Pilot`s Manual - Code F (Metric Units)

Pilot`s Manual - Code F (Metric Units)
Pilot’s Manual
SECTION V
FLIGHT CONTROL SYSTEMS &
AVIONICS
TABLE OF CONTENTS
Flight Control Systems ............................................................................. 5-1
Aileron ..................................................................................................... 5-1
Roll Disconnect ................................................................................. 5-1
Control Wheel ................................................................................... 5-2
Control Wheel (Figure 5-1) .......................................................... 5-2
Elevator ................................................................................................... 5-2
Elevator Disconnect.......................................................................... 5-3
Elevator Disconnect (Figure 5-2)................................................. 5-3
Rudder..................................................................................................... 5-4
Rudder Pedal Adjustment Switches .............................................. 5-4
Rudder Boost .......................................................................................... 5-5
Controls Gust Lock ................................................................................ 5-6
Controls Gust Lock (Figure 5-3) ..................................................... 5-6
Flaps......................................................................................................... 5-7
Flap Control Lever............................................................................ 5-8
Flap Position Indication ................................................................... 5-8
Spoiler Systems ...................................................................................... 5-9
Normal Spoiler Mode..................................................................... 5-10
Autospoilers .................................................................................... 5-10
Spoileron Operation ....................................................................... 5-11
Spoileron (Roll Disconnect Mode of Operation) ........................ 5-12
Spoiler Indications .......................................................................... 5-13
Spoiler Monitor System ................................................................. 5-14
Pitch Trim .............................................................................................. 5-15
Pitch Trim Selector Switch ............................................................. 5-15
Primary Pitch Trim ......................................................................... 5-16
Trim Switch Panel (Figure 5-4) .................................................. 5-16
Bypass Trim.................................................................................. 5-17
Secondary Pitch Trim ..................................................................... 5-18
Autopilot Pitch Trim....................................................................... 5-19
Trim-In-Motion Indication............................................................. 5-19
Pitch Trim Bias................................................................................. 5-20
Configuration Trim......................................................................... 5-20
Aileron Trim.......................................................................................... 5-21
Rudder Trim ......................................................................................... 5-22
PM-126A
V-1
Pilot’s Manual
TABLE OF CONTENTS (Cont)
Trim Indications................................................................................... 5-23
Pitch Trim Indications.................................................................... 5-23
Aileron Trim Indications ............................................................... 5-24
Rudder Trim Indications ............................................................... 5-24
Mach Trim............................................................................................. 5-25
Stall Warning System .......................................................................... 5-26
Stall Warning Indications .............................................................. 5-26
Stall Warning Operation................................................................ 5-27
Stall Vane Anti-Ice .......................................................................... 5-27
Stall System Test ............................................................................. 5-28
Angle of Attack Indicators (Optional)......................................... 5-29
Instrument Panel Layout and AOA Indicator Position
(Figure 5-5)................................................................................... 5-29
Avionics .................................................................................................... 5-30
Honeywell Primus 1000 Avionics System ....................................... 5-30
Electronic Flight Instrument System (EFIS)................................ 5-30
IC-600 Power Source ...................................................................... 5-31
Avionics Master Switches................................................................... 5-32
Electrical Control Panel (Figure 5-6)............................................ 5-32
Air Data System (ADS)....................................................................... 5-33
Pitot-Static System.......................................................................... 5-33
Pitot-Static System Schematic (Figure 5-7).............................. 5-33
Standby Pitot-Static System Schematic (Figure 5-8) .............. 5-34
Air Data Computers (ADCs) ........................................................ 5-35
ADC Reversion ............................................................................... 5-35
Standby Instruments........................................................................... 5-36
Standby Instrument Group (Figure 5-9) ..................................... 5-36
Standby Altimeter .......................................................................... 5-36
Standby Airspeed/Mach Indicator.............................................. 5-36
Standby Attitude Indicator ........................................................... 5-37
Standby Compass........................................................................... 5-37
Attitude Heading Reference System (AHRS).................................. 5-38
Attitude and Heading Comparison Monitors............................ 5-39
AHRU Power Source & Cooling .................................................. 5-40
AHRS Reversion............................................................................. 5-40
Electronic Display System (EDS) ...................................................... 5-41
Primary Flight Display (PFD)....................................................... 5-42
Primary Flight Display and Bezel Controller
(Figure 5-10)................................................................................. 5-43
Bezel Controllers............................................................................. 5-43
Multi-Function Display (MFD) .................................................... 5-44
V-2
PM-126A
Pilot’s Manual
TABLE OF CONTENTS (Cont)
EICAS Display................................................................................. 5-45
Display Controllers......................................................................... 5-46
EICAS Display and BL-871 Bezel Controller
(Figure 5-11).............................................................................. 5-46
Display Controller (Figure 5-12) ............................................... 5-47
Display Unit Reversion Panels ..................................................... 5-48
Display Unit Reversion Panels (Figure 5-13) .......................... 5-48
Data Acquisition Units (DAUs) ......................................................... 5-49
Reversionary Control Panel (Figure 5-14) ................................... 5-49
Radio Management Units (RMUs) .................................................... 5-51
Radio Management Unit (Figure 5-15) ........................................ 5-51
RMU Cross-Side Operation........................................................... 5-52
RMU Backup Pages ........................................................................ 5-52
VHF Com Tuning............................................................................ 5-53
FMS Tuning...................................................................................... 5-54
NAV Tuning..................................................................................... 5-54
ADF Tuning ..................................................................................... 5-55
Transponder/TCAS Tuning .......................................................... 5-56
TCAS (Optional) .................................................................................. 5-56
TCAS Operation.............................................................................. 5-57
System Controls and Displays ...................................................... 5-58
Traffic Display Symbols ................................................................. 5-58
Enhanced Ground Proximity Warning System (EGPWS).............. 5-59
Audio Control System......................................................................... 5-60
Digital Audio Control Panel (Figure 5-16) .................................. 5-61
Clearance Delivery Radio (CDR)....................................................... 5-63
Clearance Delivery Radio (Figure 5-17) ...................................... 5-64
Flight Guidance Control System ....................................................... 5-65
Flight Director ................................................................................. 5-65
Flight Guidance Controller (FGC)................................................ 5-66
Flight Guidance Controller (Figure 5-18)................................. 5-66
Autopilot/Yaw Damper ................................................................ 5-69
AP/YD Annunciation .................................................................... 5-70
Control Wheel Trim Switch ........................................................... 5-70
Control Wheel Master Switch (MSW).......................................... 5-70
Touch Control Steering (TCS) ....................................................... 5-71
Autopilot Engagement/Disengagement..................................... 5-71
Yaw Damper Engagement/Disengagement............................... 5-72
Mistrim Annunciation.................................................................... 5-72
Power Supply Configuration ........................................................ 5-72
Go-Around (GA) Button ..................................................................... 5-73
PM-126A
V-3
Pilot’s Manual
TABLE OF CONTENTS (Cont)
Flight Management System (FMS).................................................... 5-73
Control Display Unit (CDU) ......................................................... 5-74
UNS-1E FMS CDU (Figure 5-19) .............................................. 5-74
Data Transfer Unit (DTU).............................................................. 5-75
Configuration Module................................................................... 5-75
FMS Functions ................................................................................ 5-75
Weather Radar ..................................................................................... 5-77
WU-650 Weather Radar Control Panel (Figure 5-20) ................ 5-77
Avionics Cooling ................................................................................. 5-78
Instrument Panel Cooling ............................................................. 5-78
Miscellaneous .......................................................................................... 5-79
Cockpit Voice Recorder (CVR) .......................................................... 5-79
Clocks .................................................................................................... 5-80
Multi-Function Chronometer and Instrument Location
(Figure 5-21)................................................................................. 5-80
Hourmeter-Aircraft (Optional).......................................................... 5-81
Flight Data Recorder (FDR) (Optional) ............................................ 5-81
Emergency Locator Transmitter (Optional)..................................... 5-81
Dorne & Margolin ELT14 ................................................................... 5-81
Transmitter and Antenna .............................................................. 5-81
Transmitter Switch (ARM/OFF/ON) ......................................... 5-81
Remote Control and Monitor Unit............................................... 5-82
Operation......................................................................................... 5-82
ARTEX ELT 110-406............................................................................. 5-83
Transmitter and Antenna .............................................................. 5-83
Transmitter Switch (ON-OFF) ...................................................... 5-83
Cockpit Switch and Indicator Light............................................. 5-84
Buzzer .............................................................................................. 5-84
Operation......................................................................................... 5-84
V-4
PM-126A
Pilot’s Manual
SECTION V
FLIGHT CONTROL SYSTEMS &
AVIONICS
FLIGHT CONTROL SYSTEMS
The primary flight controls (ailerons, elevator, and rudder) are mechanically operated through the control columns, control wheels, and rudder pedals. The flaps and spoilers are hydraulically actuated and
electronically controlled. Airplane trim systems (pitch, roll, and yaw)
are electronically controlled.
AILERON
The aileron control system consists mainly of three control circuits, one
in the fuselage area and one in each of the left and right wing area. In
addition, a disconnect mechanism is incorporated into the pilot’s control wheel which allows the disconnection of the aileron control system
(in the event of a jam) and switching to spoileron system for roll control.
The fuselage control circuit connects both pilot’s and copilot’s control
wheels together, and each wing control circuit is connected to the aileron drive mechanism. The three control circuits are connected together
via a common sector assembly. In normal operation, whether by an input from the autopilot or by manual input to one of the two control
wheels, the two control circuits will move in unison to drive the two aileron panels. The aileron control system is considered the primary system for roll control and is interfaced with the spoileron system for roll
augmentation.
ROLL DISCONNECT
If ailerons become jammed, the aileron control system can be disconnected and the spoileron system can be used for roll control. The pilot’s
control wheel is disconnected from the aileron cables and copilot’s control wheel by the red lever labeled ROLL DISC located on the hub of the
pilot’s control wheel. This will also disconnect and prevent engagement of the autopilot. Safe flight can continue on spoilerons alone. For
more information on roll disconnect, see Spoileron (ROLL DISCONNECT) system.
PM-126A
5-1
Pilot’s Manual
CONTROL WHEEL
Each flight station is equipped with a U-shaped control wheel. The pilot’s control wheel is equipped with a disconnect assembly which employs a red lever labeled ROLL DISC located on the inboard side of the
control wheel hub (Figure 5-1). Each control wheel contains the following switches: Control Wheel Trim, Control Wheel Master (MSW), MIC,
IDENT, Touch Control Steering (TCS), and Checklist Line Advance.
TOUCH
CONTROL
STEERING
(TCS)
ROLL DISC LEVER
(PILOT'S SIDE ONLY)
ARM
BUTTON
CONTROL WHEEL
TRIM SWITCH
CHECKLIST
LINE ADVANCE
IDENT
SWITCH
MIC
SWITCH
(NOT SHOWN)
CONTROL WHEEL
MASTER SWITCH
(MSW)
CONTROL WHEEL
Figure 5-1
ELEVATOR
Movement of the control columns is mechanically translated into elevator surface movement through levers, bellcranks, sectors, cables, and
pushrods. The elevator control system consists of two parallel control
circuits. The two control circuits are normally connected together via
forward and aft disconnect assemblies in which either control column
moves both the left and right elevator surfaces in union. A mechanical
up/down spring is also used in the system to augment high and low
speed trim ability of the airplane. If a jam occurs in either mechanical
control circuit, an elevator disconnect feature is incorporated into the
system.
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PM-126A
Pilot’s Manual
ELEVATOR DISCONNECT
In the event of an elevator jam, the two cable circuits can be disconnected by pulling the red ELEV DISC T-handle (Figure 5-2) located at the
left-forward edge of the pedestal. The airplane will then be controlled
with the unjammed elevator. The forward and aft disconnect assemblies are dog clutch devices that are operated simultaneous by the handle being pulled. When the ELEV DISC T-handle is pulled, a cable
connected to the handle shaft is pulled which disengages the forward
clutch thereby disengaging the two control columns. Electrical switches sense the movement of the shaft and signal the aft disconnect to disengage.
E
V
E D
PIT TRIM
BIAS
SYS TEST// RESET
G F
E R
A E
R E
F
A
L I
E S
ANTI
-ICE
V C
U
SPLRN
ST
RE
P
P
C
OFF
TE
E
L
D
I
S
FIRE
DET
SS
STALL
LTS
ADC
RUDDER
BOOST
SQ
NAV
AUDIO
L
L
MODES
GEAR
FLAP
OFF
VOL
L
L
P D
U O
S W
H N
USEABLE FUEL TOTAL: 6000 LBS
RADIO CTL
HOT BUS
2404
FUSELAGE
ON
E
L
E
V
S
P
O
E D
R
E
T
A
R
M
I
L
L I
E S
R
FLAPS
E
X
T
GO AROUND
MUTE
U
P
V C
S
R WING
E
D
I
C
1798
1798
L WING
FE
IDLE
IDLE
8
CUTOFF
EMER
BRAKE
PULL
P
CUTOFF
PARK
BRAKE
PULL &
250
2
0
200
MFD JOYSTICK
D
N
150
EMERGENCY/PARKING
BRAKE
U
L
L
ELEVATOR DISCONNECT
Figure 5-2
When the handle is pulled to full extension, it must be rotated 90°,
either clockwise or counterclockwise, to lock it in the disconnect position. The elevator forward disconnect is a mechanical clutch mechanism located on the torque tube between the control columns. The
clutch is held open when the ELEV DISC T-handle is pulled and locked
in the extended position. This will disconnect and prevent engagement
of the autopilot.
PM-126A
5-3
Pilot’s Manual
The elevator aft disconnect is an electro-mechanical device located in
the top of the vertical stabilizer. When the ELEV DISC T-handle is
pulled, a two-position linear actuator on the elevator aft disconnect assembly is energized to the extended position (disconnected position),
separating operation of the two elevators. The linear actuator will remain in the extended position. When the elevator aft disconnect is actuated, the elevator disconnect sensor will send a signal to display a
message on the Crew Alerting System (CAS). Do not reconnect. Obtain
maintenance prior to next flight. Electrical power used by the elevator
disconnect system is through the ELEV DISC circuit breaker located on
the pilot’s circuit breaker panel (FLIGHT group).
The following CAS illuminations are specific to the elevator disconnect:
CAS
ELEVATOR DISC
ELEVATOR DISC
Color
Description
Amber Elevator disconnect has split the elevator
controls on the ground. Obtain maintenance
prior to flight.
White Elevator disconnect has split the elevator
controls during flight. Do not reconnect.
RUDDER
Directional control is provided by a dual closed-loop cable system with
separate parallel paths in the engine area for rotor burst considerations.
Rudder pedal movement is mechanically translated into rudder control
surface movement through cables, pulleys, and bellcranks. There is an
electrically driven rudder boost system to provide additional rudder
control power in the event of an engine-out on takeoff.
RUDDER PEDAL ADJUSTMENT SWITCHES
The pilot’s and copilot’s rudder pedals are individually adjustable with
a spring-loaded toggle switch to accommodate differences in crew size.
The pilot operated toggle switch controls a linear actuator which provides forward and aft pedal adjustment. This toggle switch is labeled
RUDDER PEDAL, and located on the lower outboard corner of the pilot’s and copilot’s switch panel. Each switch has three positions: FWD,
Off, and AFT. Only the FWD and AFT positions are labeled. The rudder
pedal adjustment is powered by 28-vdc supplied through two 3-amp
circuit breakers, L RUD ADJUST and R RUD ADJUST, on the pilot’s
and copilot’s circuit breaker panels (FLIGHTgroup).
5-4
PM-126A
Pilot’s Manual
RUDDER BOOST
The rudder boost system is provided to reduce rudder forces. Signals
from force sensors in both sets of rudder pedal mechanisms are read by
the ICs. The ICs then send rudder boost signals to the yaw damper servo. Rudder boost provides yaw servo torque proportional to rudder
pedal force, when either the pilot’s or copilot’s rudder pedal force or
the sum of their forces reaches 50 pounds. The rudder boost will override the yaw damper (if engaged) when this threshold is reached. When
the force on the rudder pedals is released, the yaw damper will resume
operation. The rudder boost system is armed when flap extension is
greater than 3° and the RUD BOOST switch is selected to On. The RUD
BOOST switch is located on the forward pedestal. When selected On,
the switch is dark and when selected OFF, OFF will be displayed in the
center of the switch. A white CAS illuminates when the switch is OFF
or the system is disabled by the IC or the yaw force interface box. An
amber CAS illuminates when the system is inoperative and not selected OFF.
Dual, redundant power inputs are provided via the RUD FORCE circuit breaker on the right essential bus and the NOSE STEER COMPUTER circuit breaker on the left essential bus. If both these power sources
should fail, the #2 IC-600 will disable rudder boost and provide appropriate annunciation.
The following CAS illuminations are specific to the rudder boost
system:
CAS
RUD BOOST INOP
RUD BOOST INOP
PM-126A
Color
Description
Amber Rudder boost is inoperative and not selected
OFF. Do not takeoff.
White Rudder boost is selected OFF. Do not takeoff.
5-5
Pilot’s Manual
CONTROLS GUST LOCK
A gust lock is provided to help prevent wind gust damage to moveable
control surfaces. The gust lock is installed on the pilot’s side only, with
control wheel rotated counterclockwise until the bend in the handle
aligns with the column, and the rudder pedals are centered. Loop
straps around bottom heel, and draw both left and right pedal straps
taut to seat control column against the primary stop. When installed,
the gust lock secures the flight controls in the rudder centered, full
aileron, and full down elevator position.
CONTROLS GUST LOCK
Figure 5-3
5-6
PM-126A
Pilot’s Manual
FLAPS
The airplane’s single-slotted Fowler flaps are electronically controlled
and operated by a hydraulic motor (flap power unit). Each flap panel,
one on each wing, has three safe-life flap tracks and is driven by two
screw jack actuators. A flexible drive shaft transfers power from the hydraulic motor to each flap actuator. The flap control lever is located on
the center pedestal and is recessed to prevent inadvertent operation.
The flap control lever has settings at 0° (up), 8°, 20° and 40° (down). To
select a new flap position, the flap control lever is moved directly to the
desired setting. Flap position is controlled by a microprocessor based
controller (Flap Control Unit). The Flap Control Unit receives position
command information and an arming signal from the flap control lever
in the cockpit. It then provides the electrical arming and control signal
to the arming solenoid valve located in the flap power unit and receives
a feed-back signal from sensors mounted on the outboard actuator of
each flap panel. When flaps are extended or retracted in flight, the configuration trim system automatically applies the appropriate amount
of pitch trim to compensate for the pitching moment caused by flap repositioning.
