1

This document describes the standard Aircraft,

including its power plant, systems and equipment.

Also included is the “Completion Description”,

describing outfitting standards and general

requirements to be used by Seller in the work to be

performed in the fabrication and installation of a

Challenger 300 model BD-100-1A10 baseline

interior, which includes interior, mechanical, electrical

and avionics equipment, and exterior paint

options and/or changes to this Completion

Description will be listed in an attachment to the

purchase agreement.

Also included in this document are descriptions of

Seller’s Customer Support services that are provided

to the Buyer as part of the sale of the Aircraft,

including warranty, technical publications, crew

training and the maintenance management system.

The Aircraft and the Completion Work may be

subject to changes during the course of the design,

manufacture and certification process or as the result

of any legislation, act, order, directive or regulation,

or, any interpretation thereof, of or by any

government or governmental body. If such changes

take place and apply to all aircraft in general or to all

aircraft of the same category as the Aircraft and are

effective after the date of the Agreement but before

Delivery Time, Buyer shall pay Seller’s reasonable

cost for such changes. If the incorporation of such

changes delays the delivery of the Aircraft, that delay

shall be an Excusable Delay under the Agreement.

While Seller employs personnel experienced in

aircraft completion work to perform the Completion

Work, unless otherwise specifically stated in any

particular section, this Completions Description is not

a guarantee of any particular result.

Capitalized terms not otherwise defined in this

Completion Description shall have the meaning

assigned in the Aircraft Purchase Agreement.

Furthermore, where the phrases “Space Provisions,”

“Structural Provisions,” “Power Provisions,” “Wiring

Provisions,” FAA,” FAR,” and “Complete Provisions”

are used in the Completion Description, their

meanings are as follows:

that space only is allocated for a defined unit or

installation. There are no brackets, bolt holes,

electrical wiring, or fluid lines.

means that the specific installation may be made in

primary structure. The brackets, bolt holes, electrical

wiring, and fluid lines are not supplied or included in

the weight of the Aircraft.

that a primary electrical power and distribution

system exist to permit later incorporation of the

specific equipment. Power provisions do not include

electrical wiring. The installation weight for the

equipment is not included in the weight of the

Aircraft.

that the primary wiring only is provided and installed

(capped and stowed) in the Aircraft. This does not

include the connector or circuit breaker installations,

unless otherwise specified. The installation weight for

the installed wire only is included in the Aircraft

weight.

equipment, means that all detail supports, brackets,

fluid lines, tubes and fittings, electrical wiring and

connectors, etc., have been provided so that the

specific equipment can be installed, without alteration

or additional parts. The system is functionally tested

prior to Delivery Time. Main components (black

boxes) are then removed and the system is disabled.

The installation weight for the equipment used for

testing is not included in the weight of the Aircraft.

2

The Aircraft is a pressurized, low-wing monoplane

with basic provisions for a typical configuration of 8

passengers and 2 crewmembers. Allowance for

luggage and optional equipment is provided.

Two (2) Honeywell HTF7000 turbofan engines are

pylon-mounted on the rear fuselage, and a

Honeywell AS GTCP-36-150BD APU is installed in

the Aircraft tailcone.

