Fueling and Test Systems Schematic . . . 5F-14

Servicing and Procedures . . . . . . . . . . . . . 5F-15

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5F-15

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

Pressure Refueling . . . . . . . . . . . . . . . . 5F-15

Testing . . . . . . . . . . . . . . . . . . . . . . 5F-15

Pressure Refueling Procedures . . . . . . . . . 5F-15

Gravity Refueling (Standby Fueling) . . . . . . . . 5F-16

Defueling . . . . . . . . . . . . . . . . . . . . . . 5F-16

Abnormal Procedures . . . . . . . . . . . . . . . . 5F-16

Low Fuel Pressure . . . . . . . . . . . . . . . . . 5F-17

High Feeder Tank Level . . . . . . . . . . . . . . 5F-17

Low Feeder Tank Level . . . . . . . . . . . . . . 5F-17

Wing Tank Level Imbalance . . . . . . . . . . . . 5F-17

Low Fuel . . . . . . . . . . . . . . . . . . . . . . 5F-17

Data Summary . . . . . . . . . . . . . . . . . . . 5F-19


5F-6

Fuel System

FuelSystem


FIRE

1

PULL

FUELC°

60

40

20- 0 +

20

40

60

PPHx 100

20

15

10

5

0

PPHx 100

20

15

10

5

0FUEL USED

POUNDS

0 0 0 0

FUELC°

60

40

20- 0 +

20

40

60

FIRE

2

PULL

0

5

1015

20

25

30

FUEL QTY

LBSx100

HI

LO

OFF

TRANSFER

HI

LO

OFF

TRANSFER

HI

LO

OFF

TRANSFER

HI

LO

OFF

TRANSFER

OUTBOARD

INBOARD

FUEL QUANTITYPROBE

R FEEDERTANK QUANTITYPROBE

FUEL P2

B BUS

REARTANK

INTERCOM

BOOSTERPUMP

PRESSURE OPERATEDINTERCOM VALVE

BOOSTERPUMP

TOT

REAR TANK

HIGH LEVELFLOAT SWITCH(1,381 LBS.)

FUEL P1

A BUS

600 LBS.FLOATSWITCH(3 MINUTETIMER)

C BUS

B BUS

BOOSTERPUMP

BOOSTERPUMP

300 LBS. FLOATSWITCH

NOTE: PRESSURE REFUELING SHUTOFFCONTROLLED BY MECHANICAL FLOATS.*SHOULD BE TESTED PRIOR TO PRESSUREREFUELING.

A BUS

TELEFORCECABLE

LO. FUEL

CROSSFEED

B BUS

A BUS

0

5

1015

20

25

30

FUEL QTY

LBSx100

PRESSUREREFUELINGMANIFOLD

Fuel System

Falcon 10/100 For SimuFlite training only 5F-7March 2001

Fuel System

The fuel storage system consists of leftand right (L/R) structural wing tankswith corresponding fuselage feedertanks that supply the left and rightengines, respectively. If one of theengine supply systems fails, cross-feeding from these tanks is possible.

Wing TanksThe wing tanks are formed by sealingthe wing structure to create integralchambers; each is divided into threeindividual tanks:� a wing center section tank� an inboard tank� an outboard tank.

Access to each tank is providedthrough panels in the underside of thewing.

The tank in the wing center section is“L” shaped. Its outer limits are thewing junction plates, which are com-pletely sealed with the exception of anorifice provided for air circulation. Thejunction plate is fitted with three flap-per valves that prevent fuel from flow-ing outward away from the center sec-tion tanks.

The inboard tanks lie between the junc-tion plate at the wing root and rib 5;the outboard tanks lie between ribs 5and 13. Rib 5 is fitted with two one-way flapper valves to allow fuel to

flow inward toward the center sectiontanks while restricting flow outward.

This particular design improves air-craft balance and response by restrict-ing fuel movement within the tank andusing outboard tank fuel first; thisallows the aircraft to maintain a for-ward center of gravity.

Feeder TanksTwo integral feeder tanks are in therear fuselage between the passengercabin and the aft compartment.Basically one unit, the structure isdivided into two tanks by a partitionthat runs along the centerline of theaircraft. A port in the upper forwardarea provides for interconnection oftank air venting and tank pressuriza-tion. Each feeder tank supplies its asso-ciated engine.

Access to the feeder tanks is via doorsin the bottom of each feeder tank. Eachfeeder tank bottom is fitted with a drainvalve (Figure 5F-1 ). Enclosed airspaces surrounding the fuel tanks arevented to prevent accumulation of fuelor fuel vapor.

Fuel CapacitiesFuel capacities are based on a fuel den-sity of 6.7 lbs per US gallon at stan-dard temperature 59°F. The wing tanks(left and right) have a capacity of 235

Fuel Storage

5F-1

Usable Fuel LimitationThe total usable fuel quantity is5,912 lbs (882 US gal).

Fuel Specification LimitationFuel used must conform to thespecifications table in AFMSection 1, Limitations.

Fuel Adjustment LimitationIf the type fuel specified in thespecifications table in AFMSection 1, Limitations, and usedin the aircraft is changed, or ifthese fuels are mixed, the appro-priate adjustment must be madeat the fuel computer inconformance with instructions inthe approved MaintenanceManual.


US gallons (1,575 lbs); left and rightfeeder tanks each have a capacity of206 US gallons (1,381 lbs). The amountof fuel left in the tanks when fuelquantity indicators reach zero is notsafely usable in all flight conditions.

Vent SystemA fuel vent system provides continu-ous ram air pressure to the wing tanksand slightly negative pressure to thefeeder tanks while the aircraft is inflight to ensure positive fuel transferduring all flight conditions. In addi-tion, the vent system prevents nega-tive pressure and overpressurizationwhile the aircraft is on the ground.

Flush-mounted, NACA-type under-wing ram airscoops admit air into thewing tanks through tubing (Figure 5F-2); the feeder tanks receive ventilationthrough slotted airscoops on either sideof the vertical stabilizer (Figure 5F-3 ).

Wing Tank Vent ValvesThe wing center section tank and twowing tanks are vented throughleakproof wing spar boxes in the wing-tips. A fuel stopper valve controls vent-ing. One valve, which is in the out-

board wing tank at rib 13, is at thehighest point in the wing tip end; itopens into the NACA vent compart-ment. The other valve, which is in theinboard wing tank between rib 0 andthe junction plate, is the second highpoint of the wing; it is connected tothe NACA vent compartment througha tube inside the wing.

Feeder Tank Vent ValveEach feeder tank is vented through atube that departs the feeder tank at thefront and opens into a vent box in thevertical stabilizer. Two reverse ramairscoops maintain a negative pressureto assist wing fuel transfer. The col-lector tank air vents also allow air to bevented during nose-up attitudes; thelowest points of the air vent tubesincorporate ball valves to prevent fuelfrom escaping.

Internal VentingAir between the internal wing tanks iscrossbled by apertures in the top edgesof the junction plate and rib 5. Aircrossbled between the feeder tanks iscarried out through ports in the topedge of the dividing central wall.

5F-35F-2


Fuel System

The fuel distribution system consistsof two independent systems: one foreach engine. Each system consists ofthe following components:� jet pumps� electric booster pump� fuel filters� fuel shutoff valve� jet pump solenoid� fuel manifold� pressure switch.

Fuel distribution is managed by thefuel control panel on the overheadpanel.

Fuel PumpsEach of the two distribution systemscontains two pumps: a jet pump andan electric booster pump. The pumpsreceive power from the followingsources:� No. 1 jet pump solenoid – A Bus� No. 2 jet pump solenoid – B Bus� No. 1 booster pump – A Bus� No. 2 booster pump – B Bus.

Jet PumpsThe jet pump inlet near the lowestpoint of the wing (in the center sec-tion tank) is submerged in fuel untilthe tanks are virtually empty. The jetpump has no moving parts and oper-ates on the venturi principle (Figure5F-4). Installed in the feeder tank, ituses motive flow from the boosterpump to transfer fuel from the wingtanks to the feeder tanks.

Electric Booster PumpAn electric booster pump is in eachdistribution system; one is in the bot-tom of each feeder tank. The boosterpump is a submerged type that is usedto provide fuel under pressure to thefuel manifold. The booster pumpsreceive power from the associated (Aor B) bus through the BOOSTER CBs.

A pressure drop in either booster pumpsystem is indicated by illumination ofa FUEL P1 or FUEL P2 annunciatoron the Failure Warning Panel. See FuelControl/Indicating System, this chap-ter, for details.

FUEL SUPPLY WING TANKSTRUCTURE

LOW PRESSURELINE

MOTIVE FLOWLINE

5F-4

FuelDistribution


Fuel FiltersLow pressure fuel filters between themanifold and the jet pump protect thejet pump from contamination.Replacement of the filters is consid-ered a maintenance procedure. If eitherfilter becomes clogged in flight, theassociated jet pump can be shut offand fuel gravity feeds from the wingtanks to the feeder tank.

Fuel ValvesFuel Shutoff ValveEach distribution system incorporatesa fuel shutoff valve in the engine feedline. The shutoff valves are two-posi-tion (open and closed) ball-and-seattype; they are mechanically operated.The valves allow the pilot to shut offthe supply of fuel to the engine if anengine fire occurs. Each FIRE PULLT-handle on the center instrumentpanel (Figure 5F-5 ) closes its respec-tive valve when the handle is pulledout; the valve opens when the handleis pushed in.

Fuel ManifoldA fuel manifold downstream of eachbooster pump inside the feeder tankincorporates a filter to protect the jetpump from contamination. Two checkvalves are also in the fuel manifold:one in the booster pump discharge line

and the other in the feeder tank inter-connection line.

A jet pump solenoid valve for wing-to-feeder tank fuel transfer and thefuel booster pump low pressure warn-ing switch are on the fuel manifold.These items, in addition to the cross-feed valve actuator, can be removedfrom outside the aircraft without open-ing the feeder tanks.

The right fuel manifold connects theright booster pump discharge line to the:� right engine feed line� crossfeed valve and line� right transfer jet pump nozzle� feeder tank interconnection valve line.

The left fuel manifold connects the leftbooster pump discharge line to the:� left engine feed line� crossfeed valve and line� left transfer jet pump nozzle� feeder tank interconnection valve line.

Pressure SwitchA pressure drop in either booster pumpsystem is indicated on the FailureWarning Panel by a FUEL P1 orFUEL P2 warning light; the light illu-minates when the pump dischargepressure is below 4.6 ± 0.3 PSI. Thepressure drop is sensed by two pres-sure switches on the fuel manifoldsoutside the feeder tanks.

5F-5


Fuel System

Fuel QuantitySystemFuel quantity in the Falcon 10/100 issensed by eight capacitance-type 28VDC probes, one in each outboard tank,inboard tank, center section tank, andrear feeder tank. Information is trans-mitted through a solid-state amplifier(bottom of rear feeder tank) to the twoFUEL QTY indicators on the centerinstrument panel (Figure 5F-6 ). Eachindicator is calibrated in pounds andreads up to 3,100 lbs.

The A and B buses supply DC powerto the left and right quantity indicat-ing systems, respectively.

ControlA two-position (TOT/REAR) fuelquantity selector switch under the fuelquantity indicators on the center instru-ment panel control fuel quantity indi-cation. The switch is used to select therespective probes to determine eitherthe overall weight of the fuel per tankgroup (TOT) or the weight of the fuelin each feeder tank (REAR). TheTOT/REAR switch normally remainsin the REAR position to monitor fueltransfer.

Feeder Tank Low-LevelWarningFloat-type level switches provide lowlevel warning for each of the feedertanks; each switch controls a commonannunciator on the Failure WarningPanel. Illumination of the LO FUELannunciator indicates fuel remainingin at least one of the feeder tanks isless than 300 lbs.

Booster PumpsTwo-position (ON/OFF) BOOSTERPUMPS switches on the engine andfuel control section of the overheadpanel (Figure 5F-7 ) are turned onprior to engine starting to providepositive fuel pressure to the engine-driven fuel pumps. The BOOSTERPUMPS switches are turned off afterengine shutdown.

Low Fuel PressureFUEL P1 or FUEL P2 annunciatorson the Failure Warning Panel illumi-nate when the pump discharge pres-sure is below 4.6 ± 0.3 PSI. The annun-ciator illuminates when the pressureswitch on the fuel manifold outside thefeeder tank in either booster pump sys-tem detects low pressure, or when theassociated booster pump is off.

5F-6

Fuel Control/IndicatingSystem

5F-7


Fuel TemperatureIndicating SystemA fuel temperature indicating systemenables inflight monitoring via a tem-perature sensor inside the right feed-er tank. The temperature indicator onthe overhead panel (Figure 5F-8 ) indi-cates from -60°C to +60°C.

Fuel FlowmetersTwo fuel flowmeters on the centerinstrument panel, powered by load-shed C bus, display fuel flow for theirrespective engines (Figure 5F-9 ). TheFUEL USED FLOWMETER (Figure5F-11) indicates total fuel used. Thissystem receives information from thefuel flowmeters.

Fire Pull T-HandlesIn case of an emergency such as anengine fire, pulling the red FIRE PULLT-handle shuts off the fuel valve in theengine pylon line. This isolates thefeeder tank from the engine.

Fuel TransferSystemThe two guarded, three-position (HI/LO/OFF) TRANSFER switches(Figure 5F-10 ) function as follows.� OFF. Solenoid valve is energizedclosed.� HI . Jet pump solenoid valve is de-energized open until the fuel level inthe affected feeder tank reaches 1,381lbs of fuel; the HI float switch thenenergizes the jet pump solenoid valveto the closed position, and fuel trans-fer from the center wing to feeder tankterminates. This position is selectedfor overwing gravity fueling.� LO . This position controls a three-minute time delay relay through theLO float switch to control the jet pumpsolenoid to maintain 600 lbs in theassociated feeder tank. During eachcycle (less than 600 lbs), the fuel levelin the feeder tank decreases for threeminutes. The time delay then de-ener-gizes and opens the jet pump motive

5F-9

5F-8


Fuel System

flow solenoid; transfer from the centerwing section to its feeder tank resumesuntil the 600 lbs level is reached. Thisposition is used during flight. If anelectrical failure should occur, thetransfer jet pumps would fail open.

Rear Tank IntercomSystemThe two-position, rotary REAR TANKINTERCOM switch between theTRANSFER switches (Figure 5F-10 )balances fuel between the feeder tanks.Rotating the REAR TANK INTER-COM switch 90° counterclockwiseopens the solenoid valve and admitsbooster pump pressure to open theintercom valve. This action allows fuelto transfer between the feeder tanks.Booster pump pressure from eitherfeeder tank is required to open theintercom valve; the valve fails closed.

Crossfeed SystemRotating the two-position X-FEEDswitch between the booster pumpswitches (Figure 5F-10 ) 90° clock-

wise opens the crossfeed valve andpermits either booster pump to supplypressurized fuel to both engines. Thissituation could arise if a booster pumpfails or is turned off. Crossfeed alsohelps balance wing fuel loads. Theswitch and valve are powered by load-shed C bus. The valve remains in thelast position selected in the event offailure.

Normal FeedIn the normal phase of fuel manage-ment, each engine is supplied fuel atbooster pump pressure from its asso-ciated feeder tank. Fuel travels throughthe open fire shutoff valve in the pylonif the following configuration is estab-lished:� TRANSFER switches in LO� REAR TANK INTERCOM valveclosed� booster pumps on� crossfeed valve closed.

Float switches in the feeder tanksmaintain automatic level control.

5F-10

5F-11

STRAINER

WING CENTER SECTION TANKS

LEFT FEEDER TANK RIGHT FEEDER TANK

TRANSFERJET PUMP

FLOAT

PERFORATEDBOX

REFUELING VALVETEST COCK

REFUELINGCOUPLING

PRESSURE REFUELING SYSTEM

REFUELING VALVE TEST SYSTEM

TRANSFER SYSTEM


Pressure Fueling and Fuel Valve Test Systems


Fuel System

PreflightDuring preflight, the following itemsare checked (See Preflight, Chapter3A, for details):� fuel quantity� fuel sumps drained, closed, and notleaking� gravity fueling ports caps secure;no evidence of tampering� no fuel leaks or damage on wingunderside� all switches and the fuel tank ventcontrol on single point fueling panelpositioned properly� single point fueling panel doorclosed; both latches secure.

RefuelingPressure RefuelingThe normal refueling mode of the air-craft is through a single-point pressurerefueling receptacle access on the rightaft section of the aircraft (Figure 5F-12). Pressure from the refueling truckis not to exceed 50 PSI. At 50 PSI (seeRefueling Pressure and Fuel ImbalanceLimitations, next page), the aircraft ispressure refueled in approximately 13minutes.

Refueling is through a 11/4 inch coup-ling connected to two fueling valves.The refueling valves are hydraulically-operated by a float valve in the top ofthe feeder tanks. The fueling valvesclose and pressure fueling is stopped assoon as the feeder tanks are filled orwhen the complete fueling is simu-lated by means of the test system.

Testing

During the initial stage of pressurerefueling, the automatic shutoff sys-tem is tested by opening the test valvenear the refueling adapter. The openvalve directs fuel to the level controlfloat valves to raise the floats, simu-late a tank-full condition, and close thefueling valves. The refueler hose stiff-ens and the refueler flowmeter showszero flow. The test valve is returned tonormal and refueling is continued.

Pressure RefuelingProcedures

When the amount of fuel required isdetermined and a fire extinguisher isavailable, the fuel truck is groundedto the ramp and then grounded to theaircraft’s right main strut. Finally, thefueling nozzle is grounded to the air-craft connector. The truck should deliv-er fuel between 30 and 50 PSI.

Servicing andProcedures

5F-12

CAUTION: If a new type offuel or a mixture of fuel is used,the engine computer must beadjusted accordingly to preserveboth the starting characteristicsand the acceleration and decel-eration characteristics of theengine.

NOTE: The fuel remaining inthe tanks when the fuel quantityindicators read zero is not usablein flight. The amber area of thefuel quantity indicators, 0 to 300lbs, is a LO FUEL caution range.

NOTE: See Quick Referencefor authorized fuel additives.

NOTE: The density of fuel orfuel mixture used is set on theengine computer by means of the10-position adjusting knob cover-ing densities from 0.61 to 0.85.

5F-16 For SimuFlite training only Falcon 10/100March 2001

Accomplish the refueling test by open-ing the test valve. If fueling stops, thetest is good. Close the test valve andresume refueling. Fill to the desirednumber of gallons/pounds. The re-fueling stops automatically when thetanks are full.

Disconnect the refueling nozzle, andcheck that the sealing valve is seatedand that there are no fuel drips.Disconnect the fueling nozzle, thenclose and latch the fueling access door.Disconnect the grounding wire.

Gravity Refueling(Standby Fueling)Ground power is required for thisprocedure.

Refueling can also be accomplishedthrough a gravity, over-wing system.Electrical power (GPU, if available)is required to operate the booster pumpsand transfer switch (HI position). Aqualified crewmember or maintenanceperson should be in the cockpit.Refueling is accomplished through thegravity refueling port in the top surfaceof each outboard wing tank.

During gravity refueling, the feedertank is partially filled by direct com-munication (approximately 600 lbs inzero pitch attitude).

The aircraft can be partially or totallyfueled by gravity. Connect the ground-ing connector (one per wing) near thefiller port. Make sure that the tank toaircraft and refueling nozzle to aircraftstatic grounding connections are made.

Turn on aircraft electrical power andverify that the wing contains a mini-mum of 132 gallons (885 lbs). Turnthe booster pumps switches to ON andturn the transfer pumps to HI. Confirmthat the feeder tank level rises.

As fuel pumps into the wing tank, fueltransfers from both wing tanks to the

feeder tanks. If the feeder tanks are notfull prior to the wing tank indicatingfull, have the refueler stand by untilthe feeder tanks are full, and then topoff the wing. Turn the booster pumpsswitches to OFF; turn the transferpumps to LO.

Select TOT on the fuel quantity indica-tor to determine any imbalance of fuel;check the total fuel quantity and addfuel as required. The fueling time foreach side is approximately 30 minutes.

If GPU power is not available, the fol-lowing procedure is recommended toconserve batteries.� Wing Tanks – FILL� Refueling – DISCONTINUE� One Engine – START� Booster Pumps – ON� Transfer Switch – HI� Engine – OPERATE (until feedertanks are full)� Transfer Switches – LO� Engine Shutdown� Booster Pump – OFF� Wing Tanks – TOP OFF

DefuelingDefueling is normally accomplishedusing the feeder tank drain valves afterswitching the transfer system on totransfer all wing tank fuel. Accomplishthe proper grounding (tank-to-aircraft,coupling-to-aircraft) and drain fuel.

AbnormalProceduresFollowing is a brief discussion of whathappens in the fuel system duringabnormal conditions. For a list of spe-cific procedural steps in abnormal con-ditions, please refer to the AFM or theSimuflite Operating Handbook.

CAUTION: If testing is notsatisfactory, stop refueling andtroubleshoot as required.

Refueling PressureLimitationMaximum pressure for singlepoint refueling is 50 PSI;minimum pressure is 30 PSI.

Fuel Imbalance LimitationNo limit for fuel imbalance(do not deliberately fuel only onewing).


Fuel System

Low Fuel PressureIllumination of the amber FUEL P1 orFUEL P2 annunciator indicates one ofthe microswitches downstream of abooster pump detected a pressure drop(below 4.6 PSI); a booster pump maybe inoperative.

The engines are supplied fuel pressurefrom the operative booster pump bypositioning the X-FEED knob toOPEN and the inoperative boost pumpswitch to OFF.

If fuel pressure annunciator extin-guishes, position the REAR TANKINTERCOM switch OPEN and mon-itor the fuel level in the feeder tanks(fuel level remains equal). If the fuelpressure annunciator remains illumi-nated, position the X-FEED knob toCLOSE and do not exceed BoosterPump Inoperative envelope outlinedin the Falcon 10/100 AFM, Section 3.The cause may be a fuel leak that couldrapidly result in an engine flameout.

High Feeder Tank LevelHigh feeder tank level indicates aninoperative or de-energized electro-valve, or a malfunctioning float switchassociated with low fuel level position.Since jet pump-assisted fuel transferis continuous, position the associatedTRANSFER switch to OFF.

If the fuel level stops rising, the floatswitch is inoperative and fuel is trans-ferred by gravity. If necessary, oper-ate the TRANSFER switch to main-tain adequate feeder tank fuel duringhigh fuel usage periods.

If the fuel level continues to increase,the jet pump electrovalve has failedopen. Position REAR TANK INTER-COM switch to OPEN to connect theleft and right feeder tanks. Position theX-FEED knob to OPEN to ensure pos-itive pressure to the engine-driven fuelpumps. Position the opposite sideTRANSFER switch to OFF to ensurefuel is used from the malfunctioning

side. Position the high side BOOST-ER PUMP switch to OFF to lessenpressure to the malfunctioning transferpump.

Low Feeder Tank LevelA low feeder tank level indicates aninoperative float switch associated withlow fuel level position or an electro-valve stuck closed. Position theTRANSFER switch to HI to de-ener-gize the electrovalve associated withthe low level condition.

If the fuel level does not rise, comparethe left and right gages after selectingthe TOT position to make sure the faultis not a feeder tank leak. If there is noevidence of a leak, position the REARTANK INTERCOM switch to OPENand reduce airspeed to below MACH0.80. If the transfer system fails, fuel istransferred by gravity.

Wing Tank LevelImbalanceAsymmetric fuel levels between thetwo wing tanks are indicated by asym-metric TOT gages and abnormalaileron trim position.

Position the X-FEED knob to OPEN toensure positive pressure to the engine-driven fuel pumps. Position REARTANK INTERCOM switch to OPENto connect the left and right feedertanks. Position the BOOSTER PUMPswitch on the low TOT gage side toOFF to lessen pressure to the mal-functioning transfer pump. Positionthe TRANSFER switch on the lowTOT gage side to OFF to ensure fuelusage from the high wing tank side.Monitor the fuel quantity gages andland normally.

Low FuelThe amber LO. FUEL annunciator onthe Failure Warning Panel illuminateswhen either feeder tank contains lessthan 300 lbs.

CAUTION: 150 lbs of fuel areunusable on each side for a pitchattitude of approximately 5°.

JP4/JP5 LimitationJP4 and JP5 is approved but notrecommended due to fuel heaterpossibly causing a vapor lockresulting from a rapid powerreduction.

Overwing RefuelingLimitationMinimum 132 gallons in wingprior to transferring fuel to feeders.


Fuel System

DataSummary

Fuel System

Power Source A busLeft FUEL QTY gageNo. 1 boost pumpNo. 1 jet pump

B busRight FUEL QTY gageNo. 2 boost pumpNo. 2 jet pump

C busIntercom valveCrossfeed valveFuel counterFuel flow gages

Distribution L/R wing tanks

L/R feeder tanks

NACA scoops (venting system)

Boost pumps

Jet pumps

Intercom valve

Crossfeed valve

Gravity feed

Suction feed

Single point refueling system

Gravity refueling

Pylon valves

Control TRANSFER switches

REAR TANK INTERCOM knob

BOOSTER PUMP switches

X-FEED knob

TOT/REAR TANK switch

Feeder tank high/low level float switches

Fueling valves/float valves

Fueling test valve

Monitor FUEL QTY gages

FUEL °C gage

FUEL P1/P2 annunciators

LO FUEL annunciator

Protection Feeder tank high/low level switches

NACA vents

Transfer switches

Falcon 10/100 For SimuFlite training only 5G-1April 2000

HydraulicSystem

Chapter 5G

Hydraulic power for the Falcon 10/100 is provided by two mainhydraulic systems, henceforth referred to as System 1 andSystem 2.

System 1 is powered by the left engine-driven hydraulic pumpand may be powered by the standby electric pump. System 1supplies pressurized hydraulic fluid to:

� one barrel of each dual-barrel actuator for the rudder,elevators, and ailerons (see Flight Controls chapter)

� rudder Arthur Q-unit (see Flight Controls chapter)

� normal (inboard and outboard) slats (see Flight Controlschapter)

� automatic (outboard) slats (see Flight Controls chapter)

� landing gear operation (see Landing Gear chapter)

� normal braking system (see Landing Gear chapter)

� airbrakes (see Flight Controls chapter)

� gust damper (see Flight Controls chapter).

System 2 is powered by the right engine-driven hydraulic pumpand does not incorporate a standby pump. System 2 suppliespressurized hydraulic fluid to:

� one barrel of each dual-barrel actuator for the rudder,elevators, and ailerons (see Flight Controls chapter)

� yaw damper (see Avionics chapter)

� outboard slat emergency system (see Flight Controls chapter)

� nosewheel steering (see Landing Gear chapter)

� emergency braking system (see Landing Gear chapter)

� thrust reversers (see Miscellaneous chapter).

� parking (emergency) brake system (see Landing Gearchapter).

5G-2 For SimuFlite training only Falcon 10/100April 2000


5G-4

HydraulicSystem andReservoir

Hydraulic System

Hydraulic System Schematic . . . . . . . . . . . . 5G-4

Hydraulic Reservoir Schematic . . . . . . . . . . . 5G-4

Hydraulic System Components . . . . . . . . . . . 5G-5

Control and Indication Panels . . . . . . . . . . . . . 5G-5

Reservoirs . . . . . . . . . . . . . . . . . . . . . . . 5G-6

Engine-Driven Pumps . . . . . . . . . . . . . . . . . 5G-7

System Filters . . . . . . . . . . . . . . . . . . . . . 5G-7

Pressure Relief Valves . . . . . . . . . . . . . . . . 5G-7

Accumulators . . . . . . . . . . . . . . . . . . . . . 5G-7

Pressure Transmitters . . . . . . . . . . . . . . . . 5G-8

Standby Pump . . . . . . . . . . . . . . . . . . . . 5G-8

Ground Service Receptacle . . . . . . . . . . . . . . 5G-8

Preflight and Procedures . . . . . . . . . . . . . 5G-11

Servicing and Preflight . . . . . . . . . . . . . . . . 5G-11

Abnormal Procedures . . . . . . . . . . . . . . . . 5G-11

Unwanted Operation of the Standby Pump . . . . 5G-11

Depressurization of Reservoir . . . . . . . . . . . 5G-11

Emergency Procedures . . . . . . . . . . . . . . . 5G-11

Loss of Left Engine Hydraulic Pump . . . . . . . . 5G-11

Loss of No. 1 Hydraulic System . . . . . . . . . . 5G-11

Loss of No. 2 Hydraulic System . . . . . . . . . . 5G-12

Loss of Both Hydraulic Systems . . . . . . . . . . 5G-12

Data Summary . . . . . . . . . . . . . . . . . . . 5G-13

Table ofContents

HYDRAULICRESERVOIR

BRAKEL

BRAKER

ON

OFF

TEST ON

OFF

HYD ST BYPUMP

THRUSTREVERSER(OPTIONAL)

HYD PRESSURIZATION

SYSTEM 1

SYSTEM 2

RETURN

HYD RESERVOIR

4

2

1

0

4/4

1/2

0

HYDR

QTY PSI 1000

OUTBOARDSLATS

RUDDERSERVO CONTROL

A BUS

NO. 1HYD PUMP

ST. BYPUMP

HYD. 1

21 PSIA FROM LPBLEED AIR

NO. 2HYD PUMP

ELEVATORSERVO CONTROL

AILERONSERVO CONTROL

RUDDERARTHUR

Q

YAWDAMPER

INBOARDSLATS STEERING

OUTBOARDSLATS

AIRBRAKES EMERBRAKE

LANDINGGEAR

NORMALBRAKES

ANTI-SKID

P. BRAKE

ANTI-SKIDSWITCH

C BUS

HYD 2

PRESSURETRANSMITTER(0 TO 3,000 PSI)

ACCUMULATOR

SYSTEM 2LOW PRESSURESWITCH(1,500 PSI)

RUDDERPRESSUREREDUCER(1,850 TO 2,000 PSI)

EMER. BRAKEPRESSURESWITCH(1,200 PSI)

EMER. BRAKEACCUMULATOR

PRESSURETRANSMITTER(0 TO 3,000 PSI)

STANDBY PUMPCONTROL PRESS.SWITCH(1,600 TO 2,150 PSI)

RUDDER PRESS.REDUCER(1,850 TO 2,000 PSI)

SYSTEM 1 LOWPRESSURE SWITCH(1,500 PSI)

ACCUMULATOR

HYD. 1 HYD. 2 ST. BYPUMP

HYD. TK1 HYD. TK2 P. BRAKE

THRUSTREVERSERACCUMULATOR

B BUS

MAIN BUS


PRESSURERELIEF CAP

PRESSURIZATION VALVE

VENT

FILLING CAP

LP AIRPRESSURESWITCH

SIGHT GAGE

MAX LEVEL(L/G EXTENDED)

MIN. LEVEL

CHLORINATEDSOLVENTARRESTERELEMENT

SUCTION PORT(ONLY 1 PORT IS USED FORTHE RH RESERVOIR)

DRAIN COCK

SUCTION PORT

PRESSURERELIEF SCREW

MAGNETIC FLOATQUANTITYTRANSMITTER

Hydraulic ReservoirHydraulic System


Hydraulic System

System 1 and System 2 hydraulic sys-tems operate independently of eachother using MIL-H-5606 B or C fluid.If the left engine-driven pump fails,System 1 incorporates an electricstandby hydraulic pump to power allcomponents of System 1.

Both systems are similar and includeall components required for storage,filtering, pressurizing, and monitoringof the hydraulic fluid.