The Flap Power Unit (FPU) is located under the center wing and contains the hydraulic motor, a servo control valve, an arming solenoid
valve and a pressure switch. The servo valve responds to electrical signals from the flap control unit and meters hydraulic pressure to the extend or retract side of the bidirectional hydraulic motor. The arming
solenoid valve must be energized open by the flap control unit before
hydraulic pressure is available to the servo valve. The pressure switch,
located upstream of the servo valve, monitors system pressure between
the arming solenoid valve and the servo valve. If pressure is not available on flap selection, flaps will be inoperative, but if pressure is available without a selection command, the FPU will show a fault and the
flaps will operate in a degraded mode, i. e. the flaps may deploy at a
reduced speed when selected. These will cause the following CAS illuminations in order described above.
The following CAS illuminations are specific to the Flap System:
CAS
FLAPS FAIL
FLAPS FAULT
PM-126A
Color
Description
Amber The flap system has failed and the flaps are
inoperative.
Amber The flap system is operating in a degraded
mode.
5-7
Pilot’s Manual
Flexible drive shafts routed along the rear wing spar transmit the rotary
motion of the flap power unit to the input shaft of each of the flap actuators. Two Rotary Variable Differential Transformers (RVDTs) mounted
on the outboard side of each outboard flap actuator provide position information to the flap control unit and the flap position indicating unit.
The flap actuators incorporate a screw jack and are attached to the rear
spar. These actuators convert the rotary input motion into linear output
motion through these screw jacks thus driving the flaps. Each actuator
has overtravel end stops. Uncommanded retraction due to airloads, vibration, etc., is prevented by the screw jack design of the flap actuators.
The flap control system operates on 28-vdc supplied through a 3- amp
FLAP CTRL breaker located on the copilot’s circuit breaker panel
(FLIGHT group).
FLAP CONTROL LEVER
The FLAP control lever will operate in one of four positions (UP, 8°, 20°,
and DN) with detents at the 8° and 20° positions. When retracting flaps,
there is a gate at the 8° position; therefore, the lever must be pulled out
slightly when raising the flaps above 8°. The lever is attached to dual
RVDTs co-located with a flap lever detent switch within the throttle
quadrant. These dual RVDTs transmit the selected position to a flap
control unit. Moving the lever between positions actuates the flap lever
detent switch and energizes a 75-second timing circuit within the flap
control unit. This circuit allows the arming solenoid valve within the
flap power unit to energize open for 75 seconds and then de-energizes.
Normal flap extension from 0° to 40° will not exceed 10 seconds with
engine-driven hydraulic pumps operating. However, this time will extend up to 60 seconds when using HYD XFLOW while lowering flaps
from 0° to 20° in flight.
FLAP POSITION INDICATION
The flap position indicating unit has two separate and independent
channels. Channel 1 provides left side equipment and Channel 2 provides right side equipment. Both channels are housed in a common
chassis. Flap position is shown full time in a digital display on the
Engine Indicating and Crew Alerting System (EICAS). The EICAS display is framed with a white box when the flaps are not in the selected
position in flight, or on the ground with flaps not set for takeoff. The
EICAS display turns red if power is advanced for takeoff and flaps are
not properly set. The display turns amber if there is a fault or failure in
the flap system. Flap selection and position are also displayed on the
right side of the FLT (flight) system schematic page. The FLT system
schematic display can be displayed on the EICAS or Multi-Function
Display (MFD).
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PM-126A
Pilot’s Manual
Selected flap position is indicated by a horizontal magenta line across
the vertical scale. Actual left and right flap position is indicated by flap
position pointers on each side of the vertical scale. When flaps have
moved to their selected position, the pointers will overlay the magenta
line. Flap position pointers turn red on the ground when power is advanced for takeoff and flaps are not properly set. Pointers turn amber
when there is a fault or failure in the flap system. A digital indication of
flap position is provided on the backup engine/systems page of the
Radio Management Unit (RMU).
SPOILER SYSTEMS
Spoilers, one on the upper surface of each wing forward of the flaps, are
provided for deceleration. The spoilers are electrically controlled and
hydraulically operated. The spoilers are extended symmetrically for
use as spoilers/speed brakes or asymmetrically for aileron augmentation. Each spoiler is hinged at four points and is extended or retracted
with a single hydraulic actuator. The spoiler control lever, located on
the left side of the throttle quadrant, is linked to two RVDTs. There are
three labeled settings for the spoiler lever that correspond to detent positions: RET (retract), ARM (autospoilers), and EXT (full extension) approximately 60° at slower airspeeds. The range between the ARM
and EXT detents allow for variable spoiler positions in flight. There are
also two unmarked detent positions between ARM and EXT which correspond to intermediate spoiler extension positions of approximately
15° and 30°. At high airspeeds the actuators cannot extend the spoilers
fully; therefore, spoileron computer commands to the actuator servos
are limited by airspeed inputs from the Air Data Computers (ADCs).
At speeds below 175 knots, spoilers will extend to 60° when the spoiler
control lever is placed to EXT; however, at higher speeds full extension
is not possible.
PM-126A
5-9
Pilot’s Manual
NORMAL SPOILER MODE
The spoilers can be extended symmetrically on the ground or in flight
by moving the spoiler lever aft of the ARM position. Placing the lever
to any position aft of ARM while on the ground will cause full extension (60°) of the spoilers. Spoiler extension on the ground requires approximately 1 second and in flight, approximately 5 to 7 seconds. When
the spoiler control lever is placed aft of the ARM position, the RVDTs
will signal the spoileron computer. The computer, in turn, energizes
torque motors on the servo valves to meter hydraulic pressure to the
extend side of the actuators. The computer receives spoiler extension
feedback from the RVDTs attached to the spoiler surfaces, and neutralizes the servo valves when the spoilers reach their selected position. In
flight, the amount of spoiler extension will depend on spoiler control
lever position and airspeed.
AUTOSPOILERS
Autospoiler mode is used to automatically extend spoilers on landing
or in an aborted takeoff. When the SPOILER lever is set to ARM, the
system will arm and CAS will illuminate. This will automatically extend spoilers when the main gear weight-on-wheels switch circuits indicate an “on ground” condition, thrust levers are in the IDLE position
and the airplane has attained 60 knots ground speed. This mode fully
extends spoilers at maximum rate (one second or less) when spoiler
control lever is in the ARM position and autospoiler deploy logic is
met. An autospoiler system is installed to automatically extend both
spoilers in order to spoil lift after landing or during an aborted takeoff.
The following CAS illumination is specific to the autospoiler system:
CAS
AUTOSPLR ARMED
Color
Description
White Autospoilers have been armed.
The main gear weight-on-wheels switch circuits are electronically
latched in the “on ground” state once the initial weight-on-wheels signal is received. This prevents inadvertent spoiler retraction in the event
the airplane should bounce during the ground roll. If either thrust lever
is moved above IDLE while autospoilers are extended, the spoilers will
immediately retract. Flap position has no effect on autospoiler operation and autospoilers are not operational when EXT or RET is selected.
The spoileron computer receives power from the L ESS BUS for operation and the spoiler indicating system receives power from the R ESS
BUS. The circuits are protected by “SPLR CTRL” circuit breaker on the
pilot’s circuit breaker panel (FLIGHT group) and the “SPLR IND” circuit breaker on the copilot’s circuit breaker panel (FLIGHT group).
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PM-126A
Pilot’s Manual
When spoilers are extended or retracted in flight, depending on the
mach number, the configuration trim system automatically applies the
appropriate amount of pitch trim to compensate for the pitching
moment caused by spoiler repositioning.
SPOILERON OPERATION
Spoilerons operate automatically on the ground and in flight to
augment the ailerons whenever either control wheel is turned more
than 5°. Rotation of either control wheel provides a roll input to the
spoileron computer via dual RVDTs inside of the pilot’s control wheel.
The appropriate spoiler, left or right, extends to the commanded angle
for the current conditions (Mach number, airspeed, AP engage and flap
setting) while the other spoiler is commanded stowed. When in the
mixed spoiler and spoileron mode, the spoiler command derived from
the spoiler lever and spoileron command derived from the control
wheel are added to form a composite position command for each
spoiler panel. The spoiler command provides a bias position command
common to both panels while the control wheel RVDTs generate a
differential command. The control wheel inputs command the angular
displacement that exists between the two spoilers regardless of the
amount of spoilers command. Spoileron commands have priority over
spoiler commands.
If the spoilers are extended, and the control wheel is turned right, the
computer mix logic retracts the left spoiler first to give differential necessary for the roll commanded. If that is not enough differential for the
roll commanded, the computer then extends the right spoiler as
required.
PM-126A
5-11
Pilot’s Manual
SPOILERON (ROLL DISCONNECT MODE OF OPERATION)
Spoilerons provide automatic roll augmentation and backup roll control. The spoilerons are electrically controlled and hydraulically actuated. Artificial friction is introduced into the pilot control wheel upon
disconnection from the mechanical aileron system to provide pilot feel
and to preclude the control wheel from free-wheeling. If ailerons become jammed, the pilot’s control wheel can be disconnected from the
aileron control cables and the copilot’s control wheel. Roll disconnect is
activated with a red lever labeled (ROLL DISC) located on the hub of
the pilot’s control wheel. In addition to mechanically disconnecting the
pilot’s control wheel from the ailerons, activation of the ROLL DISC lever trips two disconnect switches within the control wheel hub. When
the roll disconnect mode is activated within the spoileron computer, it
outputs a signal for a CAS message to illuminate. When roll disconnect
mode is activated, the autopilot will disengage.
The following CAS illuminations are specific to the spoileron
computer:
CAS
ROLL DISC
ROLL DISC
Color
Description
Amber Roll disconnect has occurred on the ground.
White Roll disconnect has occurred in flight.
The roll disconnect mode provides roll control through RVDT signals
from the pilot’s control wheel to the spoileron computer. This mode is
much the same as the normal spoileron mode but has a different gain
curve relating to control wheel input and panel deflection begins at 1°
movement instead of 5°. Spoileron operation is full time. Anytime either control wheel is turned more than 5°, there is a differential displacement of the spoiler surfaces to augment roll control. Spoilers can
be operated in conjunction with the roll disconnect mode the same as
they are with normal spoileron mode. The roll disconnect mode may be
deselected in flight by returning the ROLL DISC lever to its normal position.
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Pilot’s Manual
SPOILER INDICATIONS
Spoiler extension is indicated at the base of either the SUMRY or FLT
system schematic page on either EICAS or MFD. On the FLT system
schematic display, spoiler extension is presented as a digital display
and as a vertical analog scale with dual points (one for each spoiler).
The digital displays on the SUMRY and FLT pages only show spoiler
extension commanded by autospoilers or with the spoiler lever. They
do not reflect differential extension resulting from operation in
spoileron mode. When the airplane is on the ground with spoilers
extended, a white box will overlay the digital spoiler display. If power
is advanced for takeoff with either or both spoilers extended, this
digital display and box will turn red along with the pointers on the analog scale. In addition, a red CAS and “CONFIGURATION” voice message will activate. Spoilers should not be extended at the same time
flaps are extended while in flight except as specified in the Airplane
Flight Manual, or the following CAS message will be posted.
The following CAS illuminations are specific to the spoilers:
CAS
SPOILERS EXT
SPOILERS EXT
SPOILERS EXT
Color
Description
Red The spoilers have moved from the stowed
position, with aircraft on the ground, and
either thrust lever is advanced to MCR or
above.
Amber The airplane is in flight and spoilers are
extended with flaps extended more than 3°.
White Spoilers are not fully retracted. Spoileron
extension will not activate this CAS (flight and
ground).
The analog scale and pointers are real time showing actual spoiler
position for all conditions. When spoilers are extended as result of
spoileron operation, pointers will indicate their differential on the
analog scale. Digital spoiler indicators and analog scale pointers will
turn amber when flaps are extended 3° or more with spoilers extended.
PM-126A
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Pilot’s Manual
SPOILER MONITOR SYSTEM
The spoileron computer contains a monitor system to prevent electrical
or mechanical faults from causing uncommanded extension or retraction of the spoilers. The spoileron computer uses electrical power from
the L ESS BUS for operation and the spoiler indicating system uses
power from the R ESS BUS. The circuits are protected by the SPLR
CTRL and SPLR IND circuit breaker, respectively, located in the
FLIGHT group on the pilot’s and copilot’s circuit breaker panels. If
power to the spoileron computer is lost through the SPLR CTRL or
SPLR IND circuit breaker, the spoilers will retract and be inoperative in
all modes. The spoileron computer performs a self-test (BIT) at powerup. A test failure will trip the spoileron monitor. If the monitor detects
a self-test failure or a fault during normal operation, hydraulic pressure
is removed from the system by closing the spoiler shutoff valve. A hydraulic return is provided to blow the spoilers closed. During normal
operation, the shutoff valve is held open by an electrical solenoid. A
power failure will cause this valve to close. If the monitor does not stow
the spoilers, the crew will initiate the stow with either Control Wheel
Master Switch (MSW). When either MSW is held depressed, the spoiler
shutoff valve is depowered closed and the spoilers will blow down,
however, they may not fully retract.
A system malfunction will cause the spoileron monitor to trip and an
amber CAS display. If the malfunction clears, the system may be reset
using the “SPLRN RESET” position on the system test knob. If the monitor detects a jammed spoiler, the spoileron computer continues to operate using the spoiler that is not jammed and it applies a full retract
input to the effected actuator for 5 to 7 seconds. This will also illuminate
on the CAS.
The following CAS illuminations are specific to the spoileron monitor:
CAS
SPOILERS FAIL
SPOILER JAM
5-14
Color
Description
Amber A failure in the spoiler system is detected.
Amber The associated (L or R) spoiler is jammed.
PM-126A
Pilot’s Manual
PITCH TRIM
Pitch trim is provided by a moveable horizontal stabilizer. Operational
structural redundancy has been incorporated by using primary and
secondary sections that are independent. Primary and secondary each
have electrically and mechanically independent motors (separated for
rotor burst considerations), gear trains, and control inputs. Position
sensors in each section of the actuator, geared directly off of the main
drive screw, are monitored by both IC-600s. The computers compare
the primary position sensors to the secondary position sensors in the
actuator to annunciate to the pilot when the display position may not
be accurate. The secondary section structure, construction, and operation is the same as the primary and both sections drive a common
screwjack-type actuator to move the leading edge of the horizontal stabilizer up or down. The primary motor is actuated by manual primary
pitch trim (control wheel trim switch), configuration trim, and Mach
trim systems. The secondary motor is provided as a backup for primary
trim and is operated by the secondary pitch trim and the autopilot.
The following CAS illuminations are specific to the pitch trim system:
CAS
PIT TRIM MISCMP
PIT TRIM MISCMP
Color
Description
Red Miscompare between the primary and secondary pitch trim on the ground and either
thrust lever is advanced to MCR or above.
White Miscompare between the primary and secondary pitch trim in flight.
PITCH TRIM SELECTOR SWITCH
This switch, as shown in Figure 5-4, is located on the trim switch panel
(pedestal) and is used to select which trim system will be used. The PRI
position enables the primary trim switches in the control wheels, while
the SEC position will enable the secondary trim switches in the panel.
Selecting OFF or SEC will disable all #1 IC-600 trim functions. When set
to OFF position, the power and ground circuits for the motor command
functions in the actuator control box are disconnected and a CAS is
posted.
The following CAS illumination is specific to the pitch trim system:
CAS
PITCH TRIM OFF
PM-126A
Color
Description
White Pitch trim is selected to OFF.
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Pilot’s Manual
PRIMARY PITCH TRIM
Each control wheel has a control wheel trim switch located on the outboard horn of each control wheel (Figure 5-1). Each switch is a four position (LWD, RWD, NOSEDOWN & NOSEUP) barrel switch with a
momentary-action push button switch in the center of the barrel. This
switch is used to input trim commands for pitch and roll and autopilot
functions. Normally, the pitch trim control switch (Figure 5-4) is positioned to PRI. This position enables the control wheel trim switches and
causes commands from either of these switches to be processed by the
#1 IC. To complete the trim command circuit, the arming switch (button) on the top of the barrel must be depressed simultaneously with
movement of the barrel. Trim commands from the pilot’s control wheel
trim switch will override commands from the copilot’s. Primary trim
speed is variable and is automatically controlled by the #1 IC based on
indicated airspeed. The IC uses airspeed information from both ADCs
to schedule trim speed and Mach trim. The #1 IC sends the primary
trim commands to the primary trim actuator. The trim actuator and the
IC both monitor the trim operation. The primary trim actuator performs a power-up circuit check. If the actuator detects a fault during the
power-up check, a fault is posted on the CAS. Primary trim will still be
available with the fault displayed, however, operation may be at a low
trim rate and configuration trim and Mach trim may be inoperative depending on the malfunction. The primary trim actuator and #1 IC both
monitor primary trim operations for a number of possible malfunctions
including uncommanded trim and trim in the wrong direction. If either
of these malfunctions is detected by the trim actuator, a fail is displayed
on the CAS and primary trim is disabled.
PITCH TRIM
RUDDER TRIM
SEC
OFF
PRI
NOSE
LEFT
NDN
NOSE
RIGHT
O
F
F
NUP
SEC
TRIM SWITCH PANEL
Figure 5-4
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PM-126A
Pilot’s Manual
Electrical power for primary pitch trim is provided by the L ESS BUS
and is protected by the TRIM-PRI PITCH circuit breaker located on the
pilot’s circuit breaker panel (FLIGHT group). The dc electrical power to
the #1 IC is also required for primary pitch trim, except for primary bypass trim. The power for #1 IC is provided by the L ESS BUS and protected by IC/SG 1 circuit breaker located on the pilot’s circuit breaker
panel (INSTRUMENT/INDICATIONS group).
Bypass Trim
Primary trim reverts to bypass as a result of a detected malfunction or
#1 IC failure and cannot be selected by the crew. When in bypass trim a
fault is displayed on CAS and control wheel trim switch commands go
directly to the primary trim actuator, bypassing the IC circuits. The #1
IC trim functions (IC controller/monitored primary trim, configuration
trim, and Mach trim) are all disabled in this case. When in bypass trim,
the primary trim actuator operates at only two speeds (high or low).