Span (overall) 63.8 ft 19.4 m

Area 522 sq ft 48.5 sq m

Sweep (@25% chord) 27°

Dihedral 2.2°

Aspect Ratio 7.3

Mean Aerodynamic Chord 9.4ft 2.9 m

Span (overall) 23.7 ft 7.2 m

Area 122.5sq ft 11.4 sq m

Sweep (@25% chord) 30°

Aspect Ratio 4.6

Span 9.5 ft 2.9 m

Area 83.2 sq ft 7.7 sq m

Sweep (@25% chord) 45.6°

Aspect Ratio 1.0

Tread 10.5 ft 3.2 m

Wheel Base 27.8 ft 8,5 m

Tire Pressure – Main (max, unloaded) 159 psi 1,096 kPa

– Nosewheel (max, unloaded) 114 psi 786 kPa

Tire Size – Main 26.5 in Dia and 8.0 in wide

– Nosewheel 18 in Dia and 5.5 in wide

Height (maximum over aisle) 6.1 ft 1.86 m

Length (excluding cockpit) 28.6 ft 8.72 m

Width (centerline) 7.2 ft 2.19 m

Passenger Door Width 2.50 ft 0.76 m

Baggage Door Width 2.00 ft 0.61 m

Baggage Compartment Volume 106 cu.ft 2.99 cu. m

Maximum Ramp Weight 39,000 lb 17,690 kg

Maximum Takeoff Weight 38,850 lb 17,622 kg

Maximum Landing Weight 33,750 lb 15,309 kg

Maximum Zero Fuel Weight 27,000 lb 12,247 kg

Standard Basic Operating Weight (± 2%)* 23,500 lb 10,659 kg

Maximum Fuel Weight 14,150 lb 6,418 kg

* Standard Basic Operating Weight includes unusable fuel, oil, standard interior, standard avionics, paint and 2 crew. The weight is exclusive

of any adverse weight impacts incorporated into the production configuration from Service Bulletins (SBs) and Airworthiness Directives (ADs).

3

APPROXIMATE

DIMENSIONS & WEIGHTS

4

APPROXIMATE DIMENSIONS & WEIGHTS

5

All performance is based on a standard aircraft

certified to U.S. 14CFR Part 25 requirements,

standard (ISA) day conditions. Options, Aircraft

customization and/or foreign certification

requirements requested by Buyer may result in a

change in performance.

Noise Levels (14CFR Part 36, Stage 3)

Takeoff (thrust cutback) Flyover 75.5 EPNdB

Sideline 87.6 EPNdB

Approach 89.6 EPNdB

Long Range Cruise Speed 458 KTAS 527 mph 848 km/hr

@ 37,000 ft 476 KTAS 548 mph 882 km/hr

@ 41,000 ft 476 KTAS 548 mph 882 km/hr

@ 45,000 ft 464 KTAS 534 mph 859 km/hr

Maximum Operating Speed M 0.83

Maximum Operating Altitude 45,000 ft 13,716 m

Initial Cruise Ceiling @ MTOW, ISA + 10 ºc 41,000 ft 12,497 m

The Aircraft is certified to:

issued in accordance with the Airworthiness Manual

(AWM) Chapter 525 at Change 7 dated September

30, 1996 and Federal Aviation Administration (FAA)

Part 25 Amendment 25-85 through 25-105, excluding

25-102 and 25-104.

accordance with the Federal Aviation Regulations

(FAR) Part 25 up to and including Amendment 25-

105, but excluding Amendments 25-102 and 25-104.

The Aircraft is certified for day and night operations,

under VFR and IFR conditions, flight into known ice

and is RVSM compliant.

Seller will provide Buyer a TC Certificate of

Airworthiness for Export, which will permit the Aircraft

to qualify for FAA Standard Airworthiness Certificate.

Seller shall not be obligated to obtain any other

certificates or approvals as part of this Agreement.

6

The Aircraft structure is primarily fabricated from

aluminum alloy but also includes alloy steels,

stainless steel, titanium and composites. Materials

used are in accordance with standard US aerospace

industry specifications for aircraft quality materials.

The Aircraft structure and systems are designed and

installed to facilitate inspection, maintenance and

permit ready removal of appropriate items. Parts and

assemblies subject to ready removal from the

Aircraft are interchangeable and/or replaceable

from one (1) Bombardier Challenger 300 aircraft to

another where Seller considers this practical. The

Aircraft, and its installed equipment, operate at

ground level ambient temperatures from -40ºC (-

40ºF) to +50ºC (+122ºF).