System 1 and System 2 hydraulic com-ponents include the following:� control and indication panels� reservoirs� engine-driven pumps� system filters� pressure relief valves� check valves� accumulators� pressure transmitters� standby pump� electric pump control pressureswitches (System 1 only)� ground service receptacles.

Control andIndication PanelsThe fluid level (QTY) and pressureindicators (PSI x 1000) for System 1and System 2 (Figure 5G-1 ) are on thepilot’s side of the lower center instru-ment panel. The standby pump con-trol (STBY PUMP) three-position tog-gle switch (NORM/ON/OFF) is on thepilot’s side of the upper center pedestal.

The amber hydraulic HYD 1, HYD 2,GUST, HYD. TK1, HYD. TK2, P.BRAKE, and ST.BY PUMP annunci-ators on the master warning panel arechecked with the TEST pushbutton.

Power sources for illumination of theannunciators are as follows:� P.BRAKE, HYD. TK1, HYD. TK2,HYD 1, HYD 2, GUST, and ST.BYPUMP – A bus� System 1 hydraulic QTY indicator –B bus� System 2 hydraulic QTY indicator – B bus� System 1 pressure indicator – B bus� System 2 pressure indicator – B bus.

HydraulicSystemComponents

5G-1


ReservoirsSystem 1 (Figure 5G-2 ) and System2 reservoirs in the aft compartmentsupply hydraulic fluid to their respec-tive engine-driven pumps and receivesystem return fluid after utilization.The reservoirs have a capacity of 1.58gallons and are pressurized automati-cally with engine low pressure (LP)air through a pressure reducing/regulating valve when either or bothengines are operated.

Normal tank pressurization is main-tained at 21.0 PSIA. If pressure dropsbelow 16 PSIA, a pressure switch inthe upper section of the reservoir illu-minates the respective HYD. TK1(System 1) or HYD. TK2 (System 2)annunciator. Illumination of eitherannunciator requires the crew to mon-itor the fluid quantity/pressure indica-tors for fluctuation of system pressure.If pressure starts fluctuating, descendto an altitude below 20,000 ft as soonas conditions permit to avoid cavita-tion of the engine-driven pumps.

Additional components of the reser-voirs include:� a pressurization unit� a sight glass on the side of the reser-voir for visual checks of the fluid level� a magnetic float and quantity trans-mitter for fluid level readings� filters for the filling port and for thesupply to the pumps.

The pressurization unit consists of:� one check valve to prevent thereturn of fluid/air into the air supplyline� one pressure relief valve that pre-vents overpressure in the tank at 26.0PSIA� one pressure relief screw to relievethe pressure prior to removing the fill-ing cap� one vacuum valve set to preventnegative pressure in the tank at 0.38PSI.

5G-2

Falcon 10/100 For SimuFlite training only 5G-7March 2001

Hydraulic System

Engine-DrivenPumpsEach engine drives a self-regulatingvariable displacement pump (Figure5G-3) that produces approximately3,000 PSI of pressure. The pumps drawfluid for distribution from their respec-tive system reservoirs and return bypassfluid to the reservoir inlet. The left andright engines drive the pumps forSystem 1 and System 2 respectively.

Pump pressure output is monitoredwith a pressure switch downstream ofthe pump. This pressure switch illum-inates a HYD 1 or HYD 2 annuncia-tor when the pressure delivered by theassociated pump drops below 1,500PSI.

System FiltersBoth hydraulic systems have the fol-lowing additional filters:� discharge line filter (each pump)� reservoir filling filter

� pump supply filter (aircraft withSB 29.007)� pump pressure line filter� engine pump bypass filter.

Pressure ReliefValvesThe aft compartment contains threepressure relief valves. The valves protectthe respective system in the event ofoverpressure caused by failure of theregulation system of the correspondingself-regulating pump. The valve opensat approximately 3,630 PSI.

AccumulatorsEach system includes a nitrogen-charged accumulator downstream ofthe engine-driven pumps (Figure 5G-4). The accumulators provide a powerreserve for the hydraulic system andare preloaded with a charge of 1,450PSI. The accumulators include aninflating valve (for servicing) and apressure gage graduated in PSI.

5G-3 5G-4

NOTE: SB 29.007; HydraulicPower – Intake Filter in theReservoirs – Installation of aPaper Filter.

5G-8 For SimuFlite training only Falcon 10/100March 2001

An additional accumulator on eachpressure line dampens pressure surges;it is preloaded with a charge of 1,450PSI. Aircraft S/N 92 and subsequentor aircraft with SB 29.009have a pre-load of 850 PSI.

System 2 includes an additional accu-mulator for the thrust reverser system(see Miscellaneous chapter) and anaccumulator for the parking and emer-gency brake system (see Landing GearSystems chapter).

PressureTransmittersPressure transmitters in the aft com-partment for each hydraulic systemsignal the pressure indicators to dis-play respective system pressure.

Standby PumpSystem 1 incorporates an electricstandby hydraulic pump (Figure 5G-5 )

to power all components of the systemif the left engine-driven pump fails.

Electric power for the standby pumpis obtained from the Main bus througha current limiter and a power relay.The power relay is controlled by thestandby pump switch using powerfrom the B bus.

Switch position and system functionfor the standby pump are shown inTable 5G-A .

Ground ServiceReceptacleThe ground service receptacles forSystem 1 and System 2 are inside theaft compartment on the hydraulicequipment rack (Figure 5G-6 ). Theground receptacles are provided to sup-ply the aircraft with hydraulic powerfrom the hydraulic mule (cart). Thereceptacles are a self-sealing type andprotected with dust covers.

5G-5 5G-6

NOTE: SB 29.009;Replacement of ABEX Pumpsby Vickers Pumps.


Hydraulic System

Table 5G-A; Switch and Function

Standby SwitchPosition

AircraftConfiguration Function

ON or NORM On the ground If pressure drops below 1,600PSI, standby pump cyclespressure between 1,600 PSIand 2,150 PSI.

NORM In flight If slat/flap control handle is setto a position other than cleanor if the inboard slats are notretracted and pressure inSystem 1 falls below 1,600 PSI,standby pump cycles pressurebetween 1,600 PSI and 2,150PSI.

ON In flight Standby pump cycles pressurebetween 1,600 PSI and 2,150PSI if engine-driven pumppressure is less than 1,600 PSI.


Hydraulic System

Servicing andPreflightThe fluid reservoirs are serviced withMIL-H-5606 B or C hydraulic fluidusing a hydraulic mule through theground service connections. The reser-voirs can be serviced through the fillerinlets on the reservoirs.

Both hydraulic system reservoir areasand all components in the aft com-partment should be checked as part ofthe preflight inspection (see Preflightchapter).

AbnormalProceduresRefer to the AFM or the SimuFliteOperating Handbook for specific pro-cedures. The following is a discussionof general procedures and their effecton system operation.

Unwanted Operationof the Standby PumpThe standby pump operating time ina normal cycle is less than 30 seconds.If the ST. BY PUMP light illuminates,the pump must be turned OFF to pre-vent possible failure or fire.

Depressurizationof ReservoirIf air pressure in a hydraulic reservoirdrops below 16 PSIA, the associatedHYD. TK1 or HYD. TK2 annuncia-tor illuminates. If hydraulic pressurebegins to fluctuate on the indicatorpanel, descend to 20,000 ft as soon asconditions permit to prevent pumpcavitation.

EmergencyProceduresLoss of Left EngineHydraulic PumpThe illumination of HYD 1 annuncia-tor may indicate the loss of the No. 1engine-driven pump. The pressureswitch downstream of the pump sens-es a pressure of less than 1,500 PSIand signals the annunciator of a lowpressure condition. Check System 1QTY indicator.

With the left engine-driven pumpfailed and the standby pump operat-ing, all System 1 components operate,but at a slower speed.

Loss of No. 1Hydraulic SystemThe illumination of the HYD 1 annun-ciator along with the indicator panelfluid quantity reading empty (redrange) indicates the loss of System 1fluid. The pressure switch downstreamof the pump senses a pressure of lessthan 1,500 PSI and signals the annun-ciator of a low pressure condition. Withloss of system fluid, the standby pumpis turned OFF to prevent unwantedpump operation. The Q UNIT annun-ciator may also illuminate.

With failure of System 1 pressure, aloss of the following occurs:� System 1 flight control servo actu-ator barrel� rudder Q unit� normal slats� normal (No. 1) braking system� hydraulic operation of landing gear� airbrakes.

WARNING: Paint damage mayoccur with hydraulic fluidcontact.

CAUTION: Ensure that the hy-draulic mule has not been con-taminated by previous operations.

CAUTION: When servicingthrough the reservoir filler inlets,systems and air pressures shouldbe completely relieved beforecomplete removal of the caps.


5G-12 For SimuFlite training only Falcon 10/100March 2001

Loss of No. 2Hydraulic SystemThe illumination of the HYD 2 annun-ciator indicates the loss of the hydraulicpressure in System 2. The pressureswitch downstream of the respectivepumps senses a pressure of less than1,500 PSI and signals the respectiveannunciator of a low pressure condi-tion. A loss of System 2 results in theloss of the following:� System 2 flight control servo actu-ator barrel� yaw damper� emergency slats

� emergency and parking brake sys-tem (still available with accumulator)� thrust reversers (may be availablewith the accumulator)� nosewheel steering.

Loss of BothHydraulic SystemsComplete hydraulic system failure isindicated by the loss of hydraulicpressures and possible quantities asindicated on the panel. A pressure out-put of less than 1,500 PSI illuminatesthe HYD 1 and HYD 2 annunciatorsand requires a landing as soon as prac-tical. Q UNIT annunciator may alsoilluminate.

CAUTION: The loss of bothhydraulic systems requires higherstick forces; landing requiresincreased caution.

Airspeed LimitationWhen Hydraulic System 1,System 2, or both fail, airspeedis limited to 260 KIAS/0.76 MI.


Hydraulic System

Hydraulic Systems

Power Source Engine-driven pumps (2)

Main busStandby electric pump power

B busStandby electric pump controlHYDR QTY gagesHYDR PSI gages

Primary A busSystem annunciators

Distribution System 1Rudder, elevator, and aileron servo-actuatorsRudder Arthur Q unitNormal (inboard/outboard) slatsAirbrakesLanding gearNormal braking

System 2Rudder, elevator, and aileron servo-actuatorsYaw damperEmergency (outboard) slatsNosewheel steeringEmergency and parking brakesThrust reversers

Control ST-BY PUMP switch (standby pump)

Monitor HYDR QTY/PSI gages

AnnunciatorsHYD 1/2HYD.TK1/TK2ST-BY PUMPP.BRAKE

Hydraulic reservoir sight gages

Parking brake accumulator pressure switch


Hydraulic system pressure relief valves

Reservoir air pressure relief valves

DataSummary

Falcon 10/100 For SimuFlite training only 5H-1April 2000

Ice and RainProtection

Chapter 5H

Engine bleed air and electrical heating elements provide mostof the Ice and Rain Protection available on the Falcon 10/100.

Engine bleed air is used to anti-ice the following:

� wing leading edges

� nacelle air intake of each engine including the ellipticalspinner, if installed.

Electrically powered heating elements provide anti-icing forthe following:

� pitot probes

� static ports

� stall vane

� optional angle-of-attack probe

� Pt2 and Tt2 engine sensor

� total air temperature probe.

In addition, the pilot and copilot wipers are electrically drivenand the front windshields are demisted. Each cabin window isdemisted by an individual desiccant squib system.

A wing ice inspection light is also available.

5H-2 For SimuFlite training only Falcon 10/100April 2000


Ice and Rain Protection

Table ofContents

Engine Anti-Ice Protection System Schematic . . 5H-5

Wing Anti-ice System Schematic . . . . . . . . . 5H-6

Windshield Heat System Schematic . . . . . . . . 5H-7

Slat Anti-Ice Distribution System Schematic . . . 5H-8

Pitot Heat System Schematic . . . . . . . . . . . 5H-8

Bleed Air System . . . . . . . . . . . . . . . . . . 5H-9

Wing Anti-Ice System . . . . . . . . . . . . . . . . . 5H-9

WINGS Switch/Indicator . . . . . . . . . . . . . . 5H-9

P2 Unit/Control Valves . . . . . . . . . . . . . . . 5H-9

Supply Duct/Distribution Manifolds . . . . . . . . 5H-10

Slat Anti-Ice Air Distribution . . . . . . . . . . . . 5H-10

Thermostats . . . . . . . . . . . . . . . . . . . . 5H-10

Inboard Leading Edge Anti-Icing . . . . . . . . . 5H-10

Operations in Icing Conditions . . . . . . . . . . . 5H-10

Minimum N1 Speed . . . . . . . . . . . . . . . . 5H-10

Engine/Nacelle/Pt2/Tt2 Probe Anti-Ice System . . . 5H-11

Nacelle Anti-Ice System Components . . . . . . . 5H-11

Distribution . . . . . . . . . . . . . . . . . . . . 5H-11

Dome Spinner Anti-Ice Valve (if Installed ) . . . . 5H-11

Pt2/Tt2 Probe . . . . . . . . . . . . . . . . . . . . 5H-12

ANTI-ICE Switch . . . . . . . . . . . . . . . . . 5H-12

Anti-Ice Indicators . . . . . . . . . . . . . . . . . 5H-12

HP BLEED Switch . . . . . . . . . . . . . . . . . 5H-12

Windshield Anti-Ice System . . . . . . . . . . . . . 5H-13

Controller . . . . . . . . . . . . . . . . . . . . . 5H-13

Control and Indicating . . . . . . . . . . . . . . . 5H-13

Side Window Heat . . . . . . . . . . . . . . . . . 5H-14

Windshield Defogging System . . . . . . . . . . . 5H-14

Pitot Heating System . . . . . . . . . . . . . . . . 5H-14

PITOT Switches/Indicators . . . . . . . . . . . . 5H-15


Miscellaneous . . . . . . . . . . . . . . . . . . . 5H-17

Windshield Wipers . . . . . . . . . . . . . . . . . 5H-17

WIPER Switch . . . . . . . . . . . . . . . . . . . 5H-17

Passenger/Cockpit Window Demisting Systems . . 5H-18

Wing Ice Lights . . . . . . . . . . . . . . . . . . . 5H-18

Preflight and Procedures . . . . . . . . . . . . . 5H-19

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5H-19

Abnormal Procedures . . . . . . . . . . . . . . . . 5H-19

Inoperative Engine Anti-Ice System . . . . . . . . 5H-19

Unexpected Operation of Engine Anti-Ice Light . . 5H-19

Inoperative Wings Anti-Ice System . . . . . . . . 5H-19

Wing Anti-Ice Light Illuminates Unexpectedly . . . 5H-19

Leaking Wing Leading Edge Slat Hoses . . . . . 5H-20

Front Pane Heating (Minimum Equipment List) . . 5H-20

Side Window Heating(Minimum Equipment List) . . . . . . . . . . . . . 5H-20

Failure of Pitot/Static Probe Heating System . . . 5H-20

Data Summaries . . . . . . . . . . . . . . . . . . 5H-21

NACELLEANTI-ICING VALVEPOWER FAIL - OPEN

NACELLEPRESSURESWITCH(4:3 PSI)

ENGINE HP BLEED AIR

WING ANTI-ICING AIR

ENGINE LP BLEED AIR

A BUS

ENG 1

B BUS

ENG 2

A BUS

PT2/TT2



Engine Anti-Ice System


C BUS

WING ANTI-ICECONTROL VALVEPOWER FAILURECLOSED

ANTI-ICING HPBLEED RESTRICTOR

HP BLEEDVALVE

WING

HP BLEED VALVESPOWER FAILURE CLOSED

LP CHECKVALVES

THERMOSTAT

FLEXIBLE HOSES

A BUS

ENGINE HP BLEED

WING ANTI-ICING AIR

ENGINE LP AIR

Wing Anti-Ice System


5H-8


Slat Anti-IceDistribution

System

Pitot HeatSystem

Windshield Heat System

D BUS

AUXILIARYBUS

SIDE COPILOT

MAXNORMOFF

R. WINDSHIELDTEMPERATURE

REGULATOR

SIDE WINDOWTHERMOSTAT

L. WINDSHIELDTEMPERATURE

REGULATOR

X.FR

PILOTMAX

NORMOFF

A BUS


1

DISTRIBUTION MANIFOLD

A

A

INNER SHEET

SKIN PANEL

MACHINED SKIN PANEL RECESSES

AIR ESCAPE

AIR ESCAPE

DISTRIBUTIONMANIFOLD

SKIN

INNER SHEET

INNER SKIN PANEL

SECTION AA

1

ANTI-ICINGAIR OUTLETS

INNER SKIN PANEL

2

2

INNERSHEET

ANTI-ICINGAIR CUT-OUTS

SKINPANEL

A BUS

A BUS

PITOT 2

PITOT 1

PILOTPITOTS

COPILOT

STATIC PORTAOA VANE

D BUS

Pitot Heat SystemSlat Anti-Ice Distribution System

Falcon 10/100 For SimuFlite training only 5H-9March 2001


Pressurized bleed air for anti-icing issupplied as described below. � Low pressure and high pressure airis bled from the engines and directedto the P2 unit. Air drawn from the P2unit prevents ice formation on the slatsand leading edges of the wing whenslats are extended.� Air drawn from the high pressurebleed (HP) air system of each engineprevents ice formation in its ownnacelle air intake. HP bleed air also isused to anti-ice its own dome fan spin-ner (if installed). If engines are fittedwith a conical-shaped fan spinner, nospinner anti-ice is required or installed.

Wing Anti-IceSystemBleed air from the P2 unit anti-ices thewing leading edges, inboard and out-board slats, and wing roots (Figure5H-1) with slats extended or retracted.

Components of the wing anti-icing sys-tem include the following:� WINGS switch/indicator� P2 unit (bleed air center)� control valves� supply duct

� distribution manifolds� slat air distribution� thermostats� inboard leading edge anti-icing.

WINGS Switch/IndicatorThe WINGS two-position ON/OFFswitch in the ANTI-ICE zone on theoverhead panel (Figure 5H-2 ) controlsleading edge anti-icing.

In the OFF position, the wing anti-icecontrol valve is closed. In the ON posi-tion, the wing anti-ice control valve isenergized open and the WINGS anti-ice (green) indicator light on the over-head panel illuminates and remains onif the wing supply duct thermostats aresubjected to 176°F (80°C). See ther-mostat below.

On aircraft with SB 30 017,one ofthe thermal fuses melts and illuminatesthe amber DE-ICING annunciator if ahose is disconnected in the leadingedge.

P2 Unit/Control ValvesThe P2 unit (bleed air center) receiveslow pressure bleed air continuouslyand high pressure bleed air throughHP control valves in the pylon.

Bleed AirSystem

5H-1

5H-2

Anti-Icing LimitationThe anti-icing system is a preven-tive system and is not designedfor eliminating ice. As aconsequence, anti-icing must beturned on in flight as soon as TATdrops below +5°C in icing condi-tions, visible moisture, or if icingis anticipated.

Wing and Engine Anti-Icing LimitationThe engine and wing anti-icesystems must not be used withtotal air temperature in excess of+10°C. The wing anti-ice systemmust not be used on the groundexcept for maintenance checks.

5H-10 For SimuFlite training only Falcon 10/100March 2001

The wing anti-ice system receives alow pressure (LP) and high pressure(HP) bleed air mixture (approximate-ly 70% low pressure and 30% highpressure) from both engines. Winganti-ice air is tapped off the P2 unitand directed through the anti-ice con-trol valve to the supply duct.

Supply Duct/Distribution ManifoldsA single supply duct downstream fromthe P2 unit and wing anti-ice controlvalve splits so that each duct distrib-utes bleed air to its respective wingleading edge distribution manifold.The distribution manifold along thefront spars of the wing leading edgeducts bleed air to the inboard and out-board slats.

The supply duct is connected to thedistribution manifolds through fourflexible hoses (glass cloth reinforcedsilicone material) to permit extensionof the slats.

Slat Anti-Ice AirDistributionInside the slats, the anti-icing air out-lets distribute bleed air along the entirelength of the slat leading edge throughsmall, equally spaced holes. The airflows inside the area between the slatskin and the inner sheet through aper-tures cut in the inner sheet.

Recesses are machined in the innersurface of the skin panel to enablebleed air to circulate toward the topand bottom borders of slat skin panels.

Air escapes from the bottom border ofthe skin panel through small ducts inthe skin panels opposite the inner sheet.Air escapes from the top border of theskin through apertures in the innersheet and holes in the top skin panel.

ThermostatsTwo thermostats connected in seriesin the supply duct sense the tempera-ture of bleed air entering the slat lead-ing edges. They provide positive indi-cation of wing anti-ice. Althoughfailure of a thermostat prevents a pos-itive indication of anti-ice, wing heat isnot affected.

Normal operation is confirmed by illu-mination of the green WINGS light onthe overhead panel when the ther-mostats reach 176°F (80°C); becauseof thermal inertia, it takes one to twominutes to illuminate. The greenWINGS light extinguishes if one ofthe thermostats senses a temperatureof 158°F (70°C); this indicates inade-quate anti-icing for one of the wings.

Inboard Leading EdgeAnti-IcingA tap off the supply duct supplies bleedair to the leading edge distributionmanifold. A flexible tube connects theduct and manifold to provide inboardleading edge anti-ice protection.

Operations inIcing ConditionsMinimum N 1 speedFor adequate performance of the wingand engine anti-icing systems, the N1speed of the operative engines mustnot be lower than the minimum val-ues shown in Table 5H-A .

The temperature correction chart in theAFM must be used to derive outsideair temperature from ram air tempera-ture.

Wing anti-icing must be switched on asa preventive measure in flight whenthe indicated total temperature is below+5°C (+41°F) in visible moisture andwhen icing conditions are anticipated.

Wing Anti-Ice LimitationsDo not perform ground checks ofthe wing anti-ice stystem withoutconsulting the maintenance manual for procedures. With theaircraft on the ground, this canlead to overheating andsubsequent damage to the slats.

WINGS anti-ice systems shouldbe switched on in flight prior toentering visible moisture when-ever the TAT is +5°C or below.Do not exceed the operationalengine and wing anti-ice systemlimitations.

Wing anti-icing must not be usedwith total air temperature inexcess of +10°C.



The wing anti-icing system must notbe operated on the ground, except foroperational checks conducted at lowpower settings and under constantmonitoring of the wing leading edgetemperature.

Engine/Nacelle/Pt2/Tt2 ProbeAnti-Ice SystemEach engine provides its own highpressure bleed air to anti-ice the nacelleand the elliptical spinner (if installed).Additionally, when the engine anti-iceis selected, the Pt2/Tt2 probes are elec-trically heated. (See Maneuvers chap-ter for operating requirements).

The engine nacelle anti-ice and Pt2/Tt2probe systems have a common controlsystem for each engine.

Nacelle Anti-IceSystem ComponentsComponents of the engine nacelle airintake anti-ice systems include the fol-lowing:� nacelle anti-ice valve and pressureswitch� anti-ice supply line� restrictor and expansion sleeves� air intake circular distribution duct� internal wall, shroud, and sidelouvers.

DistributionThe respective HP bleed air port ofeach engine supplies bleed air for thenacelle anti-ice system via the electri-cally-controlled nacelle anti-ice valve.This valve fails open in the event of aDC power failure.

HP bleed air from the engine is direct-ed to the circular distribution duct thatdistributes it to the nacelle intakethrough expansion sleeves and arestrictor. The restrictor limits warmairflow to avoid excessively high tem-peratures through the anti-ice air sup-ply line. Each line is fitted with anacelle anti-ice solenoid valve that con-trols anti-ice air for each air intake. A4.3 PSI nacelle pressure switch pro-vides indication.

The bleed air in the air inlet circulatesthrough the air intake box between theair intake internal wall and shroud. Itis then evacuated outside through sidelouvers.

Dome Spinner Anti-IceValve (if installed)The dome spinner anti-icing valve onthe right side of each engine controlsthe engine bleed air to the spinner. HPbleed air is distributed into the inter-nal face of the spinner by internalplumbing; it is then discharged intothe air intake immediately upstreamof the fan.

Table 5H-A; Minimum N 1 Speed

OAT -30°C -5°C

In flight,min. N 1 speed

76% 73% 62%

-10°C-20°C

In flight,recommendedN1 speed

78% 75% 64%

In approach,min. N 1 speed

69% 69% 64%

One engine 88% 85% 79% 74%

Engine Anti-Ice LimitationThe engine anti-ice must be usedon the ground and in flight whenthe temperature is below +5°Cand icing conditions exist or areanticipated.

CAUTION: Continuous opera-tion of engine anti-ice in normalambient conditions is not rec-ommended because damage canresult.

67%

69%

69%


Each engine has an electrically-con-trolled and pneumatically-operateddome spinner anti-icing valve normallyheld closed by a spring. The valve failsclosed if DC power fails with anengine running.

HP bleed air always is available forengine anti-ice because the engine anti-ice valve is upstream of the associatedHP bleed air shutoff valve.

The fan spinner anti-icing system iseliminated on aircraft with engines fit-ted with conical-shaped fan spinners.

Pt2/Tt2 ProbeA Pt2/Tt2 sensor (Figure 5H-3 ) on eachengine provides pressure and temper-ature information to the fuel comput-er. Each engine’s probes are heatedelectrically when the respective engineanti-ice control is in ON.

ANTI-ICE SwitchEach engine anti-ice system has a two-position (ON/OFF) ENG switch belowthe corresponding indicator light onthe ANTI-ICE section of the overheadpanel.

Engine anti-icing must be switched onas a preventive means when the totalindicated temperature is below +5°C(+41°F) in visible moisture and whenicing conditions are anticipated.

Each engine anti-ice switch opens orcloses its engine nacelle anti-icingvalve and elliptical spinner anti-icingvalve (if installed). It also connectselectrical power to its Pt2/Tt2 probeheating element.

Anti-Ice IndicatorsThe ENG 1 and ENG 2 anti-ice sys-tems each have a green indicator lightabove the corresponding switch on theANTI-ICE section of the overheadpanel. When an ENG 1 or ENG 2 anti-ice switch is on, the green light illu-minates and remains on if the nacelleand spinner pressure switches are sub-jected to the proper pressure (nacelleair intake supply line 4 to 7 PSI andengine spinner supply line 6 PSI).

The indicator lights do not indicateproper Pt2/Tt2 probe heat operation.

On aircraft with SB 30 010: Whenanti-ice control switches are placed inOFF, illumination of anti-ice lightsindicates that one of the solenoidvalves failed open and that bleed airfrom the respective engine continueseven though anti-icing switch is off.

HP BLEED SwitchThe HP bleed valves are electrically-controlled, air-operated, normallyclosed valves powered by the C Bus.

5H-45H-3

Engine Anti-Ice LimitationEngine anti-icing must not beused with total air temperaturein excess of +10°C.

CAUTION: In the event ice hasalready started to accumulate, theignition system must be turnedon (start selector switches toAIRSTART) prior to selectingthe engine anti-icing systems.Select engine anti-ice ON, oneat a time with a delay of no lessthan 30 seconds. Wing anti-icingcan then be switched on.



During the initial cockpit preflight, thetwo-position OPEN/CLOSED HPBLEED switch on the pedestal bleedair panel (Figure 5H-4 ) is set toCLOSED. With the HP bleed valvesclosed, HP bleed air is not availableto the P2 unit.

WindshieldAnti-Ice SystemElectrical heating elements embeddedbetween the windshield laminationsprevent ice formation on the pilot/copilot windshields. Each windshieldis separate and independent.

The two windshield anti-icing systemsare identical and include the follow-ing components:� controller� probes/wires� automatic transfer� 3-position selector switches� XFR signal light.

ControllerEach pilot’s main windshield panel isheated electrically through a controllerwith heat sensing probes. The twoprobes in the pane (one a stand-by)and the controller automatically main-tain the panes within a design temper-ature range. The regulation tempera-ture is 30°C (86°F).

A sensing circuit between the wind-shield systems provides automatictransfer of control if either system failsor if a heat sensing probe short circuits.

Control and IndicatingThe PILOT and COPILOT WIND-SHIELD switches on the overheadpanel, labeled OFF/NORM/MAX,control windshield anti-ice operation(Figure 5H-5 ). The NORM positionprovides normal heating intensity overa maximum area; the MAX positionprovides high heating intensity over areduced area.

The NORM position is used in all nor-mal operations; the MAX position isselected in flight only when theNORM position fails to maintain thewindshields clear of ice. Selection ofMAX allows current to bypass oneheating element and thus increase cur-rent flow over a reduced area. Thisprovides greater efficiency.

If one of the regulation systems isfaulty (regulator or sensor failure), theregulator of the other system auto-matically takes over and annunciatesthe transfer by illuminating the amberX.FR light. Isolate the inoperative cir-cuit immediately by setting the wind-shield ANTI-ICE switches to OFF; theswitch that extinguishes the light indi-cates the faulty circuit.

5H-5

Windshield ElectricalHeating LimitationThe PILOT and COPILOTWINDSHIELD control switchesmust not be selected to the MAXposition, except in flight, andonly if the NORM position is notsufficient in icing conditions.


Side Window HeatThe anti-icing system for the pilot’sand copilot’s OROGLASS panes (SB30 001) is similar to the windshieldanti-ice system except there is onlyone probe in each panel. The SIDEswitch on the overhead panel controlsthe lateral rear panels heat. Normalheating temperature is 104°F (40°C).

A single controller with two overheatsensors is common to both side win-dows. If a failure occurs in one sidewindow heat system, both sides auto-matically shut down to prevent over-heating.

WindshieldDefogging System To maintain maximum visibility in theFalcon 10/100, unconditioned bleedair for defogging is supplied to the

inner windshield surface through a ductrunning along the lower part of thewindshield.

The WSHLD control lever on thepilot’s console outboard of the floorheating lever regulates windshielddefogging (Figure 5H-6 ). The leverhas CLOSED and OPEN positions.

Pitot HeatingSystemThe pitot probes, static ports, stallvanes, AOA probe, and total air tem-perature probe are anti-iced by 28VDC-powered, built-in heating ele-ments.

Components of the pitot/static anti-icing system include:� pilot/copilot switches (Figure 5H-12)� pitot probes (Figure 5H-7 )

5H-75H-6

5H-95H-8

Pitot LimitationPilot and copilot pitot heat onfor flight.



� static ports (Figure 5H-8 )� stall vanes (Figure 5H-9 )� AOA probe (if installed) (Figure5H-10)� total air temperature probe (Figure5H-11).

PITOT Switches/IndicatorsThe PILOT and COPILOT PITOTswitches control anti-icing for thepitot/static system. The switches arein the PITOT area on the overheadpanel (Figure 5H-12 ).

When the pilot’s pitot system switchis in ON, the following heating ele-ments are powered:� left pitot probe� left and right pilot static ports� left stall vane.

When the copilot’s pitot system switchis in ON, heating elements poweredinclude:� left and right copilot static ports� right pitot probe� right stall vane

The AOA switch controls the heatingelement for the AOA probe.

Printed circuits monitor the electriccurrent flowing through the heatingelements of the pitot pressure probesand the static pressure ports. When thecurrent in a monitored circuit becomeslow or nonexistent, the correspondingprinted circuit, acting as a current relay,grounds the amber PITOT 1 or PITOT2 annunciator on the Failure WarningPanel. The 28V DC power applied tothis panel then illuminates the warn-ing light to indicate abnormal operat-ing conditions.

5H-115H-10

5H-12


TAT probe and stall vane anti-icing arenot monitored. A small ammeter, nor-mally on the copilot’s instrument panel,monitors the anti-icing current to theAOA probe (if installed).