The speed depends on flap position. Dual flap position inputs are provided to the actuator electrical box to maintain redundancy. If the flap
signals do not agree, the rate of trim function is limited to slow speed.
Both flap signals must agree and must indicate flaps are greater than 3°
for the actuator to operate at a high rate. When the flaps are up (<3°),
primary bypass trim will be at a slow rate. The primary trim actuator
continues to monitor for uncommanded trim and incorrect trim direction in the bypass trim mode. It also monitors for the correct trim speed
based on flap position. If it detects a failure in any of these areas, primary trim is disabled and a fail CAS is displayed.
The following CAS illuminations are specific to the primary pitch trim:
CAS
PRI TRIM FAIL
PRI TRIM FAULT
PM-126A
Color
Description
Amber The primary pitch trim system has failed.
White The Integrated avionics Computer (IC)
detects a fault in the primary pitch trim
system.
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Pilot’s Manual
SECONDARY PITCH TRIM
Secondary pitch trim is electrically independent of the primary trim,
configuration trim, and Mach trim. In the event of primary trim failure,
secondary pitch trim is available as a backup means of trimming the
airplane in the pitch axis. The autopilot also uses the secondary trim actuator as a normal means of trimming in the pitch axis. The autopilot
can use the secondary trim actuator with the trim selector in the PRI or
SEC position.
The dual-segment SEC trim switch (Figure 5-4) is located on the center
pedestal. Manual activation of secondary trim requires that the pitch
trim selector be in the SEC position and that both segments of the
spring-loaded SEC switch be moved at the same time. When SEC position is selected, a CAS is displayed.
The secondary pitch trim actuator has a monitor function similar to the
primary actuator. It performs a power-up check and if any faults are detected, a fault is displayed on the CAS, however, secondary trim operates normally. The secondary trim actuator also monitors for
uncommanded trim, trim in wrong direction, and incorrect trim rate. If
any of these malfunctions are detected, a fail annunciation is posted on
CAS and the secondary actuator is disabled. The IC has no control or
monitor functions for manual secondary trim.
The following CAS illuminations are specific to the secondary pitch
trim:
CAS
SEC TRIM FAIL
SEC PITCH TRIM
SEC TRIM FAULT
Color
Description
Amber Secondary pitch trim has failed.
White Secondary pitch trim is selected by the crew.
White A pitch trim actuator (secondary) fault is
detected.
Electrical power for the secondary pitch trim system is provided from
the R ESS BUS and is protected by the TRIM-SEC PITCH circuit breaker
located on the copilot’s circuit breaker panel (FLIGHT group).
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PM-126A
Pilot’s Manual
AUTOPILOT PITCH TRIM
When the autopilot is engaged, it can drive the horizontal stabilizer
trim to alleviate elevator servo loading. The autopilot pitch trim function is contained in the #2 IC autopilot processor. When elevator servo
current exceeds a predetermined threshold for a given period of time,
this is considered to be a steady state error and trim will run. As the
trim runs, the horizontal stabilizer is re-positioned and the air load on
the elevator primary servo is reduced. When this load falls below the
threshold level, trim stops running.
Whenever the autopilot is engaged, #1 IC trim functions, which includes config/Mach trim, drop off-line. The autopilot commands pitch
trim based on elevator servo current demand and airspeed. Autopilot
pitch trim engagement is controlled by the autopilot engage logic. An
autopilot engage signal is provided to the horizontal trim actuator. If
the autopilot is disengaged as a result of a monitor trip, the aural tone
alert will sound until the MSW switch is pushed. A red AP will also be
displayed on the PFDs and flash for five seconds and then go steady.
The #2 IC monitors for uncommanded trim, trim direction, and incorrect trim rate. If the actuator detects one of the above faults, a CAS is
displayed.
The following CAS illumination is specific to the autopilot pitch trim:
CAS
AP ELEV MISTRIM
Color
Description
Amber Autopilot elevator servo holding excessive
torque.
TRIM-IN-MOTION INDICATION
A trim-in-motion potentiometer is installed on the secondary trim actuator. When the autopilot energizes the secondary trim actuator for
more than 2 to 3 seconds, a series of audible clacker sounds is transmitted through the audio system. A built-in time delay allows trim operation for approximately 2 to 3 seconds before the clacker sounds which
prevents a nuisance alarm on the clacker. For longer periods of continuous trim, the clacker will alert the crew. Unusual long periods of autopilot trimming may indicate trim runaway. There is no trim-in-motion
clacker for any trim operation other than autopilot trim.
PM-126A
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PITCH TRIM BIAS
The pitch trim bias system works in conjunction with the up/down
spring assembly. Its function is to assist the pilot by providing added
spring pressure against the elevator in the event the horizontal stabilizer is jammed in an out-of-trim position. Pitch trim bias is actuated by
the crew using the three-position (spring loaded to the center position)
PIT TRIM BIAS switch located at the front of the throttle quadrant.
Power for the system is provided through the PIT TRIM BIAS circuit
breaker on the copilot’s circuit breaker panel (FLIGHT group).
The following CAS illuminations are specific to the pitch trim bias.
CAS
PIT TRIM BIAS
PIT TRIM BIAS
Color
Description
Red Abnormal PIT TRIM BIAS configuration with
the aircraft on the ground and either thrust
lever is advanced to MCR or above.
White The pitch trim bias system is moved from the
normal position. PIT TRIM BIAS should only
be used for jammed stabilizer conditions in
flight.
CONFIGURATION TRIM
The configuration trim functions aid the pilot by providing automatic
relief of control column loads via the #1 IC control of horizontal stabilizer position. The configuration trim system control and monitoring
functions are provided by software contained in the #1 IC using inputs
from spoiler lever position sensors. Through these interfaces, the configuration trim provides automatic pitch control for changes in airplane
configuration. This mode is only functional when the trim selector
switch is in the PRI position and the autopilot is not engaged. Trim
commands from either control wheel trim switch will have priority
over the configuration trim commands.
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Pilot’s Manual
AILERON TRIM
The aileron trim system provides manual aileron trim tab control. The
manual trim tab control system enables the pilot, with authority, and
the copilot to eliminate out-of-trim forces which may be present in the
aileron control circuit, preventing smooth operation of the control column. This enables the airplane to be flown without either pilot having
to apply constant forces to the hand wheel to maintain the wing level.
The aileron trim system is controlled by a control wheel trim switch
mounted on the pilot and copilot’s control wheels (Figure 5-1) and incorporates two switches, trim, and trim arm. To manually trim, the pilot or copilot must press and hold the ARM button while pushing the
trim switch to the LWD or RWD position. The control wheel trim
switch induce inputs into the roll trim control electrical system which
translates commands to a rotary actuator mounted in the left aileron.
The actuator moves the aileron trim tab through dual push rods to the
command position. A trim tab position sensor is attached to the rotary
actuator shaft and provides input to the Data Acquisition Units (DAUs)
for display of aileron trim position on the cockpit Engine Indicating
and Crew Alerting System (EICAS). Driving the actuator clockwise
causes the trim tab to rise. This results in left aileron moving down and
the right aileron moving up. This results in the airplane performing a
Right Wing Down (RWD) movement. Conversely, driving the actuator
counterclockwise causes the trim tab to lower. This causes the left aileron to move up and the right aileron to move down, resulting in the airplane performing a Left Wing Down (LWD) movement.
Aileron trim is powered from the L ESS BUS and is protected by TRIMAIL 5-amp circuit breaker on the pilot’s circuit breaker panel (FLIGHT
group).
PM-126A
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Pilot’s Manual
RUDDER TRIM
The rudder trim system provides manual rudder trim tab control. The
manual trim control system enables the pilot and copilot to eliminate
out-of trim forces which may be present in the rudder control circuit.
This enables the airplane to be flown without either pilot having to
apply a constant force to the rudder pedals.
Rudder trim changes are effected through an electronically driven rotary actuator mounted in the rudder and connected to the rudder trim tab
with dual pushrods. The actuator is controlled manually by a doublepole, double-throw, center-off, momentary-action, rotary switch located on the trim switch panel (Figure 5-4) in the center pedestal. This
switch is constructed in two sections with poles that are not mechanically linked. One pole of the switch is used to provide control of the
rudder trim ARM circuit and is referred to as the ARM switch. The other pole of this switch is used to provide either nose left or nose right
trim commands and is called the rudder trim switch. These poles are independent of each other except of the fact that they are both rotated by
the same shaft. The failure of one pole will not affect the other. Since one
pole provides ARM control and the other provides the trim command
inputs, the failure of one pole will not result in a trim runaway. Setting
and holding the switch to the NOSE LEFT or NOSE RIGHT position energizes the trim tab actuator, resulting in the rudder rotating either
clockwise or counterclockwise. A trim tab position sensor is attached to
the rotary actuator shaft and provides input via the #2 data acquisition
unit for display of rudder trim position on the cockpit engine indicating
and crew alerting system.
Electrical power for rudder trim is provided from the R ESS BUS and
protected by TRIM-RUD 5-amp circuit breaker located on the copilot’s
circuit breaker panel (FLIGHT group). Rudder trim can be stopped by
depressing and holding either control wheel master switch.
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Pilot’s Manual
TRIM INDICATIONS
Pitch, aileron, and rudder trim indications are provided on the EICAS
and the MFD. A digital display of pitch trim position (PIT TRIM) is always in view below the CAS window, on the right side of the EICAS.
Pitch (PIT), aileron (AIL), and rudder (RUD) trim are digitally displayed on the SUMRY page. They are arranged in a vertical column labeled FLT on the right side of the SUMRY page. The SUMRY page is the
power-up default display on the EICAS. The SUMRY page is displayed
at the base of the MFD. Trim indications are correspondingly displayed
on the left side of the FLT system schematic page.
The following CAS illuminations are specific to the trim indications:
CAS
TAKE OFF TRIM
TAKE OFF TRIM
Color
Description
Red The aircraft is on the ground and either thrust
lever is advanced to MCR or above, and aircraft trim (pitch, aileron, or rudder) is not set
for takeoff.
White The aircraft is on the ground and aircraft trim
(pitch, aileron, or rudder) is not set for takeoff.
PITCH TRIM INDICATIONS
Pitch trim tab position is presented as both analog and digital display.
The label PITCH, in cyan, is positioned above the pitch trim tab position digital readout. The range of pitch trim is from 0 to 10, and with 0
being maximum nose down trim and 10 being maximum nose up trim.
The analog scale consists of a white vertical line with three horizontal
tick marks on the right side. The labels NDN and NUP are displayed at
the left top and bottom of the scale, respectively. The digits 0 and 10 are
displayed at the right top and bottom of the scale, respectively. The
analog scale has a white takeoff band located between 5.5 units and
8.7 units. There is a pointer which moves up and down the left side of
the scale in accordance with the digital readout of the pitch trim tab position. If the pitch trim is not within the takeoff band, and the airplane
is on the ground, the digital display of trim will have a white box
around it and a message posted to CAS. If power is advanced for takeoff (MCR or greater) and pitch trim is not within the takeoff band, the
“CONFIGURATION” voice warning will sound and the CAS message
turns red along with the digits, pointer and box in the trim position display. Invalid data will replace the digits with amber dashes, and the
pointer and box are removed.
PM-126A
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Pilot’s Manual
AILERON TRIM INDICATIONS
Aileron trim tab position is presented as both analog and digital display. The label AILERON, in cyan, is positioned above the aileron trim
tab position digital readout. The range of aileron trim position is from
L12 to R12 and with L being left wing down, and R being right wing
down. The analog scale consists of a white arc with three tic marks on
the outside of the arc. The digits 10, in white, are displayed at the left
and right ends of the scale, respectively. A white takeoff trim band is located on the outside of the scale between the values of +5 and -5. A
pointer moves along the inside of the scale in accordance with the digital readout of the aileron trim tab position. If the aileron trim is not
within the takeoff band while the airplane is on the ground the digital
display will have a white box around it. CAS messages and alerting are
the same as those described above in pitch trim.
RUDDER TRIM INDICATIONS
Rudder trim tab position is presented in both analog and digital display. The label RUDDER, in cyan, is positioned above the rudder trim
tab position digital readout. The range of rudder trim position is from
L12 to R12, and with L being nose left and R being nose right. The analog scale consists of a horizontal white bar with three tic marks on the
top of the bar. The digits 10, in white, are displayed at the left and right
ends of the scale, respectively. A white takeoff trim band is located on
the top of the horizontal scale between +5 and -5. There is a pointer
which moves along the bottom of the scale in accordance with the digital readout of the rudder trim tab position. If the rudder trim is not
within the takeoff band while the airplane is on the ground the digital
display will have a white box around it. CAS messages and crew alerting are the same as described in pitch trim. Pitch, aileron, and rudder
trim indications are available on page 2 of the backup engine/system
pages on the RMU.
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Pilot’s Manual
MACH TRIM
Mach trim is a fully automatic system installed to increase longitudinal
stability and counteract nose-down tendency at high Mach numbers. A
circuit card in the #1 IC performs all the computational aspects for
Mach trim and signals the primary trim actuator to apply trim as necessary. Airspeed information provided by the ADCs is used by the IC
in computing the trim requirement.
The pitch trim selector (Figure 5-4), located on the center pedestal, must
be in the PRI position for Mach trim to be functional and the autopilot
must be disengaged for the Mach trim to become active. If the autopilot
is engaged, it performs the pitch trim function using the secondary trim
actuator and the Mach trim is in a passive mode. Mach trim automatically becomes active at 0.725 MI. Nose up trim will be applied as Mach
increases and nose down as Mach decreases. When the horizontal stabilizer position changes, two Mach trim position sensors apply feedback signals to the IC. Mach trim is interrupted whenever the manual
trim is activated. The system resynchronizes to function about the new
horizontal stabilizer position when manual trim is released. If the IC
detects a fault within the Mach trim system function, a fail is posted on
the CAS and the overspeed cue on the airspeed indicator will also adjust to indicate a Mach limit of 0.76 to 0.78 MI.
The following CAS illuminations are specific to the Mach trim:
CAS
MACH TRIM FAIL
MACH TRIM FAIL
PM-126A
Color
Description
Amber Mach trim function has failed and aircraft
speed is greater than 0.76 to 0.78MI.
White Mach trim function has failed and aircraft
speed is equal to or less than 0.76 to 0.78MI.
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STALL WARNING SYSTEM
The stall warning system, also referred to as the Angle-of-Attack
(AOA) system, is installed to provide the crew with an indication of impending airplane stall. The stall warning system consists of two independent systems which use a dual channel computer.
Other system components include two AOA sensors, control column
shaker motors and an interface with the PFDs. Left and right AOA indicators are available as an option. Each channel of the computer generates a reference signal to the corresponding stall vane and, in return,
receives AOA information. The computer then processes this information with airspeed, altitude, flap setting, and weight-on-wheels inputs
to determine the stall warning indications. The left and right stall warning systems are powered from the left and right essential buses respectively. The circuits are protected by the L STALL WARN and R STALL
WARN circuit breakers located on the pilot’s and copilot’s circuit
breaker panels (FLIGHT group).
STALL WARNING INDICATIONS
As the airplane approaches stall speed, stall warning indications are activated. The shaker speed will be above the stall speed at the most critical weight and Center of Gravity (CG). The stall warning computer
sums inputs of AOA and altitude shift along with flap position from
the flap position indication unit. Stall warning is biased for each flap
setting. The stall warning system provides the following aural, tactile,
and visual indications when the predetermined conditions have been
reached:
(1) The left and right channels of the computer drive lowspeed cues on the pilot’s and copilot’s PFDs respectively. The low-speed cue is a vertical red bar on the
inside of the airspeed tape which rises from the bottom
of the tape as the airplane AOA increases. The point at
which the red bar reaches the airspeed pointer will
coincide with the point at which other stall warning
indications are activated.
(2) The left and right channels of the stall computer will
activate the control column shaker motors.
(3) The non-cancelable voice message “STALL” will repeat
until the AOA is decreased.
(4) The AOA indicators (if installed) will enter the red
band on the indicator.
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STALL WARNING OPERATION
The stall warning system is powered when the circuit breakers are in.
The shakers, along with other visual and aural stall indications, are inhibited until the airplane is airborne. If installed, the AOA indicators
will operate in both the air and ground modes. The stall warning system performs a power-up self-test (BIT) and monitors for a number of
possible system faults. Detection of a fault appears on CAS.
The following CAS illumination is specific to the stall warning system:
CAS
STALLWARN FAIL
Color
Description
Amber The associated (L or R) stall warning system
has failed.
STALL VANE ANTI-ICE
The stall vanes are equipped with a 28-vdc heater to anti-ice the vane
surfaces during icing conditions. The AOA vane heater of the angle-ofattack transmitter is monitored for open circuit when the vane heater
power is applied. Detection of an open circuit will result in the appropriate CAS message as well as being logged into the stall computer as
a fault for that flight. The vane heaters are controlled by the L and R
PROBE anti-ice switches located on the anti-ice section of the center
switch panel. Each vane heater is supplied power from the left and
right main bus respectively and protected by the AOA 15-amp circuit
breakers on the pilot’s and copilot’s circuit breaker panel (ANTI-ICE
group).
The following CAS illumination is specific to the stall vane anti-ice
system:
CAS
AOA HT FAIL
PM-126A
Color
Description
Amber Associated (L or R) angle-of-attack vane
heater has failed.
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Pilot’s Manual
STALL SYSTEM TEST
A self-test mode is available when the weight-on-wheels signal indicates that the airplane is on the ground and no system failure is detected. When the system test switch is rotated to the STALL position and
held down for approximately 7 to 10 seconds, the stall warning computer shall demonstrate that the stall warning system is fully operational by performing the following events in the order listed:
(1) L AOA HT FAIL message appears in CAS.
(2) Low Speed Awareness (LSA) bar will begin to sweep
up the pilot side airspeed tape, the left (pilot’s) column
will shake when the LSA bar approximately reaches
the indicated airspeed pointer, and the aural voice
warning “STALL” will be repeated through the cockpit
speakers and crew headphones.
(3) L AOA HT FAIL message extinguishes from the CAS
window, the LSA bar scrolls down the airspeed tape,
the left column stops shaking and the aural warning
stops.
(4) R AOA HT FAIL message appears in CAS. (Note: Master caution tone may not sound when the R AOA HT
FAIL is annunciated. If the master caution tone is not
heard, then the STALL aural warning will be heard as
called out in the next step below).