Vmo 320 KIAS 368mph 592 km/hr

Mmo M 0.83

Flaps 10 and 20 210 KIAS 242 mph 389 km/hr

Flaps 30 175 KIAS 201 mph 324km/hr

VLO (extend) 250 KIAS 288 mph 463km/hr

VLO (retract) 200 KIAS 230 mph 370 km/hr

VLE (maximum with gear extended) 250KIAS 288mph 463 km/hr

Forward Limit to at 23,100 (10,478 Kg) 31.0% MAC

Forward Limit at 27,000 (12,247 Kg) 26.8% MAC

Forward Limit at 33,750 (15,309 kg) 24.8% MAC

Forward Limit at 35,500 (16,104 kg) 24.4% MAC

Forward Limit at 38,650 (17,530 Kg) 25.0% MAC

Forward Limit at 38,850 (17,622 Kg) 25.2% MAC

Forward Limit at 39,000 (17,690 Kg) 25.3% MAC

Aft Limit to 23,100 (10,478 kg) 43.0% MAC

Aft Limit to 24,000 (10,886 kg) 43.0% MAC

Aft Limit at 27,000 (12,247Kg) 38.0% MAC

Aft Limit at 33,750 to 37,650 lbs (15,300 to 17,080 Kg) 33.0% MAC

Aft Limit at 38,650 lbs (17,530 Kg) 30.0% MAC

Aft Limit at 38,850 lbs (17,622 Kg) 28.6% MAC

Aft Limit at 39,000 lbs (17,690 Kg) 27.6% MAC

The fuselage consists of nose, center and aft

sections joined together, and incorporates

attachments for the wing, tailcone, engine support

structure, pylons and nose landing gear. Except for

the nose and aft sections, the fuselage cross-section

is a 92.1-inch (2.34 m) diameter circle.

The fuselage is of semi-monocoque construction with

aluminum alloy frames and stringers. Areas adjacent

to or affected by high heat sources are constructed of

fire-resistant or fireproof materials as appropriate.

A radome of composite material and designed for

use with a high resolution X-band radar is installed

7

on the fuselage nose. A system of conductive paths

along the radome is provided with the objective of

reducing the risk of damage from lightning strikes.

The fuselage is designed for internal pressurization

with the pressure-sealed area extending from the

bulkhead forward of the flight station in the nose to

the bulkhead aft of the cabin. The latter bulkhead

forms the aft face of the baggage compartment.

The wing is an all-metal swept back unit mounted

under the fuselage. It incorporates winglets,

ailerons, ground spoilers, multi-function spoilers,

single-slotted Fowler flaps, integral fuel tanks and

support structure for the main landing gear. Access

covers are provided in the lower wing skin panel to

permit access to the entire wing interior.

The wing is a one-piece unit consisting of spars

covered top and bottom with skin panels of aluminum

alloy. Shear web type ribs carry the air loads, act as

contour support and fuel tank baffles. The tanks are

compartmented and mechanically sealed using

sealing compounds. The tanks have an overcoat of

polyurethane.

The wing leading edges consist of aluminum alloy

skins, ribs and structure and incorporate hot air

thermal anti-icing.

The wing includes the following control surfaces:

from the wing root area to the inboard edge of

the aileron.

incorporates drain holes in the lower surface.

each wing immediately forward of the flap. The

inboard and outboard sections function

respectively as ground and multi-function

spoilers.

The empennage is of a T-Tail configuration,

comprising a variable incidence horizontal stabilizer

with elevators.

The horizontal stabilizer is a one-piece swept back

unit mounted at the top of the vertical stabilizer. It is

of all metal construction with spars and chord wise

ribs covered by stiffened skin panels. It incorporates

pivot and actuation mounting fittings to allow

incidence adjustments, hinges for elevators and

provision for sealing at its interface with the vertical

stabilizer. The fully cantilevered sweptback type

vertical stabilizer is attached to the aft fuselage

structure. The vertical stabilizer is constructed with

spars joined by chord wise ribs and with aluminum

alloy skin panels stiffened with span wise stringers.