The PILOT and COPILOT PITOTswitches must be ON prior to takeoff;the amber PITOT 1 and PITOT 2warning annunciators should extin-guish.

When the PILOT and CO-PILOTPITOT switches are OFF, the pitot andstatic monitoring circuits sense no cur-rent and illuminate the PITOT 1 andPITOT 2 annunciators.

Electrical power for the pilot’s pitot/static system anti-icing is provided bynon-shed 28V DC; power for the co-pilot’s pitot/static system anti-icing isprovided by load-shed 28V DC.



Windshield WipersThe pilot and copilot windshields areequipped with electric windshieldwiper systems to provide clear visi-bility during takeoff, approach, andlanding in rain or snow. The wipersare parked in a drained recess belowthe windshields when not in use toavoid any accumulation of ice or snowon the wipers.

Components of each windshield wipersystem include:� pilot/copilot wiper control switch � motor/gearbox/drive shaft� wiper arm assembly/wiper blade� relay/resistor.

WIPER SwitchA three-position switch labeled ON/OFF/PARK on the overhead panelcontrols each wiper (Figure 5H-14 ).When either the PILOT WIPER orCOPILOT WIPER control switch isset to ON, 28V DC is applied to thecorresponding wiper motor.

Both pilot and copilot windshieldwipers are powered from the A bus(Figure 5H-13 ). They have only onerate of motion of 170 cycles perminute. The wipers can be used up toa maximum speed of 190 kts. SB 30004 improves wiper return and stor-age to 150 kts.

The revolving motor drives the wind-shield wiper via a flexible drive shaft.The gearbox reduces the drive speedand transforms the rotating movementinto an oscillatory motion. This move-ment is transmitted to the wiper armassembly that drives the windshieldwiper blade. Either wiper switch auto-matically returns to OFF after thePARK position is selected.

When either wiper switch is held inPARK, 28V DC is applied to the cor-responding parking position relay.Power is then applied to the associat-ed windshield wiper motor via a resis-tor. The motor turns at a reduced speed.A ground connection is also providedfor the parking position cam of the cor-responding windshield wiper motor.

Miscellaneous

5H-14

ONOFF

PARK

ONOFF

PARK

WIPER

WIPER

A BUS

5H-13


When the cam establishes end-of-travel contact (i.e., windshield wiperarm in the low sweep position andrecessed), the cam cuts off the motorpower supply. When the control switchis released, the switch returns to thespring-loaded OFF position.

Passenger/CockpitWindow DemistingSystemsComponents of the demisting systeminclude:� the passenger window system� rechargeable silicagel cartridge� side port� plastic cap� the cockpit window system� refillable silicagel cartridge� layered windows with ports� plastic plug.

With SB 30-012, each aft window ofthe flight compartment is dried by aseparate system connected to arechargeable cartridge of silicagel des-iccant under the windows.

Without SB 30-012, a port is drilled inthe inner window pane to vent thespace between the two panes. WithSB 30-012, this port is sealed and asilicone CAF4 canister is connectedto the space at the bottom of the win-dow. The canister is covered by aleather-covered plastic cap held inplace by three locking tabs.

With SB 30-014, a refillable silicagelcartridge is included in front of framesix under the lower side of the trim-ming. A flexible hose connects the car-tridge to the space between the twopanes of the lateral glass panel.

A hole in the trimming for placementand inspection of the desiccant car-tridge is fitted with a plastic plug heldin place by three locking tabs.

Wing Ice LightsThe wing leading edges are illuminat-ed by the wing ice lights in each sideof the fuselage (Figure 5H-15 ). TheWING ICE LIGHT switch on theoverhead panel controls the wing icelights (Figure 5H-16 ).

5H-165H-15



PreflightDuring the exterior preflight inspec-tion, accomplish the following checksof the ice and rain protection systems(refer to Preflight chapter).

Check that the left and right wind-shields are clean, in good condition,and have no delamination or discolor-ation. Check that the wipers are in goodcondition and in the park position.

Check the right and left wing leadingedge slats condition.

Check the condition of the engineintake nacelle.

Ensure the wing ice inspection light(s)(if installed) are intact and in goodcondition.

Check the TAT probe, AOA sensor (ifinstalled), Pt2/Tt2 probes, pitot probes,static ports, and stall vanes for heat.Since the pitot probes and static portsare only monitored, heating of theother system components may occa-sionally be checked by touch.

If possible, perform check with a GPUconnected to prevent battery drain.

Check for engine fan or spinnerdamage; verify that the Pt2/Tt2 probe isintact.

AbnormalProceduresThe following section provides a briefdiscussion of what happens duringabnormal anti-ice situations. For a listof specific procedural steps, refer tothe AFM or the SimuFlite OperatingHandbook.

Inoperative EngineAnti-Ice SystemAn inoperative engine anti-ice systemis indicated when the ENG 1 or ENG2 ANTI-ICE switch is on but the

respective engine ANTI-ICE lightremains extinguished. Because the twoengine anti-icing pressure switches areseries-connected, a malfunction ofeither can cause the associated ANTI-ICE light to remain extinguished.

With an engine anti-icing electrovalvejammed closed, select the ENG 1 orENG 2 ANTI-ICE switch to OFF andavoid areas where icing conditionsexist.

Unexpected Operationof Engine Anti-Ice LightIf an ENG 1 or ENG 2 ANTI-ICE lightilluminates inadvertently, reduce theassociated throttle to idle and shutdown the associated engine as soon aspossible after landing is completed.

This procedure is applicable to aircraftwith conical spinners and no anti-icingper SB F10-0180or additional indi-cating system per SB F10-0121.

Inoperative WingsAnti-Ice SystemIf the WINGS anti-ice indicator lightdoes not illuminate after the WINGSanti-ice system is activated, check thatthe HP BLEED switch is selected toOPEN. This switch may have been leftin the CLOSED position.

An extinguished light may result fromeither a control thermostat malfunc-tion or a failure of bleed air supply tothe wing leading edges.

Select the WINGS ANTI-ICE switchto OFF to eliminate any risk of bleedair leakage downstream of the winganti-icing electrovalve.

The most likely problem is that thewing anti-ice electro-valve is jammedclosed. Avoid areas where icing con-ditions exist; maintain maximum speedto provide sufficient positive impacttemperature.



Wing Anti-Ice LightIlluminates UnexpectedlyIf the WINGS green light illuminatesunexpectedly with the WINGS ANTI-ICE switch in OFF, the wings anti-icing valve is jammed open.

A persistent wing anti-icing air bleedduring the flight results in a slight lossof airplane performance (e.g. thrust,speed, consumption). There also existsa risk of overheating and damaging theleading edges when the aircraft is at astop.

During the approach sequence, selectthe HP BLEED control to CLOSEDto limit the risk of damage. Wing anti-icing is then supplied only with LPbleed air. Anti-icing efficiency isreduced; therefore, areas where icingconditions exist should be avoided.

To prevent overheating of the leadingedges when the airplane is on theground and outside air ventilation andcooling are not sufficient, shut downthe engines as soon as landing iscompleted.

Leaking Wing LeadingEdge Slat HosesBleed air leakage is annunciated onlyif the aircraft is equipped with anamber DE-ICING annunciator as perSB 30-017. In S/Ns 1 to 4 and 6 to 151,this light is on the instrument panel;on S/N 152 and subsequent, it is onthe warning panel.

Illumination of the DE-ICING annun-ciator indicates airframe bleed air isleaking in the area of the flexible hosestrigged by four thermal fuses. Set theWINGS ANTI-ICE switch to OFF andavoid icing conditions.

Front Pane Heating(Minimum Equipment List)Illumination of the amber WIND-SHIELD XFR light on the overhead

panel indicates an automatic transferof windshield heat regulation from thefailed side to the operational side.Internal bleed air de-fogging providesfor complementary heating. Its oper-ation should be checked in flight.

If forward visibility is reduced, use theside window. The direct vision windowcan be removed for increased visibility.

Takeoff is authorized with a failurethat only affects copilot pane.

Side WindowHeating (MinimumEquipment List)If side window heating fails, heatingof both panes is immediately inter-rupted. A significant reduction in sidevisibility must be taken into consider-ation during operation in the trafficpattern and when maneuvering on theground.

Failure of Pitot/StaticProbe Heating SystemThe amber PITOT 1 or PITOT 2annunciator illuminates when pitot orstatic probe heating fails to operate. Itis not possible to mark the differencebetween a heating failure of the pitot orstatic probe in each circuit. Leave thepitot switch associated with the failedheating probe ON to continue heatingthe probe still operating.

If the pilot’s flight instruments indica-tions are false, use copilot’s flightinstruments and position the staticselector to EMERGENCY.

If the copilot’s flight instruments indi-cations are also false, position thepitot/static selector to PANEL ONLY,position the autopilot to the DISEN-GAGED position, and reduce airspeedto 260 kts/MACH 0.76 MAX.

CAUTION: When the HPBLEED switch is selected toCLOSED, wing anti-icing sys-tem efficiency is reduced. Avoidareas where icing conditionsexist.

CAUTION: When pitot/staticselector is selected to PANELONLY, the pilot feel force is nolonger slaved to the airspeed andthe cabin differential pressureindicator is disabled.



DataSummaries

Ice and Rain ProtectionEngine/Wing Anti-Ice System

Power Source EngineHP bleed airA bus – left engine Pt2Tt2B bus – right engine Pt2Tt2

WingHP/LP engine bleed airA bus – wing anti-ice valve

Distribution HP bleed airEngine nacelle anti-ice valvesNacelle anti-ice valvesEngine/nacelle pressure switches

HP/LP bleed airWing anti-ice valveWing leading edges

Control ENG 1/ENG 2 ANTI ICE switches

WINGS ANTI ICE switch

Monitor ENG 1/ENG 2 lights

WINGS light


Pitot/Static and Windshield Anti-Ice System

Power Source A busL/R windshield wipersPilot windshield heatPilot pitot/static heat

D busCopilot windshield heatCopilot pitot/static heat

Distribution Windshield temperature regulators

Control PILOT/COPILOT PITOT switches

A/A switch (if installed )

PILOT/COPILOT WINDSHIELD switches

Wiper switches

Monitor Window heat X.FR light

AMP meter

PITOT 1/PITOT annunciators


Windshield temperature regulators

Falcon 10/100 For SimuFlite training only 5I-1April 2000

LandingGearSystems

Chapter 5I

The Landing Gear System on the Falcon 10/100 is a standardtricycle design electrically controlled and hydraulically actu-ated. The main gear utilizes dual wheel assemblies, while thenose gear uses a steerable single wheel with a chined tire forwater and slush deflection.

In an emergency, landing gear extension is accomplished throughmechanical unlocking of the gear.

The Brake System is mechanically controlled and hydraulicallyactuated by pressure from the hydraulic systems. The system hasmultiple disc brakes on the main gear wheels. An anti-skidsystem provides maximum braking efficiency on all runwaysurfaces.

5I-2 For SimuFlite training only Falcon 10/100April 2000


Landing Gear Systems

Table ofContents

Landing Gear System Schematic . . . . . . . . . . 5I-6

Landing Gear System . . . . . . . . . . . . . . . . 5I-7

Ground/Flight Detection Proximity Switches . . . . . 5I-7

Main Gear . . . . . . . . . . . . . . . . . . . . . . . 5I-7

Main Gear Doors and Shock Strut . . . . . . . . . . 5I-8

Main Gear Uplock . . . . . . . . . . . . . . . . . . 5I-9

Main Gear Operation . . . . . . . . . . . . . . . . 5I-9

Nose Gear . . . . . . . . . . . . . . . . . . . . . . 5I-10

Nose Gear Doors and Shock Strut . . . . . . . . . 5I-10

Nose Gear Uplock . . . . . . . . . . . . . . . . . 5I-10

Nose Gear Centering . . . . . . . . . . . . . . . . 5I-10

Nose Gear Operation . . . . . . . . . . . . . . . . 5I-11

Landing Gear Control Lever . . . . . . . . . . . . . 5I-12

Landing Gear Indication . . . . . . . . . . . . . . . 5I-12

Emergency Operation . . . . . . . . . . . . . . . . 5I-13

Brake and Nosewheel SteeringSystem Schematics . . . . . . . . . . . . . . . . . 5I-14

Brake System . . . . . . . . . . . . . . . . . . . . 5I-15

Components . . . . . . . . . . . . . . . . . . . . . 5I-15

Braking Distributor . . . . . . . . . . . . . . . . . 5I-15

Brake System Operation . . . . . . . . . . . . . . 5I-15

Parking/Emergency Brake . . . . . . . . . . . . . 5I-15

Anti-Skid Brake System . . . . . . . . . . . . . . 5I-16

Servo Valves . . . . . . . . . . . . . . . . . . . 5I-16

Control Box . . . . . . . . . . . . . . . . . . . . 5I-16

Anti-Skid Operation . . . . . . . . . . . . . . . . 5I-16

Nosewheel Steering . . . . . . . . . . . . . . . . . 5I-17

Nosewheel Steering Operation . . . . . . . . . . . 5I-18


Servicing and Procedures . . . . . . . . . . . . . 5I-19

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5I-19

Anti-Skid Test . . . . . . . . . . . . . . . . . . . . 5I-19

Servicing . . . . . . . . . . . . . . . . . . . . . . . 5I-19

Inflation . . . . . . . . . . . . . . . . . . . . . . . 5I-19

Struts . . . . . . . . . . . . . . . . . . . . . . . . 5I-19

Abnormal Procedures . . . . . . . . . . . . . . . . 5I-19

Landing Gear Does Not Extend . . . . . . . . . . 5I-19

Abnormal Gear Retraction . . . . . . . . . . . . . 5I-19

Anti-Skid System Failure . . . . . . . . . . . . . . 5I-20

Low Parking Brake Accumulator Pressure . . . . . 5I-20

Brake System Failure . . . . . . . . . . . . . . . . 5I-20

Data Summaries . . . . . . . . . . . . . . . . . . 5I-21

HYDRAULICSYSTEM 1PRESSURE

EXTENSIONSOLENOID

RETRACTIONSOLENOID

RETURN

PRESSURERELIEF SLIDE

UP

LANDINGGEARPOSITION

RETRACT

EXTENSION



B BUS

MOVE

A BUS

LDG GEAR

DOWN


Landing Gear System



The Falcon 10/100 landing gear sys-tem is electrically controlled andhydraulically operated and includesdual-wheel main and single-wheelnose gear. When retracted, the nosegear is enclosed by mechanically actu-ated and linked doors (Figure 5I-1 );the main gear is partially enclosed bya door attached to the main gear strut.The landing gear normally is actuatedby hydraulic pressure from HydraulicSystem 1 (see Hydraulics chapter fordetails). In the event of electrical and/orhydraulic problems with the landinggear, a manual emergency extensionmode is available.

Ground/FlightDetection Proximity(Squat) SwitchesAn electrical proximity (i.e., squat)switch in each main gear and the nosegear sends signals to various aircraftsystems indicating whether the aircraftis on the ground or airborne. Theswitches are on the fixed portion ofeach landing gear strut.

When the struts are extended, theswitches are activated into the flight

mode by movement of a flange againstthe proximity microswitch. When thestruts are compressed, the switches areactivated into the ground mode as theflange moves away from themicroswitch.

The following systems receive inputsfrom the proximity switches:� pressurization controller� battery blower� Freon air conditioning system� Stall Systems 1 and 2� gear retract protection� airbrake warning� thrust reversers� conditioning valve� hydraulic standby pump� STAB TRIM annunciator� nosewheel steering amplifier.

Main GearThe main landing gear installation con-sists of right and left main landing gearassemblies of pneumatic/hydraulicshock struts with dual wheels (Figure5I-2, following page). The gear is

Landing GearSystem

NOSEGEAR

RH MAINLANDING GEAR

LH MAINLANDING GEAR

DUALWHEELS

NOSEGEARDOORS

MAIN GEARDOORS

5I-1


hydraulically retracted inward into thewell under the wing and fuselage, andis locked in the up position by meansof a lock assembly in each wheel well.When the gear is up, hydraulic pres-sure is removed from the landing gearsystem; each gear is held in place by itsuplock.

In the down position, the gear is lockedand braced by the landing gear exten-sion unit (operating jack). Other com-ponents of the main gear include:� trailer arm� shock absorber barrel� shock absorber (strut)� door� proximity switches� uplock unit� actuating cylinders� brake assembly� speed sensors.

The main gear includes the strut barrel,piston assembly, axle assembly, andassociated parts (Figure 5I-3 ). The

upper end of the strut main body ishinged between the forward and aftmain gear supports within the wingstructure by trunnion pins.

Main Gear Doorsand Shock StrutThe main landing gear door assemblycovers the landing gear well when thelanding gear is retracted. The door ishinged on its outboard end to the wingstructure; it is secured to the strut mainbody by an adjustable link that actu-ates the door during retraction andextension.

The main landing gear struts absorbshock forces generated during landingand taxiing. The lower portion of thestrut is filled with MIL-H-5606hydraulic fluid and the upper portion isserviced with nitrogen gas at 450 PSI.

During taxiing, shocks are absorbedlargely by compression of the nitro-gen gas within the strut; during land-ing, shock is absorbed by hydraulicaction.

ACTUATING CYLINDER

HINGE BEARING

GEAR STRUT

MAIN LANDINGGEAR FAIRINGDOOR

5I-3

5I-2



Main Gear UplockThe main landing gear uplock assem-blies contain mechanical link arms thatretain the main gear in the up position(Figure 5I-4 ). During the landing gearretraction cycle, the spring loaded hy-draulic uplock cylinder piston retractsto allow a roller to engage the latchand lock the gear in the up position.

When the landing gear extension cyclebegins, the uplock cylinder piston rodextends to push the gear uplock rolleraway from the latch to the unlockedposition. If hydraulic system pressureis lost, the uplock mechanism isunlocked by moving a mechanicallever on the uplock.

Main Gear OperationThe main landing gear is extended andretracted by two cylinder assembliesthrough the movement of the landinggear control handle on the forwardinstrument panel.

When the landing gear control isplaced in the UP position, the landing

gear selector solenoid is energized andmain hydraulic pressure is applied tothe gear up ports of the actuating cylin-ders to retract the main landing gear.The gear down ports of the actuatingcylinders are connected to thehydraulic return line.

As the landing gear is driven to itsupper limit, it actuates the link roller ofthe uplock mechanism and allows therollers to latch the landing gear in theretracted position. The uplock armactuates its uplock proximity switchon the uplock assembly, and the retrac-tion sequence is ended. During landinggear retraction, the wheel brakes areautomatically applied to stop wheelrotation.

When the landing gear handle is placedin DOWN, the landing gear selectorsolenoid valve opens and pressurizesthe landing gear extension line. Theuplatch boxes open and release thelanding gear; the actuating cylinderscontrol extension of the landing gear.Once the gear is fully down andlocked, the solenoid remains energized

TO EMERGENCYGEAR HANDLE

ROLLERHOOK

CLOSEDPOSITION

RELEASEDPOSITION

EXTENSIONPRESSURE

5I-4


to provide constant gear down linepressure. A claw type locking device inthe gear actuator provides positivemechanical locking of the strut afterengine shutdown.

Nose GearThe nose gear installation consists ofa nose gear assembly, a locking strut,an actuating cylinder assembly, a steer-ing actuator assembly, a wheel assem-bly, a tire, and attaching hardware(Figure 5I-6 ). Other componentsinclude:� piston assembly� uplock roller� doors� proximity switch� centering spring.

Nose Gear Doors andShock StrutTwo doors enclose the nosewheel wellwhen the landing gear assembly isretracted (Figure 5I-5 ). The doors arehinged to the aircraft structure andopen to either side of the gear. Thedoors are mechanically actuated bymovement of the gear and are attachedby an adjustable linkage.

The doors open and close as the land-ing assembly is extended or retractedand are dependent upon the positionof the landing gear strut. Actuation ofthe doors is accomplished by push rodsand bellcranks linked to the outer bodyassembly and the aircraft structure.

Shock forces generated during land-ing and taxiing are cushioned withpneudraulics with the nose gear in thesame manner as the main gear exceptthat the nose gear shock strut is hingedby means of a trunnion supported bythe shock strut in the nosewheel well.

Nose Gear UplockThe nose gear uplock assembly is ofthe same construction and operationas the main gear uplock assemblyexcept that it is smaller in size.

Nose Gear CenteringSprings within the nose gear center-ing device prevent the retraction of anoff-center gear wheel assembly intothe well. With nose gear extension aftertakeoff, the spring centers the lowerstrut and wheels and provides aninflight signal from a proximity switch.If centering does not occur, the landinggear does not retract.

5I-5

Falcon 10/100 For SimuFlite training only 5I-11March 2001


Nose Gear OperationThe nose landing gear is extended andretracted by a double-acting cylinderassembly similar to the main landinggear actuating cylinder assembly. Inthe retracted position, the nose land-ing gear assembly is latched by theuplock cylinder assembly in the wheelwell. In the extended position, a tele-scoping locking strut mechanicallylocks the nose gear into position witha claw-type locking device.

When the landing gear control leveris placed in the DOWN position,hydraulic pressure is applied to thegear down port of the actuating cylin-der and the gear up port is connectedto the return line. The extension of thecylinder drives the landing gear to thedown position, and the locking strutlocks the gear in place.

When the landing gear control leveris placed in the UP position, hydraulicpressure is applied to the gear up port

ACTUATOR

SHOCK STRUT

GEAR LEG

HALF-FORK

ROTARY STRUT

SHIMMY DAMPER

NOSEWHEEL STEERINGACTUATOR

5I-6

NOTE: A red override buttonabove the landing gear controllever allows overriding of thelanding gear locking device. Itsuse is discouraged.


of the nose gear actuating cylinder.The retraction of the cylinder drivesthe landing gear assembly to the fullyretracted position. Gear retractionmovement is slowed down by a one-way restrictor in the nose landing geardown line that restricts the return flowfrom the actuating cylinder.

During the last extension, hydraulicpressure unlocked the uplatch hookwhich is spring-loaded OPEN. Thisensures the uplock accepts the gearuplock roller during retraction. Theuplock roller on the nose gear joins theuplock hook, and the roller moves backinto place to lock the gear in the UPposition. The latch also actuates theuplock microswitch on the nose gearuplock assembly.

Landing GearControl LeverThe gear control lever on the centerinstrument panel controls the landinggear (Figure 5I-7 ). The control leverelectrically positions the gear selectorsolenoid valve to either retract or extend

the gear. On the ground, a lockingdevice holds the lever in the DOWNposition to prevent inadvertent retrac-tion. The landing gear electrical con-trol circuit is powered from the PrimaryA bus. If the gear fails to retract, selectgear down, verify that it is down andlocked, and then return for landing.

Landing GearIndicationThe landing gear configuration lightsabove the landing gear control panelare powered by the B bus. Three lightsand three arrows provide landing gearposition information through the prox-imity switch on the mechanism thatlocks the actuating cylinders in theextended position.

In addition, a red light in the landinggear control handle blinks when at leastone landing gear is not down andlocked, either or both power levers areretarded, and the airspeed is lower than190 knots. This light and the gearposition indicator lights are testedtogether.

5I-85I-7

Falcon 10/100 For SimuFlite training only 5I-13March 2001

5I-14

BrakeSystem

NosewheelSteeringSystem


An aural warning sounds in either ofthe following conditions.� If one or more gear is not down andlocked when the power levers areretarded to idle and the airspeed isbelow 190 knots. This warning can besilenced by means of the horn-silencepushbutton.� When one or more gear is not downand locked and the slat/flap controlhandle is selected past 30°. In this case,the gear control handle light does notblink; this warning cannot be silenced.

Illumination of the red gear panellights indicates that the landing gearposition does not agree with the selec-tor handle. During extension or retrac-tion, illumination of the red lights indi-cates that the gear is in operation. Atthe end of the sequence, the light extin-guishes.

Illumination of the green lights forboth main gear and the nose gear indi-cates that the associated gear is downand locked.

EmergencyOperationEmergency extension of the landinggear is accomplished through amechanical emergency extension sys-tem; it requires neither hydraulic pres-sure nor electrical power to operate.The emergency gear extension handleconnects to cables directly linked toall three uplock mechanisms and thelanding gear control valve.

Emergency extension of the nose andmain landing gear is accomplished bypulling the EMERG-L/G handle in abox aft of the center pedestal (Figure5I-8). The gear then extends by grav-ity and is down-latched by the effect ofits own weight and by aerodynamicforces acting on the gear. Placing thelanding gear control lever in DOWNactivates the anti-skid system and land-ing lights.

NOTE: In an emergency,sideslip may be required forextension of the main gear.


ON

OFF

TEST ON

OFF

HYDR.ST. BY PUMP

EMERGENCY

P. BRAKEACCUMULATOR

HYD SYSTEM 2PRESSURE

PRESSURE SWITCH(1,200 PSI)

ANTI-SKIDCONTROL

BOX

SKID DETECTOR (4)

PILOT/COPILOTBRAKE PEDALS

NORMAL

ANTI-SKIDVALVE

NORMALBRAKINGDISTRIBUTOR

HYDRAULICSYSTEM 1PRESSURE

RETURN

EMERGENCY BRAKECONTROL VALVE

ANTI-SKIDVALVE

HYD SYSTEM 2RETURN

C BUS

HYD SYSTEM 1 PRESSURE

HYD SYSTEM 1 RETURN

ANTI-SKID BYPASS PRESSURE



P. BRAKE A BUS

BRAKE

PARKBRAKE

BRAKEL

BRAKER

21 3 4

STEERING WHEEL ROTATION

FROM 0 TO 60°

FROM 60 TO 120°

NOSEWHEEL STEERING

FROM 0 TO 6°

FROM 6 TO 55°

SERVOVALVE

FILTER

6G2 7G 6G1

CONTROL WHEEL

CONTROLUNIT

NOSEWHEELSTEERING

PRIMARY A BUS(NON-SHED)

STEERINGJACK

SWINGINGLEVER

POSITION PICKOFFPOTENTIOMETER

NOSEWHEELCENTERINGSPRING MECHANICALLY

CONTROLLEDVALVE

HYDRAULICSYSTEM 2

PRESSURE

RETURN

STEERING SYSTEM AMPLIFIER

Nosewheel Steering SystemBrake System



The Falcon 10/100 is equipped with amultiple disc brake assembly on eachof the four main wheels. Each assem-bly is independently actuated by thenormal or the emergency braking sys-tems and is controlled by the pilot orcopilot brake pedals or emergencybrake lever. Normal braking is pro-gressive and differential. Inadvertentbrake application is prevented by deadtravel of the pedals.

An auxiliary system, supplied by land-ing gear retraction hydraulic pressure,causes the braking distributor to deliv-er pressure through Hydraulic System1, thereby braking the wheels uponretraction.

The normal braking system, includingthe anti-skid system, is supplied byHydraulic System 1 pressure. Themaximum pressure available to thebrakes with System 1 is 1,740 PSI.

Hydraulic System 2 provides pressurefor the operation of the emergencybraking system. This system, con-trolled by the PARK BRAKE lever onthe left forward portion of the centerpedestal, provides progressive but non-differential braking, and has no anti-skid protection. The intermediate posi-tion can apply regulated pressure tothe brakes to produce drag withoutskidding for initial emergency brak-ing at high speeds.

Lever positions aft of the intermedi-ate detent provide progessively morebraking. The full aft position providesmaximum braking and locks thewheels in all conditions.

ComponentsEach brake assembly is hydraulicallyoperated using Hydraulic System 1 or2. The assembly is composed of a backplate subassembly, two rotating disks,

one stationary disk, a pressure platesubassembly, brake holding plate,torque tube and two brake housingsthat each contain four pistons. Eachbrake assembly contains four normalpressure ports, four emergency pres-sure ports, and bleed screws.

Braking DistributorA braking distributor transmits to thebraking lines a pressure modulated bythe compression of the brake pedalcontrol push-rods. The control push-rods are actuated either by the pres-sure of the pedals or by the retractionpressure of the landing gear (automaticbraking upon landing gear reaction).Hydraulic System 1 pressure is sup-plied to the distributor.

Brake System OperationHydraulic pressure forces the four pis-tons in a housing against the pressureplate to produce braking action. This inturn forces the disk stacks together tocreate a braking force between therotating and stationary disks. Whenhydraulic pressure is released, thereturn assemblies pull the pressureplate and pistons to the off position toallow the disks to release and permitthe wheel to rotate.

Parking/Emergency BrakeThe parking brake system is used intwo ways: to provide main wheelemergency braking and to providemain wheel braking when the aircraftis parked. Hydraulic System 2 providespressure to the parking brake systemthrough an accumulator. The accumu-lator provides pressure to operate themain wheel brakes with the parkingbrake handle when both hydraulic sys-tems fail or when all engines are shutdown.

Brake System


The red PARK BRAKE handle on thecenter pedestal (Figure 5I-9 ) controlsthe parking brake. The handle has twodetents; pulling it to the first stop locksthe handle in the intermediate positionand applies 550 PSI pressure to thebrakes. The second detent is at the endof travel of the handle and supplieshydraulic pressure of 1,885 PSI formaximum braking.

Anti-Skid Brake SystemThe anti-skid system prevents wheelsfrom locking by limiting the pressureapplied to the brakes by the hydraulicsystem. Maximum braking efficiencyis obtained when both wheels are in aslight skid or at maximum rate ofdeceleration short of skidding.

The system electronically controls twopressure control valves to continuous-ly vary braking pressure in responseto wheel rotation. There are separateanti-skid systems on the left and rightmain wheels; these are powered by theC bus. Each main wheel utilizes atachometer generator for wheel speedinformation for the electronic controlamplifier.

Servo Valves

The servo valves modulate brakingpressure. The control amplifier createsa variable bypass pressure betweenbraking pressure and return pressureby delivering electrical signals to thehydraulic distributor, which then sig-

nals the servo valves to modulate thebraking pressure.

Control Box

The electronic control box separatelycontrols the left and right servo valves.It develops commands from the sig-nals provided by the tachometer gen-erators relative to wheel speed. Wheeldeceleration provides the skiddingsignal; at the beginning of a skid, thecontrol box sends a signal to decreasebraking pressure to both brakes on thatgear and allow the wheel to acceler-ate.

Anti-Skid Operation

The anti-skid system provides fullymodulated braking for the mainwheels. The system consists of tach-ometer generators (one per wheel) anda signal receiving control amplifier.The amplifier interprets the electricalsignals from the tachometer genera-tors and sends corresponding signalsto the servo valves, which hold orreduce brake pressure proportionally.

When the gear selector lever is in UP,the anti-skid system is deactivated. Itis energized when the gear handle isdown and the wheel speed sensorsexceed 20 kts. The tachometer gener-ator on each main wheel senses thebeginning of a skid; through signalsthen provided to the servo valve oneach main gear, braking is modulatedto prevent main wheel skidding.

5I-115I-105I-9



The C bus supplies power to the elec-tronic control amplifier through theanti-skid switch on the upper left ofthe center pedestal (Figure 5I-10 ).

A proximity switch on the nose land-ing gear provides a ground/flight sig-nal to the control amplifier. When thecontrol amplifier receives an inflightsignal, it commands the servo valves toremain closed.

Nosewheel SteeringThe lower nose gear strut can behydraulically turned to steer the air-craft when taxiing. Steering is accom-plished by action of the double actu-ating steering cylinder on the nose gearstrut. Motion of the cylinder controlsdeflection of the wheel through theshock strut.

Steering can be operated only whenthe gear is extended. When the gear isretracted, steering is disabled and thewheel is centered for retraction. Aquick uncoupling device is incorpo-rated for towing. If the device is notdisconnected for towing and the turn-ing limits are exceeded, damage to thenosewheel steering occurs.