(5) The LSA bar will begin to sweep up the copilot side
airspeed tape, the right (copilot’s) column will shake
when the LSA bar approximately reaches the indicated
airspeed pointer, and the aural voice warning “STALL”
will be repeated through the cockpit speakers and crew
headphones (if the master caution tone was not heard
in the previous step).
(6) R AOA HT FAIL message extinguishes from the CAS,
the LSA bar scrolls down the airspeed tape, the right
column stops shaking and the aural warning stops (if
the aural warning was present in the previous step).
The left and right stall warning failure discretes are not checked during
self-test. It was necessary to inhibit the output of the left and right
failure discretes in order to permit display of the LSA bar on the PFDs
during test.
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ANGLE-OF-ATTACK INDICATORS (OPTIONAL)
The optional Angle-of-Attack (AOA) indicators system consists of two
angle-of-attack indicators mounted in the instrument panel, one outboard of the pilot PFD and one outboard of the copilot PFD, Figure 5-5.
The angle-of-attack indicators display continuous angle-of-attack position to the flight crew. The AOA indicators are driven by the stall warning computer. The dual channel computer provides buffered outputs to
the indicators for protection. The pilot AOA indicator receives data
from the left channel of the stall warning computer and the copilot
AOA indicator receives data from the right channel of the stall warning
computer. The AOA indicator is adequately marked displaying .10, .20,
.30, .40, .50, .60, .70, .80, .90, 1.0 with unnumbered marks half way between each. The beginning of the red band at .80 represents shaker activation and an imminent stall condition. The AOA indicators front
plate markings are consistent with the stall warning information
shown on the PFDs, a tape type presentation at the end of the airspeed
tape. The AOA indicators are powered via the L STALL WARN and R
STALL WARN circuit breakers located on the pilot’s and copilot’s circuit breaker panels (FLIGHT group).
COPILOT AOA
INDICATOR
F40-050000-055-01
PILOT AOA
INDICATOR
F40-050000-055-01
INSTRUMENT PANEL LAYOUT AND AOA INDICATOR POSITION
Figure 5-5
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AVIONICS
HONEYWELL PRIMUS 1000 AVIONICS SYSTEM
The Learjet 45 is equipped with a Honeywell Primus 1000 Avionics system. The primary component of the Primus 1000 system is the display
flight guidance computer, or more simply, the IC-600. This computer,
together with the appropriate controllers and sensors, comprises the
Primus 1000 system. It consists of dual IC-600 (single autopilot is contained in the copilot’s IC-600), dual air data computers, PRIMUS
weather radar system and appropriate controllers. The radio sensor
package is the Honeywell PRIMUS II integrated radio system.
ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)
The PRIMUS 1000 EFIS System consists of four, 8 x 7 inch, Display
Units (DUs) driven by two Symbol Generators (SGs) resident in the two
IC-600s. The EFIS presents information to the crew in an uncluttered
format, simplifying cockpit scan, and reducing pilot workload and fatigue. The flight instruments, engine instruments, system status, navigation, TCAS, RADAR, and electronic checklist are all displayed on
these high resolution DUs. The EFIS is integrated with the Engine Indicating and Crew Alerting System (EICAS) and Crew Warning Panel
(CWP) to provide the crew with not only flight monitoring indications
but also with engine data, warning, cautionary and advisory alerts (visual and aural). Dual Primary Flight Displays (PFDs) combine attitude
and HSI formats with airspeed, vertical speed and other essential information, such as resolution advisories for the optional TCAS system. A
Multi-Function Display (MFD) offers a full spectrum of operational capabilities, from weather radar and mapping displays, to a custom programmable checklist. A digital audio control system and dual Radio
Management Units (RMUs) support the communications and navigation functions.
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The display information provided on EFIS is generated by two IC-600
computers located in the nose. Each of the IC-600s contains circuitry
that performs the symbol generation function for the EFIS. Along with
interfacing with the display units, the ICs receive data from the Data
Acquisition Units (DAUs), Air Data Computers (ADCs), Attitude
Heading Reference System (AHRS), navigation system, flight management system, autopilot and other various display controllers. The CAS
monitors the IC-600 bus interconnect, the temperature of each IC-600,
the IC cooling fans and Weight-On-Wheels (WOW). A CAS will also
illuminate if communications between the left and right ICs are invalid.
The following CAS illuminations are specific to the IC-600:
CAS
IC 1-2 OVHT
IC BUS FAIL
IC1-2 FAN FAIL
IC1-2 WOW INOP
Color
Description
Amber #1 and/or #2 Integrated avionics Computer
(IC) are/is overheated.
Amber -The off-side IC has failed.
or
- IC bus invalid
White #1 and/or #2 Integrated avionics Computer
(IC) cooling fan has failed.
White The associated (#1 or #2) Integrated avionics
Computer (IC) has tripped the weight-onwheels validity monitor.
IC-600 POWER SOURCE
The #1 and #2 IC-600s are powered from the left and right essential
buses respectively. The circuits are protected by the 7.5-amp IC/SG 1
and IC/SG 2 circuit breakers on the pilot’s and copilot’s circuit breaker panels (INSTRUMENT/INDICATIONS group).
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AVIONICS MASTER SWITCHES
Left and right avionics master switches are located on the electrical
control panel below DU 2 (Figure 5-6). When the alternate action
avionics master switches are selected to On (OFF annunciator
extinguished), contactors are closed that connect the left and right
essential avionics buses and left and right main avionics buses to the
respective generator buses.
The associated essential contactors and main bus contactors must be
closed for the avionics buses to be powered. If the avionics master
switches are on during ground start or for a starter assisted airstart, the
essential avionics buses will continue to be powered, but the contactors
for the main avionics buses will automatically open until the start is
complete. The essential avionics buses must remain powered during an
airstart since they power the critical flight display units. The emergency
bus, essential buses and essential avionics buses are all powered by the
emergency battery during a starter assisted start. The avionics equipment that must be on during a ground start is powered from the essential buses and the emergency battery bus.
ELECTRICAL
L AV
MSTR
R AV
MSTR
EMER BATT
L ESS
R ESS
EMER
OFF
OFF
OFF
L
R
L MAIN
NON-ESS
NON-ESS
R MAIN
OFF
OFF
OFF
OFF
BUS-TIE
L GEN
R GEN
MAN
OFF
OFF
EXT PWR
L BATT
R BATT
OFF
OFF
ON
AVAIL
APU GEN
ON
AVAIL
ELECTRICAL CONTROL PANEL
Figure 5-6
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AIR DATA SYSTEM (ADS)
The air data system and air data instruments depend upon pitot pressure and static pressure sensing, as well as air temperature sensing. Air
data is provided to the flight instruments and airplane systems by two
Air Data Computers (ADCs) which receive pitot and static information
from the main pitot static system. The ADCs receive total air temperature from a dual element temperature probe and barometric correction
inputs via the BARO set knobs on the corresponding PFDs.
PITOT-STATIC SYSTEM
The primary pitot-static system consists of two pitot-static probes, located one on each side of the airplane’s nose section. The pilot’s pitotstatic probe supplies the pilot’s ADC with total pressure and the copilot’s pitot-static probe supplies the copilot’s ADC with total pressure.
Each pitot-static probe has two isolated static ports. The pilot’s ADC receives static pressure from coupled static ports off the pilot’s and copilot’s pitot-static probes (Figure 5-7). The copilot’s ADC receives static
pressure from separate coupled static ports off the pilot’s and copilot’s
pitot-static probes which are isolated from the static ports used by the
pilot’s ADC.
PILOT
PITOT/STATIC
PROBE
COPILOT
PITOT/STATIC
PROBE
ADC 2
ADC 1
F40-050000-057-01
PITOT
PITOT
STATIC 1
STATIC 1
STATIC 2
STATIC 2
ADC 1 PITOT PRESSURE
ADC 2 PITOT PRESSURE
ADC 1 STATIC PRESSURE
ADC 2 STATIC PRESSURE
F40-050000-057-01
PITOT-STATIC SYSTEM SCHEMATIC
Figure 5-7
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Static 1 from the left probe connects with static 2 on the right probe to
provide static pressure to the pilot’s ADC. Static 1 from the right probe
combines with static 2 on the left probe to provide static information to
the copilot’s ADC. This crossover arrangement reduces system errors
(Figure 5-7).
A third pitot-static probe, mounted above the main probe on the right
side of the airplane, provides total and static pressure inputs to the
standby instrument group. Moisture drains are provided for the standby pitot-static lines. The two drains for the standby pitot-static system
are flush mounted on the right side of the airplane just aft of the nose
wheel door. The main pitot-static probes are physically located at the
lowest point of the primary pitot-static system plumbing and therefore,
do not require moisture drains. The pitot source on the standby probe
provides total pressure to the standby Mach/airspeed indicator. There
are two static sources on the standby probe, one provides static information to the standby altimeter and the other provides data to the
standby Mach/airspeed indicator (Figure 5-8).
STANDBY STATIC PRESSURE
STANDBY PITOT PRESSURE
STANDBY
PITOT/STATIC
PROBE
PITOT
F40-050000-058-01
STANDBY
MACH/AIRSPEED
INDICATOR
STATIC 1
STATIC 2
STANDBY
ALTIMETER
F40-050000-058-01
STANDBY PITOT-STATIC SYSTEM SCHEMATIC
Figure 5-8
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AIR DATA COMPUTERS (ADCs)
The Learjet 45 utilizes two, independent micro air data computers as
the primary source for air data. The is a self contained unit incorporating pressure sensing modules and all required processing and input/
output functions in a single unit. Each computer is independent of the
other and has independent circuit breakers.
The air data system provides the required airplane airspeed, air temperature, altitude and vertical speed data for the Electronic Flight Instrument System (EFIS) displays, Attitude and Heading Reference
System (AHRS), dual stall warning system, autopilot, transponders,
spoileron computer, cabin pressurization, Digital Electronic Engine
Control (DEEC) and landing gear warning system as required. The Air
Data System (ADS) accepts static air pressure, total air pressure, total
air temperature, various discrete signals and baro set inputs. The #1
and #2 ADCs receive power from the L and R ESS BUS respectively,
through ADC 1 and ADC 2 circuit breakers. The circuit breakers are located on the pilot’s and copilot’s circuit breaker panels (INSTRUMENT/INDICATIONS group).
ADC REVERSION
Display of the ADS data on the EFIS display is controlled by the pilot
or copilot via the ADC reversionary control switch (Figure 5-14). The
reversionary control panel, located below the EICAS display (DU#2),
incorporates an ADC reversion switch which has three positions, 1 ADC NORM - 2. In the ADC NORM position, the IC-600s receive air
data from their on-side ADC. In the “1 or 2” position both IC-600s receive air data from the selected source. If the switch is not in the NORM
position, an annunciator of the selected source is displayed above and
to the left of the ADI on both PFDs.
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STANDBY INSTRUMENTS
The standby instrument group (Figure 5-9) includes a barometric altimeter, a airspeed/Mach indicator, an attitude indicator, mounted on the
center instrument panel above the CWP and RMUs. The standby instruments have their own pitot-static probe to provide air data information. The instruments are of traditional mechanical design. If a fault
occurs which causes one of the ADCs to output misleading information
to the PFDs, the standby instruments act as a useful comparison to indicate which of the three displays is incorrect.
0
0
9
6
10
40
1
1013
12
8
In Hg
2
ALT
14
35
30
2
16
25
2 0 0
3
51,000 FT
7
20
2992
6
ULL
P
hPa
VIB
ON
5
4
TO
C
A GE
STANDBY INSTRUMENT GROUP
Figure 5-9
STANDBY ALTIMETER
The standby altimeter displays baro corrected altitude in a pointer/
counter drum display. The dial graduations are marked every 20 feet.
Above sea level the counter displays every 100 feet up to 55,000 feet of
operational range. The indicator has dual barometric correction, from
27.9 to 31 inches of mercury and 946 to 1050 hectoPascals. Back lighting
is provided by 5-vdc to illuminate the standby altimeter indicator at
night.
STANDBY AIRSPEED/MACH INDICATOR
The standby airspeed/Mach indicator provides indicated airspeed by
means of a pointer indicating against a 50- to 400-knot dial and a Mach
sub-dial ranging from 0.3 to 1.0 Mach. Maximum allowable airspeed
(Vmo) is indicated at 325 knots by a red radial mark on the airspeed dial. Maximum allowable Mach (Mmo) is indicated at 0.75 Mach by a red
and white striped radial mark on the Mach sub-dial. Back lighting is
provided by 5-vdc to illuminate the standby airspeed/Mach indicator
at night.
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STANDBY ATTITUDE INDICATOR
The standby attitude indicator provides a visual indication of the airplane flight attitude. It is located in the center of the standby instrument
group (Figure 5-9) where it can be viewed easily by both pilots. It is
powered from the emergency battery bus so that it will remain powered for at least one hour after the loss of airplane generator power. The
standby attitude indicator will continue to provide an accurate display
of aircraft attitude for a further nine minutes after the loss of all airplane
power.
The indicator is an electrically-driven gyro whose vertical attitude is
maintained by a mechanical erection system. The power warning flag
is pulled from view after the gyro has spun up to valid operating speed
and reappears if there is any interruption of source power or the unit is
in caged mode.
Back lighting is provided by 5-vdc to illuminate the standby attitude
indicator at night.
STANDBY COMPASS
The standby compass is located at the top of the windshield center post.
It is a magnetic compass that does not require any electrical power to
provide the crew with a continuous standby heading display. The only
electrical input to the compass is 5-vdc to illuminate the compass at
night.
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ATTITUDE HEADING REFERENCE SYSTEM (AHRS)
Due to the design of the Honeywell Primus 1000 Avionics System, the
various avionics systems are very integrated. The AHRS uses the PFDs
as its primary display and the MFD in the event of PFD failure. The display units, display controllers and appropriate reversion switches are
considered part of the electronic display system and are covered later
in this section.
The Learjet 45 is equipped with either a dual Honeywell (AH-800) or
dual LITEF (LCR-93) Attitude Heading Reference Units (AHRUs). Both
units contain a memory module and are located in the aircraft’s nose
section.
Each system (#1 and #2 AHRS) incorporates a flux valve located in their
respective wing tips. The AHRUs contain three Fiber Optic Gyros
(FOGs) which sense angular rotation about the three principle axis
(pitch, roll, and yaw) thus, computing the airplane’s attitude and heading. When slaved to magnetic, the flux valves provide a magnetic heading reference. The memory module stores calibration data. This data is
used to compensate AHRU inaccuracies caused from installation errors
and local disturbances to the earth’s magnetic field created by the aircraft’s structure.
The AHRUs receive true airspeed (TAS) information from the on-side
ADC. However, if a single ADC failure occurs, they will receive TAS
from the operating ADC. True airspeed information is used to compute
pitch and roll attitude. If TAS inputs to #1 AHRS or #2 AHRS are lost, a
CAS will illuminate. Although system operation will be degraded, the
AHRS still retains the same accuracy as a conventional spinning mass
type gyro. AHRU’s data output is received through their corresponding IC for attitude and heading displays on the PFDs/MFD. Attitude
and/or heading information from the AHRS is used by the Flight Guidance System (FGS), Flight Management System (FMS), weather radar
system, and the fuel quantity indicating system. In addition, AHRS #2
provides heading information through DAU #2 for the backup navigation display on the RMU.
The following CAS illumination is specific to the AHRUs:
CAS
AHRS 1-2 BASIC
5-38
Color
Description
White Attitude Heading Reference System
(AHRS 1 or 2) has reverted to basic mode
due to a loss of true airspeed from both air
data computers.
PM-126A
Pilot’s Manual
ATTITUDE AND HEADING COMPARISON MONITORS
The attitude and heading comparison monitors are functions within
the IC-600s that compare the displayed data with the cross-side or secondary source data, depending on system reversionary status. Annunciations are provided to the crew if the attitude or heading on both sides
differ.
The attitude comparison function is made of two monitors, the roll
comparison monitor and the pitch comparison monitor. If the pitch
data displayed on each side differ, the pitch comparison monitor trips
and the PIT annunciation is displayed. The comparison threshold figure for the pitch monitor is 5°. Similarly, if the roll data on both sides
differ, the ROL annunciation is displayed. The comparison threshold
figure for the roll monitor is 6°. If both the roll and pitch comparison
monitors trip, the ATT annunciation is displayed. If the heading comparison monitor trips, HDG is displayed. The normal comparison
threshold figure for the heading monitor is 6°. However, if the displayed roll information is > 6°, the heading comparison threshold figure is increased to 12°.
All comparison monitor annunciations flash for 10 seconds on activation and then remain steady. These comparison monitors provide an
extra safeguard to alert the cockpit crew in the event of any failures affecting the attitude or heading data displayed.
Other annunciations for attitude and heading which are displayed on
the PFDs, not associated with the comparison monitors, are:
ATT FAIL and HDG FAIL. These red annunciations are displayed on
the affected side’s PFD whenever the heading or attitude display from
that AHRS has failed. If an AHRS fails or both primary and auxiliary
power supplies to an AHRS fail, both the red ATT FAIL and HDG FAIL
annunciations are displayed.
ATT1/2 and DG1/2. These annunciations are displayed on the PFDs
and indicate to the crew which AHRS is the source for the attitude and
heading data on the display. If the onside AHRS is the source of display,
the annunciation is white. If AHRS reversion has been performed, the
cross-side PFD annunciation is amber. There are no crew actions
required for these annunciations.
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AHRU POWER SOURCE & COOLING
Each AHRU has a primary and a secondary dc electrical power source.
The pilot’s AHRU receives primary power from the left essential bus
and a secondary or backup power from the right essential bus. The copilot’s AHRU receives primary power from the right essential bus and
secondary power from the left essential bus. Should either essential bus
fail in flight, power to both AHRUs is uninterrupted. Separate circuit
breakers for each system, primary and secondary, are provided in the
INSTRUMENT/INDICATIONS group on the pilot’s and copilot’s circuit breaker panels. The AHRS #1 PRI and #2 SEC circuit breakers are
located on the pilot’s side, and the AHRS #2 PRI and #1 SEC circuit
breakers are located on the copilot’s side. The AHRUs are equipped
with cooling fans which operate automatically to keep the AHRU within proper temperature limits. On aircraft 45-002 thru 45-174 (Honeywell
AH-800), a CAS illuminates if the temperature exceeds predefined limits.
The following CAS illumination is specific to AHRU cooling:
CAS
AHRS 1-2 OVHT
(Aircraft 45-002 thru
45-174)
Color
Description
Amber Attitude Heading Reference System (AHRS 1
or 2) has reached an overheat condition.
AHRS REVERSION
Failure of an AHRS is apparent when the on-side horizon and pitch
lines are removed from the ADI and a red ATT FAIL annunciator
appears in the upper center of the ADI. The heading compass rose will
display a HDG FAIL annunciator on the HSI. If either AHRS fail, the
AHRS reversion switch on the reversionary control panel (Figure 5-14)
will allow the pilot to select the remaining AHRS to provide attitude
and heading information to both displays. The three-position switch is
labeled 1 - AHRS NORM - 2.
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ELECTRONIC DISPLAY SYSTEM (EDS)
Four electronic displays are used to provide the display formats for the
Primary Flight Displays (PFDs), the Multi-Function Display (MFD) and
the EICAS display in the electronic flight instrument system. The four
display units are large format 8 x 7 inch, 16 color high resolution display tubes. The display units are identical and interchangeable, except
for the bezel controllers attached to the front of the units. The bezel controllers for the outboard DUs are the same and the bezel controllers for
the inboard DUs are the same. A display controller (two) provides the
means for each pilot to control the display of the on-side PFD and to activate the EFIS test function. A display unit reversion panel, located
above the PFDs, provides reversion control capability.
The display unit configuration powers up with the following displays:
DU#1 - Pilot’s Primary Flight Display (PFD #1)
DU#2 - EICAS Display
DU#3 - Multi-Function Display (MFD)
DU#4 - Copilot’s Primary Flight Display (PFD #2)
The above configuration can be changed using the EICAS reversion
switch. This provides the ability to swap the DU #2 and DU #3 displays
between EICAS and MFD as the pilots desire.
The display units require forced air circulation for cooling which is provided by two fans mounted on the rear of each DU. If a DU fan fails, a
CAS will illuminate indicating DU 1, 2, 3, or DU 4 fail. If the temperature of the DU reaches approximately 120° F, a CAS will illuminate.
The following CAS illuminations are specific to the DUs:
CAS
DU 1-2 OVHT
DU 3-4 OVHT
DU1-2 FAN FAIL
DU3-4 FAN FAIL
PM-126A
Color
Description
Amber #1 and/or #2 Display Unit (DU) is overheated.
Amber #3 and/or #4 Display Unit (DU) is overheated.
White #1 and/or #2 Display Unit (DU) cooling fan
has failed.
White #3 and/or #4 Display Unit (DU) cooling fan
has failed.
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PRIMARY FLIGHT DISPLAY (PFD)
The PFD (DU #1 and DU #4) is a single display in which all of the required flight and navigation data is displayed for each pilot. The PFD
(Figure 5-10) format is divided into two main sections. The top half displays an Attitude Director Indicator (ADI) with an airspeed tape to the
left, and a barometric altitude tape to the right. A Horizontal Situation
Indicator (HSI) is located on the lower half of the PFD. The HSI can be
displayed in three different formats. The three options are full 360°
compass rose (HSI), a 120° compass arc display (ARC), and a 120° map
display (MAP). The MAP cannot be displayed on the PFD if the adjacent display (DU #2 or DU #3) is already displaying an MFD MAP format. Weather information can be displayed on the PFD ARC or MAP
format. To the right of the HSI, a Vertical Speed Indicator (VSI) is displayed, and to the left, navigation information is annunciated.
Comparison monitors provide indications to the pilots that there is a
difference between the data displayed on each PFD. This monitoring is
a function within the IC-600s that compares what is being displayed on
one side with either the cross-side displayed data or the secondary
source data. Should data be out of tolerance between what is being reported from the source, and what is being sent to the display units, or
if avionics related exceedances are detected, CAS will illuminate.
The following CAS illumination is specific to the PFDs:
CAS
PFD CHECK
Color
Description
Amber The associated (L or R) Primary Flight Display (PFD) is displaying invalid data.
The brightness of each PFD is controlled by the DIM control on each respective display controller. Pilot’s PFD (DU #1) receives 28-vdc power
from the left essential avionics bus by a 15-amp circuit breaker DU 1 located in the INSTRUMENT/INDICATIONS group of the pilot’s circuit
breaker panel. Copilot’s PFD (DU #4) receives 28-vdc from the right essential avionics bus by a 15-amp circuit breaker DU #4 located in the
INSTRUMENT/INDICATIONS group of the copilot’s circuit breaker
panel.
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PRIMARY FLIGHT DISPLAY AND BEZEL CONTROLLER
Figure 5-10
BEZEL CONTROLLERS
Many of the display control functions are controlled by the DU bezel
controllers and by the menus displayed on the MFD and EICAS displays. The PFDs both use the BL-870 bezel controller located at the bottom of the PFDs. Each has two push buttons and two rotary knobs
(Figure 5-10). The push buttons and rotary knobs have functions dedicated to decision height, minimum descent altitude and barometric correction.
The MFD and EICAS displays use the BL-871 bezel controllers which
have six push buttons, menu keys, and a rotary knob for menu manipulation. The push buttons allow selection of functions displayed in the
menus on the MFD. There are three functions that the MFD bezel push
buttons provide - (1) selection of a submenu, (2) toggling the selection
of a menu item and (3) selection of a variable parameter for setting. The
MFD rotary knob is dedicated to the control of the map/plan range.
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With weather radar selected for display, the MFD rotary knob will have
no function. The EICAS bezel controller provides dedicated buttons for
the displayed EICAS menu. These buttons toggle the selection of the
EICAS system page displays. The EICAS rotary knob allows for scrolling of the CAS messages on the EICAS display.
MULTI-FUNCTION DISPLAY (MFD)
The MFD (normally DU #3) provides the flight crew with a means of
displaying a variety of information. In its normal mode it can serve as
a full time weather radar display superimposed on a 120° compass arc.
There are two basic formats available on the MFD, a partial arc (Map)
display, and a plan mode (North up). Like the PFD, the MFD may have
flight plans composed of up to ten connected waypoints imposed on a
compass card. True airspeed (TAS) provided from the ADC and ground
speed (GSPD) provided from the FMS, are displayed on the MFD. Other information displayed full time include: FMS source, “TO” waypoint, distance to “TO” waypoint, time-to-go to “TO” waypoint, wind
speed and direction, Static Air Temperature (SAT), and weather radar
(WX) modes.
The MFD also provides a second source for access to EICAS systems
pages as well as providing joystick functions. The MFD (DU #3) may
serve as a backup for any other DU through pilot initiated reversionary
modes.
Other information available on the MFD includes:
• TCAS mode (optional) — Controls the display of TCAS on the
map presentation.
• MFD MENU — Activation of this key will enable the MFD SUBMENU to appear.
• CHECKLIST PAGE — This key provides entry into the normal
checklist procedure index page.
• SYSTEM PAGE — Selection of this key will access the systems
sub-menu pages which are duplicates of the EICAS system
pages.
28-vdc is provided from the right essential avionics bus by a 15-amp
circuit breaker DU #3, located on the copilot’s circuit breaker panel (INSTRUMENT/INDICATIONS group).
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EICAS DISPLAY
The Engine Indicating and Crew Alerting System (EICAS) is an integrated digital computer/display system that replaces the majority of
the traditional gauges and warning lights located throughout the cockpit. The EICAS display (Figure 5-11) is divided into four designated areas: engine instruments, CAS messages, system display pages and
menu items. The EICAS also incorporates the Crew Warning Panel
(CWP) which provides crew alerting by visual representation while the
cockpit audio system provides the aural alerting. The Crew Alerting
System (CAS) provides the crew with a visual attention getting means
to alert them to a warning that requires immediate action, a caution
alert that requires subsequent pilot or maintenance action, or an advisory indication that may require pilot or maintenance action at some
point in time. Other airplane system parameters are displayed on the
lower portion of the display via system pages and are selectable by the
bezel controller at the bottom of the DU. Normally, the airplane system
summary page (SUMRY on the menu) is in view, which provides brief
status reports of all sub-systems. Menu selectable, a system schematic
of airplane electrical, hydraulic, environmental control, flight control,
and fuel systems can be individually selected for more detailed monitoring by the flight crew.
The following CAS illumination is specific to the EICAS display:
CAS
EICAS CHK
Color
Description
Amber Available on MFD display only. EICAS wraparound monitor.
28-vdc power is provided from the left essential bus by a 15-amp circuit
breaker DU 2, located on the pilot’s circuit breaker panel (INSTRUMENT/INDICATIONS group).
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EICAS DISPLAY AND BL-871 BEZEL CONTROLLER
Figure 5-11
DISPLAY CONTROLLERS
The display controllers (two DC-550s), located on the glareshield, provide immediate access to and control of the objects on the PFDs. Each
controller is configured with seven push buttons located on the front
panel along with two rotary knobs used for reference selection for bearing source, two concentric knobs for DU dimming and a momentary
push button (located inside concentric DU knob) used to initiate a system test (Figure 5-12).
The display controllers also provide a data acquisition function, collecting inputs from sources such as the bezel controllers, guidance controller, joystick, etc. The controllers pass these inputs to the corresponding
IC-600 for processing.
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The display controller buttons are as follows:
1. IN/HPA — Inches of mercury or hectopascals.
2. CUE — Selection of single cue or cross pointer command bars.
3. FPA — Controls selection and deselection of the flight path
angle symbol and flight path acceleration display.
4. WX — Select or deselect weather radar display on the PFD.
5. HSI — Provides up to three different display options on the
HSI.
6. FMS — Allows a navigation display of FMS information (alternately FMS 1 or FMS 2 if dual) to be selected for display on the
PFD.
7. NAV — Alternately selects NAV 1 or NAV 2 as the source of
NAV data on the HSI.
In
CUE
FPA
WX
HSI
FMS
NAV
hPa
FMS
FMS
PUSH
ADF
ADF
TO
TEST
NAV
OFF
OFF
DIM
OUTBD
BRG
DU
NAV
INBD
DU
BRG
DISPLAY CONTROLLER
Figure 5-12
Each controller contains two rotary bearing source selector knobs that
are used to assign the respective bearing pointers on the HSI or ARC
displays to a particular navigation source.
Power for the display controllers, and the IC-600s, are from the left and
right essential buses and are labeled IC/SG 1 and IC/SG 2 on the pilot’s
and copilot’s circuit breaker panels in the INSTRUMENT/INDICATIONS group respectively.
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DISPLAY UNIT REVERSION PANELS
A display unit reversion panel is located on the glareshield above the
PFDs on each side of the cockpit. The panel on the pilot’s side is for controlling the display on DU #2 and the panel on the copilot’s side is for
controlling the display on DU #3. The reversion selector knob on these
panels plus the push function of the knobs allow the operators to
switch the inboard DUs (DU #2 and DU #3) to display either PFD, MFD,
or EICAS formats. With both reversion selector switches in NORM, an
EICAS format is displayed on DU #2 and an MFD format on DU #3. Depressing the selector knob on either reversion panel flip-flops the
DU #2 and DU #3 displays, reversing the MFD and EICAS display locations. Placing the reversion selector to the PFD position on either side
causes the PFD format to move to the inboard display tube on that side
and the outboard display to blank.
DU 2
NORM
DU 3
NORM
WARN
WARN
PFD
PFD
OFF
CAUT
CAUT
EICAS REV PUSH
EICAS REV PUSH
DISPLAY UNIT REVERSION PANELS
Figure 5-13
It is important to note that when selecting display unit reversion, the
bezel controllers on DU #1 and DU #4 continue to work with the PFD
display when it is transferred to an inboard display unit. The airplane
master warning/caution lights are also located on the display unit reversion panels.
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DATA ACQUISITION UNITS (DAUs)
There are two dual channel data acquisition units (DAUs) installed in
the tailcone equipment area of the airplane. The DAUs receive engine
and airplane systems sensor information and pass it, primarily, to the
IC-600 computers. Both channels of DAU #1 provide left engine data
and both channels of DAU #2 provide right engine data. For redundancy, both channels of each DAU independently convert on-side engine
information to a common ARINC 429 data bus format and send it to
both IC-600s. The IC-600s process the information and send it to the selected display unit (normally DU #2) for EICAS display. In addition to
engine information, the DAUs also collect analog data from other airplane systems such as fuel, hydraulic and accumulator pressure, dc
electrical power, flight control settings, cabin pressure settings/indications, and oxygen temperature/pressure.
A three-position DAU reversionary switch is provided on the reversionary control panel located below DU #2 (Figure 5-14). The switch positions are labeled A, DAU NORM, and B. With the switch in DAU
NORM, both IC-600s use Channel A from the left DAU and Channel B
from the right DAU for engine/systems displays. In the reversionary
positions (A or B), each IC-600 uses only the selected channel from both
DAUs.
REVERSION
IC/SG NORM
1
2
ADC NORM
2
1
AHRS NORM
2
1
DAU NORM
A
B
REVERSIONARY CONTROL PANEL
Figure 5-14
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If either channel of either DAU should fail, if either A or B reversion is
selected, if an engine or system miscompare is detected, an appropriate
CAS will illuminate.
The following CAS illuminations are specific to the DAUs:
CAS
DAU 1A-1B FAIL
DAU 2A-2B FAIL
DAU A REV
DAU B REV
DAU ENG MISCMP
DAU SYS MISCMP
LBS/KGS CONFIG
LBS/KGS CONFIG
Color
Description
Amber Channel A and/or B of the #1 Data Acquisition Unit (DAU) has failed.
Amber Channel A and/or B of the #2 Data Acquisition Unit (DAU) has failed.
White Reversion of both Data Acquisition Units
(DAUs) to Channel A is selected by the crew.
White Reversion of both Data Acquisition Units
(DAUs) to Channel B is selected by the crew.
Amber The associated (L or R) Data Acquisition Unit
(DAU) has detected a miscompare between
channel A and B involving an engine parameter (N1,N2,ITT).
Amber The associated (L or R) Data Acquisition Unit
(DAU) has detected a miscompare between
channel A and B involving a system parameter (dc voltage, Emergency Bus voltage, dc
amperage, Battery temperature, Main
Hydraulic pressure, Brake Accumulator pressure, Oxygen temperature and pressure).
Amber The configuration of the integrated avionics
computer is not compatible with that of the
data acquisition unit (i.e., one is configured
for pounds while the other for kilograms) on
the ground.
White The configuration of the integrated avionics
computer is not compatible with that of the
data acquisition unit (i.e., one is configured
for pounds while the other for kilograms) in
flight.
The DAU circuit breakers are located in the INSTRUMENT/INDICATIONS group on each circuit breaker panel. On the left side is DAU 1
CH A and CH B, and on the right side is DAU 2 CH A and CH B.
DAU 1 CH A and DAU 2 CH A are powered by the EMER BATT bus.
DAU 1 CH B and DAU 2 CH B are powered by the L and R essential
buses respectively.
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RADIO MANAGEMENT UNITS (RMUs)
The two RM-855B Radio Management Units (RMUs) provide the central controlling functions for the entire basic radio system. Each RMU
is a color, active matrix, Liquid Crystal Display (LCD) based unit. The
primary function of each RMU is to select and control the frequencies
and operational modes of each radio. Each RMU also provides access
and storage for up to twelve pre-set channels for the VHF COM and
VHF NAV functions. Cross-side operation, maintenance display, power on self-test and pilot activated self-test and optional FMS radio tuning features are also available on each RMU. Each RMU also provides
backup engine and navigation display facilities in the event of EFIS/EICAS failure. Automatic presentation of engine data occurs on RMU #1
if neither IC-600 is providing EICAS data.
There are six line select keys on each side of the RMUs (Figure 5-15).
The top key is referred to as a transfer button and has directional arrows on them. The remaining keys are referred to as line select keys.
There are also eight function keys located at the bottom of the RMUs.
The RMU main tuning page is divided into six dedicated windows.
Each window groups the data associated with a particular function.
The windows (COM, NAV, ATC/TCAS, ADF, and TCAS DSPY) each
provide for control of both frequency and operational mode of the associated function. The RMU also has other display modes, called pages,
which provide additional features and functions for the control of the
radio system. The PGE function at the bottom of the RMU is used to access additional RMU pages. RMU display brightness is adjustable using the tuning knob after the DIM function key is depressed. To return
the tuning knob to normal operation select any other key.
RADIO MANAGEMENT UNIT
Figure 5-15
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RMUs each have a primary and a secondary power source. If the primary source is not available, the RMU will automatically switch to the
secondary power source. RMU #1 primary is powered from the L
essential bus, and RMU #1 secondary is powered from EMER BATT
bus. RMU #2 primary is powered from R main avionics bus and the secondary from the R essential bus. RMU #1 and RMU #2 primary and secondary circuit breakers are located in the COMMUNICATIONS group
on the pilot’s and copilot’s circuit breaker panels. RMU #2 primary is
the only power source affected by the avionics master.
RMU CROSS-SIDE OPERATION
Should the pilot decide to tune the copilot’s set of radios, he can push
the 1/2 function key and transfer his entire RMU display and operation
to the copilot’s #2 system. Both RMU displays will be identical;
however the pilot’s RMU will show the function legends on the main
tuning and memory pages in magenta to indicate that cross-control is
being exercised. In addition to having access to the #2 system, the pilot
still has the memory frequencies in their #1 RMU available for recall for
use with the #2 system. Both pilots have this control transfer function
available. It provides flexibility in crew coordinated tuning as well as a
back up mode in the event one RMU becomes inoperative. The pilot
may change any frequency or mode on the copilot’s system using the
pilot side RMU. Any changed frequency is annunciated in yellow on
the copilot’s RMU. The frequency will be white on the pilot’s RMU. If
the pilot should push the 1/2 button again, the pilot’s side RMU will
revert to the original display.
RMU BACKUP PAGES
Either RMU can provide two pages of backup engine and systems indication and one page for a backup navigation display. These backup displays can be selected on the PAGE MENU page of either RMU. The
backup engine and systems pages can be selected by depressing the
line select key adjacent to “ENGINE PG1” or “ENGINE PG2” on the
PAGE MENU page. The backup navigation display is accessed by depressing the line select key adjacent to “NAVIGATION” on the PAGE
MENU page.
The ENGINE PG1 and ENGINE PG2 contain information such as: ITT,
O/P (oil pressure, left and right), FUEL, HYDM-B (hydraulic pressure), N1, N2, OIL ° C, FF PPH (fuel flow pounds per hour), VOLTS,
EMER VOLTS, AMPS, OXY, SAT (oxygen quantity and static air temp),
TRIM-PIT, AIL, and RUD (pitch, aileron and rudder trim).
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The NAVIGATION page provides the following data, when valid data
is available: NAV, ADF, CRS (selected course), DME (distance to tuned
station), bearing pointers for VOR and ADF, TO/FROM indication,
MARKER BEACONS, HEADING, lateral deviation (VOR and ILS),
and vertical deviation (GS only). The navigation displays on both
RMUs use AHRS#2 heading information and NAV information from
NAV 1 and ADF 1.
VHF COM TUNING
Normal operation of the RMU is with the radio tuning page displayed.
A section for COM is located in the upper left corner. The COM window displays two frequencies. The top line displays the active frequency of the COM, while the line below will display the preset frequency.
When pressing the line select key (preset frequency) adjacent to the
lower frequency, a yellow cursor encloses that frequency. This step is
not always necessary since the cursor normally “parks” over the preset
frequency box. Anytime the cursor has been moved to another area on
the main radio tuning page, it will automatically return to the COM
preset frequency after twenty seconds of inactivity on that page. When
the yellow cursor is enclosing the preset frequency, that frequency can
be changed by adjusting the tuning knobs. The preset frequency can
then be changed (flip-flopped) with the active frequency by depressing
the transfer key.
The storage function can be accomplished by pressing the STO (store)
key located at the bottom of the RMU. When the STO key is depressed,
the nomenclature below the preset COM frequency will change back to
MEMORY, and the digit following MEMORY will indicate in which
memory location the frequency is stored. With the main tuning page
displayed, the rotary tuning knob can be used to scroll through the frequencies stored in memory. As each memory location (channel) is selected, the stored frequency will be shown on the COM preset line
which can then be moved to the active position by depressing the transfer key.
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A TX will appear at the top of the COM window when the associated
radio is transmitting. Its purpose is to show that the transmitter is on
and to alert the pilot in case of a stuck microphone key. If not attended
to for approximately two minutes, a beep will sound on the audio and
a MIC STK annunciation will appear at the top of the COM window
until the mic button is released. Ten seconds after the MIC STK
annunciation appears, the selected transmitter will automatically turn
off. Depressing the SQ (squelch) function button at the bottom of the
RMU causes the COM radio to open its squelch and allows any noise
or signal present in the receiver to be heard. When selected, an SQ will
appear at the top of the COM window. Pressing the button a second
time closes the squelch.
FMS TUNING
The FMS interfaces with the RMUs for radio tuning. The FMS has a radio tuning page that can be used to control VHF COM, NAV, ADF and
transponder codes. If it is suspected that the FMS is interfering with
com/nav radio tuning, an FMS ENABLE/DISABLE selection on the
RMU NAV memory page can be toggled with the adjacent line select
key. The DISABLE selection will prevent tuning any of the radios
through the FMS CDU.
NAV TUNING
The format of the NAV window (top right corner of the RMU) is identical to the COM window in that the top frequency is active and the bottom frequency is the preset frequency. Pressing the line select key
alongside the NAV preselect window moves the cursor to that window.
This connects the tuning knobs to the NAV preset frequency. By pressing the NAV transfer key (top right) the preset and active frequencies
are exchanged. The preset frequency may be changed to a different frequency by using the tuning knobs or by pressing the line select key to
bring up the next frequency from memory. Selection of stored frequencies can be accomplished by pressing the line select key by the NAV
preset window until the tuning box encloses the memory mnemonic.
Rotating the tuning knob scrolls through the stored frequencies displaying them in the preset area.
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The memory functions and direct tuning operate the same as described
under COM operation, except the NAV window has an added function
called DME split tuning mode. Its operation is similar to the function
called DME hold. Depressing the DME function key on the RMU allows the DME frequency to be tuned independent of the NAV frequency. Depressing the DME once causes the NAV window to split into two
sections, the top one being the active VOR frequency and the lower one,
now labeled “DME”, containing the active DME frequency in VHF format. In this condition, the DME may be tuned directly by simply pressing the line select key to place the cursor box around the frequency and
retuning using the tuning knobs. The DME digital station identifier will
appear adjacent to the DME nomenclature on the top edge of the DME
window. An amber H (hold) appears in the lower DME window. This
indicates that the distance display (DME or TACAN) is not paired with
the VOR/ILS navigation data. When the H is displayed on the RMU, it
will also be displayed following the DME read-out on the PFD.
ADF TUNING
ADF operation is the same as COM and NAV tuning in that depressing
the line select key beside the ADF frequency will place the cursor over
the frequency to be changed. Rotating the small tuning knob slowly
will advance the frequency in 0.5 kHz steps. This change will increase
to 10 kHz steps when the large knob is used. The RMU has the capacity
to store one ADF frequency in memory. This is done by selecting the desired frequency, then depressing the STO function key at the bottom of
the RMU. To retrieve the stored frequency from memory, the frequency
line select key must be depressed for two seconds. ADF modes are also
controlled within the ADF window. Repetitively depressing the line select key adjacent to the ADF mode annunciator will step through the
available ADF modes of operation. This can also be accomplished by
placing the cursor over the mode annunciation, and using the tuning
knobs to step up or down through the available modes.
The ADF operating modes are as follows:
(1) ANT (Antenna) — ADF audio signal only.
(2) ADF — ADF receives signal and calculates the relative
bearing to the station.
(3) BFO (Beat Frequency Oscillator) — ADF adds a beat
frequency oscillator to detect continuous wave (CW)
signals.
(4) Voice — ADF has maximum audio clarity and fidelity,
but no bearing information.
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TRANSPONDER/TCAS TUNING
Transponder operation is similar to COM and NAV operation in that
depressing a line select key beside the function desired will move the
cursor to that location. Those aircraft without TCAS installed will have
an ATC legend at the top of the transponder window, and those
equipped with TCAS will have ATC/TCAS labeled above the window.
Either transponder 1 or 2 can be selected for use and controlled by
either RMU. A number 1 or 2 will appear in front of the transponder
mode in the ATC window on both RMUs indicating which transponder
has been selected. Transponder side selection is toggled by depressing
the 1/2 key on either RMU with the cursor anywhere within the ATC
window.
The transponder is switched from standby to an operating mode by depressing the line select key adjacent to the mode line. Once the cursor
has been selected, the mode line select key acts as a toggle to switch the
transponder between the standby mode and the active mode. Once the
transponder is in the ALT ON mode, the mode of operation is changed
using the tuning knobs. The active mode of operation can now be
changed by rotating the concentric tuning knobs. Depressing the ID
button of the RMU will initiate an approximate 18 second IDENT mode
on the transponder. This will also illuminate an ID annunciation along
the top edge of the transponder window. A reply annunciator is located
in the upper right corner of the ATC window.
TCAS (OPTIONAL)
The Traffic Alert and Collision Avoidance System (TCAS) provides the
crew with aural and visual indications of potentially dangerous flight
paths relative to other aircraft in the vicinity. The system uses the transponder to interrogate other transponder-equipped aircraft and determine their bearing, range, and altitude, if the intruder has an altitude
encoding transponder in operation. Advisories are issued to the crew
via the airplane Primary Flight Displays (PFDs), audio system and
Multi-Function Displays (MFDs) (traffic map). Two levels of TCAS are
in use today, TCAS I and TCAS II. TCAS II is the same as TCAS I with
the exception of providing Resolution Advisories (RAs) integrated
with the vertical speed indicator on the PFDs and additional aural commands through the audio system. There is no RA display on TCAS I
equipped aircraft. The TCAS system consists of a processor, two bearing antennas, and associated airplane wiring. System control is through
the radio management units. Power for the TCAS system operation is
28-vdc supplied through the 5-amp TCAS circuit breaker located on the
copilot’s circuit breaker panel (INSTRUMENT/INDICATIONS group).
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TCAS OPERATION
The Learjet 45 may be equipped with either the optional TCAS I or
TCAS II system. The controls and displays are integrated with the Honeywell Primus 1000 system. Controls are through the RMUs and
TCAS/annunciator displays are on the MFD and PFDs. On airplanes
equipped with TCAS II, the Resolution Advisories (RA) are integrated
with the vertical speed indicator display on the PFDs. The TCAS interrogates other aircraft transponders and analyzes the replies to determine range and bearing of the intruder. In addition, if the intruder’s
transponder is reporting altitude, the relative altitude is also determined. If the system predicts that safe boundaries may be violated, the
system issues a Traffic Advisory (TA) which is displayed on the MFD.
Should the TCAS II processor determine that a possible collision exists,
it issues visual and audio advisories to the crew to initiate appropriate
vertical avoidance maneuvers.
If an aircraft has a transponder, but does not have altitude reporting,
the TCAS will depict it on the TA display, but without the altitude information tag, and without the capability of providing evasive commands. TCAS II is capable of generating a TA display of traffic from
Mode A transponder-equipped aircraft, and it is also capable of generating RA signals to avoid Mode C-equipped aircraft. For similar
Mode S-equipped aircraft, the airplane’s TCAS II system coordinates
evasive maneuvers for both aircraft. TCAS I can process transponder
information from other aircraft equipped with Mode A, C, or S transponders, but does not receive altitude information to compute or coordinate a Resolution Advisory (RA). If the depicted traffic is reporting
altitude and is climbing or descending at a rate of at least 500 feet per
minute, a trend arrow is displayed beside the traffic symbol indicating
that the aircraft is climbing or descending. If the intruder is not reporting altitude, the traffic symbol appears without an altitude tag or trend
arrow. The RA displays are incorporated into the vertical speed indicator on the PFDs. Green FLY-TO zones and red NO-FLY zones are placed
on the vertical speed arc by the TCAS for collision avoidance. The zones
are not displayed on the arc until the TCAS detects an RA intruder and
computes the collision avoidance data. Synthesized voice commands
and announcements are issued by the TCAS over the airplane audio
system.
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SYSTEM CONTROLS AND DISPLAYS
Selection of the TA or TA/RA (TCAS II) modes is accomplished
through the transponder window on either RMU main radio tuning
page. After the cursor is placed over the transponder mode line, the desired mode is selected with the RMU tuning knob. The transponder selection options for TCAS I equipped aircraft will be STANDBY, ATC
ON, ATC ALT and TA. Selections available with TCAS II include the
same as TCAS I plus TA/RA. The selected TCAS mode will be annunciated in the top left corner of the TCAS display. The auto or manual
mode can be selected on the ATC/TCAS CONTROL PAGE of the RMU.
This page is accessed through the RMU PAGE MENU page. When
AUTO is selected, traffic targets display only when a TA or RA target
condition exists. When manual is selected, all traffic targets within the
viewing airspace are displayed. In either the MAP or PLAN format display, the TCAS TA display is selected by pushing the TCAS menu key
on the MFD Main Menu bezel controller. If the TCAS triggers an RA,
and TCAS display is selected OFF, the main menu is activated on the
MFD. This allows flight crew selection of TA displays with a single button push. This display is in addition to the resolution advisory on the
VSI display on the PFD (TCAS II only).
TRAFFIC DISPLAY SYMBOLS
TCAS I will display three different traffic symbols and TCAS II four
with the addition of Resolution Advisories (RA). The type of symbol selected by TCAS is based on the intruder’s location and closing rate. The
symbols change shape and color to represent increasing levels of
urgency. The traffic symbols may also have an associated altitude tag
which shows relative altitude in hundreds of feet, indicating whether
the intruder is climbing, flying level or descending. A + sign and
number above the symbol means the intruder is above your altitude.
A - sign and number beneath indicates it is below your altitude. A trend
arrow appears when the intruder’s vertical rate is 500 feet per minute
or greater. The symbology displayed on the PFDs and MFD is as follows:
(1) NON-THREAT ADVISORY (OA) TRAFFIC — An
open cyan diamond indicates that an intruder’s relative altitude is greater than ± 1200 feet, or its distance is
beyond 6 nm range. It is not yet considered a threat.
(2) PROXIMITY INTRUDER (PA) TRAFFIC — A filled
cyan diamond indicates that the intruding aircraft is
within ± 1200 feet and within 6 nm range, but is still
not considered a threat.
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(3) TRAFFIC ADVISORY (TA) TRAFFIC — A symbol
change to a filled amber circle indicates that the intruding aircraft is considered to be potentially hazardous.
Depending on your own altitude TCAS II will display
a TA when time to closest point of approach (CPA) is
between 20 and 48 seconds. An advisory voice message “TRAFFIC, TRAFFIC” may be heard through the
audio system.
(4) RESOLUTION ADVISORY (RA) (TCAS II only) — A
solid red square indicates that the intruding aircraft is
projected to be a collision threat. TCAS II calculates
that the intruder has reached a point where a resolution advisory is necessary. The time to closest approach
with the intruder is now between 15 and 35 seconds,
depending on your altitude. The symbol appears with
an audio warning and a vertical maneuver indication
on the PFD VSI.
ENHANCED GROUND PROXIMITY WARNING SYSTEM (EGPWS)
EGPWS is shown on the MFD by pushing the EGPWS menu button on
the MFD main menu. When TERRAIN is selected for display, the
EGPWS sends the terrain data directly to the MFD display unit via the
WX picture bus and replaces the WX display with terrain information.
If a potential terrain hazard is sensed by the EGPWS, terrain data automatically pops up on the MAP. This “pop-up” mode defaults to the
10 NM range. EGPWS annunciators are described in the table below
and are displayed in the upper left corner of the MFD.
Annunciator
TERR INHB or
TERR INHIB
(inhibit)
TERR FAIL
TERR TEST
TERR N/A
TERR
Description
TERRAIN displays and aurals associated with terrain are inhibited (annunciation in white)
The TAWS is inoperative.
EGPWS is in test mode.
TERRAIN map not available.
TERRAIN map selected for display.
The terrain data is displayed above the airplane symbol on the MFD in
green, yellow, and red colors that define the elevation of the terrain relative to the airplane’s current altitude. Terrain that is more than 2000
feet below the airplane is not included in the display.
A moving marker scrolls across the bottom of the EGPWS display as an
indication that the terrain is display is operational.
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AUDIO CONTROL SYSTEM
The pilot’s and copilot’s digital audio control panels are located outboard of the PFDs on each side of the flight deck. Microphone transmit
selector buttons are located in a row along the top edge of this panel
(Figure 5-16). For night flying operations, the microphone selector buttons are annunciated with a lighted bar on the switch indicating the selected microphone. When these latching buttons are pushed, they
connect the microphone (hand-held microphone, boom microphone, or
oxygen mask microphone) to the selected radio. They simultaneously
enable the audio associated with that radio, regardless of the setting on
the audio on/off buttons located below these microphone transmit selector buttons. The microphones selector buttons are mechanically interlocked so that each new selection automatically deselects the
previous selection. Depressing the PA button connects the on-side microphone to the passenger address amplifier. The audio level for the PA
is automatically adjusted for conditions and cannot be adjusted by the
crew. The pilot will use either a hand-held microphone or boom microphone for transmissions. Oxygen mask microphones are used when the
MIC/MASK selector is in the extended (unlatched) position.
An EMER switch is located in the upper right corner of each audio control panel. When the EMER switch is depressed, the microphone and
audio reception is connected directly to VHF 1 and NAV 1 and all functions of the audio control panel are bypassed except the headphone volume. In order to receive NAV 1 audio with EMER selected, the NAV
AUDIO switch on the clearance delivery radio must be selected ON.
When EMER is selected on the audio control panel and power is available to the control units, COM and NAV frequencies are set using either
RMU or the clearance delivery radio. If EMER is selected and electrical
power is still available to the audio panel, system warning audios will
still be available through the cockpit speaker and audio will be routed
to the cockpit voice recorder. Regardless of whether power is lost to the
audio control panel, the EMER switch is operational, however, system
warning audio and audio to the cockpit voice recorder are inoperative.
The audio source selector controls are located on the lower rows of the
audio control panel. When these push-on/push-off switches are
latched (in position) audio is turned off from that receiver. When unlatched (out position), the audio associated with that button is connected to the headphone and also to the speaker, if it is selected on. The
audio level can be adjusted by rotating the button, counterclockwise to
decrease, and clockwise to increase the volume.
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One knob, labeled DME, controls the audio reception for both DME 1
and DME 2. When the DME knob is unlatched (out position) and the
arrow on the knob is centered straight up, the audio level is at a minimum. Rotating the control knob in either direction toward 1 or 2 will
increase the volume for that corresponding channel only. The audio
level pointers on the knobs are displayed for night flight. There are separate controls for speaker volume and headphone volume which adjust
the volume level for all audio buttons selected. The speaker push-on/
push-off selector is combined with the sidetone knob. When the speaker switch is extended, it turns on audio to the on-side speaker. The
speaker sidetone audio is controlled by the speaker SIDETONE volume
control and the SPEAKER volume control for both on-side and off-side
transmit conditions.
SPKR
ON
DIGITAL AUDIO CONTROL PANEL
Figure 5-16
The ID/BOTH/VOICE switch is located on the right side of the audio
panel. In the ID position, the VOR and ADF audio is filtered to enhance
the Morse Code identification and eliminate the voice signal. In the
VOICE position, the ident audio is filtered to pass the voice content
only and in the BOTH position, voice and ident signals will be heard
simultaneously.
The controls for the marker beacon receiver are located at the bottom of
the audio panel. They include the marker audio volume control (MKR),
marker sensitivity control (LO SENS/HI SENS) and marker mute control (MUTE). The sensitivity is controlled by the rotation of the MUTE
control. If either audio panel MKR sensitivity control is set LO, then
both MKR receivers are set to LO, regardless of the position of the other
audio panel controls. Either pilot can temporarily mute the marker beacon receiver by depressing the MUTE/HI/LO switch.
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The INPH (interphone) volume control adjusts the on-side headset audio level when the interphone function is used. The interphone operates on a “hot microphone” basis. The interphone is not available over
the cockpit speaker except when the oxygen mask audio is selected.
The MIC/MASK control allows for microphone audio switching between the boom/hand-held microphone (MIC) and the oxygen mask
microphone (MASK). When the switch is latched (depressed position),
MIC is selected and when the switch is unlatched (out position), MASK
is selected.
The MASK intercom feature provides interphone audio to the on-side
cockpit speaker. Audio is available regardless of the SPKR ON/OFF
button position. Selecting INPH allows adjustable volume control of
the off-side MASK intercom on the speaker.
Warning system audio signals are input to the audio panel for dissemination to the flight crew over the headphones and speaker. The audio
output from the headphone, speaker, and microphone are recorded by
the Cockpit Voice Recorder (CVR). The CVR microphone is the input
for the AGC circuit and if the CVR microphone becomes disabled, or
the CVR circuit breaker is pulled, then the aural warnings will be at the
fixed HIGH volume level. If the Crew Warning Panel (CWP) has detected a fault in any one of the audio output channels, or in the Automatic
Gain Control (AGC) input, a CAS annunciation will be posted.
The following CAS illuminations are specific to the Crew Warning Panel audio:
CAS
WARN AUDIO
WARN AUDIO
Color
Description
Amber The audio function of the Crew Warning
Panel (CWP) has failed.
White • A Crew Warning Panel (CWP) audio output
channel fault is detected.
or
• A problem exists with the Automatic Gain
Control (AGC).
The pilot’s audio panel receives 28-vdc from the left essential bus and
is protected by a 3-amp circuit breaker labeled AUDIO 1/CLR DLY on
the pilot’s circuit breaker panel (AVIONICS [COMMUNICATIONS]
group). The copilot’s audio panel receives 28-vdc from the right essential bus and is protected by a 3-amp circuit breaker labeled AUDIO 2 on
the copilot’s circuit breaker panel (AVIONICS [COMMUNICATIONS]
group). The passenger address amplifier receives power from the left
essential bus and is protected by the CABIN PA 5-amp circuit breaker
on the pilot’s circuit breaker panel (AVIONICS [COMMUNICATIONS]
group).
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CLEARANCE DELIVERY RADIO (CDR)
The Clearance Delivery Radio (CDR), Figure 5-17, is located on the
right, front corner of the center pedestal. The CDR provides an alternative capability for tuning the VHF COM 1 transceiver and the VHF
NAV 1 radio. The CDR will tune the VHF COM radio prior to applying
electrical power to the airplane. The CDR control head is normally
powered by the left essential bus through the AUDIO 1 circuit breaker;
however, it and other communication related equipment can be powered from the right forward hot bus, prior to applying electrical power
to the aircraft.
With airplane batteries OFF, depressing the momentary action RADIO
CTL HOT BUS switch on the center pedestal applies power from the
right hot bus to the left audio control panel, CDR control panel, COM
section of the integrated communications unit, and NAV section of the
integrated navigation unit. The display on the CDR is liquid crystal
type with white letters on a black background. The push button, display, and control identifier legend are on a black background displayed
with electroluminescent lighting.
In the emergency mode, RMU and FMS tuning capabilities are inhibited and the COM and NAV units are tuned exclusively by the CDR. An
AUX ON caption replaces the NB or WB annunciators on the RMUs to
indicate that tuning through the RMUs is inhibited. Tuning through the
CDR is no different when EMRG is selected, but the CDR does not look
at the radio bus data to check the echoed frequency.
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123 RMT
COM
NAV
NAV AUDIO
PRE
NAV
TX
SQ
SQ
E
M
R
G
NAV
AUDIO
F40-050000-517-01
MODES
F40-050000-517-01
CLEARANCE DELIVERY RADIO
Figure 5-17
The CDR controls are as follows:
• Transfer Key — Alternately selects either the COM (top) or
NAV (bottom) frequency to be connected to the tuning
knobs.
• Tuning Knobs — Used to change the frequency indicated
by the tuning cursor.
• Normal/Emergency Mode Switch — This rotary knob provides alternate selection of Normal and Emergency modes.
• NAV AUDIO On/Off Switch — This alternate action push
button switch is used to toggle NAV audio ON or OFF
when in the EMER audio mode on the audio control panel.
• Squelch (SQ) Switch — Used to toggle COM squelch on or
off.
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FLIGHT GUIDANCE CONTROL SYSTEM
The Primus 1000 system includes an autopilot, yaw damper and dual
synchronized flight directors. These are all co-located in two IC-600 display flight guidance computers located in the nose avionics bay. Each
IC-600 houses a flight director; however, only the copilot IC-600 is connected to the pitch, roll and yaw servos for the autopilot/yaw damper
and rudder boost functions. A flight guidance controller, located in the
center of the glareshield, provides the means of engaging the autopilot/yaw damper and controlling both Flight Director (FD) systems. It
also contains a transfer (XFR) switch that allows the crew to select either the left or right flight director as master and for autopilot coupling.
The autopilot is a single channel with a fail passive design. The monitor
system provides for automatic disconnect in the event of a malfunction
in autopilot, yaw damper, or rudder boost. All automatic disconnects,
which result from monitor trips, will be stored in a non-volatile memory for later recall by technicians.
FLIGHT DIRECTOR
The flight director system, utilizing two separate IC-600 computers,
provides dual flight director computations, either of which can be coupled to the autopilot. Only one flight director can be coupled to the autopilot at a time. In this case the coupled flight director is classified as
the master flight director, and the other flight director is classified as the
slave flight director. The flight director can couple to Short Range Navigation (SRN) units (e.g. VOR, ILS), dependent upon which SRN is being displayed as a NAV data source on the Primary Flight Display
(PFD). The flight director will use the displayed SRN data on the associated PFD for command-bar control computations. The flight director
can be coupled to optional Long-Range Navigation (LRN) units (e.g.
flight management system) if it is installed and selected as the NAV
data source and displayed on the associated PFD. The flight director
utilizes the lateral and/or vertical steering commands from the LRN in
the command control computations. Each flight director uses the displayed on-side Air Data Computer (ADC) data for all vertical modes
and gain programing. The flight director modes are synchronized in a
manner that allows selection of the modes to be accomplished by the
single set of FD mode select buttons on the flight guidance controller
(FGC), see Figure 5-18. The FD mode select buttons on the FGC panel
are momentary action and each has a vertical green bar that illuminates
whenever the mode is selected.
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FLIGHT GUIDANCE CONTROLLER (FGC)
The Flight Guidance Controller (Figure 5-18) is the prime controller for
the flight director and the autopilot/yaw damper. Located on the center glareshield, the FGC provides the means, via push button switches,
for flight director mode selection, couple status and autopilot/yaw
damper engage selection. Flight director modes engage status is indicated to the crew via a green light on the right edge of each mode
switch, which is illuminated when the mode is active and extinguished
when the mode is inactive or dropped. The controller also has several
rotary controls enabling selection of IAS, MACH, and VS targets, altitude, course and heading. All push button selections are signaled to
both IC-600s via a discrete output from the FGC. The IC-600s then provide the drive to illuminate the appropriate light on the FGC.
FLIGHT GUIDANCE CONTROLLER
Figure 5-18
The FGC annunciations and controls are as follows;
FD 1/2 buttons — The flight director buttons (FD 1 and FD 2) are located on the upper left/right corners of the flight guidance controller. Depressing these buttons alone will not bring the FD command bars into
view. Any FD mode selection causes the FD command bars to appear
on both PFDs. When the FD command bars are in view on both PFDs
and the autopilot is not engaged, depressing the master side FD button
will disengage all FD modes and remove the command bars from both
sides. Pressing the slave side FD button will remove the command bars
from the PFD on that side acting as a flight director clear function.With
the autopilot engaged, the FD command bars will be in view at all times
on the coupled side and cannot be removed from the PFD. The opposite
side FD command bars can be removed from view by depressing the
appropriate FD 1 or FD 2 button.
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Course set knobs — A course set knob (CRS 1 and CRS 2) is located at
each end of the FGC. These knobs are used to individually set the
courses on the left and right PFD HSI displays. They are used primarily
to set the course for a VOR radial or LOC course. The course knobs
have a push button in the center to synchronize the display to the aircraft’s “direct-to” course.
Heading set knobs — Heading is selected via a rotary knob, with a
“Heading Bug” symbol on the face of the knob. The heading knob controls the heading bug and digital display on both PFDs and the bug on
the MFD MAP display. Depressing the HDG knob will synchronize the
HDG bug on all display units to the aircraft’s current heading.
HDG (heading) button — Depressing the HDG button engages the
heading mode and displays a green HDG annunciation on the PFDs.
The flight director command bars will command a turn in the direction
the heading bug was moved to achieve the set heading. Heading select
is used to maintain a magnetic heading. The heading bug is positioned
to the desired heading on the HSI using the HDG knob on the FGC. The
heading select mode is canceled when any armed lateral mode captures
or if GA is selected.
NAV (navigation) button — Pressing the NAV button alternately selects and deselects the navigation mode. The NAV mode is normally
used to intercept route segments identified with VOR radials and to intercept and fly desired FMS tracks (SIDs, routes, holding and STARS).
APP (approach) button — The intended function of the APP mode is
that APP be used for all approaches, regardless of nav source or whether a vertical mode is also associated with the approach. The APP mode
is normally used to select lateral and vertical steering for ILS and FMS.
The VOR approach mode is selected by pressing the APP mode button
with the navigation receiver tuned to a VOR frequency and selected as
the active nav source. Pressing the APP button arms both localizer and
glideslope modes when the navigation receiver is tuned to an ILS frequency and ILS is selected as the active navigation source. Selection of
APP mode when the nav source is FMS engages the FMS lateral mode
the same as described for NAV and also arms VNAV for approach.
BNK (bank) button — Pressing this button alternately selects or deselects a reduced maximum bank angle of 14° (for all lateral modes, except roll) on both FDs. When selected, a green low bank arc appears on
the top of the ADIs and BNK is annunciated on the PFD.
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AP (autopilot) button — Depressing this button engages the autopilot.
Depressing a second time disengages the autopilot.
XFR (transfer) button — Located in the center section of the FGC. The
XFR button is used to select the desired flight director (left or right) to
command the autopilot.
YD (yaw damper) button — Depressing this button engages yaw
damper. The YD can be engaged independent of the AP, but the autopilot system will not engage, or remain engaged, without the YD.
SPD (speed) knob — The rotary SPD knob is used to change the IAS/
Mach speed reference (SPD mode) and the vertical speed reference (VS
mode). The speed knob changes the bug airspeed at any time, as long
as VS is not selected. When VS mode is engaged, rotation of the SPD
knob changes the digital vertical speed reference and the vertical speed
bug position. The integral PUSH IAS/M button within the SPD knob is
used to toggle the airspeed tape between IAS and Mach. The master
flight director computes the airspeed reference, and the slave flight director synchronizes to this reference.
SPD (speed) button — Depressing the SPD button engages the speed
hold mode (IAS or Mach) on both FDs. The speed select mode is used
to fly to a selected airspeed or Mach number, and to provide limited
overspeed/underspeed protection during climbs and descents. When
speed select mode is active, a green IAS or Mach annunciation is displayed in the captured vertical mode field on the PFDs.
FLC (Flight Level Change) button — Depressing the FLC button once
engages the normal climb/descent profile on both PFDs. Depressing it
a second time selects the high speed climb/descent schedule. A third
depression deselects the mode. The FD chooses between the climb and
descent schedule based upon the aircraft’s present altitude and preselected target altitude. The FD annunciation on the PFD is FLC for the
normal profile and FLCH for the high speed profile.
VS (Vertical Speed) button — Depressing the VS button engages the
vertical speed hold mode on both FDs. When VS is selected, the speed
bug disappears and reference goes to dashes.The FD commands pitch
changes to hold the vertical speed that existed at the time of engagement. Once engaged, the vertical speed bug positions on the inner side
of the vertical speed scale and a digital readout appears above the vertical speed indicator.
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VNV (Vertical Navigation) button — Depressing the VNV button
arms, then captures the FMS pitch steering commands of the FDs if
FMS is selected as the NAV source, the FMS is programmed for a vertical navigation profile, the altitude preselector is set below existing altitude and the aircraft is within the TOD (top of descent) window. When
the VNAV mode is armed, a white VNAV is annunciated on the PFDs
in the FD vertical mode annunciation field and will turn green upon
capture.
ASEL (Altitude Select) knob — The preselected altitude is set via the
ASEL rotary knob on the FGC. The altitude preselect mode provides a
means for FD/AP to climb or descend to a preselected altitude and
then level off and maintain the preselected altitude. The ASEL knob is
used to set the altitude preselect function, and also provides the altitude reference for the altitude alerter function.
ALT (Altitude Hold) button — Altitude hold may be engaged by depressing the ALT button on the FGC. When ALT is engaged, the FD
commands pitch to hold the existing altitude at the time the ALT button
was depressed, or at the ASEL reference altitude if ALT automatically
engages.
AUTOPILOT/YAW DAMPER
The autopilot is a single-channel autopilot which may be coupled to
either flight director. The autopilot function is contained within the #2
IC-600 located in the nose. The IC-600 will fly the aircraft based on the
selected control mode and guidance inputs from the coupled flight director, via servo control of the elevator, aileron and rudder. When engaged, the autopilot will also automatically command trim changes as
required to alleviate the aerodynamic loading on the elevator (pitch
trim), and will allow control surface commands to be entered via the
control wheel trim switches for the ailerons and elevator. There is also
a Touch Control Steering (TCS) feature which allows the cockpit crew
to manually maneuver the aircraft with the autopilot engaged when
the TCS switch is pressed. The autopilot provides aircraft control in response to pitch and roll steering commands from either flight director.
The yaw damper software computes servo commands based on sensor
input data only. the yaw damper control software provides yaw rate
damping that holds rudder force to zero. The servo position reference
is synchronized to zero at engagement and is constantly washed out to
ensure that the steady-state rudder forces are zero. The yaw damper
can be engaged independent of the autopilot but the autopilot cannot
be engaged without the yaw damper.
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AP/YD Annunciation
At the very top, center section of the PFDs is an area dedicated to flight
director and autopilot annunciations. A horizontal arrow appears at
the top center of each PFD between the flight director vertical and lateral mode annunciators. This arrow points left or right, as selected on
the FGC XFR switch, to indicate which flight director the autopilot will
couple to when engaged. It also indicates which flight director has priority. Just below the arrow is a line reserved for autopilot and yaw
damper annunciation. A green AP and YD appear in this area when the
autopilot and/or the yaw damper is /are engaged.
Control Wheel Trim Switch
Manual primary pitch trim or roll trim commands are initiated by
pressing and holding of the arm switch and actuation of the control
wheel trim switch. Pressing and holding of the arm switch with control
wheel trim switch input will result in immediate autopilot disengagement.
The yaw damper and flight director modes are not affected by manual
pitch or roll trim commands. Actuation of the control wheel trim switch
without pressing the arm switch allows manual autopilot commands.
During autopilot engagement, basic attitude commands (pitch and roll)
can be entered through either the pilot’s or copilot’s control wheel trim
switch (dependent upon which flight director is coupled to the autopilot). With the autopilot engaged, activation of the control wheel trim
switch (without arm switch depressed), on the side coupled to the autopilot, causes the flight director roll and/or pitch hold mode to be
activated.
Control Wheel Master Switch (MSW)
The MSW switch immediately disconnects all autopilot and yaw
damper servo drives. The selected flight director modes are not affected. In normal operation, a 28-vdc signal is routed through the normally
closed contacts of each MSW and then to the onside IC-600. This input
to the IC-600s is the autopilot disengage discrete, and if 28-vdc is removed from this discrete on either IC-600, the autopilot will disengage.
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Touch Control Steering (TCS)
TCS allows the pilot to manually fly and retrim the aircraft without disengaging the autopilot. To use the TCS function, the pilot will press the
TCS button on the control wheel, maneuver the aircraft to the desired
condition, and release the TCS. Operation of the TCS button has no effect on flight director mode of operation. While the TCS button is held
depressed (AP engaged), a white “TCS ENG” annunciator appears in
the autopilot display area at the top of the PFDs.
Autopilot Engagement/Disengagement
Engagement of the autopilot is achieved via the AP push button on the
FGC. Each button has a vertical green bar that illuminates when engaged. Engagement of the autopilot will cause the yaw damper to automatically engage. When engaged, the autopilot will couple to the
master flight director, and will follow the guidance commands from the
master FD. If no flight director is active, engagement of the autopilot
will cause the master FD to default to the basic PIT and ROL modes.
When engaged, the appropriate annunciation will be provided on both
PFDs, and the green bar on the AP push button will be illuminated.
Disengagement of the autopilot via the AP switch will cause the AP
annunciation to be removed from both PFDs and the green bar on the
AP push button will extinguish. Other actions that will cause autopilot
disengagement include:
(1) Control wheel master switch activation.
(2) Pilot initiated trim commands with control wheel trim
switch depressed.
(3) Yaw damper disengagement.
(4) AP switch on the FGC disengaged.
(5) Autopilot primary or secondary monitor trip.
For a normal disconnect, AP flashes red for 5 seconds, then is removed.
For a monitored disconnect, it flashes red for 5 seconds, then steady,
and the aural alert is continuous until the crew cancels it with the MSW.
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Yaw Damper Engagement/Disengagement
Selection of the autopilot via the AP push-button will automatically engage the yaw damper. Alternatively the yaw damper may be selected
via the YD push button on the FGC. When engaged, the green YD annunciation will be provided on both PFDs, and the green bar on the YD
push button will illuminate. Manual disengagement of the yaw damper via the YD button will cause the YD annunciation to flash amber then
be removed from both PFDs and the green bar on the YD push button
will extinguish. In the case of a monitored disconnect, the YD annunciation will turn amber and flash for five seconds and then remain steady.
Other actions that will disengage yaw damper include:
(1) Control wheel master switch activation.
(2) YD Switch on the FGC disengaged.
(3) Yaw damper monitor trip.
Mistrim Annunciation
When the autopilot is engaged and the roll or pitch servo remains energized for a longer than normal period, this condition will be annunciated with a CAS. The autopilot does not have a capability to trim in
the roll axis; therefore, if there is a mistrim in the roll axis, this will also
display a CAS.
The following CAS illuminations are specific to the autopilot:
CAS
AP AIL MISTRIM
AP ELEV MISTRIM
AP ELEV MISTRIM
Color
Description
Amber Autopilot is engaged and autopilot aileron
servo is holding excessive torque. Disengage
autopilot.
Amber Autopilot is engaged and autopilot elevator
servo is holding excessive torque. Disengage
autopilot.
White Autopilot is engaged and autopilot elevator
servo is holding torque. Disengage autopilot.
Power Supply Configuration
The power supply for all the AFCS servos is supplied from the L MAIN
bus, and is routed through the #2 IC-600 to the servos. The system is designed so that power failure to any major component in the system will
result in the system reverting to the safe, disconnect mode. The AFCS
SERVOS circuit breaker is located in the FLIGHT group of the pilot’s
circuit breaker panel.
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GO-AROUND (GA) BUTTON
The GA button is located on the outboard side of the left thrust lever.
Selection of the GA button disconnects the autopilot (but not the yaw
damper) and cancels all other vertical and lateral modes except automatic altitude preselect arm. The flight director provides a wings level
command bar display in the lateral axis and a fixed pitch-up vertical
command. The pitch command does not guarantee that the go around
airspeed will be achieved. If used on takeoff, the pitch attitude will not
guarantee achieving the V2.
FLIGHT MANAGEMENT SYSTEM (FMS)
The UNS-1E is a fully integrated Flight Management System (FMS).
The FMS provides centralized navigation sensor control, flight planning, lateral and vertical flight guidance, steering enroute, terminal and
approach modes of operation. Database management, fuel management, and maintenance functions are also provided by the FMS.
The UNS-1E accepts position information from up to five long-range
navigation sensors as well as DME, VOR, or TACAN sensors. The data
from these sensors is used to determine the best computed position.
The UNS-1E incorporates an internal GPS sensor with Receiver Autonomous Integrity Monitoring (RAIM) as a standard part of the FMS configuration. The FMS interfaces with the IC-600s for a transfer of
information to the FMS and lateral and vertical steering commands
back to the EFIS and FD/AP.
If a single FMS is installed, it can receive ADC, AHRS, EFIS, and AP
data from either IC-600 but will use the #1 IC-600 as the primary source.
If dual FMSes are installed, FMS 1 will use the #1 IC-600 as the primary
source and FMS 2 will use #2 IC-600 as its primary source. With dual
FMS installation, the pilots can use either FMS as the navigation source
on their PFDs/MFD. In the material that follows, the FMS unit will be
referred to as the Control Display Unit (CDU). The CDU contains a color, flat panel display, ten Line Select Keys (LSK) and dedicated function
keys.
The FMS is powered from the left essential bus by a 5-amp circuit
breaker labeled FMS located on the pilot’s circuit breaker panel (AVIONICS [NAVIGATION] group).
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CONTROL DISPLAY UNIT (CDU)
The CDU (Figure 5-19) is the primary interface to the pilot. It provides
keypad input for selection of NAV modes and entering of waypoints
and a display to indicate current operational modes. The CDU contains
all the components required for the FMS functions. Other functions that
are provided include:
• Radio tuning and RMU interface — The CDU allows
remote tuning of the VHF COMM, VOR/ILS, ADF and
ATC radios.
• Flight planning — The CDU can store a fixed number of
flight plans into the FMS database.
DATA
NAV
VNAV
DTO
LIST
FUEL
FPL
PERF
TUNE
MENU
PREV
NEXT
1
2
3
4
5
6
A
B
C
D
E
F
G
7
8
9
H
I
J
K
L
M
N
BACK
0
MSG
O
P
Q
R
S
T
U
V
W
X
Y
Z
ON/OFF
DIM
ENTER
UNS-1E FMS CDU
Figure 5-19
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DATA TRANSFER UNIT (DTU)
The DTU allows updating of the FMS database, storing of pilot data
and can also be used to record flight data. Database updates can be
obtained on a subscription basis. The data transfer unit is a drive (e.g.
3.5-inch floppy or zip disk) designed for mounting in the aircraft.
CONFIGURATION MODULE
The configuration module is used to store configuration data that is
specific to the aircraft in which the FMS is installed.
At start-up, the CDU checks the version number stored in FMS memory against the version number stored in the configuration module. Any
discrepancy between version numbers results in an FMS message,
“CONFIG UPDATE REQUIRED”, on the CDU.
FMS FUNCTIONS
FLP (flight plan) — Before using the FMS for navigation, an active
flight plan must be defined within the FMS. The operator may select a
previously stored route or create a new one to load as the active flight
plan. A route or active flight plan can be created on the FMS as well.
Once the departure airport is identified, the UNS-1E will present tailored lists from which the current runway, SID and transition can be selected. Also, both low and high altitude airways can be accessed for
route or flight plan creation using the LIST function. Routes or the flight
plan may also be constructed waypoint by waypoint, or by combining
stored route segments. When en route and nearing the destination, a
progression of smart prompts similar to those used on departure may
be utilized to input a STAR, the approach and landing runway.
Upon selection of NAV on the MFD (FMS MENU), the IC-600 will display the closest eight navaids (VORs or NDBs) received from the FMS
on the MFD MAP as background data. Selection of APT on the MFD,
will result in the IC-600 displaying the first four airports received from
the FMS on the MFD MAP as background data.
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FUEL — Before takeoff, the fuel quantity signal conditioner provides
the FMS with fuel quantity on-board. The pilot must accept or change
the transmitted fuel quantity on the CDU fuel page. After engine start,
the FMS receives real time fuel flow information independent of the aircraft indicating system. APU fuel is not included in fuel flow. Specific
range and endurance are provided along with fuel, time and distance
predictions for the destination.
PERF (performance) — A performance program in the FMS can compute takeoff speeds, takeoff N1 , takeoff distance and landing speed.
The operator enters pertinent data such as takeoff configuration and
environmental conditions mostly through menu selections. V Speeds
and balanced field data is calculated and displayed. For landing, Vref
is calculated along with approach speeds for different flap settings.
NAV (navigation) — All pertinent en route navigation data is displayed on the first NAV page of the CDU. This page along with the
PFD/MFD displays provide complete integrated real time information
on flight progress.
DTO (direct to) — A dedicated function key can be used to navigate
from present position directly to any point on or off the present flight
plan.
VNAV (vertical navigation) — Vertical navigation pages allow the operator to define waypoints with altitudes or flight levels. Features such
as computed top-of-descent, target vertical speed and vertical direct-to
are included. The FMS outputs vertical commands to the flight director
when selected.
HOLDING PATTERNS — Holding patterns may be programmed at
any waypoint on or off the flight plan or stored in the navigation data
base as part of a SID, STAR or approach procedure. The holding pattern
page provides a graphic depiction of the holding pattern. The pattern
is defined with some crew inputs, and when the ACTIVATE line select
key is pressed, the aircraft will proceed from its present position directly to the fix and make the appropriate entry (direct, parallel or teardrop), all automatically calculated and flown by the FMS.
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WEATHER RADAR
The Primus 650 is the standard radar installed in the Learjet 45. It is an
X-band radar system designed for weather detection and analysis. The
radar can also be used for ground mapping. The WU-650 is an integrated unit which incorporates the receiver, transmitter and antenna (RTA)
into a single unit, located in the nose of the aircraft. The only remaining
radar component is the cockpit control panel which is mounted on the
center pedestal. The antenna is a 12-inch flat plate that is stabilized with
inputs from the #1 AHRS. Weather patterns can be displayed on both
PFDs, on the ARC or MAP format, and on the MFD MAP format. The
radar generates high-level RF pulses and should be operated with caution while on the ground. When operating on the ground, position the
nose of the airplane so that the antenna scan sector is free of large metallic objects such as hangers or other aircraft for a distance of at least
100 feet.
OFF
STAB
RCT
PULL
VAR
WX
SBY
GMAP
FP
OFF
MIN
MAX
GAIN
SECT
TGT
+
0
TST
15
-
RADAR
SLV
TILT
WU-650 WEATHER RADAR CONTROL PANEL
Figure 5-20
The weather radar is controlled from the WC-650 weather radar controller, Figure 5-20. WX is selected for display on the PFDs by selecting
the WX button on the display controller, the HSI automatically switches
to the ARC mode. WX is selected for display on the MFD MAP by depressing the WX bezel button on the MFD main menu.
A six-position rotary switch is provided on the radar control panel for
selecting the different radar modes. They are:
• OFF — removes electrical power from the system.
• SBY (standby) — in this position the RTA is powered up but
does not radiate any RF energy nor does the antenna scan.
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• WX (weather) — selects the weather radar main operating
mode.
• GMAP (ground map) — in the ground mapping mode, the
system internal parameters are set to enhance returns from
ground targets.
• FP (flight plan) — selecting this position places the radar in the
flight plan mode.
• TST (test) — when this mode is selected the weather depiction
will be a special colored test pattern to allow verification of system operation.
Power is provided to the RTA and cockpit controller from the right avionics main bus with a 7.5-amp circuit breaker located in the INSTRUMENT/INDICATIONS group of the copilot’s circuit breaker panel.
AVIONICS COOLING
INSTRUMENT PANEL COOLING
The instrument panel cooling system is located forward of the throttle
quadrant and is provided to draw flight deck ambient air through the
instrument panel and thus prevent overheating of avionics displays
and instruments. The system consists of an avionics cooling fan, an on/
off thermostat switch and an overtemperature thermostat circuit. The
avionics cooling fan is activated when the temperature sensing switch
reaches 90° F (32° C). The fan automatically turns off when the temperature has been reduced below 70° F (21° C). If the temperature reaches
an extreme of 135° F (57° C) the overtemperature circuit is energized
and an annunciation is posted on the CAS. The CAS remains displayed
until the temperature has been reduced to 125° F (51.7° C).
The following CAS illumination is specific to the avionics cooling
system:
CAS
INSTR PNL TEMP
Color
Description
White The temperature at the instrument panel is
higher than normal.
The instrument panel cooling system is powered from the L MAIN bus
and protected by the INSTR FAN 3-amp circuit breaker located on the
pilot’s circuit breaker panel (ENVIRONMENTAL group).
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MISCELLANEOUS
COCKPIT VOICE RECORDER (CVR)
A solid state Cockpit Voice Recorder (CVR) is installed in the Learjet 45.
The standard CVR is a three-channel unit providing 30 minutes of recording. An optional unit is available which provides 120 minutes of
recording. Two of the channels are used to record pilot and copilot
audio. The third channel is used for the area microphone. Located in
the tailcone, the CVR is painted international orange with reflective
tape added to aid in recovery following a mishap. It also has an underwater locator beacon installed on one end of the unit. The recording is
converted to a digital format and stored in crash protected memory.
The area microphone is located in the upper center area of the instrument panel. An erase button and headphone jack are located on the
CVR panel just beneath the copilot audio control panel. The CVR performs a self-test at power-up and has a continuous self monitor. If a
fault is detected at any time, an annunciation is posted on CAS.
The following CAS illumination is specific to the CVR:
CAS
CVR FAIL
Color
Description
White The cockpit voice recorder has failed.
The erase function is initiated by pressing the erase button on the CVR
panel. An interlocking device only allows this function to work when
the airplane is on the ground and parking brake is set. When erase function is complete, a three-second tone is output to the headphone jack.
Voice recorder system power is 28-vdc supplied through a 3-amp CVR
circuit breaker located on the pilot’s circuit breaker panel (INSTRUMENTS/INDICATIONS group).
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CLOCKS
Each instrument panel is equipped with a multi-function chronometer
to display GMT, local time (LT), flight time (FT), and elapsed time (ET).
The SEL button selects what is to be displayed and the CTL button controls what is being displayed (Figure 5-21). Pressing SEL sequentially
selects GMT, LT, FT or ET for display. FT starts counting when the main
gear weight-on-wheels switches transition to the air mode and stops
counting when they transition back to ground mode. The CTL button
resets FT back to zero when held down for three seconds. ET is started
and reset when the CTL button is pushed momentarily. Depressing the
SEL and CTL buttons simultaneously enters the set mode and GMT or
LT can be set. The CTL button is then pressed to increment the flashing
digit to the desired value. Pressing the SEL button enters that value and
toggles to the next digit to be set.
F40-05000-521-01
Power for the chronometers is 28-vdc supplied through a 1-amp L and
R CLOCK circuit breaker located on pilot’s and copilot’s circuit breaker
panels respectively (INSTRUMENT/INDICATIONS group).
M850
CHRONOMETER
GMT
PILOT
LT
FT
ET
SELECT
CONTROL
SEL
CTL
COPILOT
F40-050000-521-01
MULTI-FUNCTION CHRONOMETER AND INSTRUMENT LOCATION
Figure 5-21
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HOURMETER-AIRCRAFT (OPTIONAL)
An optional hourmeter may be installed to measure aircraft accumulated time. The typical location for the hourmeter is on the RH side of the
pedestal, just aft of the SELCAL decoder. It is wired to the right hand
main gear weight-on-wheels switch, through a switch in the lower
entry door frame. It will measure accumulated time as soon as the
aircraft lifts off. The hourmeter receives 28-vdc from a 1-amp HOURMETER circuit breaker located in the INSTRUMENTS/INDICATIONS
group of the copilot’s circuit breaker panel.
FLIGHT DATA RECORDER (FDR) (OPTIONAL)
The flight data recorder, which may be installed, is a 25-hour SolidState Flight Data Recorder (SSFDR) with Underwater Locator Beacon
(ULB) and remote mounted tri-axial accelerometer.
The following CAS illumination is specific to the flight data recorder:
CAS
FDR FAIL
Color
Description
White The flight data recorder has failed.
EMERGENCY LOCATOR TRANSMITTER (OPTIONAL)
Dorne & Margolin ELT14
The Dorne & Margolin ELT14 system simultaneously transmits distress
signals on the frequencies of 121.5 and 243.0 MHz. The system will automatically activate under emergency conditions or may be manually
activated with a cockpit-mounted switch. The system consists of a
transmitter, antenna, remote control and monitor unit, and associated
airplane wiring.
TRANSMITTER AND ANTENNA
The transmitter and antenna are installed in the airplane tail section.
Power for the transmitter is provided by an internal battery pack. The
transmitter incorporates a three-position switch (ARM/OFF/ON). Access to the transmitter is through an access cover placarded “ELT LOCATED HERE.” The antenna is externally mounted and connects to the
transmitter with antenna cable.
Transmitter Switch (ARM/OFF/ON)
Because of its location, this switch is not generally used by the crew. In
the OFF position, the transmitter will not transmit distress signals. This
position is normally used only while servicing the airplane. In the ON
position, distress signals will be transmitted continuously. In the ARM
position, the transmitter will automatically activate if the airplane stops
abruptly. The switch should be in the ARM position for flight.
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REMOTE CONTROL AND MONITOR UNIT
The remote control and monitor unit is installed in the cockpit. Power
for this unit is provided by an internal coin cell. A three-position ON/
ARM/RESET switch provides the remote control for the ELT transmitter. The ON and ARM positions function the same as described for the
transmitter switch. Once activated, the transmitter may be returned to
an armed status using the RESET function. The ELT can be reset but not
switched off from this control unit.
A red LED, mounted in the end of the switch handle, provides the crew
with the ELT status. The LED indicates ELT status as follows:
LED is:
On continuously
Flashing slowly (80 times per minute)
Flashing quickly (5 times per second)
Extinguished
ELT Status:
The ELT is transmitting.
The ELT transmitter is switched
OFF or the transmitter battery
needs replacement.
The remote control/monitor unit
coin cell needs replacement.
The ELT is armed.
OPERATION
To arm the transmitter for automatic activation the ON/ARM/RESET
Switch is placed in the ARM position. If the red LED flashes slowly,
check that the transmitter switch is in the ARM position. If the transmitter switch is in the ARM position and the LED continues to flash, the
transmitter battery needs replacement. To manually activate the transmitter, place the ON/ARM/RESET Switch to the ON position and
check that the red LED is on continuously. To reset the transmitter, momentarily place the ON/ARM/RESET Switch to RESET and check that
the red LED extinguishes.
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ARTEX ELT 110-406
The ARTEX ELT 110-406 transmits on 121.5, 243.0 and 406.025 MHz.
The ELT may be manually activated with a cockpit mounted switch or
will automatically activate during an impact. Once activated, the ELT
transmits the standard swept tone on 121.5 and 243.0 MHz. During that
time the 406 MHz transmitter turns on and an encoded digital message
is sent to the satellite. The information contained in the message includes:
• Serial number of the transmitter.
• Country code.
• Manufacturer.
• Position coordinates (optional).
The information sent to the satellite is programmed at the factory and
contains a unique number that can be used to identify the beacon. The
ELT 110-406 system consists of a transmitter, antenna, cockpit switch
and indicator light, buzzer (aural monitor), and associated airplane
wiring.
TRANSMITTER AND ANTENNA
The transmitter and antenna are installed in the airplane tail section.
Power for the transmitter is provided by an internal battery pack which
consists of 4 D size lithium manganese dioxide cells connected in series.
The ELT unit incorporates an ON-OFF switch. Because of its location,
this switch is not generally used by the crew. Access to the transmitter
is through an access cover placarded “ELT LOCATED HERE.” The antenna is externally mounted and is connected to the transmitter using
antenna cable.
Transmitter Switch (ON-OFF)
In the ON position, the transmitter will transmit distress signals continuously. In the OFF position, the transmitter is armed to activate either
automatically (impact) or manually (remote control from the cockpit
switch). If removed from its mounting rack, the transmitter will be deactivated.
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COCKPIT SWITCH AND INDICATOR LIGHT
The ELT 110-406 remote control (cockpit panel switch) provides manual On, Armed, and Reset modes. A 28-vdc indicator light, powered
through the ELT WARN circuit breaker on the options circuit breaker
panel in the tailcone, flashes continuously if the ELT has been activated
and is transmitting.
BUZZER
The buzzer (aural monitor) provides a distinct signal (loud, siren-type
sound) enabling a search and rescue team to locate an aircraft with a
transmitting ELT in a confined area with a large number of aircraft
(such as an airport). The buzzer is installed in the tail section and is
powered by the ELT battery pack. The buzzer does not operate continuously, but sounds at predetermined intervals, and runs for shorter periods toward the end of battery life.
OPERATION
Under normal operation the cockpit switch is in the ARM position. The
switch on the ELT unit will be positioned OFF. With these switch settings, the ELT will automatically activate on impact. To manually activate the ELT, set the cockpit switch to the ON position. When the ELT
is activated, the presence of the emergency swept tone, a flashing cockpit indicator light, and the buzzer in the tail indicates a normally functioning unit. If the ELT is activated, it can be reset. This is done by
moving the cockpit switch to ON and then immediately back to ARM.
The 406 MHz transmitter will operate for 24 hours and shut down automatically. The 121.5 and 243.0 MHz transmitter will continue to operate until the unit has exhausted the battery power. The 406 MHz
transmitter transmits a digital message that allows search and rescue
authorities to retrieve information from a database. Information contained in the database that may be useful include:
• Type of aircraft.
• Address of owner.
• Telephone number of owner.
• Aircraft registration number.
• Alternate emergency contact.
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