It incorporates horizontal stabilizer pivots and

actuating mounting fittings, rudder hinges and a

fairing to provide a sealing interface with the

horizontal stabilizer. The vertical stabilizer leading

edge incorporates the HF antenna.

The Aircraft undercarriage is composed of one (1)

steerable nose landing gear (NLG) and two (2)

trailing-arm type main landing gear units (MLG)

equipped with hydraulically operated carbon brakes

and anti-skid systems.

The landing gear is of the tricycle type. The two (2)

MLG assemblies, one (1) located on the inboard

portion of each wing, retract inboard and up into the

MLG bay located in the central fuselage/wing area.

The NLG assembly is located beneath the cockpit

and retracts forward into the nose section of the

fuselage. A landing gear control handle, accessible

to both the pilot and co-pilot, controls normal

hydraulic extension and retraction of the landing

gear. A separate control handle, accessible to both

pilot and co-pilot, is located in the center pedestal

and allows manual lowering of the gear in case of

8

failure of the normal operating system. Gear

downlock is achieved when the gear is fully deployed.

Full extension and retraction of the gear in indicated

in the cockpit. The gear and its related linkages and

mechanisms are replaceable without removing

adjacent parts of the Aircraft structure. Shock strut

inflation valve and grease fittings are accessible

without jacking the Aircraft.

The Aircraft can make a 180º turn within a 58-ft. wide

taxiway without using differential braking. The Aircraft

is equipped with a Brake-By-Wire (BBW) system.

Each main wheel is equipped with carbon

composite multiple disk brakes.

Landing gear position/status is provided to the crew

via the EICAS and the landing gear panel. In

addition, an aural warning is provided whenever the

Aircraft is in a landing configuration and any gear is

not down and locked.

Two (2) Honeywell HTF7000 medium-bypass

turbofan engines are installed on the Aircraft. The

HTF7000 engine produces 6826 lb of thrust up to

86°F (30°C) (ISA+15°C) at sea level. The Aircraft is

equipped with a dual channel Full Authority Digital

Engine Control (FADEC) system. The FADEC

system provides engine, thrust reverser, starting and

ignition control. The FADEC system also provides

fault diagnostics and supports engine trend

monitoring. In addition, there is a thrust lever “end-ofstop”

position that is referred to as “Max. Thrust”

position. This position provides selected Auto Power

Reserve (APR) thrust. Detents are provided for

takeoff thrust and climb, and maximum continuous

thrust.

The thrust reverser system is hydraulically

operated.

The HTF7000 engines have been certified for oncondition

maintenance and cater to on-wing

maintenance.

A Honeywell 36-150BD gas turbine APU is installed

in the tailcone providing bleed air and electrical

power on the ground and in flight. The APU can

operate in-flight up to 37,000 ft, and up to 30,000 ft

with concurrent electrical load. It can also be

restarted up to 18,000 ft. The APU is automatically

controlled via an Electronic Control Unit (ECU).

The flight controls architecture consists of cables and

pulley assemblies for roll, pitch and yaw axes control.

The pitch and yaw axes are hydraulically powered

with manual reversion capability. The auxiliary

hydraulic system provides hydraulic power to one (1)

rudder PCU when required. The primary flight control

system consists of two (2) ailerons, two (2) elevators

and a rudder. The secondary flight control system

consists of multi-function and ground spoilers, flaps

and a variable incidence tailplane. Each elevator and

aileron surface is mass balanced.

Pitch control is accomplished through conventional

manual control (cables and pushrods), which transmit

pilot input to hydraulically powered actuators. In the

absence of hydraulic power, the surfaces can be

actuated via manual reversion. Yaw control is

accomplished through conventional manual

controls (cables & pushrods), which transmit pilot

input to hydraulically powered actuators. In the

absence of hydraulic power, the surfaces can be

actuated via manual reversion. Roll control is

accomplished through the use of one (1) aileron

and two (2) multi-function spoilers per side. The

pilot inputs are transmitted to the ailerons via fully

manual conventional cables, pulleys, and pushrods

arrangement, and electrically to the multi-function

spoilers.