The electrically controlled andhydraulically operated nosewheelsteering system receives electricalpower from the A bus and hydraulicpower from Hydraulic System 2. Itconsists of the following components:� steering handwheel� control potentiometer� wheel position potentiometer� control amplifier� electro-distributor (solenoid valve)� mechanically operated valve� hydraulic servo system� steering cylinder� centering device.

The handwheel (Figure 5I-11 ) turns120° left or right of center. The first60° of handwheel movement in eitherdirection results in 6° of nosewheeldeflection. The subsequent 60° ofhandwheel deflection (full 120° travel)results in 55° of nosewheel deflectionin the selected direction. The hand-wheel is depressed against a spring toinitiate operation and, when released,returns to the up position.

A pair of potentiometers controls steer-ing when the handwheel is pressed androtated. The potentiometer on the con-trol wheel senses the angular positionof the steering wheel; the secondpotentiometer on the landing gear strutsenses the deflection of the wheelassembly. The potentiometers provideinputs to a control amplifier.

The control amplifier compares thesteering wheel position signal to thesignal from the wheel assembly anddelivers a signal to the servo valve tocommand the wheel to deflect accord-ing to the steering control commands.

Hydraulic System 2 pressure is ini-tially supplied through an electrodis-tributor (solenoid) valve. The solenoidvalve opens with:� power from the A bus� steering wheel pressed� all gear actuators locked down.

Closing the solenoid valve portshydraulic fluid to the return.

After leaving the solenoid valve, thehydraulic fluid enters the mechanicaldistributor that opens when the nosegear is downlocked. The distributorcloses when the landing gear deflects15° from the down-and-locked posi-tion and hydraulically disengages thesteering system as the gear retracts. Arod on the shock strut controls thismechanical valve.


Hydraulic fluid from the mechanicalvalve next enters the hydraulic servosystem. The servo system is on thenose landing gear strut next to the actu-ating cylinder and includes the steeringservo valve. This two-stage valve con-verts electrical command signals fromthe control amplifier, through a torquemotor and slider valve, to hydraulicpressures in the actuating cylinder. Theservo system also includes an auxil-iary system for hydraulic shimmydamping.

The steering cylinder, on the nose land-ing gear strut, is double-actuating. Itcontains a rack that engages a toothedgear integral to the shock absorber.

Nosewheel SteeringOperationWhen the aircraft is on the ground withA bus power and Hydraulic System 2pressure available, the nosewheel steer-ing system is armed for operation.

Holding the steering handwheeldepressed and rotating it creates dif-ferential signals in the potentiometersthat are sent to the control amplifier.Simultaneously, the electro-distribu-tor opens to supply hydraulic fluidthrough the open mechanical valve tothe hydraulic servo unit. The controlamplifier commands the servo unit tohydraulically reposition the cylinder,thus turning the lower nose gear strutand wheel until the two potentiome-ters sense the same position. The nosewheel then holds the deflected posi-tion until the handwheel is again turnedor released. If the handwheel isreleased, the nose gear wheel returns tocenter.

During takeoff when the steering hand-wheel is released, the electro-distribu-tor closes and system hydraulic pres-sure is removed. Upon gear retraction,the mechanical valve closes, whichalso releases system pressure.



PreflightDuring the exterior preflight inspec-tion, accomplish the following checksof the landing gear and brake system(see Preflight chapter for details):� Conduct visual inspection of nosegear assembly for fluid leaks, doorcondition and security, tire wear, nosegear wheel well condition, and tirepressure.� Ensure the steering pin is installedand locked.� Check main gear and doors for gen-eral security, fluid leaks, strut exten-sion, brake condition, and main geartire pressure.

During the cockpit preflight inspec-tion, test the landing gear annuncia-tors and aural warning with the land-ing gear TEST button on the landinggear control panel. Test the anti-skidannunciators with the same button.

Check that the landing gear controlhandle is down; check that three green“gear down” lights and no red lightsare illuminated.

Anti-Skid TestDuring taxiing and when clear of otheraircraft, depress one brake pedal, thenpush and release the ANTI-SKID testbutton; observe that the brakes releaseand engage and that the anti-skid lightsilluminate during release. Repeat forthe other brake pedal.

ServicingInflationTire inflation should be accomplishedaccording to the following:� Nose Wheel – 94 PSI� Main Gear Wheels –

135 PSI (18,740 lbs GW)

145 PSI (19,300 lbs GW)

StrutsThe landing gear struts are filled withMIL-H-5606 hydraulic fluid and inflat-ed with nitrogen. See MaintenanceManual for information on landinggear strut static extension and the fol-lowing:� Nose Strut – 1.48 to 2.56 IN� Main Struts – 5.26 to 6.15 IN

AbnormalProceduresThe following is a brief discussion ofvarious abnormal procedures for thelanding gear systems. For a list of spe-cific procedural steps, refer to the AFMor the SimuFlite Operating Handbook.

Landing GearDoes Not ExtendWhen all three green LOCKED DNlights fail to illuminate, hydraulic orelectrical problems are interfering withgear extension. After ensuring that thecontrol handle is correctly positionedin DOWN, that the CB is engaged, andthe lights are operative, the landing gearcan be extended using the followingemergency gear extension procedure.

Pull the manual extension handle (ina box aft of the center pedestal) tomechanically unlock the gear. The gearunits then extend by force of gravityand by aerodynamic forces. Once thegear is extended, verify their positionby observing green lights on the land-ing gear panel and no gear handle light.

Abnormal Gear RetractionIf the landing gear fails to properlyindicate UP after gear retraction ontakeoff (gear handle UP), it is possi-ble that ice or snow has accumulatedon a proximity switch. If this is sus-pected, maintain speed below 190KIAS and recycle the gear to shake orblow off the contaminants.



If the landing gear handle does notmove up, it is possible that the nosegear did not center; therefore, retrac-tion is not recommended. Return forlanding.

Anti-Skid System FailureIf either of the ANTI-SKID lights failsto indicate properly during a test, or ifthe lights indicate a malfunction dur-ing brake application, ensure that theCBs are set and the ANTI SKIDswitch is ON. Use landing perfor-mance data adjusted for anti-skid offoperations. Use caution when brakingwithout anti-skid protection to preventlocking a wheel.

Low Parking BrakeAccumulator PressureIf the amber P BRAKE annunciatorilluminates, parking brake accumula-tor pressure is less than 1,200 PSI.Under these conditions, the parkingbrake may not hold the aircraft in posi-

tion. Hydraulic System 2 should becharged by starting or motoring theright engine. Chock the aircraft’swheels to prevent rolling before start-ing the right engine.

Brake System FailureIf the normal brake system fails, theNo.2 (emergency) braking system,through the parking brake handle, canbe used to stop the aircraft. Duringemergency braking, the anti-skid sys-tem is inoperative. Landing distancecorrections must be determined withanti-skid system inoperative.

When using the parking brake handle,the first detent provides approximate-ly 550 PSI to the brakes while the aftdetent provides about 1,885 PSI. Usecaution when pulling the handle aft ofthe first detent during landing since noanti-skid protection is available. Usethe first detent for initial braking; whenthe aircraft has slowed sufficiently,increase braking by slowly moving thehandle aft of the detent.



DataSummaries

Landing Gear/Brakes/Steering SystemsLanding Gear System

Power Source A bus – gear selector and handle light

B bus – position lights

Hydraulic System 1 – landing gear actuators

Control Landing gear selector

Emergency gear T-handle

Horn silence button

Landing gear handle

Override pushbutton (overrides ground down-lockdevice; its use is discouraged

Monitor Hydraulic System 1 pressure/quantity gages

HYD.1 annunciator

Down-and-locked lights (3 green)

LDG GEAR MOVING lights (3 red)

Gear handle light (blinking red)

Warning horn

Protection Mechanically controlled emergency gear valve,

Ground down-locking device

Nose gear centering after takeoff

Proximity (squat) switches

Proximity (Squat) Switches

Power Source Primary A bus

Distribution Nose gear switchGear retraction protection

Left main switchAirbrake warningBattery blowerGear retraction protectionFreon condenser blowerPressurization controllerStall system 1Thrust reversers

Right main switchConditioning valveHydraulic standby pumpSTAB TRIM annunciatorStall system 2

Control Automatic via extension/compression of three landinggear struts


Brakes/Nosewheel Steering System

Power Source A bus – nosewheel steering

C bus – anti-skid

Hydraulic System 1 – normal brakes

Hydraulic System 2Emergency brakePark brakeAccumulatorNosewheel steering

Control Brake pedals

PARK BRAKE leverFirst detent – emergency brakesSecond detent – parking brake

Anti-skid switch

Anti-skid test button

Nosewheel steering handwheel

Monitor Hydraulic Systems 1/2 pressure/quantity gages

AnnunciatorsHYD 1/2 (amber)ANTI-SKID BRAKE L/R (amber)BRAKE (red)P. BRAKE (amber)

Protection Steering prevention via gear-extend microswitches

CAUTION: Nosewheel steer-ing control wheel deflection doesnot correspond to nosewheeldeflections.

Falcon 10/100 For SimuFlite training only 5J-1April 2000

MiscellaneousSystems

Chapter 5J

The Miscellaneous Systems chapter provides information inthe following areas: Oxygen System, Thrust Reverser System,Emergency Equipment, and Warning System.

The Oxygen System provides supplementary oxygen for thecrew and passengers.

The Thrust Reverser System provides additional decelerationforce to assist braking.

The Emergency Equipment on the aircraft includes safety equip-ment, fire extinguishers, and life vests.

The Warning System provides warning of aircraft equipmentmalfunctions, indication of a condition requiring immediateattention, or indication that a system is operating.

5J-2 For SimuFlite training only Falcon 10/100April 2000


Miscellaneous Systems

Table ofContents

Oxygen System Schematic . . . . . . . . . . . . . 5J-6

Oxygen System . . . . . . . . . . . . . . . . . . . 5J-7

Components . . . . . . . . . . . . . . . . . . . . . 5J-7

Oxygen Bottle . . . . . . . . . . . . . . . . . . . . 5J-7

Filling Adapter . . . . . . . . . . . . . . . . . . . . 5J-7

Filling Pressure Gage . . . . . . . . . . . . . . . . 5J-7

Shutoff Valve . . . . . . . . . . . . . . . . . . . . 5J-8

Discharge Indicator . . . . . . . . . . . . . . . . . 5J-8

Right Console Pressure Gage . . . . . . . . . . . 5J-8

Pressure Reducer Valve . . . . . . . . . . . . . . 5J-8

Crew Masks . . . . . . . . . . . . . . . . . . . . . 5J-8

Passenger Mask Box and Mask . . . . . . . . . . . 5J-9

Passenger Control and Distribution . . . . . . . . . . 5J-9

Passenger Oxygen Control Unit . . . . . . . . . . 5J-10

Robertshaw Regulator . . . . . . . . . . . . . . 5J-10

EROS . . . . . . . . . . . . . . . . . . . . . . 5J-10

Therapeutic Masks . . . . . . . . . . . . . . . . . 5J-10

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5J-11

Complete Test . . . . . . . . . . . . . . . . . . . 5J-11

Crew . . . . . . . . . . . . . . . . . . . . . . . 5J-11

Passenger . . . . . . . . . . . . . . . . . . . . 5J-11

Short Test . . . . . . . . . . . . . . . . . . . . . 5J-11

High Altitude Airports . . . . . . . . . . . . . . . 5J-12

Servicing . . . . . . . . . . . . . . . . . . . . . . . 5J-12

Abnormal Procedures . . . . . . . . . . . . . . . . 5J-12

Masks Do Not Drop Out Automatically . . . . . . . 5J-12

Air Conditioning Smoke/Electrical Smoke . . . . . 5J-13

Oxygen System Data Summary . . . . . . . . . . 5J-15


Thrust Reverser Schematic . . . . . . . . . . . . 5J-16

Thrust Reverser System . . . . . . . . . . . . . . 5J-17

Components . . . . . . . . . . . . . . . . . . . . . 5J-17

Thrust Reverser Lever . . . . . . . . . . . . . . . 5J-17

Hydraulic Actuator . . . . . . . . . . . . . . . . . 5J-17

Secondary Latch Solenoid . . . . . . . . . . . . . 5J-18

Indicating and Warning System . . . . . . . . . . 5J-18

Warning Horn . . . . . . . . . . . . . . . . . . . 5J-18

Accumulator and Control Valve . . . . . . . . . . 5J-18

Operations . . . . . . . . . . . . . . . . . . . . . . 5J-18

Deployment . . . . . . . . . . . . . . . . . . . . 5J-18

Stowage . . . . . . . . . . . . . . . . . . . . . . 5J-19

Throttle Snatch . . . . . . . . . . . . . . . . . . . 5J-19

Emergency Stow Switches . . . . . . . . . . . . . 5J-19

Reverser Operating Characteristics . . . . . . . . 5J-19

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5J-19

Emergency Procedures . . . . . . . . . . . . . . . 5J-19

Inadvertent Thrust Reverser Deployment . . . . . 5J-19

Thrust Reverser Deployment – Climb or Cruise . . 5J-20

Thrust Reverser Data Summary . . . . . . . . . . 5J-21

Emergency Equipment . . . . . . . . . . . . . . . 5J-23

Safety Equipment . . . . . . . . . . . . . . . . . . 5J-23

Overwater Equipment . . . . . . . . . . . . . . . . 5J-23

Warning Systems . . . . . . . . . . . . . . . . . . 5J-25

Visual Warning Systems . . . . . . . . . . . . . . . 5J-25

Altitude Alerter . . . . . . . . . . . . . . . . . . . 5J-25



Aural Warning System . . . . . . . . . . . . . . . . 5J-25

Components . . . . . . . . . . . . . . . . . . . . 5J-25

Operation . . . . . . . . . . . . . . . . . . . . . 5J-26

Horn Silence Cutoff Pushbutton . . . . . . . . . 5J-26

Landing Gear Warning . . . . . . . . . . . . . . 5J-26

Cabin Pressure Warning . . . . . . . . . . . . . 5J-26

Fire Warning . . . . . . . . . . . . . . . . . . . 5J-27

Stall Warning . . . . . . . . . . . . . . . . . . . 5J-27

VMO/MMO Warning . . . . . . . . . . . . . . . . 5J-27

Altitude Deviation Warning . . . . . . . . . . . . 5J-27

Horizontal Stabilizer Warning . . . . . . . . . . 5J-27

Test . . . . . . . . . . . . . . . . . . . . . . . . . 5J-27

Voice Advisory System . . . . . . . . . . . . . . . 5J-28

Vocabulary . . . . . . . . . . . . . . . . . . . . . 5J-28

Annunciator Cross Reference . . . . . . . . . . . 5J-33

LOW PRESSURE (>70 PSI)

HIGH PRESSURE

OVERPRESSURE DISCHARGE

O

NSURG

E

USE NOOIL

PU

SH

TO

TU

RN

OFF MAN

AUTO+

PASSENGEROXYGEN

PASSENGER OXYGEN

EMER NORM

TURN TO NORM (WHEN INDICATEDOXYGEN PRESSURE LESS THAN 600 PSI)

OXYGEN PSI x 100

5 10 15 20

FILLINGPRESSURE

GAGE

INSTRUMENT PANEL OXYGENPRESSURE INDICATOR

DILUTION OR 100%OXYGEN CONTROL SWITCH

MASKSTOWAGEBOX

PILOT'S ORCOPILOT'SCONSOLE

ROBERT SHAW CONTROLLER

FILLINGADAPTER

VENT LINEDISCHARGEINDICATOR

OXYGENBOTTLE

PASSENGEREMERGENCYOXYGEN OVERRIDE(ROBERTSHAWSYSTEM ONLY)

COPILOT'SMASK

TOILET MASK-BOX (OPTIONAL)

AUTOMATIC DROP-OUTMASK BOX (2 MASKS)

PILOT'SMASK

PRESSUREREDUCER

1

0

50 100

150

OXYGENEROS

BAR PSI

5001000

1500

22000°C

-40

-30

-20

-10

0

PSI

1,465

1,530

1,590

1,650

1,715

°C

10

20

30

40

50

PSI

1,775

1,835

1,900

1,960

2,020

MAXIMUM BOTTLE PRESSURE

1


Oxygen System



Oxygen is available to the crew ondemand anytime and is available to thepassengers either manually throughcockpit selection or automatically oncabin depressurization. The oxygensystem meets requirements for oxy-gen during emergency descent due topressurization loss or if smoke is inthe cabin. The system permits oxygento be used for therapeutic purposes.

ComponentsThe oxygen system consists of:� oxygen bottle� brass filler valve/non-return valve� servicing cap/filling pressure gage� shutoff valve� discharge indicator/ejection disc� (HP) safety valve� right instrument pressure gage� reducer valve� EROS masks/mask boxes� passenger oxygen controller unit� constant flow passenger masks� therapeutic masks� Robertshaw/EROS regulators.

Oxygen BottleA removable oxygen bottle on the rightside of the nose cone (Figure 5J-1 )supplies oxygen to the crew and pas-

sengers. The oxygen bottle has acapacity of 49.8 cubic feet with a nor-mal operating pressure of 1,850 PSIat 21°C (70°F).

Two small holes in the oxygen bottlecompartment provide moisturedrainage and prevent pressurization byoxygen in the event of a leak. For ser-vicing, an external access panel on theright side of the nose, forward of thebottle, provides access for inspectionand shutoff valve operation.

Filling AdapterBottle servicing is accomplishedthrough a brass filler valve (Figure 5J-2). The filler valve is fitted with a built-in non-return valve that ensures tight-ness of the oxygen bottle when theoxygen supply unit is disconnected. Aservicing cap protects the fillingadapter. A data plate adjacent to thefilling adapter gives pressure vs. temp-erature information.

Filling Pressure GageNext to the filling adapter is a fillingpressure gage that indicates bottle pres-sure when the high pressure shutoffvalve is open. The pressure gageincludes a circular dial with tworanges, the first being a white rangewith up to 2,176 PSI (150 bar) in 145PSI (10 bar) graduations, and the sec-ond a green range in 500 PSI gradua-

OxygenSystem

5J-1 5J-2


tions to 2,200 PSI and expandedbetween 1,850 and 2,200 PSI.

Shutoff ValveThe oxygen bottle is fitted with a slow-opening shutoff valve which is nor-mally in the open position and safety-wired; it is not accessible in flight.

Discharge IndicatorThe discharge indicator (Figure 5J-3 )has a green ejection disc connected tothe high pressure (HP) safety valve onthe oxygen bottle. When the safetyvalve cracks open for a thermal over-pressure exceeding 2,500 to 2,700 PSI,the disc is ejected when the pressureupstream of the indicator reaches 39to 145 PSI.

Right ConsolePressure GageA rectangular gage on the copilot’sinstrument panel (Figure 5J-4 ) is grad-uated from 0 to 2,000 PSI with tworanges: the red range (below 250 PSI)indicates a low oxygen content; theyellow range (above 2,000 PSI) indi-cates that the bottle is charged abovethe rated pressure.

A pressure of 1,850 PSI at 21°C (70°F)corresponds to 2,000 PSI at 118°F.This range is the maximum system

pressure with the aircraft on theground. Bottle pressure must thereforebe checked and if necessary partiallyrelieved to prevent the safety dischargedisk from blowing out.

Pressure Reducer ValveA 70 PSI piston-type HP reducervalve, fitted with a venturi device,between the oxygen bottle (HP) andthe low pressure (LP) lines maintainsconstant pressure in the oxygen distri-bution system. The pressure reducervalve prevents high pressure withinthe bottle from reaching the crew orpassenger masks.

Crew MasksThe crew system consists of two EROSmasks inside the associated pilot/copilot mask boxes (Figure 5J-5 ) andthe connecting lines from the pressurereducing valve. The mask boxes are onthe left and right consoles.

Oxygen supply and mask microphoneconnections are in the bottom of themask boxes. A portion of the face ofthe mask to protrudes from an open-ing in the mask box doors. Two redtabs allow the mask to be grasped andremoved from the box while simulta-neously inflating the harness for quickdonning (Figure 5J-6 ).

5J-3 5J-4

Falcon 10/100 For SimuFlite training only 5J-9March 2001


After the mask is donned and the redtabs released, the inflated harnessdeflates. A miniature flow indicator inthe mask oxygen line indicates a pos-itive pressure with a green stripe.

A self-contained regulator on the faceof the mask has a two-position selec-tor control tab. When the control tabis pushed, the regulator supplies 100%oxygen on demand. When the controltab is pulled out (in the N position),the regulator supplies on demand oxy-gen mixed with cabin air proportionalto the cabin altitude up to 30,000 ft, atwhich point either position provides100% oxygen. Above approximately33,000 ft, the regulator automaticallyprovides pressure breathing of 100%oxygen. Oxygen consumption forflight above 41,000 ft is based on thecontrol tabs being in the N position.

A red test button on the face of themask allows a test of the pressurebreathing feature. When it is pushedin, listen for oxygen flow hiss.

Oxygen is metered at 70 PSI to thecrew at all altitudes. At higher alti-tudes, passenger oxygen is metered at70 PSI. At lower altitudes, passengeroxygen is metered at 27.5 PSI.

Passenger Mask Boxand MaskA passenger mask box is above eachseat position (Figure 5J-7 ); a mag-netically latched cover normally con-ceals the box. When sufficient pres-sure is in the passenger oxygendistribution system, an actuator drivesthe cover clear of the box and releas-es an internal door to allow the maskto fall clear of the box. The lanyardattached to the mask must be pulled tomake oxygen available at the mask.At high altitudes, oxygen is deliveredat 70 PSI at 3.2 liters per minute. Atlow altitudes, oxygen is delivered at27.5 PSI at 1.025 liters per minute.

Passenger Controland DistributionA typical installation of the oxygencontrol and distribution system is com-posed of:� passenger oxygen controller unit� seven constant flow passengermasks� optional passenger mask with theretractable toilet� optional therapeutic mask (EROSonly).

5J-5 5J-6 5J-7

Oxygen LimitationWhen the aircraft altitude isabove 41,000 ft, one pilot mustwear his oxygen mask at alltimes.

5J-10 For SimuFlite training only Falcon 10/100March 2001

Passenger OxygenControl UnitThe copilot controls oxygen delivery tothe passenger masks through an oxy-gen controller on his instrument panel.With an oxygen pressure of 43 PSI orgreater, the latch opens the internaldoor and the passenger masks fall,forcing the magnetic door to open.

Two brands of controllers are used forthe Falcon 10/100: a Robertshaw oran EROS. Both deliver oxygen to thepassenger masks, but pressures andflow rates differ slightly.

Robertshaw RegulatorRobertshaw type regulators (Figure5J-8) are equipped with a mode selectlever labeled AUTO/OFF/MANUALand a pressure indicator, labeledSURGE SYSTEM ON.

Regulator function is based upon thecabin altitude of 10,700 ft and modelever position. With the lever inAUTO, the mask boxes open and oxy-gen is delivered at a pressure depend-ing upon cabin altitude (the pressuremay range from 8 PSIG to 42 PSIG).With the lever in MAN, the masksdeploy and the oxygen is deliveredregardless of cabin altitude. Pressureschedule remains in effect.

EROSEROS regulators have a four positionmode lever labeled NORMAL/OVERRIDE/FIRST AID/CLOSEDand an ARMED and SUPPLY indica-tor. Operation/indications from thelever and indications are listed in Table5J-A .

The RESET button enables the auto-matic circuit to be reset if the pressuresensitive valve is functioning (controlswitch on NORMAL). After any 70PSI operation following a cabindepressurization, set the control switchto NORMAL and press the RESETbutton.

Delivery pressure is either 72.5 PSIabove 18,000 ft cabin altitude or 27.5PSI below 18,000 ft. The use of twopressure schedules and mechanicalindicators enhances the simplicity andeffectiveness of regulator operation.

Therapeutic MasksOn aircraft with the EROS system,a therapeutic mask and outlet are inthe coat rack. The therapeutic maskmay be connected to the therapeuticoutlet without opening any doors orreleasing any latches. When the con-troller selector is in FIRST AID, oxy-gen flows from the therapeutic maskat reduced pressure (27.5 PSI).

5J-8

Oxygen LimitationAbove flight level 410, allpassengers must wear an oxygenmask secured to the face, andconnected to the oxygen system.

CAUTION: On EROS regula-tors, do not press the RESETbutton when the system isfunctioning because bleeding thecontrol chamber immediatelystops oxygen flow.



PreflightThe high pressure valve is slowlyopened (5 to 10 seconds) or checkedopen during preflight. Perform thecomplete test of the crew and passen-ger oxygen systems before the firstflight of the day and the short test forsubsequent flights.

Complete TestCrew

Before the first flight of the day, per-form the following procedures.� Remove the crew masks from theirboxes and don.� Breathe oxygen with the regulatorin N.� Select 100%, then press the testbutton to provide a mask pressure ofapproximately 70 PSI above baro-metric.� If pressure read by the gage dropsabruptly during the test, check that theoxygen system shutoff valve is open,if closed, the oxygen consumed dur-ing the test comes from the high pres-sure circuit, which is soon emptied.� Check the microphone; selectMASK and C’PIT.� Set the regulator to 100%.

Passenger� Check that the passenger oxygencontroller is in NORMAL or AUTO.� Check the ARMED and SUPPLYindicators; if the ARMED indicator isvisible, press the RESET button toblack it out.

Short TestPerform the procedures listed belowto test the crew masks in their boxes.� Switch the audio control panel toMASK and C’PIT and briefly pressthe red TEST button on the mask.Oxygen flow in the mask causes a hiss-ing noise that can be heard in the head-set or cockpit speakers.� Check regulator is in 100%.

To avoid passenger mask deployment,do not perform a test of the passengeroxygen system prior to takeoff. Whencabin altitude has decreased to lessthan 10,000 after takeoff, move thepassenger oxygen selector to normal(counter-clockwise).

High Altitude AirportsBefore landing on an airfield at analtitude of approximately 10,000 ft orhigher, move the controller to CLOSEDprior to decompressing the cabin toprevent automatic drop of the masks.

CAUTION: The bottle shutoffvalve must always be openedvery progressively (in at least 5to 10 seconds).

WARNING: Personnel shouldalways remember: Oxygen +Hydro carbons = Explosion.

Table 5J-A; EROS Operation/Indications

MODE Operation

CLOSED Passenger supply is shut off.

OVERRIDE Delivery circuit overrides NORMAL to supply masks.

NORMAL Delivery circuit functions as a result of cabindecompression (over 10,700 ft). Masks deployand oxygen is supplied automatically.

ARMED Indicator Opens to show passenger masks are receivingoxygen.

FIRST AID Delivery circuit provides oxygen to the therapeuticmasks if used; normal function for cabin remainsactive.

SUPPLY Indicator Opens to show passenger masks are receivingoxygen.


ServicingCertain safety precautions must betaken when servicing the oxygen sys-tem. No greasy substances should be inthe vicinity. Check that the filling valveis clean. Set the oxygen bottle cartabout 7 ft (2 meters) from aircraft oxy-gen filling valve.

The high pressure valve and the groundcart valve must be actuated slowly.Filling pressure is monitored with pres-sure gages. Pressure corrections fortemperature must be taken into accountand can be read on the placard on theoxygen bottle access door.

Unscrew the filling valve cap and con-nect the filling line and adapter to thefilling valve. Slowly open the shutoffvalve on the oxygen bottle on the cart.Then slowly open the shutoff valveson the oxygen bottles. Graduallyincrease bottle pressure by turning thepressure adjustment screw of the pres-sure reducing valve on the cart. Checkand compare data given by the oxy-gen pressure indicator on the instru-ment panel (copilot side) and HP oxy-gen bottle pressure gage. Rated aircraftoxygen bottle pressurization is 1,850PSI (127 bar) at 70°F (21°C).

When rated pressurization is reached,close the shutoff valve and oxygen cartbottle shutoff valves. Disconnect theoxygen filling line and replace the capon the filling valve. Use a leak detec-tor fluid to ensure there are no leaks.Close the access doors.

To comply with regulations for suffi-cient oxygen in case of cabin fire, aminimum pressure of 965 PSI is

required to dispatch a flight below andabove 10,000 ft MSL. For flightsabove 41,000 ft MSL, a chart or tableis provided for minimum pressure inthe system for dispatch.

AbnormalProceduresThe following section provides briefdiscussion of what happens to the oxy-gen system during abnormal condi-tions. For a list of specific proceduralsteps in abnormal conditions, pleaserefer to your SimuFlite OperatingHandbook.

Masks Do Not DropAutomaticallyPosition the passenger oxygen selectorto OVERRIDE or MANUAL to over-ride NORMAL/AUTO and deploy thepassenger masks. After the passengermasks are deployed, ensure passen-gers don and check the masks for prop-er operation. Position the passengeroxygen selector to FIRST AID todeliver oxygen to the therapeutic out-let and to maintain normal delivery ofoxygen to the cabin passenger masks.

With the Robertshaw system,oxy-gen does not flow if the cabin altitudeis below approximately 10,000 ftunless the bypass valve on the copilot’sside panel is opened (Figure 5J-9 ).

The AUTO/OFF/MANUAL selectorselects passenger mask drop mode anda flow rate dependent on altitude. Openthe bypass valve on the copilot’s sidepanel to ensure maximum flow rate.

WARNING: During oxygenfilling, no one should be insideaircraft. Oxygen filling is to beperformed in a well ventilatedarea. Aircraft-installed oxygenbottle shutoff valve should beopened slowly. Make sure thataircraft is ground-connected,aircraft power system is de-energized, and no refueling ordefueling is under way.

CAUTION: Rated oxygen bottlepressurization depends on tem-perature. Refer to the OxygenCylinder Pressure as a Functionof Temperature chart in theMaintenance Manual beforerefilling oxygen; the safety valverupture disk will burst if pres-sure reaches range of 2,553 to2,698 PSI.

WARNING: Filling valve capmust be closed properly toensure tightness.

CAUTION: Excessive move-ment of oxygen controller maycause wear and leaks.

CAUTION: Do not use theRESET pushbutton in flightwhen oxygen system is operat-ing. To stop the flow of oxygen:position the passenger oxygenselector to the CLOSED position.



Air Conditioning Smoke/Electrical SmokeThe initial steps for all procedures deal-ing with smoke start by neutralizingthe life threatening situation. The crewdons oxygen masks and selects 100%to breathe 100% oxygen. The pilotsthen select MASK and C’PIT on theaudio panel to ensure communication.With MASK and C’PIT selected, themicrophone in the oxygen mask is acti-vated to a hot interphone system; the

hot interphone is heard over the speak-er (if it is selected) or over the headset.

If there are no flames in the cabin, thepassenger oxygen masks are deployedby selecting MANUAL or OVER-RIDE on the oxygen control unit. Thisaction directs 8 to 42 PSI (Robertshaw)or 28 to 72 PSI (EROS) to the passen-ger oxygen system for mask deploy-ment and a maximum flow of up to9.0 liters per minute. To ensure maxi-mum flow, open the bypass on thecopilot’s side panel.

5J-9

CAUTION: Oxygen must notbe used when there are flamesin the cabin or cockpit.