In normal operation, both control wheels are

interconnected through the pitch and roll

disconnect mechanism and move in unison.

9

Pulling a roll disconnect lever located on the left hand

wheel can disconnect the control wheels. In this

mode, roll control is available using the MFS via the

LH wheel, or ailerons using the right hand wheel.

Pulling the pitch disconnect handle allows each pilot

to operate one (1) elevator panel independent of the

other.

The Aircraft is equipped with a stall protection system.

The stall protection system consists of two angle of

attack sensors, a dual channel computer, stick

shakers located on each control column, and a single

pusher.

Secondary flight controls comprise trailing edge flaps,

flight and multi-function/ground spoilers and pitch

trim. The Aircraft is equipped with a single trailing

edge Fowler flap panel per wing. The system is

electrically controlled by a Flap Control Unit and

powered via a central hydro-mechanical Flap Power

Drive Unit. The PDU is normally powered by one (1)

hydraulic system, with an alternate mode powered by

the other hydraulic system. Each flap panel is

actuated through a combination of flexible and hard

drive shafts connected to two irreversible ballscrew

actuators.

Eight spoiler panels are installed on the wings, four

(4) per side. The two (2) inboard panels function in

an on/off manner to provide ground lift dumping

only.

The two (2) outboard panels are multi-function

spoilers and function as roll control spoilers,

proportional flight spoilers, and ground lift dumpers.

They are electronically controlled and hydraulically

actuated by one (1) PCU per surface. The

proportional flight spoiler function is controlled by a

pedestal mounted lever accessible to both pilots. A

dual channel resolver connected to the roll control

system controls the spoileron function. The ground

lift dumping function is fully automatic.

All flight control positions are indicated on the

EICAS.

Estimated fuel quantity is 2,096 US gallons (7,941

liters) and is all contained in the wings.

Wing fuel is contained in a wet wing box structure

sealed to form two (2) separate wing tanks, split at

the wing centerline. The fuel is contained between

the front and rear spars in the left and right tanks.

Feed/collector tanks are integrally contained within

each tank. Two (2) baffles in each wing tank restrict

fuel sloshing thus limiting center of gravity shifts with

changes in Aircraft attitude. Swing check valves in

the main tanks allow fuel flow in the inboard direction.

Flush, self-closing dual seal water drain valves are

located at all low points of the system.

All pressure fueling and defueling operations are

controlled from the fuel/defuel control panel located

adjacent to the single point fueling adapter. Gravity

fueling via an overwing filler point is also provided

for each wing tank. Suction defueling is

accomplished at the single point adapter with

suction provided by the ground defuelling unit.

Cockpit control is provided on the fuel system panel

and fuel quantity and warnings are displayed on

EICAS.

The Aircraft has two (2) main independent hydraulic

systems and one (1) auxiliary hydraulic system.

Both main systems are powered by engine driven

pumps supplemented by DC electric motor driven

pumps. The auxiliary hydraulic system is powered

by an accumulator supplemented by a DC electric

motor driven pump.

Synthetic Phosphate-Ester base hydraulic fluid is

used. Each system operates at a nominal pressure

of 3,000 psi. The status of each hydraulic system is

indicated on the EICAS.

10

The primary electrical power for the Aircraft is 28

Vdc provided by three (3) identical 400 Amp Direct

Current (DC) brushless generators (two (2) enginemounted

and one (1) APU-mounted units). Ground

DC electrical power can be supplied by the APU

generator or via an external power cart.

Two (2) NiCad batteries are provided. The batteries

perform the following functions; provide power on

the ground (with engines and APU shutdown),

provide power for APU start (both on ground and in

flight), and provide emergency power in flight in case

of total loss of generators.

The DC system has power provision for the addition

of baseline interior completion 115V AC electrical

requirements. Failure indications, system

configuration and status information required to

troubleshoot and rectify failures are provided by the

EICAS.

The Environmental Control System (ECS) provides:

the cockpit, ventilation at each crew station and

passenger station

supply.

Engine bleed air from each engine supplies the air

conditioning system. The APU can also supply the

air conditioning system up to 20,000 ft. (6,096 m).

One (1) air conditioning pack conditions the air

supplied to the cockpit and cabin. In the event of

loss of this pack, an auxiliary system will pressurize

the cabin. If both these systems fail, a ram air

supply can be opened to ventilate the cockpit and

cabin. Ozone converters are installed in the bleed

air supply to the air conditioning pack.

The design of the ducting system assures even air

distribution and minimizes temperature stratification.

The cabin is supplied both direct & indirect

conditioned air as follows:

through the Passenger Service Unit (PSU). Air

gaspers located in the overhead PSU, four (4) left

hand and four (4) right hand over each passenger

seat, connect to the direct supply duct with flexible

duct hoses. One (1) gasper is installed in the

lavatory.

the left hand and right hand outboard sections of the

cabin headliner and the lower registers located

underneath the left hand and right hand lower dado

panels, supply conditioned indirect air. The upper

indirect registers are an integral part of the headliner

and run the length of the cabin section. The upper

conditioned indirect air is deflected from the PSU

into the cabin aisle way.

The oxygen system contains 77 cu. ft. (2.18 cu. m) of

usable oxygen and is sized for minimum requirement

for overland flights (North American). The flight

profile for rapid decompression condition and head

winds are considered in calculating supply.

Quick-donning demand type masks with mask

mounted regulators are provided for each pilot. A

crew oxygen low pressure warning annunciation is

incorporated on the EICAS.

The passenger oxygen system is gaseous type

supplied from the same cylinder(s) as the crew

system. For extended range and mission critical

flights, an additional 115 cu.ft oxygen bottle is

available as part of an overwater flight kit option. If

11

the oxygen system is activated, the cabin upwash

lights will come on bright to alert passengers.

In auto mode, the system automatically deploys the

passenger masks when the cabin altitude reaches

14,500 ft (4,420 m). The cockpit panel also allows

pilots to manually deploy the passenger masks or

shut off the oxygen supply to the passengers. Drop

out boxes are located over each cabin seat and in

the lavatory with extra boxes as required to comply

with the regulatory requirements. Each box is

equipped with constant flow masks, a lanyard valve,

a door opening valve and proper length oxygen

hoses.

externally primed. The tubing is double-flared and of

a size for the oxygen flow requirements. Flexible

hoses are used locally for main system to PSU

connections. A bleed orifice is installed in the system

to prevent mask deployment due to thermal pressure

build-up.

Electrical Pneumatic Actuating Valve (EPAV) is

installed to brighten the upwash lights, and provide

an aural warning in the cockpit when oxygen masks

are deployed.

will be located in the entryway headliner, over each

cabin seat, and in the lavatory with extras as required

to comply with the regulatory requirements. Each

container will be equipped with constant flow masks,

a lanyard valve, a door opening valve and proper

length oxygen hoses. Each container will have an

instruction placard on the inside for stowing the

oxygen masks.

and gauge is installed for servicing the oxygen

cylinder. A filler gauge is installed to indicate

pressure in the cylinder. The cylinder may be filled

without turning the pressure reducer shut-off valve.

outlet is installed in the cabin and is controlled by a

switch in the cockpit.

Both wing leading edges and engine inlets are antiiced

using engine bleed air.

The forward vision areas of the flight compartment

windshield are electrically heated for anti-icing

protection, defog, and demist. The side windows are

electrically heated for defogging and demisting.

Electrical heaters are used on the pitot probes and

angle-of-attack transducer vanes. An ice detection

system with two (2) detectors mounted on the lower

portion of the nose provides crew alert when icing

conditions exists. A passive hydrophonic coating

provides windshield rain repellant.