5J-16


Thrust ReverserSystemData SummaryOxygen System

Power Source Oxygen cylinder (1,850 PSI at 70°F)

Distribution Crew/passenger oxygen masks

Passenger oxygen system

Therapeutic oxygen mask (EROS)

Control Oxygen cylinder shutoff valve

Crew masks (N/100% PUSH)

Passenger oxygen control unitAUTO/OFF/MANUAL (Robertshaw)NORMAL/OVERRIDE/FIRSTAID/CLOSED (EROS)

Monitor Bottle pressure gage

Cockpit oxygen pressure gage

Bottle safety discharge disc

Protection Bottle safety valve (2,500 to 2,700 PSI)

WARNING: During oxygen fill-ing, no one should be inside air-craft. Perform oxygen filling in awell-ventilated area. Open bottleshutoff valve slowly, taking at least5 to 10 seconds. Valve cap mustbe closed properly to ensure tight-ness. Ensure aircraft is ground-connected, aircraft power systemis de-energized, and no refueling ordefueling is underway.

WARNING: Oxygen contactwith hydrocarbons causes anexplosion.

CAUTION: Excessive move-ment of oxygen controller maycause wear and leaks.

CAUTION: On EROS control-ler, do not use the RESET push-button in flight when oxygen sys-tem is operating. To stop the flowof oxygen, position the selectorto the CLOSED position.

CAUTION: Rated oxygen bottlepressurization depends on temp-erature. Refer to the Oxygen Cyl-inder Pressure as A Function ofTemperature Chart in the Main-tenance Manual before refillingoxygen. The safety valve rupturedisk can burst if pressure reach-es range of 2,553 to 2,698 PSI.

REVERSETHROTTLE

MAIN THROTTLE

REVERSEPIGGYBACKUNLATCH

MICROSWITCHES

HORN

HORNTEST

BUTTON

CONTROLRELAYS

TELEFORCE CABLE - NORMAL & REVERSE THROTTLE

EXTEND - STOW

STOW

SECONDARY LATCH(PRIMARY WITHIN HYDRAULIC CYLINDER)

UNLATCH SOLENOID

DEPLOYED

REVERSECONTROLVALVE(ELECTRIC)

ACCUMULATOR

HYDRAULICSYSTEM 2

EMERGENCYSTOW

SWITCHEMERGENCYSTOW

T/RCONTROL

SQUATSWITCH

HYDRAULIC SYSTEM 2 DEPLOY PRESSURE

HYDRAULIC SYSTEM 2 RETRACT PRESSURE

NITROGEN CHARGE



RETURN

MAINBUS

BAT 1

A BUS

TRANS

STOWED

REVERSE

FUELCONTROL

TELEFORCECABLE

THROTTLE SNATCH


Thrust Reverser System



Many Falcon 10/100 aircraft have beenequipped with the Grumman Aero-space conventional target-type designthrust reverser system (Figure 5J-10 ).The system is electrically controlledand hydraulically actuated. It provideseffective braking force and can be usedinstead of brakes on longer runways.Since the reverser is quite loud in thecabin at higher power settings, fullreverser power is rarely used.

The thrust reverser doors replace theaft cowling on the nacelle and formthe upper and lower skins of the aftcowl when stowed. The doors openusing Hydraulic System 2 pressure.The doors direct engine exhaust up,down, and forward.

ComponentsThe thrust reverser system is composedof:� thrust reverser levers� reverser power levers� SQUAT switch� control valve and hydraulic actuator� stowing latch and secondary latchsolenoid� warning horn

� hydraulic accumulator with pres-sure gage� microswitches� throttle snatch cable� teleforce cable.

Thrust Reverser LeversThrust reverser levers (piggybacklevers) on the engine power levers(Figure 5J-11 ) control the operationof the thrust reverser system. The thrustreverser release latches are below thethrust reverser levers.

The thrust reversers are enabled onlywhen the main power levers are at idle.Lifting the latch and moving the leversup about 1/2 inch contacts a micro-switch. If the left hand main squatswitch is closed, power is routed tothe deploy side of the reverser controlvalve. After the reverser is deployed,the power lever can be moved furtheraft to increase engine RPM andreverse thrust.

Hydraulic ActuatorA hydraulic actuator provides motionto the doors and contains the primarystowing latch. Deploy hydraulic pres-sure is required to unlatch, as well asextend the doors.

ThrustReverserSystem

5J-10 5J-11

Thrust Reverser LimitationUse of the thrust reversers islimited to ground operations onlyand may only be used for 60seconds total within a 30 minuteperiod. Using thrust reversers toback up is not permitted. Powermust be retarded to idle at speedslower than 60 knots. The powerlever retarder must be operative.


Secondary LatchSolenoidAs a backup to the hydraulic actuator,a solenoid-controlled latch is in eachreverser; the latch consists of a hookto hold the doors closed and an electricsolenoid that is energized to removethe hook. Unless the thrust reverserlever is in REVERSE and the revers-er door is fully deployed, the amberTRANS light illuminates when thishook is not latched.

Indicating andWarning SystemThe thrust reverser indicator panel onthe instrument panel includes an amberTRANS advisory light and green REVadvisory light (Figure 5J-12 ). TheTRANS light illuminates when thedoors are unlocked but not fullydeployed, and the green REV adviso-ry light illuminates when the doors arefully deployed. The thrust reverserHORN TEST button next to the thrustreverser indicator panel powers thehorn directly to verify operation.

Warning HornA non-silenceable independent warn-ing horn in the thrust reverser systemsounds whenever there is disagreementbetween power lever position andthrust reverser door position; the hornsounds if the main power levers are

forward and the doors are deployed orif the thrust reverser levers are in thereverse power range with the doorsstowed.

Accumulator andControl ValveHydraulic System 2 provides hydraulicpressure for thrust reverser actuation.The thrust reverser accumulator (partof Hydraulic System 2) in the aft com-partment is fitted with a pressure gageand charged with nitrogen (1,200 PSIwhen hydraulic pressure is zero). Theaccumulator allows emergencystowage of the reverser doors if theHydraulic System 2 pressure is notavailable. The reverser control valveis electrically controlled by micro-switches in the power lever quadrant.

OperationsDeploymentThrust reverser deployment requiresthe aircraft to be on the ground uponlanding or before V1 for an accelera-tion stop. Power levers must be set toidle. A mechanical latching systemprevents operation of the thrust revers-er when these conditions are not sat-isfied. To operate thrust reverser, movethe main power levers to idle. Graspreverse thrust power levers and therelease latches and lift the latches toallow the reverser levers to be moved

5J-12

CAUTION: Do not move thethrust reverser levers beyondreverse idle until the REV lightilluminates. If the horn sounds,do not add reverse thrust.



aft. The first 1/2 inch of travel contactsmicroswitches that deploy the revers-er doors. Further aft movement ofreverse power levers increases engineRPM and reverser thrust.

StowageThe thrust reversers are stowed bymoving the thrust reverser levers tothe idle reverse position and lifting thelatches to permit restow of the powerlevers and reverser doors.

Throttle SnatchAn automatic power lever retardingdevice, or throttle snatch, returns eithermain power lever or reverser lever toidle if the reverser doors move to thewrong position. If the main powerlever is in forward thrust position anda door opens, the power lever moves toidle. If the reverser lever is in thereverse position and the doors close,the reverser lever moves to idle. Bothare accomplished mechanically by ateleforce cable that connects the doorand the power lever quadrant.

Emergency Stow SwitchesThe guarded red EMERGENCYSTOW switches on the thrust reverserindicator panel bypass the power leversquadrant controls and normal thrustreverser control system to apply elec-trical power directly to the stow sideof the reverser control valve and hy-draulic pressure directly to the actuatorto stow the respective thrust reverser.

An independent source of power isprovided for these switches throughseparate circuit breakers on later thrustreverser systems or systems modifiedwith GAC F10 TR 78-019.

Reverser OperatingCharacteristicsThe aircraft weathercocks away froma crosswind on a slippery runway. Thethrust reverser causes a noticeablepitch-up tendency when deployed. Itis possible to scrape the tail andreversers on the runway if nose-downpressure is not applied.

PreflightFor a list of specific preflight proce-dural steps, refer to the SimuFliteOperating Handbook.

During the exterior inspection, checkthe general condition of the thrustreverser doors. Verify that the doorsare fully stowed. Check for hydraulicleaks.

EmergencyProceduresFor a list of specific emergency pro-cedural steps, refer to the SimuFliteOperating Handbook.

Inadvertent ThrustReverser DeploymentDuring TakeoffTake immediate action after V1 to con-trol the aircraft by positioning the nosefull nose-down to counter the pitch-up tendency. Control roll and yaw.Activate the EMER STOW switch. Ifthe thrust reverser stows, leave theemergency stow switch in EMERSTOW and land. If the thrust reverserdoes not stow, leave emergency stowswitch in EMER STOW and land assoon as possible.

Thrust ReverserOperational LimitationsDo not exceed: 60 seconds ofcontinuous or accumulatedoperation within a 30 minuteperiod; 87% N1 for speedsexceeding 60 KIAS at idle; 40%N1 for speeds below 60 KIAS;50% N1 for ground checks of theautomatic power lever retarder.

Flight with the automatic powerlever retarder inoperative is onlyauthorized with the thrustreversers stowed and bolted.

Thrust ReverserAsymmetrical OperationLimitationUse of the thrust reverser on oneengine only is authorized on dryrunways with the nosewheelsteering system available. In thiscase, observe the same limitationsand operational recommendations,with one control lever, as fornormal use of the thrust reversers.


At higher gross weights with slats andflaps at 15°, it may not be possible tomaintain level flight with the engineoperating at idle reverse thrust.

Thrust ReverserDeployment –Climb or CruiseDuring inflight deployment, the auto-matic power lever retarder on therespective lever automatically retardsthe affected lever to idle. The aircrafttakes a nose-up attitude with possiblewing roll, abnormal buffeting, andvibrations. Immediate action must betaken to control the rolling and yaw-ing. Set the emergency stow switch toEMER STOW. Maintain airspeedbelow 190 KIAS, and ensure theaffected engine is set to idle. If thethrust reverser stows, leave the emer-gency stow switch in EMER STOWand maintain engine thrust as required.Land as soon as possible.

If the thrust reverser does not stow,leave the emergency stow switch inEMER STOW and land as soon aspossible. Shutdown of the affectedengine is optional, but idle reversedecreases aircraft performance.

To silence the warning horn, pull thethrust reverser system L and R IND 5CBs. This also deactivates the thrustreverser indicator lights. Do not usereverse thrust for landing under thiscondition.

Follow the procedures for approachand landing with one engine operative,with the following differences:� VREF – adjust + 20 KTS� slats and flaps 15°� delay landing gear extension untilfinal approach� when landing is assured, set slatsand flaps 30°.



DataSummary

Thrust Reversers

Power Source Hydraulic System 2 pressure

Aircraft without SB 0154A bus (28V DC)No. 1 battery

Aircraft with SB 0154A busMain bus

Control No. 1 battery switch (aircraft without SB 0154 )

HORN TEST button

EMERGENCY STOW switch

Main throttle lever

Thrust reverser throttle lever

Monitor Reverser warning horn

Reverser TRANS/REV lights

Hydraulic System 2 pressure gage

Battery voltmeter (Main bus voltage)

Protection Secondary latch

Left main squat switch

Throttle auto retard (Snatch)

Reverser maximum RPM limit stop



Emergency equipment on the aircraftincludes safety equipment, fire extin-guisher (see Fire Protection chapter),and life vests.

Safety EquipmentThe following items of safety equip-ment are standard:� axe (most aircraft)� first-aid kit � three pairs of smoke goggles incockpit.

OverwaterEquipmentA life vest (Figure 5J-13 ) is storedunder each seat in the cabin. Pilot andcopilot life vests are under each seat.

To make passenger evacuation throughemergency exits easier, survival cordsare included in the emergency equip-ment. They are stowed in the closetbehind the copilot’s seat. The cords areattached to the seat rails, then stretchedto the wing leading edges where theycan be attached to the anchor point.

EmergencyEquipment

WHISTLE

LOCATINGLAMP

BATTERY

STRAPS

INFLATINGMOUTH-PIECE

CO BOTTLE (FORINFLATING VEST)

INFLATINGPULL BALL

2

5J-13



Visual WarningSystemsThe Failure Warning Panel alerts thecrew of a configuration or operationmalfunction in a system. Early aircraftwere equipped with a 23-annunciatorpanel (22 amber, 1 red) and three otherred annunciators on the instrumentpanel. The annunciators on later air-craft consist of 26 or 28 individualannunciators (24 amber, 4 red), eachis illuminated by two bulbs. Engravedmarkings allow the corresponding sys-tem to be identified. For a list ofannunciators, see end of chapter.

The Failure Warning Panel is equippedwith a TEST pushbutton; when pressedall of the annunciators on the failurewarning panel illuminate.

The T/O CONFIG warning is notdirectly testable. To check its opera-tion, it is necessary to reproduce theconditions required for T/O CONFIGlight illumination.

Altitude AlerterThe altitude alerter (Figure 5J-14 ) pro-vides the pilot with visual and auralwarnings when the aircraft deviatesfrom a preselected altitude. The alert-er is equipped with a warning light, an

aural warning output, and a windowin which altitude is set by means of acontrol knob.

The alerter receives corrected altitudeinformation from the pilot’s altimeter.When the aircraft comes within 1,000ft of the indicator setting, the auralwarning sounds for about one secondand the warning light illuminates. Thelight extinguishes when the aircraftcomes within 300 ft of the setting andremains off so long as the aircraft stayswithin 300 ft of the setting.

As soon as this tolerance is exceeded,the aural warning is heard for aboutone second and the warning light illu-minates. The ALT ALERT light canbe turned off at any time by pressingthe light itself or by making a new alti-tude setting.

Aural WarningSystemComponentsThe audio warning system consists ofthe following:� buzzer unit � loudspeaker/headsets � potentiometer� test pushbutton/horn silence button.

WarningSystems

5J-14


OperationThe audio warning system alerts thecrew to configuration anomalies orindicates certain operational condi-tions. The audio warnings are producedby an audio warning buzzer unit for-ward of the pedestal console. Theaudio warning buzzer unit receivesfailure information from the differentsystems and sends out discrete audiosignals that correspond to the differ-ent type of failures. These discrete sig-nals are transmitted to the crew throughloudspeakers in the aircraft andthrough the crew’s headsets. The audiowarning system includes three poten-tiometers: one adjusts the audio warn-ing unit loudspeaker output; oneadjusts the pilot’s headset audio output;and one adjusts the copilot’s headsetaudio output. The HORN SILENCEpushbutton cannot silence the audiowarning.

Horn SilenceCutoff Pushbutton

In a number of cases, the aural warn-ing sound can be cancelled with theHORN SILENCE cutoff switch on theflight compartment control pedestal(Figure 5J-15 ).

Landing Gear Warning

The landing gear warning is a contin-uous, 285 Hz, low pitch sound whilethe landing control handle flashes red.The horn sounds when both of the fol-lowing occur:� airspeed lower than 190 kts� either 1) one or more power leversin the reduced power position with thelanding gear control handle in the upposition, or 2) the landing gear con-trol handle in the down position withone or more gear not downlocked.

Press the HORN SILENCE pushbut-ton to silence the audio warning.However, with the same conditions asabove, and with flaps 52°, the HORNSILENCE pushbutton cannot silencethe audio warning.

Cabin Pressure Warning

The cabin pressure warning is an inter-mittent, 250 Hz, low pitched soundthat sounds for 600 msec, then is silentfor 200 msec; the red CABIN annun-ciator on the Failure Warning Panelalso illuminates.

The warning indicates the cabin alt-meter reading is higher than 10,000 ft.The HORN SILENCE pushbutton cansilence the audio warning.

5J-175J-165J-15



Fire Warning

The fire warning is a continuous, sharp,500 to 550 Hz alternated sound thatcycles on and off in 150 msec cycleswith illumination of at least one FIRElight on the fuel shutoff handle whenfire is detected.

The HORN SILENCE pushbutton cansilence the audio warning.

Stall Warning

The stall warning is a fast, intermit-tent, sharp 1660 Hz (BIP-BIP) soundthat is on for 100 msec and then offfor 100 msec. There is simultaneousillumination of the two IGN lights andgreen slat light.

This warning is triggered by slats notextended and the left local AOAgreater than 17° (which correspondsto an aircraft AOA greater than 11°)or the right local AOA greater than 19°(which corresponds to an aircraft AOAgreater than 12°).

With the slats extended, warnings areactivated when the left or right localAOA greater than 27° (which corres-ponds to an aircraft AOA greater than17°). The HORN SILENCE push-button cannot silence the audiowarning.

VMo/MMo Warning

The VMO/MMO warning is a continu-ous, sharp variable sound (modulat-ed signal) between 660/3,330 Hz for1 second. Both Mach-airspeed indi-cator readings are above the VMO/MMO red line and on the EADI (ifinstalled). This situation is causedwhen the VMO/MMO limits areexceeded and the IAS is higher than350 kts at sea level to 10,000 ft or 370kts at 10,000 ft (indicated MACHhigher than 0.87 at altitudes greaterthan 25,000 feet).

Altitude Deviation Warning

The altitude deviation warning is asharp, continuous sound when flyingthrough the preselected altitude. Thetwo second warning occurs when theaircraft comes within 1,000 ft of thepreselected altitude; the light on theindicator illuminates.

Once the preset altitude is reached, theaudio warning sounds for two seconds;the light illuminate when the airplanedeviates from the preselected altitudeby more than 300 ft. The HORNSILENCE pushbutton cannot silencethe audio warning.

Horizontal Stablizer Warning

The horizontal stablizer warning is acontinuous rattle sound (clacker) in12.5 Hz pulses to indicate the hori-zontal stabilizer actuator is moving.The warning sounds when the hori-zontal stabilizer moves in both normaland emergency operation.

TestThe audio warnings for the flaps andlanding gear are tested by depressingthe LANDING GEAR INDICATORTEST pushbutton (Figure 5J-16 ); suchaction activates the audio warning andcauses the landing gear control handleto flash red.

The audio warnings for the pedestalconsole (Figure 5J-17 ) are tested bypressing the appropriate pushbuttons.

Pressing the VMO/MMO pushbuttonperforms an operational test of theaudio warning by air data computer;a continuous variable frequency soundis issued.

Pressing the CAB pushbutton activatesthe audio warning; the CABIN lighton the overhead panel illuminates.Silence the warning with HORN SILpushbutton).


Pressing the Stall 1 or 2 pushbuttonsactivates the audio warning; the IGNlights on the overhead panel illuminate.

The audio warnings for the fire panelare tested by pressing the appropriateTEST pushbutton to activate the fireaudio warning (Figure 5J-18 ). All thered lights in the fire pull handle illu-minate if the detection systems areoperative. Silence the warning withthe HORN SILENCE pushbutton.

Voice AdvisorySystemThe FPA-80 flight profile advisory sys-tem is a solid-state aural advisory andwarning system. The FPA-80 is com-pletely automatic and requires no con-trols or visual displays. All advisoryand warning information is conveyedto the pilot with a natural soundingvoice. The FPA-80 flight profile advi-sory CPN 622-4730-001 uses a malevoice; the FPA-80 CPN 622-4730-002uses a female voice.

The FPA-80 is used with a radioaltimeter system such as the Collins

ALT-50/55 or equivalent, with anARINC radio altimeter system such asthe Collins AL-101, or any other sys-tem with ARINC analog output.

One of the main functions of the FPA-80 is to announce radio altitude anddecision height. The FPA-80 informsthe pilot when the aircraft enters theoperating range of the radio altimetersystem. At 1,000 ft and continuing to100 ft, radio altitude is announced in100 ft intervals. Decision height isannounced when the aircraft reachesthe selected altitude.

A second function of the FPA-80 is toannounce messages of a warning oradvisory nature. Such messages arerepeated three times. Messages areincluded for glideslope and localizerdeviations, trim failure, attitude andbarometric altitude deviations, andlanding gear.

VocabularyThe FPA-80 callouts and conditionsthat cause an advisory are listed inTable 5J-B .

5J-18



Table 5J-B; Voice Advisory Vocabulary

MINIMUM

a. Radio altitude decreasing

b. Decision height crossed

Radio altitude exceeds decision height

RADIO ALTITUDE andindividual altitudes


b. Specific altitude is crossed

Radio altitude exceeds that specific altitude

GLIDESLOPE

a. Localizer frequency tuned

b. Gear down

c. Radio altitude between200 and 700 ft

d. One dot below glideslope

e. Not in back localizermode (difference betweenairplane heading andcourse arrow does notexceed 105°)

Correction of glideslope deviation

OR APA-80 of APS-80

a. Glideslope annunciatorlamp lights

LOCALIZER

a. Localizer frequency tuned

b. Gear down

c. Radio altitude between200 ft and 700 ft

d. One dot belowglideslope

e. Not in Back Localizermode (difference betweenairplanes heading and course arrow does notexceed 105°)

Correction of glideslope deviation

OR APA-80 of APS-80

a. Glideslope annunciatorlamp lights

Announcement/Condition

Reset


Table 5J-B; Voice Advisory Vocabulary (continued)

CHECK TRIM

a. Trim annunciator lamplights

Correction of trim failure

CHECK ATTITUDE

a. An attitude annunciatorlamp lights

Correction of attitude deviation

CHECK BARO ALTITUDE

a. PRE-80 aural warningsounds to indicatealtitude deviation greaterthan 300 ft frompreselected altitude

b. Not in Approach mode(optional )

Automatically resets after completionof callout

CHECK GEAR


b. Radio altitude between500 and 100 ft

c. Gear not down

Radio altitude exceeds 500 ft, the gear isdown, or radio altitude increases morethan 32 ft

Announcement/Condition (continued)

Reset


5J-32


Annunciators


STAB TRIMAIRBRAKE BRAKE

P. BRAKEHYD. TK2

HYD. 1 HYD. 2 ST. BYPUMP

PITOT 1 PITOT 2 Q. UNIT

GENE. 1 GENE. 2 BATT.

HYD. TK1 LO. FUELGUST

OIL 1 OIL 2 COND.O. HEAT

CABINFUEL P2FUEL P1

ENG. C1 ENG. C2 HOT BAT

LH

NOSE

RH

DOWN

FIRE

1PULL

FIRE

2

PULL

OFF

180160140120

F

BAT. TEMP.1 2

WARM HOT

FLAPS

52 30

0

DN

UP

15

WINGS INVERTERSNO. 1 NO. 2

X. FRIGNITER ON IGNITER ON

DE-ICING

REV REV

A B

C D

MOVING

LDG GEAR

AIRBRAKE

FLAPASYM

TEST

BRIGHT

DIM

BRAKE

BRAKEL

BRAKER

TESTON

OFF

ON

OFF

HYD.ST. BY PUMP

AIL RUD

E F

G H

TRIMTRIM

TRIMTRIMFFAILAIL

TRIMTRIMFFAILAIL

APAPDISCDISC

YDYDFFAILAIL

VLFVLFDRDR

VLFVLFWPTWPT

VLFVLFBBAATTTT

HDGHDG

APAPXFERXFER

GSGSARMARM

LALATTARMARM

ALALTTSFISFI

ALALTTHOLDHOLD VERVERTT OO

MM

A

FGCFGCDRDR

BBAACKCKLOGLOG

LALATTCAPCAP

GSGSCAPCAP GAGA SMOKESMOKE

DETDET

A

D E

G

C

F

B

OFF

180160140120

F

H

TRANS TRANS

ENG 1 ENG 2

AC FAIL

DE-ICING

Annunciators



AnnunciatorCrossReference

Major annunciators, with brief explanations, are listed alpha-betically to correspond with alphabetical designations on theillustration on page 5J-32. Specific information about whatcauses each annunciator to illuminate is in the appropriatesystem chapter.

A

AC FAIL NO. 1AC FAIL NO. 2

The illumination of a red AC FAIL light indicates failure of the respective 115V ACInverter bus.

ENG 1ENG 2

The green ENG 1/2 lights illuminate when HP bleed air is available to the engineinlet lip.

IGNITOR ONThe green advisory ignitor lights illuminate when power is available to the ignition box.A faulty ignitor does not prevent the light from illuminating.

WINGSThe green WINGS anti-ice light illuminates when both thermostats connected in series in the wings reach 80°C. It extinguishes when either thermostat senses lessthan 70°C.

X.FRThe amber windshield X.FR annunciator on the overhead panel illuminates wheneither windshield temperature control system fails. The windshield with the failedtemperature control system is then controlled by the system of the opposite side.

B

DE-ICINGOn S/Ns 001 to 151 with SB F10-208 , the amber DE-ICING annunciator illuminatesto indicate airframe anti-ice air is leaking in the area of the flexible hoses. Fourthermal fuses trigger the indication.

REVThe green REV lights in the thrust reverser panel on the instrument panel indicate therespective thrust reversers are in the deployed position.

TRANSThe two amber TRANS lights in the thrust reverser panel indicate their respectivethrust reverser doors, actuator lock, and/or secondary latch are unlocked and thereverser is not fully deployed.

YDFAIL

Illumination of the amber YD FAIL light indicates yaw damper failure.

C

FIRE 1 PULLFIRE 2 PULL

The red FIRE 1/2 PULL handles illuminate when the associated engine sensingelement reaches an average temperature of 232°C or 427°C at a single point.Pulling a handle mechanically closes fuel valves in the associated engine pylon.


D

AIRBRAKEThe red AIRBRAKE annunciator illuminates on the ground if the airbrakes are notretracted and either throttle is in the TAKEOFF position. In flight,, the annunciatorcomes on anytime the airbrakes are not retracted and the slats/flaps handle is out ofthe CLEAN position.

BATT. Illumination of the amber BATT. annunciator indicates at least one of the two batteriesis not connected to the Main bus.

CABINIIllumination of the red CABIN annunciator in the air accompanied by an aural warninghorn indicates the cabin altitude pressure altitude is more than 10,000 ft. Illuminationof the CABIN annunciator without the warning horn indicates the cabin access door,the rear compartment door, or the pressure-fueling door is not properly closed.

COND.O. HEAT

The COND. O. HEAT annunciator illuminates when the temperature of the airdischarged into the air-conditioning system exceeds 250°C.

DE-ICINGOn S/N 152 and subsequent , the amber DE-ICING annunciator illuminates toindicate airframe anti-ice air is leaking in the area of the flexible hoses. Four thermalfuses trigger the indication.

ENG.C1ENG.C2

The amber ENG.C1/C2 annunciator illuminates when the associated enginecomputer switch is off, the associated engine computer fails, or power to theassociated computer is lost.

FUEL P1FUEL P2

Illumination of the amber FUEL P1/P2 annunciator indicates pressure below 4.6 PSIin the fuel manifold between the booster pump and the engine.

GENE. 1GENE. 2

Illumination of an amber GENE.1/2 annunciator indicates the left or right generatoris not connected to the Main bus. Possible faults are: differential/switch tripped,overvoltage/switch tripped, reverse current, mechanical.

GUST.The amber GUST. annunciator illuminates when Hydraulic System No. 1 pressure isless than 400 PSI. Illumination of the amber GUST annunciator indicates closing ofthe automatic control valve to provide gust damping to the rudder.

HOT BATThe HOT BAT annunciator illuminates when the internal temperature of one of thebatteries reaches 135°F (160°F with SAFT batteries).

HYD. 1HYD. 2

The amber HYD.1/2 annunciator illuminates whenever the associated hydraulicsystem is 1,500 PSI or less.

HYD. TK. 1HYD. TK. 2

The amber HYD.TK.1/2 annunciator illuminates when the respective hydraulicreservoir tank LP air pressure is less than 16 PSIA.

LO. FUELThe amber LO.FUEL annunciator illuminates when either feeder tank has less than300 lbs of fuel.

OIL 1OIL 2

Illumination of the amber OIL 1/2 annunciator with normal oil pressure indicates metalchips in the associated engine oil system. The amber OIL 1/2 annunciator illuminateswhen the oil pressure in the associated engine drops below 25 PSI.



P. BRAKEThe amber P. BRAKE annunciator illuminates when the pressure in theemergency/park brake accumulator is less than 1,200 PSI.

PITOT 1PITOT 2

Illumination of the PITOT 1/2 annunciator indicates a power failure to either a pitotprobe or static port.

Q. UNIT

Illumination of the Q. UNIT annunciator indicates a failure of rudder or aileronArthur Q unit. When the annunciator illuminates, the malfunction can be furtherinvestigated by checking the copilot’s circuit breaker panel for illumination of theaileron or rudder Q failure lights. Illumination while accelerating indicates theapplicable Q Unit failed in Low Speed mode. Illumination while decelerating indicatesthe applicable Q Unit failed in High Speed mode.

ST.BYPUMP

The amber ST.BY PUMP annunciator illuminates when the hydraulic standby pumpruns for more than 30 seconds consecutively.

STAB TRIMThe STAB TRIM annunciator illuminates on the ground if the stabilizer is out of the 4°to 8° aircraft nose-up range and either power lever is in the takeoff range.

E

AIRBRAKE

Illumination of the amber AIR BRAKE annunciator indicates the airbrakes are notfully retracted.

FLAPASYM

The red FLAP ASYM annunciator illuminates when a position difference of more than5° exists between the flaps.

LDG GEARLH/NOSE/RH

The three green LDG GEAR lights (LH, NOSE, and RH) on the configuration panelilluminate when the respective gear is in a locked down position.,

LDG GEARMOVING

The three red LDG GEAR lights on the configuration panel illuminate when therespective gear is not in a locked down or up position.

Gear Control HandleThe blinking red light in the control handle indicates that at least one gear is not downand locked, one or both throttles are retarded, and the airspeed is less than 190 kts.

Slat LightIllumination of the red slat light indicates the slats are not in agreement with theslats/flaps handle, stall input, or emergency slat extend switch. With emergencyslats selected and the slat/flap handle out of the CLEAN detent, a red slat indicationis normal.

F

BRAKEOn S/N 152 and subsequent , the red BRAKE annunciator illuminates whenthe parking brake lever is moved from the OFF position (fully forward). It is on themain annunciator panel. Prior to SN 152 , the annunciator is either on the centerinstrument panel forward of the emergency brake lever or above the mainannunciator panel.

BatteryTemperature

Indicator

The amber light on the battery temperature indicator illuminates when a battery internal temperature reaches 120°F. The red light on the battery temperatureindicator illuminates when a battery internal temperature reaches 150°F.


G

ANTISKIDBRAKE L/BRAKE R

The amber anti-skid lights illuminate any time the anti-skid system is functioning.

H

AIL

The amber aileron/elevator Arthur signal failure light on the copilot’s circuit breakerpanel illuminates when accelerating and the aileron/elevator Q unit fails in the LowSpeed mode or when decelerating and the aileron/elevator Q unit fails in the HighSpeed mode. Do not pull the warning horn CB until the landing gear is down andlocked.

RUD

The amber rudder Arthur signal failure light on the copilot’s circuit breaker panelilluminates when accelerating and the rudder Q unit fails in the Low Speed mode, orwhen decelerating and the rudder Q unit fails in the High Speed mode. If the rudderQ unit fails in high, the flap overspeed warning sounds when flaps are selected.Do not pull the warning horn CB until the landing gear is down and locked.

Falcon 10/100 For SimuFlite training only 5K-1April 2000

PneumaticSystems

Chapter 5K

This chapter describes the systems that extract, distribute, andcontrol engine bleed air (for anti-icing bleed air, see Ice andRain Protection chapter). The Freon Air Conditioning System,Pneumatic Air Conditioning, and Pressurization Systems arecombined in this chapter to present the flow of engine bleedair and its use throughout the aircraft.

The Pneumatic System extracts bleed air from the engines,collects it, and then transfers it to various other systems (i.e., airconditioning, ice and rain protection, and pressurization). Thesection covering the Pneumatic System begins on page 9.

The Air Conditioning System routes engine bleed air collectedby the Pneumatic System through a conditioning valve for heat-ing and cooling. The air conditioning section, which begins onpage 13, discusses aircraft heating and cooling.

The Air Conditioning System also supplies conditioned air forthe pressurized vessel. Pressurization is controlled by meter-ing the outflow of conditioned air through the outflow valves.The operation of the outflow valves is controlled by thePressurization System, which is discussed on page 21.

5K-2 For SimuFlite training only Falcon 10/100April 2000


Pneumatic Systems

Table ofContents

Bleed Air System Schematic . . . . . . . . . . . . 5K-6

Pneumatic System Components Schematic . . . 5K-8

Pneumatic System . . . . . . . . . . . . . . . . . 5K-9

Distribution . . . . . . . . . . . . . . . . . . . . . . 5K-9

LP Bleed Air Supply . . . . . . . . . . . . . . . . . 5K-9

HP Bleed Air Supply . . . . . . . . . . . . . . . . 5K-9

Pneumatic System Components . . . . . . . . . . . 5K-9

Flow Limiter Venturis . . . . . . . . . . . . . . . . 5K-9

HP Valves . . . . . . . . . . . . . . . . . . . . . . 5K-9

P2 unit (Bleed Air Center) . . . . . . . . . . . . . . 5K-9

HP Air Injector . . . . . . . . . . . . . . . . . . . 5K-10

Conditioning Valve . . . . . . . . . . . . . . . . . 5K-10

Anti-Icing Electro-Valve . . . . . . . . . . . . . . 5K-10

Air Conditioning System Schematic . . . . . . . 5K-12

Air Conditioning System . . . . . . . . . . . . . 5K-13

Primary Heating and Cooling System . . . . . . . . 5K-13

Heat Exchanger . . . . . . . . . . . . . . . . . . 5K-13

Temperature Regulating Valve . . . . . . . . . . 5K-13

Windshield and Pilot’s Floor Heat . . . . . . . . . 5K-14

Footwarmer/Defog Valves . . . . . . . . . . . . . 5K-14

Temperature Regulation Control Box . . . . . . . 5K-14

Freon System Schematic . . . . . . . . . . . . . 5K-16

Freon System . . . . . . . . . . . . . . . . . . . 5K-17

Compressor Motor . . . . . . . . . . . . . . . . . 5K-17

Compressor . . . . . . . . . . . . . . . . . . . . 5K-17

Underpressure Switches . . . . . . . . . . . . . 5K-18

Overpressure Switches . . . . . . . . . . . . . . 5K-18

Condenser . . . . . . . . . . . . . . . . . . . . . 5K-18

Condenser Fan . . . . . . . . . . . . . . . . . . 5K-18

Reservoir Tank . . . . . . . . . . . . . . . . . . . 5K-18


Pressure Reducing Valves . . . . . . . . . . . . 5K-18

Evaporator . . . . . . . . . . . . . . . . . . . . . 5K-18

Evaporator Fans . . . . . . . . . . . . . . . . . . 5K-19

Gasper Air . . . . . . . . . . . . . . . . . . . . . 5K-19

Pressurization System Schematic . . . . . . . . 5K-20

Pressurization Systems . . . . . . . . . . . . . . 5K-21

IDC Pressurization System . . . . . . . . . . . . . 5K-21

Cabin Pressurization Indicating System . . . . . . 5K-21

Cabin Pressurization Failure Warning System . . 5K-22

Pressurization Controller . . . . . . . . . . . . . 5K-22

Pressurization Mode Selector . . . . . . . . . . . 5K-23

Electropneumatic Outflow Valves . . . . . . . . . 5K-23

Manual Controller . . . . . . . . . . . . . . . . . 5K-24

Jet Pump . . . . . . . . . . . . . . . . . . . . . 5K-24

Automatic Mode . . . . . . . . . . . . . . . . . . 5K-24

Dump Mode . . . . . . . . . . . . . . . . . . . . 5K-24

Emergency Pressurization . . . . . . . . . . . . . 5K-24

ABG – SEMCA Pressurization System . . . . . . . 5K-25

Cabin Pressurization Indicating System . . . . . . 5K-25

Cabin Pressurization Failure Warning System . . 5K-25

Pressurization Controller . . . . . . . . . . . . . 5K-25

Pressurization Mode Selector . . . . . . . . . . . 5K-26

Main Regulating (Electropneumatic)Outflow Valve . . . . . . . . . . . . . . . . . . . 5K-26

Pneumatic (Emergency) Outflow Valve . . . . . . 5K-26

Manual Pressurization Controller(UP/DN Knob) . . . . . . . . . . . . . . . . . . . 5K-27

Jet Pump . . . . . . . . . . . . . . . . . . . . . 5K-27

Automatic Mode . . . . . . . . . . . . . . . . . . 5K-27

Dump Mode . . . . . . . . . . . . . . . . . . . . 5K-27

Manual Mode . . . . . . . . . . . . . . . . . . . 5K-27


5K-6

Pneumatic Systems

Bleed AirSystem

Servicing and Procedures . . . . . . . . . . . . . 5K-29

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5K-29

Abnormal Procedures . . . . . . . . . . . . . . . . 5K-29

COND’G OVHT Warning . . . . . . . . . . . . . 5K-29

Conditioning Valve Fails to Open on Takeoff . . . 5K-29

Cabin Pressure Too High . . . . . . . . . . . . . 5K-30

Improper Cabin Vertical Speed . . . . . . . . . . 5K-30

Pressurization Loss (High Cabin Altitude) . . . . . 5K-30

Emergency Procedures . . . . . . . . . . . . . . . 5K-31

Air Conditioning Smoke . . . . . . . . . . . . . . 5K-31

Rapid Decompression/Emergency Descent . . . . 5K-31

Smoke Removal . . . . . . . . . . . . . . . . . . 5K-31

Data Summaries . . . . . . . . . . . . . . . . . . 5K-33


WINDSHIELDDEMISTINGMANIFOLD

INSTRUMENTPANELGASPERS

SILENCERS

WINDSHIELDDEFOG VALVE

PILOT'S FLOOR HEATAIR SUPPLY VALVE

RETURN AIR

CONDITIONINGAIR INLET

PASSENGERSGASPERS

OPTIONALAUX. HEAT

FLOODDUCT

RECIRCULATINGJET PUMP

SILENCERS

RAM AIRINLET

COLD

COLD

HOT

HOTMANUAL

HP BLEEDVALVE

TEMPREGULATINGVALVE

C BUS

PRESSCONT.

HYDPRESS

COND'GVALVEAUTO

OFF

B BUS

FLOWLIMITER

PRESSUREREGULATOR

CONDO'HEAT

FLOWLIMITEREMERPRESS

FAN

RECIRCULATINGAIR

WING ANTI-ICEVALVEHP BLEED VALVE

TO WINGANTI-ICING

NACELLEANTI-ICING VALVE

MECHANICAL FLOODDUCT VALVE(IF INSTALLED)

FREONEVAPORATOR

DOME SPINNERANTI-ICING VALVE(IF INSTALLED)

LP AIR

ONON

COLD AIR/RECIRCULATED

RECIRCULATED AIR

LP AIR

HP AIR

RAM AIR

CONDITIONED AIR

EMER PRESSVALVE

P2 UNIT

HP AIR

LP AIR

HP AIR

Bleed Air System


Pneumatic Systems

JET PUMPPRESSURIZATIONVALVE CONTROL

HYDRAULIC RESERVOIRPRESSURIZATION

EMERGENCYPRESSURIZATION

COCK

EMERGENCY MANUAL CONTROLELECTRO VALVE (OPTIONAL)SB - 36 - 003

TO WING ANTI-ICING

TO CONDITIONING SYSTEMP2 UNIT (BLEED AIRCENTER)

HP AIR INJECTOR

CONDITIONING VALVE

WING ANTI-ICINGELECTRO-VALVE

ANTI-ICING HPRESTRICTOR

HP BLEED VALVES

TO NACELLE AIRINTAKE ANTI-ICING

LOW PRESSURE BLEED

HIGH PRESSURE BLEED

FLOW LIMITERVENTURI

ENGINE LP AIR

ENGINE HP AIR

WING ANTI-ICE AIR/CONDITIONINGSYSTEM AIR


Pneumatic System Components


Pneumatic Systems

DistributionLow pressure (LP) and high pressure(HP) bleed air from each engine sup-plies air through check valves to a P2unit (bleed air center) in the rear com-partment. HP bleed air is controlledby HP bleed control valves.

The P2 unit supplies bleed air to thefollowing systems:� air conditioning� pressurization� wing anti-ice� emergency pressurization� hydraulic reservoirs.

LP Bleed Air SupplyThe left and right engines continuouslysupply LP bleed air. LP air is extract-ed through two bleed ports on theengine at the 3 and 9 o’clock positions.After exiting through the bleed ports,LP bleed air passes through one-waycheck valves, which block the path ofbleed air trying to enter a non-operat-ing engine.

LP bleed air is available through thebleed air center to all systems except:� air conditioning and pressurizationwhen the conditioning valve is closed� wing anti-ice when the wing anti-ice control valve is closed.

HP Bleed Air SupplyHigh pressure (HP) air from the 12o’clock HP port on the engine is drawnfrom both engines. HP bleed air flowsfrom each engine through electricallycontrolled, air-operated shutoff valvesin the pylons prior to the P2 unit. If LPbleed air is insufficient to pressurizeor heat the cabin due to reduced enginepower at altitude, HP bleed air flowsinto the P2 unit through a mixing valve(HP air injector). HP air is availablefor anti-ice through a restrictor andinlet hose.

Pneumatic SystemComponentsComponents of the pneumatic systemthat distribute and regulate the bleedair supply include the following:� flow limiter venturis� HP valves� HP air injector (mixing valve)� P2 unit (bleed air center)� conditioning valve� anti-icing electro-valves.

Flow Limiter VenturisHP air passes through a flow limiterventuri that limits the amount of bleedair drawn from the HP port to 2.5% ofthe maximum engine compressor out-put. The venturi provides a safety fea-ture that prevents excessive bleed airdrawn from the compressor in theevent of a ruptured HP line.

HP ValvesEach engine supplies HP bleed airthrough HP valves powered by theLoad Shed C bus. These valves areelectrically controlled and pneumati-cally operated. They fail closed withloss of electrical power.

Both HP valves are controlled withone switch labeled HP BLEED OPEN/CLOSED on the forward pedestal.� In OPEN position, the valve allowsHP bleed air to enter the P2 unit andmix with LP bleed air to support bleedair demand.� In CLOSED position, HP valvesclose to shut off HP bleed air flow.

The P2 unit distributes bleed air foranti-icing and air conditioning.

P2 Unit (Bleed Air Center)The P2 unit regulates the volume andpressure of air available for the air con-ditioning and wing anti-ice systems.

PneumaticSystem


The P2 unit includes check valves forone-way airflow to prevent bleed airreversal during single engine opera-tions (e.g. starting or engine failure).

The air conditioning system utilizesLP air only in normal operation.However, under conditions of high alti-tude and low power settings (begin-ning of descent), insufficient LP air isavailable. In this case, the HP injectorvalve injects HP air into the air con-ditioning system as required to main-tain pressure. Because HP air is muchhotter than LP air, a temperature sen-sor downstream of the conditioningvalve monitors the duct temperature.If temperature exceeds 250°C, aCOND O’HEAT annunciator illumi-nates on the Failure Warning Panel.

At takeoff power, LP air is approxi-mately 230°C at 62 PSI; HP air is notinjected until pressure drops to approx-imately 16 PSI. Above 16 PSI, HP airis restricted from entering the cabinair system.

The wing anti-ice system is providedair from both HP (30% by volume)and LP (70%) manifolds. Because HPair is much hotter than LP air, most ofthe heat required for airframe anti-icingis provided by HP air with additionalvolume being provided with LP air(see Chapter 5H, Ice and Rain).

HP Air InjectorThe HP air injector infuses HP bleedair into the P2 unit to supplement thedemand of bleed air to the cabin. Theinjector is an auto-adjusting orifice thatsenses upstream pressure changes. Thepressure reducer starts opening at 15.83PSI to permit the air conditioning sys-tem to be properly supplied in all flightconditions.

Conditioning ValveThe conditioning valve controls all air-flow into the cabin except during emer-gency pressurization. The B bus pow-ers the valve through the COND’GVALVE switch. The switch positionsfunction as shown below:� OFF – airflow is cut off.� ON – the conditioning valve opensto allow bleed air flow.� AUTO – the conditioning valvecloses with weight on the gear andthrottles advanced for takeoff. Aftertakeoff, the valve automatically opensslowly to prevent pressure bumps.Approximately two minutes arerequired for complete opening.

Anti-Icing Electro-ValveFor information on the anti-icing elec-tro-valve, see Ice and Rain chapter.


5K-12

Pneumatic Systems

Air ConditioningSystem

C BUSON

OFF

AUXILIARYBUS

OFF

AUTO

FREON

EVAPORATOR

EXPANSIONVALVE (3 EA.)

REGULATING BULB

TEMP. REGULATINGVALVE - SWITCHCLOSED - 90% FULL COLDOPEN - 76% FULL COLD

EVAPORATOR FANS

TO OPERATING GPU

TO OPERATING GEN

COMPRESSOR SUCTIONPRESSURE SENSITIVESWITCH (MINIMUMPRESSURE)

FREON TANK

COMPRESSORDISCHARGE PRESSURESWITCHES (MAXIMUMAND MINIMUM PRESSURES)

COMPRESSORDRIVING MOTOR

COMPRESSOR

LEFT SQUATSWITCH

TO OPERATING GEN

B BUS

CONDENSER

CONDENSER BLOWERHP FREON GAS

HP FREON LIQUID

LP FREON GAS

TO CABIN

RECIRCULATING AIR

ON

FLIGHT

GROUND

MAIN BUS


Air Conditioning System

ACSystem

AuxiliaryBus

FreonSwitch

CompressorMotor

CondenserBlower

LH RH

Power Supply Evaporator Fans

“OFF” AUTO ONOFF

STOP STOP STOP STOP

AC Batteries orDuring StartingSequence

“ON” AUTO STOP STOP OPERATES

“ON” ON STOP STOP OPERATES OPERATES

“ON” OFF STOP STOP STOP STOP

STOP

GPU or At LeastOne Generator

“ON” AUTO Demand Operation:TemperatureRegulating Valve inFull Cold Positionwithin SafeEnvelope

Demand Operation:on the Ground:TemperatureRegulating Valvein Full ColdPosition

OPERATES

“ON” ONOperates if withinSafe Envelope

Operates onthe Ground OPERATES OPERATES

“ON” OFF STOP STOP OPERATES OPERATES

OPERATES


Pneumatic Systems

The air conditioning system providesconditioned air to the cabin and cock-pit. The conditioning system is com-posed of the following systems:� primary heating and cooling–engine bleed air system� Freon system� auxiliary heat (optional).

Primary Heatingand CoolingSystemThe primary heating and cooling sys-tem provides conditioned air to be dis-tributed through the aircraft’s cabinand cockpit. Components are asfollows:� heat exchanger� temperature regulating valve� footwarmer/defog valves� temperature regulation control box.

Heat ExchangerThe heat exchanger cools engine bleedair ducted directly from the engine ven-turis (flow limiters). Ram air, obtainedwith an outside ram air scoop on thetop side of the right engine pylon,directs ambient air across the heatexchanger to cool the bleed air.

Temperature RegulatingValveThe gate-type temperature regulatingvalve is operated with a variable speedDC motor; it is controlled with anelectronic cabin temperature controlsystem. The regulating valve in linewith the heat exchanger is automati-cally controlled with a electronic temp-erature regulator. Load Shed C busdistributes power for the regulatingvalve.

An electronic temperature controllerdetermines the amount of hot enginebleed air routed through the heatexchanger to provide cooling air. Thesystem may be operated automaticallyby the temperature controller or man-ually by the pilot.

To operate the system manually thepilot controls the system with theCOLD/HOT switch selected to the 6o’clock position, then rotating theswitch clockwise for cooler selectionor counterclockwise for a warmerselection. Some aircraft have a gageindicating the position of the temper-ature regulation valve.

In the automatic mode, the valve closesto pass a greater percentage of the airthrough the heat exchanger as thetemperature in the cabin increases.

AirConditioningSystem

5K-1


If the Freon switch is in AUTO, theFreon system turns on when the valvereaches 90% full cold.

Windshield Defog andPilot’s Floor HeatWith the COND’G VALVE switchselected to ON or AUTO, the con-ditioning valve allows engine bleed airto flow directly from the P2 unit tothe cockpit (Figure 5K-1 , previouspage). The bleed air passes throughthe following before reaching thecockpit:� windshield defog valve (see Ice andRain chapter)� windshield demisting manifold (seeIce and Rain chapter)� pilot’s floor heat air supply valve.

The bleed air heats the floor and pro-vides windshield demisting throughthe windshield defog valve air supplyvalves.

Footwarmer/Defog ValvesThe footwarmer and defog valves onthe pilot’s side console provide uncon-ditioned air (hot) to the cockpit foradditional heat.

The air source of both valves is down-stream of the conditioning valve andprovides an excellent check after take-

off to confirm that the conditioningvalve is fully open.

The defog valve also directs uncondi-tioned air to the windshield. TheWSHLD lever on the pilot’s left con-sole next to the FLOOR lever controlsthe valve. In OPEN, the levers controlwindshield and foot warmers.

Temperature RegulationControl BoxThe temperature regulation control boxat the bottom center section of theinstrument panel consists of a rotatingbutton and a circular stop (Figure 5K-2). Display on the upper scale indi-cates an automatic temperature regu-lation between 59°F (12°C) and 84°F(30°C).

The control box compares the tem-perature instructions preset on its dialwith the cabin temperature sensed byprobes; it then controls the openingand closing of the valve. The follow-ing probes send an electrical signal tothe temperature control box:� cabin temperature probe – (signalto temperature control box proportionalto the cabin air temperature)� duct air temperature probe – sensestemperature in the air duct; (signal tothe control box proportional to the airtemperature in the duct).

5K-2


Pneumatic Systems

RECIRCULATINGAIR

RECIRCULATINGAIR

BLOWERS

HEAT EXCHANGER

THERMOSTAT (139H) DELETEDON A/C WITH SB 21-007

RECIRCULATINGCOOLED AIR

PRESSURE REDUCINGVALVE

LP VAPORCOMPRESSOR

HP VAPOR

VENTILATIONNOZZLE CONDENSER

BLOWER (WORKSON THE GROUND)

HP FLUID

VAPORIZED HEATCARRYING FREONUNDER PRESSURE

CONDENSED FREON,COOLED TO A LIQUIDAND UNDER PRESSURE

FREON DROPLETSABSORBING CABINHEAT UNDER SUCTION

SUCTION RETURNHEATED FREONHOT AIR

GRD

AIRFLIGHT

FREONRESERVOIR

AIRFLIGHT

EXTERIOR AIRINTAKE GRD


Freon System


Pneumatic Systems

The self-contained air conditioningsystem uses Freon (refrigerant 12) tocool the cabin and cockpit on theground and in the air. The temperaturecontroller used to regulate conditionedair further controls the Freon system.The Freon system operates when thetemperature regulating valve is morethan 90% full cold with the systemswitch in AUTO (Figure 5K-3 ).

With the Freon system operating, recir-culation fans draw warm cabin airacross the evaporator to cool it andthen deliver the cooler air back to thecabin and cockpit distribution system.Either external power or at least onegenerator must be powering the Mainbus for the Freon compressor tooperate. The compressor shuts downautomatically when an engine start isinitiated.

Components of the Freon system are:� compressor motor� compressor� underpressure switches� overpressure switches� condenser� condenser fan� reservoir tank� pressure reducing valves� evaporator� evaporator fans� gasper air vents.

Compressor MotorA 28V DC electric motor receivespower directly from the Main bus anddrives the automotive-type two cylindercompressor.

The current consumption to drive theelectric motor is between 150 and 170AMPs, depending on ambient tem-perature. A three-position FREONswitch powered through the Load ShedC bus controls power for the com-pressor (Main bus) and the Freon sys-tem components. The switch positionsand functions are:� OFF – removes electricity from theFreon system� ON – actuates the Freon system byturning on the compressor motor aftera 30 second delay� AUTO – allows the temperatureregulation valve to control the com-pressor motor, turning the system onwhen the temperature regulation valvereaches 90% full cold� ON or AUTO – (on the ground withbatteries only), Freon switch is a fanswitch only.

CompressorThe Freon compressor is an automo-tive-type two-cylinder compressor. Itcompresses Freon drawn in as a gasfrom the evaporator. The compressorcompresses the vapor and delivers itunder pressure to the condenser wherethe gas returns to a liquid state.

Freon System

5K-3


The electric motor drives the com-pressor with a belt.

Underpressure SwitchesA switch on the pressure discharge lineof the compressor senses a low pres-sure condition of 26.1 ± 7.2 PSI. Theswitch stops the compressor motorwith the opening of the control circuitof the compressor. These switches alsoprotect against a Freon leak or opera-tion in a cold environment.

The inlet (suction) line of the com-pressor incorporates a pressure switchset at 14.5 PSI, which shuts the com-pressor motor down if the followingoccurs:� the intake line is clogged (the ther-mostatic expansion valve needle isjammed shut).� the Freon system is set into opera-tion when the temperature is too low.

Overpressure SwitchesA switch on the discharge line of thecompressor turns off the system if dis-charge pressure is 283 ± 14.5 PSI. Theswitch stops the compressor motorthrough opening of the control circuit.The switch protects against head pres-sure which is too high for the motorto overcome. This also protects againstoperation at an excessive temperature.

CondenserThe condenser in the tail of the aircraftconverts high-pressure and tempera-ture gaseous Freon into a liquid. A con-denser fan draws air across the con-denser on the ground consequentlycausing the Freon to give up its heatand cool to a liquid. When airborne,air is scooped into the tail opening bythe tail screens and exits the aircraftthrough a screen on the aft compart-ment door.

Efficiency is improved on the groundwith the opening of the aft compart-ment door to allow an increased airvolume.

Condenser FanA condenser fan in the tail of the air-craft with the condenser unit is a 28VDC operated electric fan that pullscooling air across the condenser finsto cool vaporized Freon back to aliquid. The fan operates only when:� the aircraft is on the ground� 28V DC aircraft system is suppliedby one aircraft generator or with aGPU� the FREON switch is in the ON orAUTO selection (in AUTO, the temp-erature control valve must also be ina position close to the full cold posi-tion, greater than 90% cold).

Reservoir TankA reservoir tank in the rear compart-ment has a capacity of 1/8 gallon. Thereservoir incorporates a sight glass thatallows maintenance to check for air inthe system with the Freon system oper-ating. The flow of bubbles across thesight glass indicate low Freon chargeor air in the system.

The reservoir contains a desiccator thatabsorbs traces of water. This absorptionprevents moisture damage of thesystem.

The maximum Freon pressure is 326PSI at 80°C (176°F); normal operat-ing pressure is 247 PSI at 38°C(100°F).

Pressure Reducing ValvesFreon pressure reducing valves incor-porate jets in which the cross section-al area is variable. The jets:� meter the flow of liquid Freon sup-plying the three bundles of tubes ofthe Freon evaporator� spray the Freon in fine droplets toallow for faster evaporation in theevaporator.

EvaporatorThe evaporator in the passenger cabincontains liquid Freon that boils after


5K-20

Pneumatic Systems

PressurizationSystem

leaving the expansion valves. TheFreon in the evaporator absorbs heatfrom the cabin as it passes through theevaporator core. This Freon heatabsorption turns the boiled liquid intoa vapor that carries the heat away fromthe aircraft cabin.

Cabin air is blown across the evapor-ator using jet pump action and Freonevaporator fans. The Freon enters eachlayer of the evaporator through a pres-sure relief valve and vaporizes due tothe low pressure level established inthe tubes by the compressor suction.Cabin air, brought in with the fans,circulates from top to bottom betweenfins brazed onto the tubes. This aircools by collecting cold units dissi-pated by Freon evaporation and trans-mitting them to the cabin.

Evaporator FansTwo electric evaporator fans in thepassenger cabin draw cabin air aroundthe evaporator fins and distribute itinto the air conditioning distributionsystem. The left fan supplies the gasper

system and the right supplies the rightconditioning air distribution duct.

Both fans operate automatically withthe aircraft electrical system supply-ing 28V DC unless the engines arebeing started. If batteries only are on,the left fan that supplies air to thepilot’s gaspers operates with theFREON switch placed in AUTO. Ifthis switch is ON, both fans operate.The fans stop with the selection ofAUX BUS switch to OFF.

Gasper AirEach cold air duct divides into twoducts to supply the gasper manifold.The gasper manifold, at the top of theframe at the rear of the passengercabin, supplies the distribution ducts.

Ducts are in the passenger cabin ceil-ing and supply adjustable outlets abovethe passenger seats (Figure 5K-4 ). Theduct routed through the right cabinsidewall and into the cockpit suppliestwo adjustable outlets on the outboardedge of each pilot and copilot instru-ment panel.

5K-4


.5 11.5

2

1.51.5

0

UP

DOWN

VERTICALSPEED

X1000 FT.PER MIN.

21

34 5 6

789

10CABINPSI

0

4 8

12

16FT x 1000

CABIN

1000

M.

9

8

76

5

4

3

2

10

FLIG

HT

ALTITUDE

29 30 31

2325

30

3540

45

50

I.D.C

LIFT

ALTITUDESWITCH(10,000)

CABIN

A BUS

PRESS STATICPORTCOPILOT

NORMALSTATIC

UP

DOWN

MANUAL PRESSURIZATIONCONTROL (CHERRY PICKER)

THROTTLESRELAY

TO CABIN

OFF

COND'GVALVE

AUTO

FILTER

FILTER

FILTER

RIGHT GEARSHOCK ABSORBERMICROSWITCH 9G2

OVERPRESSURERELIEF VALVE

CLOSINGSOLENOIDVALVE

CALIBRATED JET

NEGATIVE PRESSRELIEF VALVE

OPENINGSOLENOIDVALVE

LP BLEED AIR

REGULATINGVALVE

PRESSUREREGULATOR+200 MB

JETPUMP-200 MB

BLEEDAIRCENTER

CLOSINGSOLENOIDVALVE

LP BLEED AIROPENINGSOLENOIDVALVE

CALIBRATED JET

OVERPRESSURERELIEF VALVE

AC W BUS

A BUS

B BUS

LEFT GEAR SHOCKABSORBERMICROSWITCH 9G1

LP BLEED

OPENING REFERENCE PRESSURE

CLOSING REFERENCE PRESSURE

CABIN PRESSURE

AMBIENT PRESSURE

MANUAL PRESSURIZATION ONLY

REGULATINGVALVE

ON

Pressurization System


Pneumatic Systems

Conditioned air entering the cabinthrough the air distribution ducts pres-surizes the cabin. The amount of airexhausted from the cabin controls pres-surization.

If cabin pressurization and air refresh-ing are not sufficient, an emergencyhand-controlled or electrically oper-ated valve permits emergency pressur-ization to supply the cabin with freshair. Pressurization may be achieved,but cabin air temperature is not regu-lated.

IDC PressurizationSystemIn normal operation, the pressuriza-tion system maintains a cabin altitudeof 8,000 ft at a flight altitude of 45,000ft with maximum differential of 8.8PSI. Pressure relief is set for 9.1 PSI.

The pressurization system has twooperating modes:� automatic – (outflow valves arecontrolled electropneumatically)� manual – (outflow valves are con-trolled pneumatically).

Components for the IDC system are:� cabin pressurization indicatingsystem

� cabin pressurization failure warn-ing system� automatic pressure regulator� pressurization mode selector� electropneumatic outflow valves� manual controller� jet pump.

Cabin PressurizationIndicating SystemThe cabin pressurization indicatinginstruments at the bottom of the centerinstrument panel monitor operation ofthe cabin pressurization system(Figure 5K-5 ). The instruments indi-cate the following:� cabin differential pressure indica-tor – indicates readings from 0 to 10PSI; includes a green arc between 8.8and 9.0 PSI and a red arc from 9.1 to10 PSI� cabin altitude indicator – indicatescabin altitude (inside cabin) in thou-sands of feet from -1,000 to +16,000feet; the dial displays a red line at+10,000 ft� cabin vertical speed indicator –indicates vertical speed of cabin climbor descent in FPM from -2,000 to+2,000 FPM.

PressurizationSystems

5K-5


Cabin PressurizationFailure Warning SystemA pressurization warning system alertscrew of cabin altitude of 10,000 ft. Thesystem incorporates an altitude sensi-tive switch in the copilot’s side con-sole. The system responds as followsat 10,000 ft.

� a horn sounds (horn is silenced withthe HORN SILENCE button on pilot’sside of the center console) (Figure 5K-6)

� CABIN annunciator illuminates inthe Failure Warning Panel.

Depress the test button on the pedestalto check operation of the pressuriza-tion failure warning system.

Pressurization ControllerThe automatic pressure regulator (con-trol box) transmits electrical altitudeand rate of altitude change signals tothe outflow valves (Figure 5K-7 ). Itreceives 115V AC from the pilot’s ACbus. The regulator casing is tapped toreceive ambient cabin pressure andcontains:

� a cabin-altitude sensor – that sensesaltitude inside the cabin

� cabin-altitude variable-speed cap-sule – that detects cabin rate of alti-tude change

� rate-adjusting potentiometer – thatadjusts the rate of cabin climbs anddescends

� servo amplifier – that actuates theopening and closing of the solenoidvalves of the outflow control valves.

The outer scale of the dial on the pres-sure regulator indicates cabin/altitude.The inner scale indicates aircraft/flight/altitude.

The PULL for BARO knob on theright side of the regulator enables thepilot to rotate the inner scale and selecta flight altitude while concurrently set-ting a cabin altitude on the outer scalethat results in a cabin altitude differ-ential of 8.8 PSI.

Pulling the knob allows the altitudescales on the dial to reference baro-metric pressure in inches-of-mercury.

The RATE knob on the left side of theregulator adjusts the cabin altitude rate-of-change. The RATE knob is adjustedon the ground only. It should not beturned during flight.

The RATE knob position results in thefollowing:

� setting of the RATE knob gives a+650 FPM climb and -450 FPMdescent in the detent position.

� pulling the RATE knob and rotatingit full counterclockwise accomplishesa +300 FPM climb and -200 FPMdescent.

5K-75K-6


Pneumatic Systems

� pulling the RATE knob and rotat-ing it full clockwise accomplishes arate of +2,000 FPM climb and -1,300FPM descent.

For a typical flight, the automatic pres-sure regulator is set as shown below:� takeoff – displays cruise altitude onaltitude selector and 29.92 on the barowindow (IN.HG)� landing – displays field elevationon altitude selector and field baromet-ric pressure on the baro window(IN.HG).

Pressurization ModeSelectorThe pressurization mode selector onthe pressurization and conditioningcontrol panel contains a three-positionAUTO/MAN/DUMP toggle switch(Figure 5K-8 ). The positions functionas shown below:� AUTO – pressurization regulationof AC electrical commands sent tovalves through the control valveamplifier� MAN – valves are electrically iso-lated from the control bus and can onlybe controlled pneumatically by themanual control valve� DUMP – the valve receives A busDC electrical power to operate thesolenoid and dump cabin pressure.

The toggle switch uses a guard to pre-vent inadvertent dumping of pressur-

ization. The switch guard is first turnedclockwise, which permits the switchto move from AUTO to MAN. Toselect DUMP from MAN, move theswitch guard down.

Electropneumatic Outflow ValvesTwo electropneumatic outflow valveson the pressure bulkhead respond tosignals from the automatic pressureregulator. The valves establish ametered flow of air from the cabin tomaintain the selected cabin altitude orto establish a pressure rate of changecorresponding to the aircraft’s rate ofclimb or descent.

Each valve chamber contains a cali-brated jet that communicates betweenthe servo chamber and pressure source(closed by a solenoid-operated valve).Pressure sources are the cabin and thenon-pressurized zone set at 3.6 PSI.

A second calibrated jet communicatesbetween the servo chamber and thevacuum source (closed with a solenoid-operated valve). The vacuum source isa jet pump in the non-pressurized zonesupplied with engine bleed air; thisvacuum source provides air at belowambient air static pressure. Withengines failed, the jet pump no longersupplies vacuum; instead, the suctionpipe communicates with the outsidefor satisfactory performance.

5K-8

5K-24 For SimuFlite training only Falcon 10/100March 2001

The valve chamber also has a reliefvalve that communicates between theservo chamber and outside air as soonas an overpressure in cabin of 9.1 PSIis reached.

The chamber contains a fitting with acalibrated orifice that provides forinterconnection of both exhaust con-trol valves servo chambers and theirconnection with the three-way man-ual control valve.

In addition, the outflow valve has ananti-nicotine filter on the cabin air inletthat protects the valve against air con-tamination.

Manual ControllerManual operation is required if theautomatic pressure regulator mal-functions or AC power to the unit fails.Selecting the manual controllerremoves electrical power from theautomatic pressure regulator andreverts the system to pneumatic controlof the electropneumatic outflow valves.The cabin altitude (cherry picker) knobon the pedestal (Figure 5K-8 , previ-ous page) is used for manual control.

To initiate a cabin descent, cabin pres-sure is introduced into the controlchamber of each outflow valve tomove the valves toward the closedposition. Admitting ambient pressurefrom a static port into the controlchamber of each outflow valve movesthe valves toward the open position.This results in a higher cabin altitude.

Jet PumpThe jet pump in the rear compartmentuses motive flow from engine bleedair to establish a stable pressure in thecontrol chamber of each outflowvalve.

Automatic ModeThe automatic pressure controller reg-ulates the outflow valves to maintaincabin pressure according to:� commands set by the pilot� cabin pressure� rate of change in cabin pressure� aircraft configuration (on ground orin flight).

The outflow valves are closed for take-off when all of the following occur:� the power levers reach 82%� conditioning valve is closed� weight is on the gear� pressurization is in AUTO.

With power levers below 82% andmicroswitches in ground mode, theoutflow valves are commanded to a+500 FPM cabin climb rate (pressuredecrease) to relieve any pressure.

Dump ModeIn dump mode, DC power from the Abus is applied directly to the operat-ing solenoid with the pressurizationselector to DUMP. The power signalsthe outflow valves to open and dumpthe cabin to flight altitude.

Emergency PressurizationThe emergency pressurization systemprovides unconditioned warm air fromthe P2 unit directly to the cabin throughthe aft pressure bulkhead.

The emergency valve is controlledmanually with a mechanical knob onthe left-hand side of the aircraft abovethe couch. With SB F10-0003, an elec-tric valve replaces the mechanicalvalve, and a switch on the console nextto the pressurization controller con-trols the valve.

NOTE: The Freon systemshould be used to reduce cabintemperature when emergencypressurization is required.


Pneumatic Systems

ABG – SEMCAPressurizationSystemIn normal operation, the ABG –SEMCA pressurization system main-tains a cabin altitude of 8,000 ft at aflight altitude of 45,000 ft with maxi-mum differential of 8.8 PSI. Pressurerelief is set for 9.1 PSI.

Pressurization is maintained by twooutflow valves. One is an electro-pneumatic valve controlled by theautomatic (i.e., electric) pressure con-troller; the second is a pneumatic valveslaved to the electropneumatic valve.System components for the ABG –SEMCA are:� cabin pressurization indicatingsystem� cabin pressurization failure warningsystem� automatic pressurization regulator� pressurization mode selector� main regulating (electropneumatic)outflow valve� pneumatic (emergency) outflowvalve� manual pressurization controller(UP/DN knob)� jet pump.

Cabin PressurizationIndicating SystemThe cabin pressurization indicatinginstruments at the bottom of the cen-ter instrument panel monitor opera-tion of the cabin pressurization system.The instruments provide the follow-ing indications:� cabin differential pressure indica-tor – indicates readings from 0 to 10PSI; includes a green arc between 8.8and 9.0 PSI and a red arc from 9.1 to10 PSI

� cabin altitude indicator – indicatescabin altitude (inside cabin) inthousands of feet from -1,000 to+16,000 ft; the dial displays a red lineat +10,000 ft� cabin vertical speed indicator – indi-cates vertical speed of cabin climb ordescent in FPM from -2,000 to +2,000FPM.

Cabin PressurizationFailure Warning SystemA pressurization warning system alertsthe crew of a cabin altitude of 10,000ft. The system incorporates an altitude-sensitive switch in the copilot’s sideconsole. The system responds at10,000 ft as follows:� a warning horn sounds (horn issilenced with the HORN SILENCEbutton on pilot’s side of center con-sole)� the CABIN annunciator illuminatesin the Failure Warning Panel.

Depress test button on the pedestal tocheck operation of the pressurizationfailure warning system.

Pressurization ControllerThe automatic pressurization regula-tor on the lower right portion of thecenter instrument panel generates elec-trical commands to the electropneu-matic outflow valve. Pressurizationcontrol is accomplished with manip-ulation of the ALT, RATE, and BAROknobs.

The ALT knob allows for selection ofcabin pressure altitude from -1,000 ftthrough 10,000 ft (an inner scale showscorresponding flight level).

The RATE knob allows for selectionof rate-of-change for cabin climb anddescent. The detent provides a +650FPM rate-of-climb and a -450 FPMrate-of-descent for the cabin.


Movement of the RATE knob fullcounterclockwise allows a rate-of-climb of +200 FPM and a rate-of-descent of -150 FPM; full clockwiseallows a rate-of-climb of +1,200 FPMand a rate-of-descent of -830 FPM.

The BARO knob rotates the altitudedisplay ring to adjust for barometricpressure.

Following a comparison between actu-al vs. commanded cabin altitude andactual vs. commanded rate-of-change,the controller signals the electro-pneumatic valve torque motor to posi-tion the outflow valves as desired. Fordeviations between preset and truecabin altitude greater than 100 ft, therate-of-change is set by the RATEknob. For deviations less than 100 ft,the rate-of-change decreases as the alti-tude deviation decreases.

Pressurization ModeSelectorThe pressurization mode selector onthe pressurization and conditioningcontrol panel contains a three-position(AUTO/MAN/DUMP) toggle switch.The positions and functions are shownbelow:� AUTO – pressurization regulationof DC electrical commands (A bus)are sent to valves through the controlvalve amplifier� MAN – the valves are isolated elec-trically from the control bus and canonly be controlled pneumatically bythe manual control valve� DUMP – the valve receive A busDC electrical power to operate thesolenoid to dump cabin pressure.

The toggle switch uses a guard to pre-vent inadvertent dumping of pressur-ization. The switch guard is first turnedclockwise to permit the switch to movefrom AUTO to MAN. Next, the switchguard is moved down to select DUMPfrom MAN.

Main Regulating(Electropneumatic)Outflow ValveThe electropneumatic outflow valveon the top left side of the aft cabinbulkhead receives commands from theautomatic pressurization controller.

The valve is spring-loaded toward theclosed position. A combination of sta-tic, cabin, and negative pressures posi-tions the outflow valves.

The outflow valve is protected againstoverpressure. A pressure sensorattached to the valve body senses cabinstatic pressure and relieves cabin pres-sure higher than 9.1 PSI. The outflowvalve’s negative pressure relief deviceon the valve assembly causes the valveto open when ambient pressureexceeds cabin pressure.

Finally, cabin altitude is limited with asealed aneroid attached to the outflowvalve assembly; this aneroid expandsto open a port that allows cabin pres-sure to close the outflow valve whencabin altitude exceeds 12,500 ft.

In addition, the outflow valve has ananti-nicotine filter on the cabin air inletthat protects the valve against airbornecontamination.

Pneumatic (Emergency)Outflow ValveThe pneumatic (emergency) outflowvalve on the right side of the aft cabinbulkhead is identical to the electro-pneumatic valve except the torquemotor is replaced by a pneumatic relay.

In AUTO mode with the manualpressurization controller in DN (fullcounterclockwise), the relay essen-tially is inoperative; the pneumaticvalve slaves pressure equal to that ofthe electropneumatic valve becauseof a crossfeed pipe.

In MAN mode, the slaving pressureof the pneumatic outflow valve is con-trolled by the manual pressurizationcontroller (UP/DN knob).


Pneumatic Systems

The slaving pressure in the electro-pneumatic outflow valve is equal tothe other outflow valve because of thecrossfeed pipe. Electric control of theelectropneumatic outflow valve isdisabled.

Manual PressurizationController (UP/DN Knob)The manual pressurization controlleron the right lower portion of the cen-ter instrument panel controls the oper-ation of the pneumatic relay on thepneumatic outflow valve.

The controller is an adjustable orificethrough which cabin pressure and jetpump negative pressure are combinedto create a reference pressure at a pres-sure relay (bellow) which, in turn,causes operation of the outflow valve.At DN (full counterclockwise), a cabinrate-of-descent of -1,000 FPM is com-manded. At UP (full clockwise), acabin rate-of-climb of +2,500 FPM iscommanded.

Jet PumpThe jet pump in the rear compartmentuses motive flow from engine bleedair to establish a stable pressure in thecontrol chamber of each outflow valve.

Automatic ModeThe automatic pressure controllerregulates the outflow valves to main-tain cabin pressure according to thefollowing:� commands set by the pilot� cabin pressure� rate of change in cabin pressure

� aircraft configuration (ground or inflight).

With the power levers forward of 92%and pressurization in automatic mode,the outflow valves close.

With power levers below 92% andflight/ground relays in ground mode,the outflow valves are commanded toa +450 FPM cabin climb rate (non-adjustable) to relieve any pressure.

Dump ModeIn dump mode, Primary A bus poweris applied directly to the torque motoron the electropneumatic (left) outflowvalve to drive it open. The pneumatic(right) outflow valve follows (opens)as a slave. If cabin altitude approaches12,500 ft, a separate cabin altitude lim-iter on each outflow valve causes theoutflow valves to close pneumatically.

Manual ModeThe pressurization manual modeconsists of a manual (emergency)pressurization controller and a pneu-matic outflow valve. The manual pres-sure regulator:� sets the required climb or descentrate� holds cabin altitude at a constantvalue.

The manual pressurization controllerknob UP or DOWN position is usedin relationship to cabin altitude. If ahigher cabin altitude is desired, rotatethe knob toward UP; if a lower alti-tude is desired, rotate toward DOWN.For a level cabin altitude, the knobshould be near 12 o’clock.


Pneumatic Systems

PreflightPreflight of the pneumatic, air condi-tioning, and pressurization systems isaccomplished in accordance with thePreflight chapter of this manual.Normal operation of these systems isaccomplished in accordance with theExpanded Normals and Maneuverschapters of this manual.

Abnormal ProceduresThis section discusses what happenswithin the pneumatic, pressurization,and air conditioning system duringabnormal or situations. For a list ofspecific procedural steps, refer to theAFM or the SimuFlite OperatingHandbook.

COND’G OVHT Warning

The COND’G OVHT annunciator onthe Failure Warning Panel illuminatesif a temperature of 250°C or greater isdetected in the conditioned air supplyduct.

The likely cause of a COND’G OVHTis a faulty HP injector. HP air is beingadded to LP air at too high an engineRPM so that the air in the duct betweenthe conditioning valve and the heatexchanger is too hot. The maximumtemperature of LP air is 230°C; there-

fore, the excess heat is coming fromHP air. Turn the HP bleed OFF andthe problem should be solved; the lightshould extinguish.

If annunciator remains illuminated,manipulate the throttles to see if oneof the HP valves may be stuck open. Ifthe annunciator remains on, land assoon as possible, open the emergencypressurization valve (Figure 5K-9 ),and close the conditioning valve. TheFreon system may be used to cool thecabin while using emergency pressur-ization.

Conditioning Valve Failsto Open on Takeoff

The conditioning valve closes as poweris advanced on takeoff (at 82°). Onceairborne, a separate delayed timer cir-cuit associated with the conditioningvalve governs the progressive open-ing of the valve. The valve is fullyopen within two minutes followingtakeoff. Should a conditioning valvefail to open, conditioned air is notavailable to the cabin or cockpit. Thiscondition may be due to a failure ofthe timer circuits. This condition hap-pens shortly after takeoff. If bleed aircannot be restored, flight at high alti-tudes is not possible.

5K-9



If there is no bleed air directed to thecabin, select the conditioning valve toON and monitor pressurization. If nochange occurs, return for a landing.

Cabin Pressure Too High

Cabin pressure too high is recognizedby referencing the cabin pressure andcabin vertical speed indicators.

Initially, select a higher altitude on theautomatic pressurization controller toincrease the cabin altitude referencefor the pressurization system. This low-ers the cabin pressure. If this action issuccessful, continue the flight with thehigher altitude.

If the cabin pressure does not decrease,shift the pressurization from automat-ic control to manual control. Move theUP/DN knob using brief actuations toavoid a “bump” in pressurization.Control aircraft pressurization with theUP/DN knob.

Selection toward UP increases cabinaltitude (by decreasing cabin pressure),and selection toward DN decreasescabin altitude. Monitor cabin verticalspeed.

If manual pressurization does not solvethe overpressure problem, return thepressurization switch to AUTO andmove the UP/DN knob to UP asbefore. If the cabin pressurizationremains too high, close the condition-ing valve and use oxygen as required.Land as soon as practical.

Improper Cabin Vertical Speed

Improper cabin vertical speed is detec-ted on the vertical speed needle of thecabin vertical speed indicator to theright of the center panel.

First, check that the pressurization con-trols are positioned properly. Adjustthe rate knob on the automatic pres-surization controller. If this solves theproblem, continue the flight withincreased surveillance of the verticalspeed indicator.

If the problem is not solved, shift con-trol of the pressurization system fromautomatic to manual. Move the UP/DN knob using brief actuations to avoida “bump” in pressurization. Continuethe flight controlling aircraft pressur-ization with the UP/DN knob. Selectiontoward UP increases cabin altitude (bydecreasing cabin pressure), and selec-tion toward DN decreases cabin alti-tude. Monitor the pressurization sys-tem with reference indicators.

If improper indications continue, selectMANUAL on the pressurizationswitch and move the UP/DN knob asbefore.

Pressurization Loss(High Cabin Altitude)

A loss of cabin pressure is indicatedby a high cabin altitude on the CABINFT x 1000 indicator and/or illumina-tion of the red CABIN annunciator onthe Failure Warning Panel and sound-ing of the cabin altitude warning horn(which can be silenced with theHORN-SILENCE button on thepedestal).

Check that the COND’G VALVEswitch is in the proper position.Attempt to gain control of the pres-surization system by switching fromautomatic mode to manual. To lowerthe cabin altitude, move the UP/DNknob toward DN. Monitor the cabinvertical speed indicator.

Don oxygen masks, select cockpitcommunications, place the NOSMOKE switch to ON, and deploy thepassenger oxygen masks by selectingMANUAL/OVERRIDE on the pas-senger oxygen controller.

If the condition continues, selectAUTO on the pressurization switchand use the UP/DN knob again. If thealtitude remains too high, open theemergency pressurization valve; if thecabin altitude is uncontrollable, con-tinue the flight below 10,000 ft.


Pneumatic Systems

Emergency ProceduresThis section discusses what happenswithin the pneumatic, pressurization,and air conditioning system duringemergency situations. For a list of spe-cific procedural steps, refer to the AFMor the SimuFlite Operating Handbook.

Air Conditioning Smoke

The Phase I (memory) procedures aredesigned to initially take care of a lifethreatening situation. Subsequent stepsidentify and eliminate the source ofthe smoke.

Air conditioning smoke normally isassociated with oil entering the bleedair/conditioning system. Close the HPbleed valves and reduce engine thrust.Select WSHLD/FLOOR valves COLDand open pilot gaspers. Pull the FreonCB and bring either engine to idle.

If smoke continues, manipulate thepower levers to identify the engineproducing the smoke though the LPbleed air ducting. If smoke stops, leavethat engine at idle. If smoke stillcontinues, select passenger oxygenMANUAL/OVERRIDE, don masks,take the aircraft below 12,000 ft, andselect the conditioning valve closed.

At less than 200 kts, open the DV win-dow for smoke evacuation, turn offpassenger oxygen when the cabinsmoke clears.

Rapid Decompression/Emergency Descent

The Phase I (memory) proceduresdesigned to take care of life and deathsituations begin with the crew donningoxygen masks, selecting communica-tions and executing a maximum per-formance descent.

Passenger oxygen masks are donned,ignition is turned on to prevent engineflameouts, ATC is advised, and thetransponder is set to emergencysquawk. The aircraft descent must berapid to minimize the risk of losingconsciousness at high altitude.

Smoke Removal

While no smoke removal procedure ispublished, several procedures in theFlight and Operating Manuals do givespecific instructions dealing with thisproblem. The Electrical Fire or SmokeProcedure calls for opening the wind-shield and floor valves to move thesmoke to the rear. Further ventilationis automatically ensured by the air con-ditioning system. The air conditioningSmoke and Odor Procedure suggestsopening the DV window, however, thecabin must be completely depressur-ized. Smoke evacuation will occurwhile using the pressurization systemto depressurize and will continue afteropening the DV window. The airspeedmust be below 200 kts to open the DVwindow and it can be removed forbetter vision.


Pneumatic Systems

DataSummaries

Pneumatic SystemsPneumatic System

Power Source LP/HP bleed air – either engine

Distribution Air conditioning via sidewall ducts and cockpitdistributors

Wingshield defog outlets

Cockpit footwarmer

Emergency pressurization

Outflow valve control pressure and vacuum

Control HP BLEED switch


Emergency pressurization valve

Cabin temperature regulator

Footwarmer defog valves

Anti-ice switches

Engine power setting

Monitor ENG 1/2 anti-ice lights

WINGS anti-ice light

AnnunciatorsCOND O’HEATHYD. TK. 1/2CABIN (cabin pressure over 10,000 ft)DEICING (some aircraft )

Cabin temperature gage (if installed )

Protection HP check valves

LP check valves

Check valves – aft pressure bulkheads


Pressurization Systems

Power Source HP/LP bleed air from either or both engines

W bus (115V AC) (aircraft with IDC controller )

A bus (28.5V DC) (aircraft with ABG-SEMCAcontroller )

Distribtion Pressurized fuselage area

Control CABIN ALTITUDE controllerRATE knobPULL FOR BARO knob

AUTO/MAN/DUMP pressurization mode selector

UP/DOWN manual pressurization knob (cherry picker)

HP BLEED switch


Monitor CABIN annunciator (10,000 ft)

IndicatorsCabin rate of climbCabin altitudeCabin differential pressure

Protection Overpressuration controller (9.1 PSI)

Squat switch (landing depressurization)

Freon Air Conditioning System

Power Source Main bus via operating generator or GPU

Distribution Air conditioning ducts

Control SwitchesFreon systemAuxiliary busPressure

Temp regulation valve

Start buttons

30-second timer

Monitor Generator ammeter

Cabin temperature gage

Protection Suction pressure switches

Discharge pressure switches

Starter button shutoff

Generator/GPU power source relay

Falcon 10/100 For SimuFlite training only 5L-1April 2000

Powerplant

Chapter 5L

The Powerplant chapter contains information on several areas:

� the turbofan engine, including its components

� instrumentation and operation

� engine oil and lubrication

� ignition

� engine fuel and fuel control

� power control.

Two Garrett TFE731-2-1C turbofan engine power the Falcon10/100. Each engine produces 3,230 lbs of static takeoff thrustat sea level. The TFE731 is a lightweight, low noise, two-spool,front fan engine with a medium bypass ratio. The engine’smodular engine design allows for ease of maintenance andrepair.

5L-2 For SimuFlite training only Falcon 10/100April 2000


Powerplant

Table ofContents

Garrett TFE 731-2-1C Engine Schematic . . . . . . 5L-6

Basic Operation . . . . . . . . . . . . . . . . . . . 5L-7

Turbofan Engines . . . . . . . . . . . . . . . . . . . 5L-7

Engine Components . . . . . . . . . . . . . . . . . 5L-7

Fan . . . . . . . . . . . . . . . . . . . . . . . . . 5L-7

Compressors . . . . . . . . . . . . . . . . . . . . 5L-7

Spools and Indicators . . . . . . . . . . . . . . . . 5L-8

Low Pressure Spool (N1) . . . . . . . . . . . . . 5L-8

N1 Indicator . . . . . . . . . . . . . . . . . . . . 5L-8

High Pressure Spool (N2) . . . . . . . . . . . . . 5L-9

N2 Indicator . . . . . . . . . . . . . . . . . . . . 5L-9

Annular Combustor . . . . . . . . . . . . . . . . . 5L-9

Interstage Turbine Temp (ITT) Indicator . . . . . . . 5L-9

Transfer and Accessory Gearboxes . . . . . . . . 5L-10

Engine Bleed Air . . . . . . . . . . . . . . . . . . . 5L-10

Engine Oil System Schematic . . . . . . . . . . . 5L-12

Oil Flow Schematic . . . . . . . . . . . . . . . . . 5L-14

Powerplant Systems . . . . . . . . . . . . . . . . 5L-15

Engine Oil System . . . . . . . . . . . . . . . . . . 5L-15

Oil Level Sight Gage . . . . . . . . . . . . . . . . 5L-16

Dual Oil Pressure/Temperature Indicator . . . . . 5L-16

Oil Warning . . . . . . . . . . . . . . . . . . . . 5L-16

Ignition System . . . . . . . . . . . . . . . . . . . 5L-16

Start Selector Switch . . . . . . . . . . . . . . . . 5L-17

Start Pushbutton . . . . . . . . . . . . . . . . . . 5L-17

Start Sequence . . . . . . . . . . . . . . . . . . 5L-17

Fuel Control System . . . . . . . . . . . . . . . . . 5L-18

Fuel Pump . . . . . . . . . . . . . . . . . . . . . 5L-18

Hydro-Mechanical Fuel Control Unit . . . . . . . . 5L-18


Electronic Engine Control . . . . . . . . . . . . . 5L-18

Surge Bleed Valve . . . . . . . . . . . . . . . . 5L-19

Computer Failure . . . . . . . . . . . . . . . . . 5L-20

Fuel Density Adjustment . . . . . . . . . . . . . 5L-20

Fuel Flow Divider Assembly . . . . . . . . . . . . 5L-21

Fuel Nozzles . . . . . . . . . . . . . . . . . . . . 5L-21

Fuel Flow Indicators . . . . . . . . . . . . . . . . 5L-21

Start Pressure Regulator . . . . . . . . . . . . . 5L-21

Engine FIRE PULL Handles . . . . . . . . . . . . 5L-22

Power Control System . . . . . . . . . . . . . . . . 5L-22

Preflight and Procedures . . . . . . . . . . . . . 5L-25

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5L-25

Servicing . . . . . . . . . . . . . . . . . . . . . . . 5L-25

Abnormal Procedures . . . . . . . . . . . . . . . . 5L-26

Oil Pressure . . . . . . . . . . . . . . . . . . . . 5L-26

Engine Computer Light ON . . . . . . . . . . . . 5L-26

Engine Overspeed . . . . . . . . . . . . . . . . . 5L-26

Engine Airstarts . . . . . . . . . . . . . . . . . . 5L-26

Engine Failure During Takeoff . . . . . . . . . . . 5L-27

Engine Failure – Shutdown in Flight . . . . . . . . 5L-27

Emergency Procedures . . . . . . . . . . . . . . . 5L-28

Engine Fire . . . . . . . . . . . . . . . . . . . . . 5L-28

Limitations . . . . . . . . . . . . . . . . . . . . . 5L-29

Temperature, RPM, and Time Limitations . . . . . . 5L-29

Data Summary . . . . . . . . . . . . . . . . . . . 5L-31


5L-6

Powerplant

GarrettTFE731-2-1C

Engine


ACCESSORYGEARBOX

PT2/TT2

N2 MONOPOLE

TRANSFERGEAR BOXASSEMBLY

4-STAGELOW-PRESSURECOMPRESSOR

HIGH-PRESSURECOMPRESSOR

OIL SCAVENGELINE

HIGH-PRESSURETURBINE

FUEL MANIFOLD(PRIMARY/SECONDARY)

3-STAGELOW-PRESSURETURBINE

N1 MONOPOLE

FUELNOZZLEBYPASS

DUCT

PLANETARYGEARBOX

FAN

Garrett TFE731-2-1C Engine


Powerplant

Turbofan EnginesJet engines accelerate air to producethrust. Thrust output can be generat-ed two ways: a small volume of airaccelerated to a high velocity or a largevolume of air accelerated to a lowvelocity.

The turbofan engine utilizes bothmethods. Only a portion of the incom-ing air is compressed, mixed with fuel,combusted, and exhausted at a highvelocity. The fan compresses andaccelerates a large volume of air at alow velocity and bypasses it aroundthe core of the engine without mixingit with fuel or using it for combustion.

The relation of the mass of bypassedair to the mass of air going through thecombustion chamber is known as thebypass ratio. The TFE731-2-1C is con-sidered a medium-bypass engine witha bypass ratio of 2.8 to 1, rated at 3,230lbs static thrust.

EngineComponentsThe TFE731 engine consists of fivemajor components:� spinner and fan� low pressure spool (N1)� high pressure spool (N2)� annular combustor� transfer and accessory gearboxes.

FanThe fan is an axial flow fan that moveslarge quantities of air into the bypassstator and low pressure (core) inlet.Energy is translated into pressure bythe rearward acceleration of air.Approximately 280% more air passesthrough the fan bypass duct thanthrough the engine core. At sea level,the fan produces approximately 70% ofthe total thrust. The low pressure spooldrives the fan through a planetary geardrive system.

CompressorsAxial and centrifugal compressors areused in the TFE731 engine. Axial flowcompressors accelerate air rearwardwith increasing velocity through eachstage where kinetic energy is translat-ed into pressure. A stage is a consec-utive pair of rotors (rotating blades)and stators (non-rotating blades).

Centrifugal compressors consist of animpeller (rotor), a diffuser, and a com-pressor manifold. Air picked up andaccelerated outward toward the dif-fuser causes the accelerating air’skinetic energy to be translated intopressure. The diffuser maintains themaximum amount of energy impartedby the impeller.

Engine thrust on the Falcon 10/100begins with the acceleration and com-pression of inlet air by the front axialfan (Figure 5L-1 ). Air is then split into

BasicOperation

5L-1


two streams; one is passed around theengine core to the exhaust nozzle bythe bypass duct, and the other is com-pressed by the four-stage LP compres-sor. The single-stage HP centrifugalcompressor further compresses the airbefore it enters the combustion chamber.

Spools and IndicatorsThe two spools of the engine are theN1 (or LP) spool and the N2 (or HP)spool. The LP compressor and LP tur-bine are connected by a common shaftto form the N1 spool. The center sec-tion of the N1 shaft passes through theinterior of a much shorter shaft. Theouter, concentric shorter shaft connectsthe HP compressor and HP turbine toform the N2 spool.

The HP spool compressor and turbineare between the LP compressor andLP turbine. The N1 shaft rotates inde-pendently of the N2 shaft. The mostforward end of the N1 shaft extendsinto and drives the planetary reductiongearbox. A tower shaft, perpendicularto and driven by the N2 shaft, drivesthe transfer gearbox. A shaft from thetransfer gearbox drives the accessorygearbox. During start, the starter drivesthe accessory section to rotate the N2spool through the transfer gearbox.

Low Pressure Spool (N 1)

The low pressure spool (N1) consists ofa four-stage, low pressure, axial flowcompressor and a three-stage, lowpressure turbine. Both low pressurecompressor and turbine stages aremounted on a common shaft. Air isaccelerated rearward through eachstage (i.e., rotor and stator pair) withincreasing pressure. Low pressurespool rotational speed is expressed asN1. The N1 shaft drives the fan throughthe planetary gears.

N1 Indicator

Low pressure compressor speed foreach engine is measured by a pickupon the aft end of the low pressurerotor shaft. The pickup produces dualindependent signals; one set of sig-nals is used by the electronic enginecontrol (EEC) for its operation.

The other set of signals is displayed aspercent N1 RPM (graduated from 0to 100) on an analog/digital indicatoron the center instrument panel (Figure5L-2). Power for the indicating cir-cuit is supplied through CBs from theassociated A or B bus or, during lowtemperature starts, from the Batterybus.

5L-35L-2

Falcon 10/100 For SimuFlite training only 5L-9March 2001

Powerplant

High Pressure Spool (N 2)

The high pressure spool consists of asingle stage centrifugal compressordriven by a single stage turbine throughan outer concentric shaft. Air is accel-erated outward by the impellers withincreasing force; kinetic energy istranslated into pressure by the action ofthe impeller and diffuser. The high pres-sure spool also drives the accessorysection through a tower shaft and gearreduction system. High pressure spoolrotational speed is expressed as N2.

N2 Indicator

HP compressor (N2) speed is sensedfrom a pickup on the transfer gearboxthat provides dual independent signals.One set of the speed signals is sent tothe EEC for use in its operation. Theother set of signals is sent to the cock-pit indicator on the center instrumentpanel (Figure 5L-3 ). The indicator dis-plays percent RPM (graduated from 0to 100) on an analog/digital indicator.Power for the N2 indicating circuit issupplied from either associated A orB bus or, during low temperature starts,from the Battery bus.

Annular CombustorTo decrease the length of the engine, acompact, reverse flow, annular com-

bustor is used. In the combustionchamber, fuel is introduced into thereverse flow annular burner by 12duplex fuel spray nozzles. The air andfuel are mixed, ignited, heated, andexpanded. Hot gases pass through thehigh and low pressure turbines, driveboth rotating compressor assemblies,and exit the exhaust nozzle with thebypassed air (Figure 5L-4 ). Fuel isintroduced upstream of the primaryignition zone to allow premixing of airand vaporized fuel in the combustor.

Two igniter plugs at the six and seveno’clock positions in the combustionplenum provide a discharge spark of18,000 to 24,000 volts at a rate of oneto five sparks per second.

Interstage TurbineTemperature (ITT)IndicatorTen Chromel-Alumel thermocouplesin the gas path between the high pres-sure turbine and first stage of the lowpressure turbine sections measureInterstage Turbine Temperature (ITT).An average temperature is presentedin analog and digital format on a cock-pit indicator on the center instrumentpanel (Figure 5L-5 ) and to the EEC.The indicator shows ITT in degrees

5L-55L-4

Engine Operating LimitationsRefer to the end of the chapter forengine operating limitations.

ITT LimitsRefer to the end of the chapter forITT limits.


centigrade. Power for the analog ITTindicating system is from the thermo-couple; digital-type gages receivepower from the A or B bus.

Transfer and AccessoryGearboxesThe transfer and accessory gearboxesare mounted under the engine (Figures5L-6 and 5L-7). The transfer gearbox,which powers the accessory gearbox,is driven by the high pressure spool(N2) through a tower shaft. The N2monopole pickup, which senses HProtor speed, is also mounted on thetransfer gearbox.

A hydraulic pump and starter/genera-tor are mounted on the forward sideof the accessory gearbox; an enginefuel pump/control and oil pump aremounted on the aft side of the acces-sory gearbox.

Engine Bleed AirEach of the engines on the Falcon 10/100 has three bleed air ports. Two LPbleed ports at 3 o’ clock and 9 o’ clocktake bleed air aft of the last stage ofthe axial compressor; one HP bleedport at 12 o’ clock takes air fromdownstream of the centrifugal com-pressor.

Each bleed air system is equipped withcheck valves that prevent air fromreturning to the engines when they areshut off or running at low RPM.

All of the engine bleed ports providethe aircraft pneumatic system with air.In addition, each engine also has aninternal HP bleed that is used for en-gine anti-icing. For more informationon bleed air, see the Pneumatics chap-ter; for more information on engineanti-icing, see the Ice and Rain chapter.

SPARE

STARTERGENERATOR

HYDRAULIC PUMP

FUEL PUMP AND HYDROMECHANICAL FUEL CONTROL REGULATOR

CONNECTION TOTRANSFERGEARBOX

OILFILTER

LUBRICATINGPUMPS

FRONT VIEW

REAR VIEW

ACCESSORY DRIVE GEARBOX

5L-7

5L-6


5L-12

Powerplant

Engine OilSystem


BREATHERPRESSURIZINGVALVE 3.5 PSIA

TRANSFERGEARBOX

ASS'Y SCAVENGECOMMONSCAVENGE

FILTER

FILTER BYPASSVALVE

ANTI-SIPHONORIFICE

OIL TANK

VENT

OILCOOLER

(HALF SECTION)

BYPASS VALVE TEMPAND PRESSURE

FUEL IN

FUEL OUT

OIL COOLER(QUARTER SECTION)

CHECKVALVE

FUELIN

FUELOUT

BYPASSVALVE

FUEL HEATER

METAL PARTICLEDETECTOR

TEMPERATURESENSING PORT

PT PRESSURE TRANSMITTER

LUBE ANDSCAVENGE

OIL PUMP

PLANETARYGEARBOX ASS'Y

NO 1, 2 AND 3BEARING

NO.6 BEARINGSUMP

HIGH PRESSURE OIL

SCAVENGE OIL

VENT LINE

OIL SUPPLY

A BUS

OIL 1

A BUS

B BUS

LEFT

RIGHT

REGULATOR ANDRELIEF VALVE

100

150

20

60

40

80

50

0 0

°C

OIL

PSI

P INDICATOR

OIL PUMP INLET

BYPASS VALVE

FUEL/OIL COOLER

TRANSFERGEARBOX ASS'Y

NO.4 AND 5BEARING CAVITY

ACCESSORY DRIVEGEARBOX ASS'Y

Engine Oil System


Powerplant

SCAVENGEPUMPS

CHIP DETECTOR

OIL PUMP (HIGH PRESSURE)

PUMP RETURNFLUID

PRESSURE

OIL SUPPLY

RETURN

FUEL/OILCOOLER

AIR/OILCOOLER

OIL TANK

OILFILTER

FUEL HEATER


Oil Flow


Powerplant

Engine Oil SystemOil is provided under pressure to thebearings, transfer, accessory, and plan-etary gearboxes. The system consistsof:� oil reservoir� oil pump� filter and housing� filter bypass valve� fuel/oil cooler and bypass valve� three surface air/oil coolers� oil temperature control and bypassvalve� breather pressurizing valve� air/oil separator� oil sight gage� chip detector� pressure and temperature trans-mitters.

Oil is drawn from the oil tank by therotation of the engine oil pump on thebottom of each engine. Oil pump out-put is directed through a pressure reg-ulator then through a paper cellulousfilter before delivery to the enginebearings, the transfer gearbox, theaccessory gearbox, and the fan reduc-tion planetary gear assembly. Thepressure regulator maintains oil pres-sure at 42 ± 4 PSIG.

Filter restriction of more than 25 PSIDis indicated by projection of a red pinon the right side of the engine acces-sory gearbox. The indicator pin is vis-ible through a small access hole in theengine cowling. Once the pin pops out,it remains out until reset. The pin pro-vides no indication to the cockpit dur-ing flight; therefore, it should bechecked on pre- and post-flight inspec-tions. If the pin is observed to beprotruding, an impending bypass isindicated; maintenance action isrequired.

In the event of filter blockage, engineoil lubrication is rerouted past the fil-ter by a bypass valve when there is apressure drop of 30 to 40 PSID acrossthe filter, and the indicator pin popsout.

After oil leaves the filter, it passesthrough a fuel heater and oil-to-air heatexchangers in the fan bypass duct. Atemperature control valve allows theoil to bypass the oil-to-air heatexchangers until oil temperature risesagain. Oil at temperatures higher than65°C passes through the three seg-mented oil-to-air heat exchangers. Ifthe coolers become obstructed, thetemperature control valve reroutes theoil around them.

Part of the oil is then distributed to theNo. 6 bearing sump, No. 4 and No. 5bearing sumps (HP rotor shaft), trans-fer gearbox bearings and gear meshes,and accessory gearbox and gearmeshes. Oil is also routed through atemperature control valve at the fuel/oil cooler assembly. If oil temperatureexceeds 99°C, the temperature controlvalve at the fuel/oil cooler opens toroute oil through the cooler.

The oil then lubricates fan shaft bear-ings No. 1 and No. 2, front low pres-sure spool bearing No. 3, and all plan-etary gearbox bearings and meshes.

After traveling to all main sump areas,oil drains by gravity to the lowest pointin each sump. The sumps are con-nected to four scavenge pumps in thepump pack. A magnetic particle chipdetector is in the scavenge pump oilreturn line; if metallic particles collecton the detector, the OIL 1 or OIL 2light illuminates.

Spectrometric oil analysis program(SOAP) and oil filter residue analysesare used as maintenance procedures tomonitor the condition of the engine.This is done by determining the sus-pended particle content and composi-tion in the lubrication system.

PowerplantSystems


The condition of engine parts contactedby engine oil can be monitored, andimpending failure of specific parts canbe predicted from the amounts of metalin the oil.

Venting for the oil tank and lubricat-ing system is via the accessory gear-box. The accessory gearbox is ventedto the atmosphere through a normallyopen breather pressurization valve. Aspressure altitude increases above27,000 ft, the breather pressurizationvalve closes to maintain an engine casepressure of approximately 3.5 PSIA.

Oil Level Sight GageThe sight gage for engine oil level isbuilt into the oil tank on the right sideof the engine (Figure 5L-8 ). The sightgage may be observed through smallopenings in the cowling for eachengine. There is no oil level indicationin the cockpit. Oil quantity indicationsare not accurate unless checked with-in one hour of engine operation; nor-mally, the check is conducted duringthe post-flight inspection.

Dual Oil Pressure/Oil Temperature IndicatorDual oil pressure/oil temperature indi-cators display engine oil pressure andengine oil temperature for each engine(Figure 5L-9 ). An oil pressure trans-mitter in the oil line to the planetarygearbox transmits the operating pres-sure to the indicator. Normal operat-ing pressure is 38 to 46 PSI, with a

minimum (red line) of 25 PSI and amaximum (red line) of 55 PSI for 3minutes or less.

Information for engine oil temperatureindication consists of two resistancebulbs that provide a signal to the temp-erature side of the dual indicator onthe instrument panel. Normal operatingtemperature is 30 to 113°C up to analtitude of 30,000 ft with a maximumoperating temperature of 132°C above30,000 ft. A transient limit of 140°Cis permitted for a maximum of twominutes.

Power for the oil pressure and oil tem-perature indicating systems is providedthrough CBs from the A and B busesfor their respective engines.

Oil WarningOIL 1 and OIL 2 warning lights on thefailure warning panel are operated bya pressure switch in the oil line to theplanetary gearbox. The lights illumi-nate for an oil pressure less than 25PSI, or if the metal particle detector inthe oil return line detects metal chips.The lights serve as a backup to the oilpressure indicator.

Ignition SystemEach engine has an independent igni-tion system that consists of the fol-lowing components:� ignition exciter box� ignition leads� igniter plugs

5L-105L-95L-8

CAUTION: Do not exceed idleRPM with oil temperature lessthan 30°C unless ambient temp-erature is so low as to preventwarmup, when slightly higherRPM may be used.


Powerplant

� ignition switch� circuit breakers.

An ignition exciter box on the upperleft side of the fan bypass housing is ahigh voltage, capacitance discharge,noise-suppressed, intermittent spark-ing-type unit utilizing 10 to 30V DC.It provides 18,000 to 24,000 volts tothe igniter plugs. Continuous opera-tion is possible when flight conditionsrequire it.

The igniters, mounted at the six andseven o’clock positions on the annu-lar combustor, operate independentlyof each other. Each plug fires at a rateof 1 to 5 sparks per second when trig-gered by the ignition unit.

The IGN lights on the overhead panelabove the start selector switches illu-minate when power is supplied to theexciter box; the lights do not indicatethat the igniters are actually firing.Continuous ignition is recommendedfor all takeoffs and landings and incontaminated runway conditions.

Start Selector SwitchA three-position start selector switchfor each engine is on the overheadpanel (Figure 5L-10 ).

In the GND START position, the start-ing circuitry is armed for an enginestart if the power lever is in the cutoffposition. After the start pushbutton ispressed and the power lever is subse-quently moved to idle, power is sup-plied to the ignition system as fuel issupplied to the combustion chamber.When the engine N2 RPM exceeds50%, the EEC terminates the start andturns the ignition off.

The MOTOR START-STOP positionof the switches provides the controlcircuits to motor the engine withoutignition system operation. This posi-tion can also be used to end a startsequence if the EEC fails to terminatethe start. With the switch in MOTORSTART-STOP, the start pushbutton canbe pressed; the starter/generator motors

the engine as long as the start push-button is held. The ignition systemdoes not function with the switch inMOTOR START-STOP.

Continuous ignition is available in theAIR START position, which activatesthe ignition circuit. This position pro-vides ignition for takeoffs, landings,and contaminated runway operationsbut does not provide power for starteroperation.

Normally supplied by A and B busesrespectively, engine ignition will alsoreceive Battery bus power in the eventof Main bus failure or generator failurewhile in flight.

Start PushbuttonThe momentary-type start pushbuttonsfor each engine are on the overheadpanel. Momentarily pressing the startpushbutton for approximately one totwo seconds begins the engine startsequence if the power lever is in thecutoff position, the respective genera-tor switch is in ON, both battery switch-es are in ON, and the start selectorswitch is in GROUND-START. Oncethe start sequence begins, the start push-button need not remain pressed becausethe start circuit is a self-latching cir-cuit. With the preceeding conditions,the ignition system automatically func-tions when the power lever is movedto the idle position.

Start SequenceNormal engine start is accomplishedwith a DC-powered starter on theaccessory section of the engine. Thestarter turns the high pressure spoolthrough the tower shaft. The HP com-pressor provides enough airflow andfuel pressure for engine starting at 10%N2. The pilot monitors the N2 RPMand, when in excess of 10% N2 andN1 rotation is indicated, advances thepower lever from cutoff to idle. Withthe power lever in idle, fuel is avail-able from the fuel control unit (FCU).The EEC begins to meter the fuel flow

5L-18 For SimuFlite training only Falcon 10/100March 2001

to the combustion section, and igni-tion is activated. Once the engine pas-ses approximately 50% N2, the EECturns off the igniters and the starter.

Fuel ControlSystemThe Falcon 10/100 fuel control sys-tem consists of:� fuel pump assembly� hydro-mechanical fuel control unit(FCU)� electronic engine control (EEC)� fuel flow divider assembly� fuel nozzles� engine FIRE PULL handles.

Fuel is pumped by the engine fuelpump, filtered by the fuel filter,metered by the hydro-mechanical fuelcontrol, delivered by the fuel mani-fold, and atomized by the fuel nozzles.

Fuel PumpThe engine-driven fuel pump, mount-ed on the accessory gearbox, is actuallytwo pumps: a centrifugal, low pres-sure boost pump and a vane-type, highpressure pump. The pump also con-tains a filter element, filter bypassvalve, pressure relief valve, and dif-ferential pressure indicator.

The low pressure pump draws fuelfrom the aircraft fuel system, whichthen passes through a fuel filter and isdelivered to the high pressure pumpelement. To ensure continuous engineoperation if the filter clogs, the filterbypass valve opens to deliver fuel tothe high pressure pump. If the filterstarts to clog, a red indicator pin onthe valve pops out (6 to 8 PSID) towarn of potential failure. Once the pinpops out, it remains out until reset. Thepin provides no indication to the cock-pit during flight; therefore, it shouldbe checked on pre- and post-flightinspections. If the pin is observed to

be protruding, maintenance action isrequired.

The high pressure pump supplies highpressure fuel to the fuel meteringvalves in the FCU. The FCU metersthe fuel according to power lever posi-tion and EEC inputs when the EEC isfunctioning. When the EEC is func-tioning normally, the capacity of thefuel pump exceeds the needs of theengine; therefore, all of the excess fuelis returned to the high pressure pumpinlet.

Hydro-MechanicalFuel Control UnitThe fuel control unit (FCU) is attachedto the rear of the fuel pump and isdriven by the fuel pump shaft. It incor-porates a metering valve, a fuel meter-ing torque motor, pressure regulatorvalve, bypass valve, mechanical fly-weight governor, compressor dischargelimiter (P3 air), shutoff valves, ulti-mate overspeed solenoid, and powerlever potentiometer.

The FCU receives commands from theEEC in normal operation. In the MAN-UAL mode, the fuel flow is com-manded by throttle control of P3 pres-sure. To protect the engine from amalfunctioning EEC, a governor isbuilt into the FCU. The governor,mounted on the main shaft, protectsthe engine from exceeding 105% RPMregardless of EEC commands; it doesnot shut the engine off, but limits theRPM to less than 105%.

Two shutoff valves are in the FCU.One is the power lever actuated shut-off valve, and the other is a shutoffvalve commanded by the EEC if ulti-mate overspeed is sensed.

Electronic Engine ControlThe electronic engine control (EEC)(P/N 2101142) in the baggage com-partment provides efficient fuel sched-uling in addition to overspeed, surge,

NOTE: Some very early aircraftmay still have the early EEC(P/N 949572-5 or -8.


Powerplant

and temperature protection for theTFE731 engine (Figure 5L-11 ). TheEEC receives inputs of engine inletpressure and temperature, ITT, N1 andN2 RPM, and power lever position viaa potentiometer on the FCU. With thisinformation, the EEC provides start,idle through maximum thrust sched-uling, acceleration, deceleration, andminimum fuel scheduling.

The ENGINE COMPUTER switch foreach engine on the overhead panel con-trols power to the EEC. Power for theleft engine EEC is available from theA bus; power for the right engine isfrom the B bus. Each receives powerthrough CBs behind the copilot seat.If bus power fails, the power to theswitch is supplied directly from thegenerator of the respective enginewhen the generator switch is on. CBsare connected to the starter/generatorcircuit at a point where, during start,power is supplied through an alternatesource for the computer switches.

With the engine computer switch on,the EEC is on and the metering valveis held open by compressor dischargepressure (P3 air) and regulated fuelpressure. The fuel control torque motorand bypass valve controls fuel deliveryrate. The EEC determines the amountof fuel required for the power leverposition and returns excess fuel to thehigh pressure pump inlet. Through nor-mal metering of the fuel flow, the com-

puter keeps the engine within its thrust,temperature, and RPM limits.

If the engine exceeds normal RPMlimits, the EEC provides an ultimateoverspeed protection that shuts off fuelwhen N1 exceeds 109% or N2 exceeds110%.

With the engine computer switch inOFF, the EEC is not powered. Controlof the engine is maintained by powerlever mechanical linkage to the FCU.With the EEC not powered:� the surge bleed valve assumes anintermediate position of 1/3 open� engine acceleration is slower� fuel consumption is approximately5% higher� idle thrust at low altitudes may behigher� normal limit RPM governing andITT limiting is not available.

Surge Bleed Valve

Under certain conditions, gas turbineengines have a tendency to surge andstall. For each compressor RPM, thereis a relationship between its amountof airflow and its pressure increase.When this balance is disturbed, enginesurges occur. A surge bleed valve pro-vides protection against this problem.

The EEC controls the position of thesurge bleed valve, which is betweenthe LP compressor and the HP com-

5L-11


pressor. If the valve is open, it releas-es compressed air into the bypass duct.By controlling the surge bleed valve,the EEC prevents compressor stallsand surges. The EEC normally posi-tions the surge bleed valve fully openfor start and idle conditions and fullyclosed for high RPM conditions. Fortransient RPM conditions, however,the EEC modulates the surge bleedvalve in response to impending stalls.With the fuel EEC OFF or failed, thesurge bleed valve stays 1/3 open.

Computer Failure

If a computer fails, the monitor sys-tem automatically switches to manualmode and disables the fuel meteringfunction of the computer. An amberENG C1 or ENG C2 warning light onthe failure warning panel indicates thatthe associated EEC is inoperative. Ifeither light illuminates and the engineinstruments are within limits, reducethe power lever to idle, then cycle theengine computer switch for that engineto reset the EEC.

If reset is unsuccessful, the switchshould be turned off. N1 and N2 ulti-

mate overspeed protection (109% to110%) is lost; the pilot must controlengine operation to ensure that limitsare not exceeded.

Both EECs must be ON for takeoff. Ifan EEC is inoperative, a ferry flightcan be authorized by appropriateauthorities.

When the aircraft is flown in accor-dance with the ferry flight procedure,a two-position switch on the face ofthe fuel computer is moved fromNORMAL to MANUAL (Figure 5L-12). The cockpit fuel computer switchis then placed in ON. Ultimate over-speed protection for N1 and N2 is pro-vided (N1 and N2 inputs must be func-tioning); however, all other functionsof the fuel computer are disabled.

Fuel Density Adjustment

Engine operating parameters for dif-ferent fuel types are compensatedthrough the fuel computer. If fuel spe-cific gravity adjustments are notadhered to, there is the possibility ofengine surging and higher turbine tem-peratures in start, acceleration, anddeceleration.

FUEL ADJUSTMENT MANUAL MODE SWITCH

AIR FILTER

IDLE

INCR

NORMAL

MANUAL

FR/MINPT2

5L-12


Powerplant

The fuel density knob on the front ofthe EEC is calibrated for AV GAS,JET B, and JET A fuel (Figure 5L-13). The fuel adjustment knob may berotated one “click” either side of theappropriate setting to compensate forindividual engine characteristics orinstallation effects. Consult the Main-tenance Manual for proper settings.

Fuel Flow DividerAssemblyThe fuel flow divider is between thefuel control unit and the fuel nozzles.Fuel flow to the primary and secondaryfuel nozzles is controlled during enginestart, operation and shutdown. Duringengine start, the divider routes fuel atreduced pressure to the primary noz-zles. As the start sequence continues,the fuel flow and pressure differenceacross the divider orifice increase; fuelis allowed to pass into the secondarylines that supply the fuel nozzles asRPM increases.

Fuel NozzlesEach engine uses 12 duplex (primaryand secondary) fuel atomizers that are

mounted to two manifold assemblies;each manifold contains six duplexatomizers (Figure 5L-16 , page 5L-22).Fuel swirls and breaks into micro-scopic droplets as it passes through theatomizer orifice into the combustor.The primary and secondary fuel atom-izers provide a finely atomized fuelspray pattern.

Fuel Flow IndicatorsEngine fuel flow is indicated by a gagefor each engine on the center of theinstrument panel.

A signal proportional to the fuel flowof each engine is derived from a fuelflow transmitter in the fuel supply lineof that engine. This signal is sent tothe associated engine’s fuel flow indi-cator, which is calibrated in poundsper hour times 100 (lbs/hr x 100)(Figure 5L-14 ).

The fuel flow indicating systemrequires power for operation fromAuxiliary C bus.

Start Pressure RegulatorTwo momentary-type start pressureregulator (SPR) switches are the over-

JET

JETPOSITION 8

POSITION 7

POSITION 6POSITION 5

POSITION 4

POSITION 3

POSITION 2

POSITION 9

POSITION 10

POSITION 11POSITION 1

AV

5L-13

5L-14

5L-22 For SimuFlite training only Falcon 10/100March 2001

head panel add extra fuel during start.During starts, the EEC normally addsextra fuel without the SPR switchbeing pressed up to an ITT of 200°C.During starts in cold ambient temper-atures (-15°C or less), the SPR switchmay be pressed and held up to an ITTof 400°C. Additional fuel is added aslong as the switch is pressed.

SPR is a function of the EEC; the SPRswitch provides a signal to the EECthat responds by opening the torquemotor metering valve more than itwould in a normal start. The EEC pro-vides fuel enrichment beyond 200°ITT as long as the SPR switch is heldto ON. To avoid a hot start, the switchshould not be held on longer than thatrequired for the engine temperature toreach 400°C. An overheat conditionis likely to occur if the switch is actu-ated while the engine is in operation.The SPR does not function if the EECis off or failed.

Engine FIRE PULLHandlesTwo engine FIRE PULL handles onthe upper center of the instrumentpanel are mechanically linked to thefuel shutoff valves. Pulling an engineFIRE PULL handle cuts off fuel to theassociated engine via a shutoff valve atthe engine pylon. There is enough fuelin the fuel line between the shutoffvalve and the engine-driven pump toallow the engine to continue to run atidle for approximately 20 seconds afteran engine FIRE PULL handle ispulled. Information on engine firedetection and extinguishing is provid-ed in the Fire Protection Chapter.

Power ControlSystemEngine power is regulated by the pilotwith power levers on the centerpedestal (Figure 5L-15 ). The power

FUEL MANIFOLDS

FUEL NOZZLEFUEL PUMP ANDFUEL CONTROL

LOCATIONON ENGINE

5L-16

5L-15

NOTE: Some early aircraft maystill have the early EEC installed.This early EEC does not have anautomatic SPR (manual only).


Powerplant

levers are mechanically connected tothe fuel control units. With the EECfunctioning, the pilot’s control of theengines is modified and there is lin-earity of thrust with power lever posi-tions. If an EEC is not functioning, thepilot directly controls that engine.

The power lever positions are listed inrelationship to the angle of rotation ofthe control shaft on the FCU. The fullaft (0°) position is the engine fuel cut-off position. Forward at 20° FCU isthe idle (or engine start) position. Tomove the power lever either from theidle to the cutoff position or from thecutoff to the idle position requires posi-tioning a release lever forward or aftof the power lever. Forward movement

from the idle position is unrestricteduntil reaching maximum limit of travelof the power lever.

During cold day conditions, requiredengine thrust is achieved by setting theN1 RPM specified in the flight manual.This thrust is based on internal pres-sures, temperatures, and N1 and N2RPM. Using N1 RPM higher thandesign specifications can result indamage.

Under hot day conditions, operationof the engine at higher than specifiedsettings can also cause excessive N1RPM This can reduce the life of orcause damage to the main shaft, plan-etary bearings, and combustor plenum.

NOTE: Power is normally setby moving the power levers for-ward to the maximum position;the EEC controls the engine andthe pilot monitors N1 and ITTfor normal operation.


Powerplant

PreflightDuring the external preflight, theengines are physically inspected toensure that the engine inlets andexhausts are clear of foreign objectsand the fan, turbine blades, and exhaustducts are not bent, nicked, or cracked.The engine fuel bypass indicatorsshould be checked; they should not beextended. Finally, nacelle conditionshould be checked; no fuel leaksshould be evident.

Engine oil level should be checkedwithin one hour after shutdown.Otherwise, start the engine and allowit to stabilize at idle before shuttingdown and checking oil level. Checkthat the oil bypass indicator is not pro-truding (right side of the engine)(Figure 5L-17 ).

The engine oil filler cap and accessdoors on each side should be checkedfor security (Figure 5L-18 ). Checkengine cowlings for security.

ServicingThe oil system capacity is 12 quarts(11.36 liters); the oil tank holds 6quarts (5.68 liters).

Maximum allowable engine oil con-sumption is 0.05 US gallons per hour.

This is the equivalent of 1 quart per 5hours of operation.

Approved engine oils conforming toGarrett EMS 53110, Type II, includethe following and those listed in theAFM:� Aeroshell/Royco Turbine Oil 500(Type II)� Castrol 5000 (Type II)� Exxon/Esso 2380 Turbo Oil(Type II)� Mobil Jet Oil II (Type II).

The listed brands of approved oil maybe mixed. Other types of oil are notapproved.

The following procedures should beused when adding oil.

1. Aircraft engine cowling must be opened. This is a maintenance procedure.

2. Remove oil tank filler plug bypushing down and rotating counter-clockwise.

3. Fill tank with oil until sight gage ordipstick indicates full.

4. Install filler plug by pushing downand turning clockwise.

5. Close cowling.

5L-17 5L-18


CAUTION: Synthetic lubricantsused in this engine contain addi-tives that are readily absorbedthrough the skin and are consid-ered highly toxic. Prolongedexposure to these lubricantsshould be avoided.

NOTE: Access to the oil tank isdifficult for No. 1 engine.Therefore, remove the cap at theend of the oil refill tube on theleft side of the engine. This capis fitted with a dipstick scaled inmissing quarts. Do not use thedipstick for checking; use thesight glass on the tank.

NOTE: To prevent falseindications, check the oil levelwithin one hour after engine shutdown.


AbnormalProceduresThe following is a brief discussion ofabnormal procedures for the enginesand their associated systems. For a listof specific procedural steps, refer tothe AFM or the SimuFlite OperatingHandbook.

Oil PressureThe illumination of the amber OIL 1(or 2) annunciator signals that eithermetal chips are detected in the oil scav-enge line or the oil pressure is less than25 PSI. If the indicated oil pressureexceeds 25 PSI, the most probablecause is metal chips. If possible, reducethrust, monitor engine operation andland as soon as practical.

If the indicated oil pressure is less than25 PSI, the crew action is to verify oilpressure in the associated engine andaccomplish a precautionary engineshutdown.

Engine ComputerLight ONAn engine computer light illuminateswhen one of the following indicationsoccurs.� The associated engine computerswitch is OFF.� Power to the associated computeris lost.� The associated computer fails inter-nally.

Cycle the computer switch to reset thecomputer. If the light remains on aftercycling the computer, the exhaust tem-perature of the operating engine is usedto determine proper setting for theengine with the inoperative computer.Avoid rapid power lever movements;maintain N1 RPM at 70% above20,000 ft. Thrust reversers are notauthorized for landing.

Engine OverspeedIf N1 or N2 exceeds normal operatingspeeds, retard the power lever on theaffected engine. Do not turn the affect-ed engine EEC off because the EECprovides additional engine overspeedprotection.

If the engine does not respond to powerlever movement, shut it down usingnormal procedures.

Engine AirstartsAll engine airstarts should incorporatethe following procedures.� Do not attempt an airstart withoutindication of fan rotation.� Do not attempt an airstart follow-ing an engine failure where the possi-bility of internal engine damage or fireexists.� If ITT is approaching limits and ris-ing rapidly, immediately abort the startby placing the power lever in CUT-OFF.� If a malfunction in the EEC is iso-lated, use the Computer Off Start pro-cedure and continue the flight with fuelcomputer off.

One of two airstart procedures isimplemented, depending on whetherthe EEC is on or off.

1. Maintain airspeed and altitude with-in airstart envelope (Figure 5L-17 ).

2. Abort airstart if any of the followingsituations occur.� Oil pressure does not rise within 10seconds of lightoff.� ITT does not rise within 10 secondsof fuel introduction.� ITT rise rapidly approaches 860°Climit.� N1 RPM remains near 0% when N2is 20%.� N2 RPM does not rise normallyafter lightoff.

WARNING: Engine fire mayresult if airstart is attemptedfollowing engine failure accom-panied by indications of internalengine damage.


Powerplant

� N1 RPM exceeds 80% with powerlever at IDLE.

With the fuel computer off, overspeedprotection and temperature protectionare not provided. Exercise care to pre-vent reaching engine temperature orspeed limits.

Refer to the AFM or the SimuFliteOperating Handbook for the restartenvelope and airstart procedures.

Engine FailureDuring TakeoffIf speed is below V1, abort the take-off by bringing power levers to IDLE,extending the airbrakes, applyingbrakes, and applying reverse thrust toslow aircraft. Accomplish EngineShutdown procedure when clear of therunway.

During braking, apply continuous max-imum pressure to the brake pedals toobtain best performance from the anti-skid system. Takeoff field lengthsassume application of maximum

braking at scheduled V1 speed duringaborted takeoff and assume no thrustreversers are available. Refer to theManeuvers chapter for specific proce-dure for engine failure after V1.

After stabilizing the aircraft, performEngine Shutdown procedure for failedengine.

Engine Failure –Shutdown in FlightSeveral conditions (e.g., abnormallyhigh or low oil pressure, rising or highITT, engine vibration, fan/turbine RPMfluctuations, high oil temperature, orerratic fuel flow) could necessitate anengine shutdown. Usually the EngineShutdown procedure is part of anoth-er abnormal or emergency procedure.

Shut down the affected engine by plac-ing its power lever in CUTOFF andadjust power on the operating engine.Call for the Engine Failure/ Shutdownin Flight checklist.

30,000

20,000

10,000

0

200 250 300 350

START ASSIST AIRSTARTN2 BELOW 15%

IAS (kt)

ALTITUDE(ft)

400

WINDMILLINGAIRSTARTN2 ABOVE 15%

M = 0.60

VMO

5L-19


EmergencyProceduresThe following is a brief discussion ofemergency procedures for the enginesand their associated systems. For a listof specific procedural steps, refer tothe AFM or the SimuFlite OperatingHandbook.

Engine FireA warning horn and an illuminatedFIRE PULL handle light indicate firein the associated engine. The horn issilenced by pressing the HORNSILENCE button on the pedestal.

When the problem is identified, silencethe warning horn. Retard the powerlevel to IDLE. The fire shutoff valve inthe pylon is closed when the FIREPULL handle is pulled. Reducing air-speed to 250 KIAS or below alsoreduces airflow to the nacelle.

Move the appropriate extinguisherswitch to position 1 to discharge thecontents of one fire bottle into theengine nacelle. If the condition per-sists, move the extinguisher switch toposition 2 to discharge the contentsof the second fire bottle into thenacelle.


Powerplant

LimitationsTemperature, RPM and Time Limits

Time Limit

Starting

N2%RPM

N1%RPM

ITT(°C)

860

Takeoff 100 100

860 5 minutes

Max. Continuous 30 minutes100 100 832

Transient 1 minute103 103

105 105 5 seconds

Ground Start/Starter AssistAirstart From10% N2 to Lightoff

10 seconds

Ground Start fromLightoff to Idle

50 seconds

25 secondsWindmilling Airstartfrom WindmillingN2 to 60% N2


Powerplant

DataSummary

Powerplant System

Power Source A bus – left engineN1, N2, oil pressure, oil temperature, boost pump, ignition, fire detection, fuel computer,fire XTING switch position 1

B bus – right engineN1, N2, oil pressure, oil temperature, boost pump, ignition, fire detection, fuel computer

Auxiliary busesFuel flow indicators and totalizer

Battery busIgnition backupFire XTING switch position 2

GeneratorsDirect backup power for engine computers (GEN 1, GEN 2)

Control Thrust lever

PRESS TO START buttons

SwitchesIGNITER ONSPRSTART/STOPAIR/GROUND/MOTOR-STARTFuel computer ENGINE 1/ENGINE 2

Monitor Engine operationFuel computer

IndicatorsN1

N2

Oil temperatureOil pressureFuel flow

AnnunciatorsOIL 1/2ENG. C1/C2FUEL P1/P2GENE. 1/2IGNITER ON (2)

Protection Computer ONFlat rating - 3,230 lbs thrustN1, N2, 100%EGT limiting 860°CUltimate overspeed protectionN1 – 109%; N2 – 110%

MechanicalOverspeed 105%

Falcon 10/100

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