A fire detection and warning system with fault

determination is installed with fire detectors located

in each engine nacelle, the APU compartment and

the main wheel well areas. A fire warning system

test switch installed in the cockpit permits testing of

each detector and fire warning circuit. Fire warning

is provided via indication on the PBA, EICAS, and by

an aural warning.

An element in the wheel well warns of an overheat

condition in either of the main wheel wells.

Indication and test switches are provided in the

cockpit. The baggage compartment is equipped with

a smoke detector connected to the EICAS system.

A two-shot fire extinguishing system for the engine

and a single shot system for the APU are installed in

the fuselage aft section. The system permits

electrical checkout of the circuitry without

discharging the extinguishing agent.

One (1) Halon type hand-held fire extinguisher is

installed in the flight crew compartment. The cabin

also includes 2 hand-held fire extinguishers.

12

The Aircraft is equipped with a four-screen display,

all EFIS, Collins Pro Line 21 avionics suite. The four

12” x 10” Liquid Crystal Display (LCD) displays are

laid out in a four-across configuration. The EFIS

presents a pilot and co-pilot Primary Flight Display

(PFD) and a Multi-Function Display (MFD). The PFD

incorporates the following functions into a single

display: Attitude Director Indicator, Airspeed/Mach

Indicator, Altitude Indicator, Vertical Speed Indicator,

Horizontal Situation Indicator, full or arc compass,

and Flight Director and auto-pilot mode annunciator.

The MFD allows crew selectable functions.

The EFIS integrates the EICAS functions (engine

instruments, crew alerting and synoptic pages)

described earlier in this document. It provides

reversionary modes that ensure that all flight

essential information can be presented to the pilot or

co-pilot with any one (1) EFIS display failed.

The Maintenance Diagnostic System (MDS)

provides a centralized means of collecting and

interpreting failure data from participating Aircraft

systems, downloading recorded data, and initiating

and controlling operator initiated tests.

Table 1 lists all the components of the standard

avionics package.

VHF Communications Radio (8.33 KHz

channel spacing)

2

Mode S Transponder 2

Cockpit Audio Control Panel 2

Cockpit Audio Jack Panel 3

Cabin/Ground/Crew Audio Jack Panel 2

Cockpit Speakers 2

VOR/ILS/MKR Navigation Receivers 2

DME Transceiver 2

ADF Receiver 1

GPS Receiver 2

Radio Interface Unit 2

Integrated Avionics Processing Cabinet 1

Cockpit Display Unit 4

Clock 1

Integrated Standby Instrument 1

Display Control Panel 2

Reversion Control Panel 1

Data Concentrator Unit 1

Remote Data Concentrator 1

Cursor Control Panel 2

Emergency Locator Transmitter (121.5,

243.0 and 406.025 MHz)

1

Automatic Flight Control Function 2

Flight Guidance Panel 1

Servos 3

FMS Function 2

Maintenance Function 1

Radio Altimeter 1

Radio Altimeter Antenna 2

Cockpit Voice Recorder 1

Air Data Computers 2

Pitot Probe 1

Pitot Static Probes 2

Attitude Heading Reference Units 2

Flux Detectors 2

Standby Compass 1

Enhanced GPWS 1

Static Ports 2

TCAS II (Change 7.0) Receiver/Transmitter

System

1

14” WXR Transmitter/Receiver w/dual scan 1

Stall Computer 1

Stick Shaker 2

Stick Pusher 1

AOA Vanes 2

13

A flexible communication system is provided to

accommodate the short and long-term evolution of

aeronautical two-way communications. Refer to

Table 1, previous, for details.

An intercommunication system is also installed. The

system provides access for the pilot and co-pilot to

all the onboard radios. The pilot and co-pilot are

provided with a combination

microphone/headset. Intercommunication jack

boxes are installed in appropriate locations in the

aircraft (two are installed: one (1) in the nose wheel

voice communication, pilot and co-pilot audio. The

unit records continuously and retains the last 120

minutes of voice data.

The navigation system includes the following

components: