FSI - CE-560XL - PTM - Pilot Training Manaul Volume 1 XL and

FlightSafety
international
CITATION XL/XLS
PILOT TRAINING MANUAL
VOLUME 1
OPERATIONAL INFORMATION
FlightSafety International, Inc.

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Copyright © 2006 by FlightSafety International, Inc.
All rights reserved.
Printed in the United States of America.
CITATION XL/XLS PILOT TRAINING MANUAL
INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES
LIST OF EFFECTIVE PAGES
Dates of issue for original and changed pages are:
Original............0 ................. January 2006
NOTE:
For printing purposes, revision numbers in footers occur at the bottom
of every page that has changed in any way (grammatical or typo-
graphical revisions, reflow of pages, and other changes that do not nec-
essarily affect the meaning of the manual).
THIS PUBLICATION CONSISTS OF THE FOLLOWING:
Page *Revision Page *Revision

No. No. No. No.
Cover ...................................... 0 LIM-i—LIM-ii .......................... 0
Copyright ................................ 0 MAP-i—MAP-44 .................... 0
LEP-1—LEP-2 ........................ 0 WB-i—WB-18 ........................ 0
i—iv .......................................... 0 PER-i—PER-80 .................... 0
EC-i—EC-ii ............................ 0 CRM-i—CRM-6 ...................... 0
NP-i—NP-ii ............................ 0 SRE-i—SRE-100 .................. 0
AP-i—AP-ii ............................ 0 SRX-i—SRX-102 .................. 0
EP-i—EP-ii ............................ 0 MW-i—MW-24 ........................ 0
*Zero in this column indicates an original page.
FOR TRAINING PURPOSES ONLY LEP-1

FOR TRAINING PURPOSES ONLY
NOTICE
The material contained in this training manual is based on
information obtained from the aircraft manufacturer’s Pilot Manuals
and Maintenance Manuals. It is to be used for familiarization and
training purposes only.
At the time of printing it contained then-current information. In the

event of conflict between data provided herein and that in
publications issued by the manufacturer or the FAA, that of the
manufacturer or the FAA shall take precedence.
We at FlightSafety want you to have the best training possible. We

welcome any suggestions you might have for improving this manual
or any other aspect of our training program.

CONTENTS
EXPANDED CHECKLIST
Normal Procedures
Abnormal Procedures
Emergency Procedures
LIMITATIONS
MANEUVERS AND PROCEDURES
WEIGHT AND BALANCE
PERFORMANCE
CREW RESOURCE MANAGEMENT
RECURRENT
Syllabus
Systems Review—Excel
Systems Review—XLS
Master Warning
EXPANDED CHECKLIST
CONTENTS
Page
NORMAL PROCEDURES ................................................................ NP-i
ABNORMAL PROCEDURES........................................................... AP-i
EMERGENCY PROCEDURES......................................................... EP-i
FOR TRAINING PURPOSES ONLY EC-i

Information normally contained in this chapter will be
provided in the Aircraft Flight Manual.
FOR TRAINING PURPOSES ONLY NP-i

FOR TRAINING PURPOSES ONLY AP-i

FOR TRAINING PURPOSES ONLY EP-i

FOR TRAINING PURPOSES ONLY LIM-i


CONTENTS
Page
V-SPEED DEFINITIONS .............................................................. MAP-1
PREFLIGHT AND TAXI PROCEDURES .................................... MAP-1
TAKEOFF DATA ........................................................................... MAP-2
Emergency Return Information ............................................ MAP-3
LANDING DATA........................................................................... MAP-3
STANDARD CALLOUTS (IFR AND VFR)................................. MAP-4
TAKEOFF LIMITATIONS (FLAPS “TAKEOFF AND
APPROACH” AND FLAPS “TAKEOFF”) ................................... MAP-8
TAKEOFF BRIEFING................................................................... MAP-8
Static vs. Rolling Takeoff ..................................................... MAP-8
Flap Setting........................................................................... MAP-8
Normal Callouts.................................................................... MAP-9
Emergencies.......................................................................... MAP-9
Takeoff Briefing—Example.................................................. MAP-9
TAKEOFF ROLL......................................................................... MAP-10
Normal Takeoff................................................................... MAP-10
Engine Failure at or Above V1............................................ MAP-10
Obstacle Clearance (Loss of Engine at V1) ........................ MAP-11
Takeoff Flight Profiles ........................................................ MAP-12
ENROUTE LIMITATIONS ......................................................... MAP-16
HOLDING SPEEDS .................................................................... MAP-16
MINIMUM MANEUVERING SPEED....................................... MAP-16
STEEP TURNS............................................................................ MAP-16
Procedure............................................................................ MAP-16
APPROACHES TO STALL......................................................... MAP-18
UNUSUAL ATTITUDES ............................................................ MAP-22
Recovery Procedures .......................................................... MAP-22
EMERGENCY DESCENT.......................................................... MAP-24
FOR TRAINING PURPOSES ONLY MAP-i

APPROACHES AND LANDING PROCEDURES..................... MAP-26

Flight Deck Discipline........................................................ MAP-26
APPROACH BRIEFING ............................................................. MAP-28
Scan Transfer ...................................................................... MAP-30
CIRCLING APPROACHES ........................................................ MAP-30
MISSED APPROACH OR GO-AROUND.................................. MAP-31
LANDING PROCEDURES......................................................... MAP-34
Adjustments to Landing Distance ...................................... MAP-36
Hydroplaning Speeds ......................................................... MAP-36
LANDING LIMITATIONS ......................................................... MAP-37
CROSSWIND LANDING ........................................................... MAP-38
Method No. 1:..................................................................... MAP-38
Method No. 2:..................................................................... MAP-38
FLAPS INOPERATIVE LANDING
(NOT IN LANDING POSITION) ............................................... MAP-38
PRACTICAL TEST ..................................................................... MAP-40
Preflight Preparation........................................................... MAP-40
Preflight Procedures ........................................................... MAP-40
Takeoff and Departure Phase.............................................. MAP-41
In-Flight Maneuvers ........................................................... MAP-41
Instrument Procedures ........................................................ MAP-42
Landings and Approaches to Landings .............................. MAP-42
Normal and Abnormal Procedures ..................................... MAP-43
Emergency Procedures ....................................................... MAP-43
Postflight Procedures.......................................................... MAP-43
Parking and Securing.......................................................... MAP-43
PTS Tolerances ................................................................... MAP-44
MAP-ii FOR TRAINING PURPOSES ONLY

ILLUSTRATIONS
Figures Title Page
MAP-1 Takeoff and Landing Card ........................................ MAP-2
MAP-2 Takeoff Climb Profile.............................................. MAP-11
MAP-3 Takeoff—Aborted.................................................... MAP-13
MAP-4 Takeoff—Normal .................................................. MAP-14
MAP-5 Takeoff Engine Failure at or Above V1 .................. MAP-15
MAP-6 Steep Turns.............................................................. MAP-17
MAP-7 Approach to Stall—Enroute Configuration ............ MAP-19
MAP-8 Approach to Stall—Takeoff Configuration ............ MAP-20
MAP-9 Approach to Stall—Landing Configuration ............ MAP-21
MAP-10 Emergency Descent ................................................ MAP-25
MAP-11 Approach Plate (Typical) ........................................ MAP-27
MAP-12 ILS Approach—Normal/Single Engine .................. MAP-28
MAP-13 Nonprecision—Normal/Single Engine .................. MAP-29
MAP-14 Circling Approach .................................................. MAP-31
MAP-15 Missed Approach—Normal .................................... MAP-32
MAP-16 Missed Approach—Single Engine .......................... MAP-33
MAP-17 VFR Approach—Normal/Single Engine ................ MAP-35
MAP-18 Visual Approach and Landing
with Flaps Inoperative ............................................ MAP-39
TABLES
Tables Title Page
MAP-1 Standard Callouts ...................................................... MAP-5
MAP-2 FAR Part 25 Climb Profile ...................................... MAP-12
MAP-3 Minimum Maneuvering Speeds .............................. MAP-16
MAP-4 Landing Limitations ................................................ MAP-37
FOR TRAINING PURPOSES ONLY MAP-iii


V-SPEED DEFINITIONS
V1 Decision speed—This speed is obtained from the performance
charts in the AFM and varies with aircraft weight, engine bleeds,
altitude and temperature. It must always be less than or equal
to V R .
VR Rotation speed—This speed is a function of weight and aircraft
configuration. It must always be equal to or greater than V 1 . If
V 1 is greater than V R for a particular set of takeoff conditions,
V 1 must be lowered to equal V R .
V2 Safety climb speed—V 2 is also a function of weight and aircraft
configuration. It is obtained from the performance charts in the
AFM or from the abbreviated check-list. V 2 gives the best angle
of climb (altitude vs distance).
VFR Flap retraction speed—Flap retracting speed (V 2 + 10 knots).
Also used as minimum final segment climb speed.
V ENR Single-engine enroute climb speed—This speed can be used for
a variety of purposes and is obtained from the AFM:
• Best single-engine rate-of-climb (altitude vs time)
V REF Minimum final approach speed—This speed is 1.3 V SO and is
the minimum speed to be used on final approach. It is the air-
speed that is used for the threshold crossing speed with full flaps
and landing gear extended.
V APP Minimum landing approach climb speed—The landing ap-
proach climb (1.3 V S1 ) with 15° flap position, landing gear up.
Also used as minimum go-around speed.
PREFLIGHT AND TAXI PROCEDURES

NOTE
With the gust lock on, the flight controls are locked in
neutral and the throttles are locked in the off position.
CAUTION
Do not tow with the control lock engaged, to prevent
damage to the nosewheel steering mechanism. After
completing the initial flight planning and preflight
checks, takeoff data should be computed to obtain cor-
rect takeoff thrust setting, V 1 , V R , V 2 , and the emer-
gency return V REF , V APP speed.
FOR TRAINING PURPOSES ONLY MAP-1

TAKEOFF DATA
A Takeoff Data Card is shown in Figure MAP-1.
TO N 1 & CLB N 1 —Maximum fan settings for takeoff and climb based on ex-
isting temperature and pressure altitude taken from the Flight Manual or check-
list. With EECs in manual mode an adjustment must be made for anti-ice.
V 1 , V R , V 2 , V FR & V ENR —Calculated V 1 , V R , V 2 and V ENR based on exist-

ing temperature, pressure altitude and aircraft weight and flap setting taken
from the Flight Manual or checklist. (V FR is V 2 + 10 knots)
CLEARANCE—Space provided for copying ATC clearances and other per-

tinent airport information.
ARPT—Name of airport or ICAO identifier.
ELEV—Airport elevation or runway elevation if significantly different than

airport elevation.
RWY—Runway in use for departures.
FlightSafety CITATION FlightSafety CITATION

international international
TAKEOFF DATA LANDING DATA

TO N1 CLB N1 VREF VAPP
V1 VR V2 GA N1 RWY REQ'D
VFR VENR FLAPS CLEARANCE
CLEARANCE
ARPT________ELEV_________RWY________
ATIS________WIND___________VIS________
CIG________________TEMP/DP______/_____ ARPT________ELEV_________RWY________
ALT________RMKS______________________ ATIS________WIND___________VIS________
RWY LENGTH__________RWY REQ'D______ CIG________________TEMP/DP______/_____
ZFW___________T.O. WT._________________ ALT________RMKS______________________

EMERGENCY RETURN
ZFW_____________RLDG WT_____________
VREF________VAPP_________MSA________
Figure MAP-1. Takeoff and Landing Card
MAP-2 FOR TRAINING PURPOSES ONLY

ATIS—Current ATIS information code.
WIND—Wind direction and speed as reported by ATIS.
VIS—Visibility as reported by ATIS.
CIG—Clouds and significant weather as reported by ATIS.
TEMP/DP—Temperature and dew point as reported by ATIS.
ALT—Altimeter setting as reported by ATIS.
RMKS—Any additional information provided by ATIS.
RWY LENGTH—Actual length of runway to be used for takeoff.
RWY REQ’D—Charted takeoff field length. If actual runway is less, reduce

gross weight to equal the actual runway length. Adjust for runway conditions.
ZFW—Zero Fuel Weight. This is the basic operating weight (BOW) plus
weight of passengers and cargo (or BEW plus crew, stores, passengers and
cargo). Fuel is not included.
T.O. WT.—The actual weight of the airplane at the beginning of takeoff roll
(does not include taxi fuel).
EMERGENCY RETURN INFORMATION

V REF - V APP —Calculated approach speeds corresponding to the appropriate
flap settings and based on landing weight.
MSA—Minimum Safe Altitude required for obstacle clearance. May be taken

from MSA circle on approach plate, ATC clearance or if in VMC, the VFR
pattern altitude.
LANDING DATA
A Landing Data Card is shown in Figure MAP-1.
V REF - V APP —Calculated approach speeds corresponding to the appropriate

flap settings and based on landing weight.
GA N 1 —Obtained from Flight Manual for go-around (TO N 1 ). It is based on

the approach climb configuration.
RWY REQ’D—Landing distance adjusted for: aircraft configuration (flaps–an-

tiskid), landing gross weight, runway conditions.
CLEARANCE—Space provided for copying ATC clearances and other per-

tinent airport information.

ARPT—Name of airport or ICAO identifier.
ELEV—Airport elevation or runway elevation if significantly different than

airport elevation.
RWY—Runway in use for departures/arrivals.
ATIS—Current ATIS information code.
WIND—Wind direction and speed as reported by ATIS.
VIS—Visibility as reported by ATIS.
CIG—Clouds and significant weather as reported by ATIS.
TEMP/DP—Temperature and dew point as reported by ATIS.
ALT—Altimeter setting as reported by ATIS.
RMKS—Any additional information provided by ATIS.
ZFW—Zero Fuel Weight. This is the basic empty weight or basic operating
weight plus weight of passengers and cargo. Fuel is not included. (This fig-
ure should be the same as the takeoff ZFW.)
LDG WT—Actual weight for landing at the destination airport. ZFW plus fuel
remaining.
NOTE
When using the charts to determine the V speeds, re-
member VREF and VAPP speeds are functions of weight
and flap configurations.
STANDARD CALLOUTS (IFR AND VFR)

NOTES:
1. Check for appearance of warning flags and gross
instrument discrepancies. See Supplement 4 (XLS)
or Supplement 19 (XL) for more information.
2. Care must be exercised to preclude callouts which

can influence the pilot flying and result in pre-
mature abandonment of instrument procedures.
3. It is recommended that all aircraft utilize avail-

able electronic/visual systems as an aid in main-
taining glide slope.
Table MAP-1 is a good example of standard crew calls on a typical flight. It

describes the aircraft position, the duties and callouts of both pilots. It will
be referred to in this section.

Table MAP-1. STANDARD CALLOUTS
LOCATION CONDITION CALLOUTS-PF CALLOUTS-PNF

Cleared for takeoff PF calls for ìbelow the lineî checklist then PNF 1. ìBelow the line itemsî 2. ìBEFORE TAKEOFF
completes checklist and reports complete. CHECKLIST completeî
Takeoff PF sets throttles to TAKEOFF detent. PNF 1. ìVerify powerî 2. ìPower Setî

verifies power at target N1.
First indication of airspeed (both PFDs and SFD). ìAirspeed aliveî

Airspeed indication of 80 KIAS. ì80 knotsî
Airspeed indicators at computed V1. ìVee-Oneî
Airspeed indicators at computed VR. ìRotateî
Takeoff and climb Positive rate of climb. 1. ìPositive rate, gear UPî 2. ìGear UP selectedî
Single engine at V2 airspeed. 2. ìFLCî 1. ìVee-Twoî
3. ìFLC selectedî
Landing gear confirmed UP. ìGear is UPî
V2 + 10 KIAS at Safe Altitude. ìVee-Two plus 10î
V2 + 10 KIAS. 1. ìFlaps up ñ Yaw damper ONî 2. ìFlaps UP selected ñ Movingî
3. ìYaw damper ONî
Flaps confirmed up. ìFlaps are UPî
Cruise climb Passing 10,000 feet (or lower if level off at ìAFTER TAKEOFF/CLIMB ìAFTER TAKEOFF/CLIMB
lower altitude) PF calls for, then PNF CHECKLISTî CHECKLIST completeî
completes the AFTER TAKEOFF/CLIMB
checklist and calls complete.
Crossing transition altitude 2. ìAltimeter 29.92 (or 1. ìAltimeter 29.92 (or 1013
MAP-5
(before TA for JAA) right seat pilot sets 1013 as required) set twiceî as required) set right sideî
primary and SFD altimeters.
MAP-6
Table MAP-1. STANDARD CALLOUTS (Cont.)

Climbs/descents Altitude changes set by PNF 2. “Roger, (present altitude) 1. “Altitude preselect set
and confirmed by PF. for (assigned altitude)” for (assigned altitude)”
Cruise Level off PF calls for, then PNF “CRUISE CHECKLIST” “CRUISE CHECKLIST

completes the CRUISE CHECKLIST. complete”
Navigation When CDI comes off full deflection. “Course alive”

Descent PF calls for, then PNF completes “DESCENT CHECKLIST” “DESCENT CHECKLIST
DESCENT CHECKLIST. complete”
Passing through transition altitude 2. “Altimeter (QNH or QFE 1. “Altimeter (QNH or
(transition level for JAA) right seat as required) set twice” QFE as required) set right side”
pilot sets primary and SFD altimeters.
Configuring for approach Prior to FAF, PF calls for, then the PNF “APPROACH CHECKLIST” “APPROACH CHECKLIST
and landing completes the APPROACH CHECKLIST. complete”
One dot below (precision) or prior 1. “Landing Gear DOWN, 2. “Landing gear selected
to FAF (non-precision). BEFORE LANDING CHECKLIST” DOWN”
3. “Landing gear is DOWN,
three green, no red”
After landing gear extended 1. “Flaps full down” 2. “Landing flaps
(two-engine) or after landing selected DOWN”
assured (single-engine). 3. “Flaps are full DOWN”
Prior to landing, confirm landing configuration “BEFORE LANDING
and BEFORE LANDING CHECKLIST complete. CHECKLIST complete”
During approach, 1,000 feet above DA or MDA. “1,000 feet above
radar vectors or minimums”
procedure turn (IMC)
Table MAP-1. STANDARD CALLOUTS (Cont.)

500 feet above DA or MDA. “500 feet above minimums”
100 feet above DA or MDA. “100 feet above minimums”
Approach lights/runway in sight “Approach lights/runway

(at any time sighted). in sight at (o’clock position)”
Precision approach At DH with no visual contact. “Minimums – Go-around”

Nonprecision approach At MDA. “Minimums”
At missed approach point (or VDP) “Missed approach point
with no visual contact. (or VDP), no contact, go-around”
Departing DH/MDA/VDP Visual for a landing. 1. “Visual for Landing” 3. “Sink rate (feet per minute)”
2. “Departing MDA (nonprecision)” 4. “REF + (amount)”
Go-around Called by either pilot. PF calls for and PNF 1. “Going around, flaps fifteen, 2. “Flaps fifteen selected”
completes GO-AROUND CHECKLIST. GO-AROUND CHECKLIST” 3. “Flaps are fifteen”
4. “Positive rate, Gear up. Go- 5. “Gear-up selected”
around checklist”
Single engine. 2. “FLC” 1. “Vee APPROACH is ___”

VFR approach 500 feet above field elevation PNF calls 500 1. “500 Above”
feet above, speed (Vref plus), sink rate. 2. “Speed – Ref plus ________”
3. “Sink _________ hundred”
In flight Transfer of aircraft control. “You have control, autopilot is “I have control, heading _____,
(ON / OFF), Flight Director altitude ______.”
(NAV / HDG), heading ______,
altitude ______.”
Significant deviation Significant deviations in airspeed, heading, Calls and/or acknowledges the Calls the deviation.
MAP-7
altitude, course, vertical speed, or bank angle deviation then states

may be called by either pilot. “CORRECTING” when doing so.
TAKEOFF LIMITATIONS (FLAPS “TAKEOFF AND

APPROACH” AND FLAPS “TAKEOFF”)
The takeoff weight is limited by the most restrictive of the following
requirements:
1. Maximum certified takeoff weight (structural).
2. Maximum takeoff weight permitted by takeoff climb requirements
(normally, 2nd segment climb requirement).
3. Maximum takeoff weight permitted by takeoff field length.
Takeoff field length ensures a rejected takeoff can be completed on the ex-
isting runway and it allows for the takeoff to be continued, ensuring the air-
craft reaches a height of 35 feet dry, 15 feet wet, (reference zero) by the time
it reaches the end of the takeoff distance.
These requirements are operating limitations and must be complied with.

Additionally, obstacle clearance capability may be an actual physical neces-
sity, if not a legal requirement, and may further limit the takeoff weight.
The pilot should also consider the landing weight restrictions at the destina-
tion airport. The limited landing weight plus the expected fuel to be burned
enroute may be more limiting than any restrictions at the departure airport,
especially if the trip is of short duration.
TAKEOFF BRIEFING
Prior to takeoff, the pilot-in-command should review with the copilot the stan-
dard callouts, the departure procedures and also the emergency procedures
for a rejected takeoff prior to V 1 or a continued takeoff after V 1. Considerations
should be given to a minimum of the following items.
STATIC VS. ROLLING TAKEOFF

All performance data is based on a static takeoff (engines producing takeoff
thrust prior to releasing the brakes. However, this type of takeoff is highly
uncomfortable for the passengers, therefore, runway length permitting, it is
advisable to perform a rolling takeoff.
FLAP SETTING
Review and check the flap setting. This will be based on the performance cri-
teria required for the airport departure procedure. Anti-ice will affect perform-
ance, therefore, it is advisable to brief whether anti-ice will be on or off.

NORMAL CALLOUTS
With EECs operational, setting power is just a matter of advancing the throt-
tles to the takeoff detent. Power only needs to be verified within the normal
range of fan speed. Standard calls during the takeoff roll may vary, but,
should be standard within each flight department.
Monitoring engine instruments and flight instruments for any irregularity is

essential for safety of flight. Any such irregularity prior to the specified
speed for abort, e.g.,V 1 , should be called out as “ABORT” with a simple ex-
planation why, e.g., “CABIN DOOR OPEN.” The pilot-in-command will
have final authority to abort. After an abort, the problem may be sorted out
once safely stopped and clear of the runway.
EMERGENCIES
A plan of action should be discussed in the event of an emergency. The plan
should consist of safety items, such as safe altitudes and headings, emergency
checklists, airplane handling, and a safe return to the departure airport or de-
parture alternate, all based on weather conditions.
TAKEOFF BRIEFING—EXAMPLE
The following is an example of a standard takeoff briefing. The briefing
should be accomplished prior to requesting takeoff clearance. Although your
exact phraseology may differ, the main ideas should remain in the briefing.
1. “This will be a (static or rolling) takeoff with flaps set at (state flap
position).”
(Mention Anti-Ice if required).
2. “I will set the throttles, and you verify the takeoff power.”
3. “Call: ‘Airspeed Alive,’ ‘80 knots, cross-check,’ ‘Vee One,’ ‘Rotate,’

‘Positive Rate,’ and ‘Vee two plus ten.’
4. “Monitor all engine instruments and the annunciator panel during takeoff,
cross-check both airspeed indicators at 80 knots.”
5. “In the event of a serious malfunction prior to V1, call ‘Abort’ and I will
execute the abort.”
6. “If a malfunction occurs at or after V1, we will continue the takeoff. After
safely airborne, advise me of the malfunction and we will handle it as an
in-flight emergency.”
7. “In the event of a thrust reverser deployment, I will fly the aircraft and you
will do the emergency stow.”

8. “In the event of an engine failure or fire, do not identify the engine, only
advise if it is a failure or a fire.”
9. “Minimum safe altitude for emergencies will be (state altitude). Plan to fly
(type of approach).” Fly V2 until altitude is reached.
10. “Departure instructions are (Inst. Depart., route, altitude, etc.).”
11. “The navaids are set to (__________________).”
12. “Any questions?”
TAKEOFF ROLL
The pilot will steadily advance the throttles to the takeoff detent. The copi-
lot will check and verify the N 1 gages and make the standard calls while mon-
itoring all instrument indications.
If an abnormal situation, annunciator light, system failure, etc., occurs dur-

ing the takeoff roll, the copilot notifies the pilot-in-command, who makes the
final decision to take off or abort. If the decision to abort is made, the mem-
ory items for ENGINE FAILURE, OR FIRE, OR MASTER WARNING DUR-
ING TAKEOFF—Speed Below V 1 , should be performed. Once stopped, or if
able, clear of the runway, notify ATC of your actions.
NORMAL TAKEOFF
When “ROTATE” is called (V R ), the pilot should apply steady back pressure
and allow the aircraft to rotate to a 10° noseup pitch attitude on the ADI. When
a positive rate of climb is indicated, retract the gear. As the airspeed increases
through a minimum of V 2 + 10 knots (VFR), retract the flaps. Continue to ac-
celerate to normal climb speed and complete the After Takeoff—Climb items.
ENGINE FAILURE AT OR ABOVE V1

If an engine fails at or above V 1 , the takeoff will normally be continued. At
V R , steadily apply back pressure to allow the aircraft to rotate the nose to 10°
noseup pitch attitude. Do not attempt to “pull” the aircraft off the runway.
Perform a “normal” rotation to allow the aircraft to fly off the runway. After
establishing a positive rate of climb, raise the landing gear. Maintain V 2 until
reaching a safe altitude, or 1,500 feet above airport elevation, whichever is
higher, then lower the nose, without losing altitude to accelerate to V ENR . As
the airspeed reaches V 2 + 10 knots (VFR), retract the flaps and accelerate to
V ENR . When V ENR is achieved or the single-engine 10-minute limitation for
takeoff power is reached, reduce power to the climb detent. Continue climb
at V ENR to assigned or amended altitude. When time and cockpit duties per-
mit, complete the appropriate Emergency Procedures checklist and the After
Takeoff—Climb checklist.

WARNING
If rudder bias in inoperative, it will be necessary to
apply greater rudder pressure to maintain directional
control. The amount of rudder pressure will depend
on several factors, i.e., airspeed, power setting, and
flap or gear configuration. Maintain sufficient rud-
der pressure to keep the ball centered. Remember, as
speed changes, the rudder pressure will also change.
NOTE
• Don’t let the emergency distract from flying the
airplane. Wait until safety air borne, at a safe al-
titude, before performing the emergency and the
After Takeoff—Climb checklist. Some memory
items may require a more immediate action.
• If engine time limits at takeoff power is reached

prior to reaching V ENR (clear of obstacles) main-
tain attained airspeed, reduce power to the climb
detent, and climb to the enroute altitude.
• If it becomes necessary to maneuver the airplane

during the single-engine departure climb before
attaining minimum maneuvering speed, limit the
bank angle to 15°.
OBSTACLE CLEARANCE (LOSS OF ENGINE AT V1)

FAR 25 requires that the aircraft manufacturer display a Takeoff Profile be-
ginning at reference zero and ending at 1,500 feet AGL (Figure MAP-2).
TAKEOFF THRUST*
T
MEN
L SEG
FINA
3RD SEGMENT
NT
GME
SE
REFERENCE ZERO D
2N 1,500 FEET AGL
GMENT
1S T SE GEAR UP
* SEE TABLE MAP-1, THRUST SETTING
REFERENCE ZERO: = 35 FEET ABOVE TAKEOFF SURFACE FOR A DRY RUNWAY

= 15 FEET ABOVE TAKEOFF SURFACE FOR A WET RUNWAY
Figure MAP-2. Takeoff Climb Profile

Second segment is generally the most limiting segment. When at an airport

that requires a minimum climb gradient to an altitude that is higher than
1,500 feet AGL, the second segment is extended to that minimum “safe” al-
titude. In order to meet the second segment climb, all conditions must be met,
particularly, climbing at V 2 .
Table MAP-2 shows FAR PART 25 climb profile.
Table MAP-2. FAR PART 25 CLIMB PROFILE
1ST SEGMENT 2ND SEGMENT 3RD & FINAL

SEGMENT
SPEED V2 V2 V2 + 10 Flaps
transitioning to UP-
accelerating to VENR
Thrust Setting: Takeoff Takeoff Takeoff
10 Minutes for (One Engine - (One Engine - (One Engine -
single engine. Anti ice On/Off) Anti ice On/Off) Anti ice On/Off)
Flap Position: 7° or 15° 7° or 15° Transitioning from
(As Required) (As Required) Takeoff to UP
Gear Position: Down Up Up
Required Climb Positive 2.4% Gross 1.2% Gross
Gradient: (Greater than Zero) (1.6% Net) (0.1% net)
* Refer to the Aircraft Flight Manual for limitations on takeoff thrust time limitations (normally
5 minutes, but may be extended to 10 minutes if required).
TAKEOFF FLIGHT PROFILES

Figures MAP-3, MAP-4 and MAP-5 demonstrate takeoff flight profiles.

EVALUATE SITUATION *
1. CLEAR RUNWAY
OR
EMERGENCY EVACUATION

DECISION TO ABORT
1. CALL "ABORT"
2. BRAKES—MAXIMUM EFFORT
3. THROTTLES—IDLE
4. THRUST REVERSERS—DEPLOY
ON UNAFFECTED ENGINE(S)
5. SPEED BRAKES—EXTEND
CLEARED FOR TAKEOFF

1. THROTTLES—T/O N1 SET
2. BRAKES—RELEASE
BEFORE TAKEOFF
1. TAKEOFF CHECKLIST/
BRIEFING—COMPLETED
* NOTE: CONSIDER BRAKE ENERGY PRIOR TO SUBSEQUENT

MAP-13
OPERATION OF THE AIRCRAFT.
Figure MAP-3. Takeoff—Aborted

MAP-14
AFTER TAKEOFF/CLIMB
1. ACCELERATE TO NORMAL CLIMB SPEED
2. THROTTLES—MCT, OR AS REQUIRED
3. AFTER TAKEOFF/CLIMB CHECKLIST—
COMPLETED

ROTATE
1. VR—SMOOTHLY ROTATE GEAR/FLAP RETRACTION
TO 10˚ NOSE UP ATTITUDE
1. POSITIVE RATE OF CLIMB—GEAR UP
2. AT A PREDETERMINED ALTITUDE
CONSIDERING TERRAIN, AND AT A
MINIMUM AIRSPEED OF V2 + 10 KT—
CLEARED FOR TAKEOFF FLAPS UP
2. BRAKES—RELEASE
BEFORE TAKEOFF
Figure MAP-4. Takeoff—Normal

AFTER TAKEOFF/CLIMB
1. CLIMB AS REQUIRED AT VENR
2. THROTTLES—MCT, OR AS REQUIRED
3. AFTER TAKEOFF/CLIMB/ENGINE
FAILURE CHECKLISTS—COMPLETED
GEAR RETRACTION/INITIAL CLIMB

1. POSITIVE RATE CLIMB—GEAR UP
2. AIRSPEED—V2
3. CLIMB AT V2 TO 1,500' AGL OR CLEAR OF

OBSTACLES, WHICHEVER IS HIGHER
ROTATE
1. AT VR—SMOOTHLY ROTATE FLAP RETRACTION
TO 10˚ NOSE UP ATTITUDE
1. AT V2 + 10 KT (MINIMUM)—
FLAPS UP
2. ACCELERATE TO VENR
CLEARED FOR TAKEOFF
2. BRAKES—RELEASE
ENGINE FAILURE
1. LOSS OF ENGINE AT
OR ABOVE V1
BEFORE TAKEOFF
MAP-15
Figure MAP-5. Takeoff Engine Failure at or Above V1

ENROUTE LIMITATIONS
The AFM chart, “Enroute Net Climb Gradient: Single Engine,” is not an op-
erating limitation of the airplane. However, it allows the pilot to calculate the
maximum enroute altitude that the airplane will climb to on one engine or drift
down to if an engine fails at a higher altitude. The chart depicts the actual or
gross gradient of climb reduced by 1.1% net.
HOLDING SPEEDS
Based upon approximately 200-220 KIAS depending upon altitude for a
20,000 pound Citation Excel/XLS with a 5-knot decrease for each 1,000
pound of weight decrease, if the angle-of-attack indicator is used for hold-
ing, .38-.40 will provide optimum specific range or miles per gallon of fuel.
If fuel is critical, flying 0.6 on the angle-of-attack indicator will provide best
endurance or maximum flight time per gallon of fuel.
MINIMUM MANEUVERING SPEED

This speed is the minimum speed that will provide an adequate margin above
stall while maneuvering the aircraft. Table MAP-3 lists the factors to be added
to full flap V REF for the Citation Excel/XLS minimum maneuvering speeds.
Table MAP-3. MINIMUM MANEUVERING SPEEDS
FLAP CONFIGURATION VREF

CLEAN +30
FLAPS T.O. (7°) +20
FLAPS T.O. AND APPR (15°) +20
FLAPS LAND (35°) +10
STEEP TURNS
Figure MAP-6 demonstrates a steep turn profile.
PROCEDURE
• AIRSPEED—200 KIAS
• BANK ANGLE— 45°
• MAINTAIN ALTITUDE
• INCREASE THRUST PASSING THROUGH 30° BANK ANGLE (AP-
PROXIMATELY 3% N).
• PLAN ROLLOUT SO THAT WINGS ARE LEVEL AS THE AIR-
CRAFT REACHES THE DESIRED HEADING.

EXIT
1. PLAN ROLLOUT SO THAT WINGS
ARE LEVEL AS THE AIRCRAFT
REACHES THE DESIRED HEADING
1. INCREASE THRUST PASSING

THROUGH 30˚ BANK ANGLE
(APPROX. FLOW OR 3% N1)
ENTRY
1. AIRSPEED—200 KIAS
2. BANK ANGLE—45˚
3. MAINTAIN ALTITUDE
Figure MAP-6. Steep Turns

APPROACHES TO STALL
Prior to any planned approaches to stall (Figures MAP-7 through MAP-9),
clear area visually. All recoveries will be made with power and minimum loss
of altitude.
At least one approach to a stall shall be accomplished while in a turn using a

constant bank angle of 15° to 30°.
For Citation Excel/XLS aircraft, stall warning is normally provided by a

stick shaker attached to the control columns. It is activated by an angle-of-
attack indication of approximately .82 (gear down, full flaps). Additionally,
stall strips on the inboard section of each wing leading edge provide aerody-
namic stall warning during high angles of attack, which causes disruption of
airflow over the horizontal stabilizer, resulting in a prestall buffet. Stall re-
covery should be initiated at the onset of either indication (AOA warning or
aerodynamic prestall buffet).
Prior to stalls, the following items should be completed. The acronym ICEY
will aid in remembering the items:
1. Ignition................................................................................................... ON
2. Compute.......................... VREF (landing configuration) for aircraft weight
3. Engine Synchronizer ............................................................................ OFF
4. Yaw damper.......................................................................................... OFF
NOTE:
Limitations: No intentional stalls are permitted above
25,000 feet.

BEGINNING OF MANEUVER RECOVERY COMPLETION OF MANEUVER
1. LEVEL FLIGHT—CLEAN AIRCRAFT 1. APPLY MAX POWER 1. ACCELERATE

2. POWER—IDLE 2. MAINTAIN PITCH ATTITUDE
3. MAINTAIN ALTITUDE 3. KEEP WINGS LEVEL

4. TRIM—AS REQUIRED 4. IT MAY BE NECESSARY TO RELAX
PITCH ATTITUDE SLIGHTLY
AERODYNAMIC BUFFET OR
STICK SHAKER (IF APPLICABLE),
WHICHEVER OCCURS FIRST
MAP-19
Figure MAP-7. Approach to Stall—Enroute Configuration

MAP-20
1. LEVEL FLIGHT 1. APPLY MAX POWER 1. ACCELERATE TO VAPP + 10 KT

2. FLAPS—TAKEOFF & APPROACH 2. CHECK FLAPS AT TAKEOFF & APPROACH 2. RETRACT FLAPS
3. ROLL INTO 20˚ BANK 3. MAINTAIN PITCH ATTITUDE

4. SET POWER TO IDLE 4. ROLL WINGS LEVEL *
5. MAINTAIN ALTITUDE
6. TRIM—AS REQUIRED
STICKSHAKER (IF APPLICABLE),
* USE RUDDER TO AID IN LEVELING THE WINGS. THESE WILL

MINIMIZE THE ADVERSE YAW PRODUCED BY DOWN AILERON.
Figure MAP-8. Approach to Stall—Takeoff Configuration


1. LEVEL FLIGHT 1. APPLY MAX POWER 1. MAINTAIN ATTITUDE UNTIL A
2. GEAR—DOWN 2. MAINTAIN 5˚ - 10˚ NOSE UP PITCH ATTITUDE POSITIVE RATE OF CLIMB IS

3. FLAPS—LAND (35˚) 3. MAINTAIN WINGS LEVEL OBTAINED
4. SET POWER TO 45% - 50% N1 4. CALL FOR FLAPS TO TAKEOFF & APPROACH 2. RETRACT THE GEAR
5. TRIM—AS REQUIRED 3. CLIMB TO DESIRED ALTITUDE
AT VAPP THEN ALLOW AIRSPEED
TO INCREASE TO VAPP + 10 KT
4. RETRACT FLAPS
STICKSHAKER (IF APPLICABLE),
MAP-21
Figure MAP-9. Approach to Stall—Landing Configuration

UNUSUAL ATTITUDES
An unusual attitude is an aircraft attitude occurring inadvertently. It may re-
sult from one factor or a combination of several factors, such as turbulence,
distraction from cockpit duties, instrument failure, inattention, spatial dis-
orientation, etc. In most instances, these attitudes are mild enough for the pilot
to recover by reestablishing the proper attitude for the desired flight condi-
tion and resuming a normal cross-check.
Techniques of recovery should be compatible with the severity of the un-

usual attitude, the characteristics of the airplane and the altitude available
for recovery.
The following aerodynamic principles and considerations are applicable to

recovery from unusual attitudes:
• The elimination of a bank in a dive aids in pitch control.
• The use of bank in a climb aids in pitch control.
• Power and speedbrakes, used properly, aid in airspeed control.
RECOVERY PROCEDURES
Attitude Indicator(s) Operative
Normally, an attitude is recognized in one of two ways: an unusual attitude
“picture” on the attitude indicator or unusual performance on the perform-
ance instruments. Regardless of how the attitude is recognized, verify that
an unusual attitude exists by comparing control and performance instrument
indications prior to initiating recovery on the attitude indicator. This precludes
entering an unusual attitude as a result of making control movements to cor-
rect for erroneous instrument indications.
• If diving, adjust power and/or speedbrakes as appropriate, based on

indicated airspeed while rolling to a wings-level, upright attitude, and
correct to level flight on the attitude indicator.
• If climbing, use power as required, and bank to the “nearest” horizon

as necessary to assist in pitch control and to avoid negative G forces.
As the airplane symbol approaches the horizontal bar, adjust pitch,
bank and power to complete the recovery and establish the desired
aircraft attitude.
If there is any doubt as to proper attitude indicator operation, then recovery

should be made using attitude indicator inoperative procedures:

Attitude Indicator(s) Inoperative

With an inoperative attitude indicator, successful recovery from unusual at-
titudes depends greatly on early recognition of attitude indicator failure. For
example, attitude indicator failure should immediately be suspected if con-
trol pressures are applied for a turn without corresponding attitude indicator
changes. Another example is satisfactory performance instrument indica-
tions that contradict the “picture” on the attitude indicator.
If an unusual attitude is encountered with an inoperative attitude indicator,

the following procedure is recommended:
• Check other attitude indicators for proper operation and recover on the
operative attitude indicator.
• If unable to determine a reliable attitude indicator, use the following

procedures based on indicated airspeed.
Airspeed High and Increasing

1. If airspeed is high and increasing, decrease power and extend the
speedbrakes to prevent speeds in excess of VMO or MMO.
2. Level the wings based on movement of the heading indicator. Example, if

the heading indicator is turning clockwise, the aircraft is in a left bank, rotate
the yoke clockwise until the heading indicator stops turning.
3. Level the pitch attitude based on the movement of the altimeter / VVI. If
the altitude is decreasing, gently but firmly pull on the yoke until the
altitude is constant and/or the VVI is reading zero. Adjust yoke pressure to
maintain a constant attitude.
4. Once the airspeed has reached a comfortable level, adjust power and
retract the speedbrakes to maintain a safe airspeed while using the heading
indicator for bank control and altimeter for pitch control.
Airspeed Low and Decreasing

1. If the airspeed is low and decreasing, increase power as necessary.
2. Level the pitch attitude based on the altimeter/VVI. If the altitude is

increasing, gently push on the yoke, avoiding any negative g, until the
altitude is constant or the VVI is reading zero. Adjust yoke pressure to
maintain a constant altitude.
3. Level the wings based on the heading indicator. If the heading indicator is
turning counterclockwise, the aircraft is in a right bank, rotate the yoke
counterclockwise until the heading indicator stops turning.

NOTE
In a nose high situation, without the use of an atti-
tude indicator, it may be risky to roll the aircraft to
reduce the vertical lift to bring the nose down to a level
attitude. Accurate monitoring of the heading indica-
tor is necessary to ensure the aircraft does not go into
an overbank situation. If the heading indicator is
turning slowly, let the climb rate decrease to zero be-
fore leveling the wings.
EMERGENCY DESCENT
1. Start maneuver at an altitude of 35,000 to 45,000 feet (Figure MAP-10).
2. The initial entry into the descent begins when the throttles are brought to
idle and the speedbrakes are extended. The aircraft will begin a pitch down
movement. Allow the nose to drop to about 20° nosedown pitch avoiding
any negative g forces on the airplane. As the speed approaches MMO/VMO,
adjust nosedown pitch to maintain this speed and trim to maintain the
desired speed.
3. Call out periodic altitude checks during descent.
4. Copilot calls 2,000 feet above level-off altitude; start level-off 1,000 feet
above altitude and retract speedbrakes.

INITIAL DESCENT LEVEL-OFF
IF THIS DECISION IS A RESULT

OF CABIN DECOMPRESSION

1. PERORM MEMORY ITEMS FOR “
CAB ALT” “CABIN DECOMPRSSION”

2. COMPLETE THE CABIN DECOM-
PRESSION CHECKLIST ITEMS AS
TIME PERMITS.
3. PERFORM THE EMERGENCY
DESCENT AS REQUIRED
DURING DESCENT
1. ATC—NOTIFY
INITIATE DESCENT 2. ATC TRANSPONDER—7700 APPROACHING
1. PERFORM THE MEMORY ITEMS (IF NECESSARY) DESIRED ALTITUDE
“EMERGENCY DESCENT”. 3. ALTIMETER SETTING—REQUEST 1. LEVEL OFF—INITIATE 1,000' PRIOR
2. COMPLETE THE EMERGENCY 4. DETERMINE MINIMUM SAFE ALTITUDE TO DESIRED ALTITUDE
DESCENT CHECKLIST ITEMS 5. PRESSURIZATION—RESET, IF ABLE 2. SPEED BRAKES—RETRACT
AS TIME PERMITS. 3. CREW OXYGEN—NORMAL
4. DETERMINE WELL BEING
* NOTE:
IF SMOKE IS PRESENT IN THE COCKPIT,
PERFORM THE “ELECTRICAL FIRE OR
SMOKE” OR THE “SMOKE REMOVAL”
CHECKLISTS AS REQUIRED.
MAP-25
Figure MAP-10. Emergency Descent

APPROACHES AND LANDING PROCEDURES

FLIGHT DECK DISCIPLINE
Good operating practices are essential for precise execution of approach pro-
cedures, whether on instruments or visual. By constantly maintaining an
awareness of the progress along the approach profile, the crew provides for
an orderly transition to the landing runway. Cross-checking must be thorough
and continuous.
Approach planning begins sufficiently in advance of the approach, with a re-

view of the approach charts and attention given to alternative courses of action
(Figure MAP-11).
Flight information redundancy improves the ability to cross-check, which in

turn provides for a continuous validation of one information source against
another. It also decreases the affect of overconcentration on a single instru-
ment display.
The cross-check on final approach is, therefore, enhanced by tuning both pilot
navigation aids to the same frequencies.

FIgure MAP-11. Approach Plate (Typical)

APPROACH BRIEFING
Prior to completing the Before Landing Checklist, a thorough briefing should
be given by the pilot flying. Items to cover should include, but not be limited
to, type of approach and transition, radio frequencies, courses and altitudes,
timing and missed approach procedures along with the standard calls as out-
lined in Table MAP-1.
Approach profiles are shown in Figures MAP-12 and MAP-13.
The following is an example of a standard approach briefing:
1. “This will be the ILS approach to runway 1L at Wichita, chart number 11-
1, dated eleven September, XXXX.”
2. “Localizer frequency is 109.1. set in NAV 1 with the inbound course of

013° set on the HSI. Set 113.8 in NAV 2 with 149° course selected to
identify CHITO. Identify all navigation aids.”
ABEAM FAF OR
IAF (OR DOWNWIND VEC TORS) PROCEDURE TURN OUTBOUND
1. APPROACH CHECKLIST—INITIATE 1. FLAPS—15˚
2. AIRSPEED—160 - 180 KIAS 2. AIRSPEED (MIN)—MINIMUM
MANEUVERING SPEED *
GLIDESLOPE INTERCEPT (NORMAL)

1. ONE DOT FROM G/S INTERCEPT—GEAR DOWN
2. G/S INTERCEPT—FLAPS 35˚
3. AIRSPEED (MIN)—VREF + 10 KT
4. BEFORE LANDING CHECKLIST—COMPLETED
GLIDESLOPE INTERCEPT (SINGLE ENGINE)

1. GEAR DOWN
2. AIRSPEED (MIN)—VAPP + 10 KT
3. SINGLE ENGINE APPROACH AND LANDING
CHECKLIST—COMPLETED
DECISION HEIGHT
1. RUNWAY VISUAL REFERENCES IN SIGHT:
a. MAINTAIN GLIDESLOPE
b. LANDING ASSURED (NORMAL)—
VREF CROSSING THRESHOLD
c. LANDING ASSURED (SINGLE ENGINE)—
FLAPS 35˚ AND VREF CROSSING
THRESHOLD
2. RUNWAY VISUAL REFERENCES NOT IN SIGHT:
a. ACCOMPLISH MISSED APPROACH
NOTE:
IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE
GUST FACTOR IN EXCESS OF 5 KT.
* MINIMUM MANEUVERING SPEED IS:

VREF + 30 (FLAPS 0˚)
VREF + 20 (FLAPS T.O. & APP)
Figure MAP-12. ILS Approach—Normal/Single Engine

3. “Start timing at CHITO, using two minutes, three seconds for 140 knots
ground speed. After crossing CHITO, set the ILS frequency in NAV 2 and
set your HSI to match mine.”
4. “Missed approach point will be a decision height of 1514 with 200 set in
the radar altimeter” (XL). Baro minimums 1520 (XLS).
5. “In the event of a missed approach, I’ll start a climb to 3,600 feet. At 3,000
feet, I will turn left direct to ICT VOR and hold.”
6. “We will observe all standard callouts. Any questions?”
ABEAM FAF OR
IAF (OR DOWNWIND VEC TORS) PROCEDURE TURN OUTBOUND
1. APPROACH CHECKLIST—INITIATE 1. FLAPS—15˚
2. AIRSPEED—160 - 180 KIAS 2. AIRSPEED (MIN)—MINIMUM
MANEUVERING SPEED *
INBOUND TO FAF (NORMAL)

1. APPROX. 2 MILES PRIOR TO
FAF—GEAR DOWN
2. FLAPS—35˚
3. AIRSPEED (MIN)—VREF + 10 KT
4. BEFORE LANDING CHECKLIST—
COMPLETED
INBOUND TO FAF (SINGLE-ENGINE)
1. APPROX. 2 MILES PRIOR TO FAF—
GEAR DOWN
2. AIRSPEED (MIN)—VAPP + 10 KT
(WITH FLAPS 15˚)
3. SINGLE ENGINE AND LANDING
MINIMUMS
MINIMUM DESCENT ALTITUDE

1. RUNWAY VISUAL REFERENCES IN SIGHT:
a. CONTINUE APPROACH
b. BEGIN DESCENT AT VISUAL DESCENT POINT
c. LANDING ASSURED (NORMAL)—
VREF CROSSING THRESHOLD
d. LANDING ASSURED (SINGLE ENGINE)—
FLAPS 35˚ AND VREF CROSSING THRESHOLD
2. RUNWAY VISUAL REFERENCES NOT IN SIGHT:
a. CONTINUE TO MISSED APPROACH POINT
b. ACCOMPLISH MISSED APPROACH
NOTE:
IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR
IN EXCESS OF 5 KT.

Figure MAP-13. Nonprecision—Normal/Single Engine

SCAN TRANSFER
The transfer from instruments to visual flight differs with the approach
being made.
Noncoupled Approaches:
• The pilot flying remains on instruments. When reaching DH or MDA
and being advised of continuous visual reference, he progressively
adjusts his scan to visual flight, announces “I am visual,” and lands.
• The pilot not flying, when approaching DH or MDA, adjusts his scan
pattern to include outside visual clues. When the pilot flying announces
that he is “visual,” the pilot not flying assumes the responsibility for
monitoring the instruments and provides continuous advice of warning
flags and deviations from approach tolerances (sink rate, airspeed,
glide slope and localizer) to touchdown.
Coupled Approaches:
• The pilot flying adjusts his scan pattern to include outside visual cues.
When reaching DH and having assured himself of continuous visual
reference, he announces, “I am visual” and lands.
• The pilot not flying concentrates on instruments to touchdown,
advising of warning flags and deviation from approach tolerances.
CIRCLING APPROACHES
A circling approach may follow any authorized instrument approach (Figure MAP-
14). Although the Citation Excel aircraft are in approach category B, category
C minimums are used during the circling approach due to the higher maneuver-
ing airspeeds. A normal instrument approach is flown down to the circling MDA
until visual contact with the airport environment is made. With the airport in sight,
the approach becomes a visual reference approach with a continued cross-check
of the flight instruments. Since it is primarily a visual approach at this point,
configuration and speeds will be the same as for a normal visual approach.
Leaving the final approach fix, minimum maneuvering speed with the flaps
in the LAND position and the landing gear down, reduce the power to pro-
vide a 1,000 fpm rate of descent. When approaching MDA, power should be
added to maintain airspeed while leveling off, thereby reducing the rate of
descent and ensuring that the aircraft does not go below MDA. There are many
recommended circling procedures once the airport is in sight. Any procedure
is acceptable, provided the following criteria are met:
• The airport environment is always in sight.
• A safe and controllable airspeed is maintained.
• MDA is maintained until the aircraft is in position to perform a normal

descent to a landing on the landing runway without excessive
maneuvering.

ABEAM FAF OR
DOWNWIND VEC TORS PROCEDURE TURN OUTBOUND
OR APPROACHING THE IAF 1. FLAPS—15˚
2. AIRSPEED (MIN)—MINIMUM
1. APPROACH CHECKLIST—INITIATE
MANEUVERING SPEED *
2. AIRSPEED—160 - 180 KIAS
INBOUND TO FAF
1. APPROX. 2 MILES PRIOR TO FAF—
MINIMUM DESCENT ALTITUDE GEAR DOWN
1. IF AIRPORT ENVIRONMENT IS IN SIGHT: 2. AT FAF—FLAPS 35˚ (NORMAL) OR
a. CIRCLE/MANEUVER TO LAND FLAPS 15˚ (SINGLE-ENGINE)
b. SPEED—MINIMUM MANEUVER SPEED * 3. AIRSPEED (MIN)—MINIMUM
c. MAX BANK ANGLE—30˚ MANEUVERING SPEED *
2. IF AIRPORT ENVIRONMENT IS NOT IN SIGHT: 4. LANDING CHECKLIST—COMPLETED
a. CONTINUE TO MISSED APPROACH POINT
b. ACCOMPLISH MISSED APPROACH
90˚
ON FINAL
1. AIRSPEED (MIN)—VREF (NORMAL)
OR VAPP (SINGLE ENGINE)
2. IF SINGLE ENGINE—FLAPS 35˚ AND
AIRSPEED VREF WHAN LANDING IS
ASSURED
KE
EP
AIR
PO
RT
E NV
IRO
NM
EN
T IN
SIG
HT

NOTE: TURN TO FINAL

IN GUSTY WIND CONDITIONS, INCREASE VREF/VAPP 1. AIRSPEED (MIN)—MINIMUM
BY 1/2 GUST FACTOR IN EXCESS OF 5 KT. MANEUVERING SPEED *
2. MAX BANK ANGLE—30˚
Figure MAP-14. Circling Approach
MISSED APPROACH OR GO-AROUND

In the event of a missed approach or a go-around, simultaneously push the
throttle levers to the TO detent, while pressing the go-around button (Figures
MAP-15 and MAP-16). Pressing the go-around button will cancel all modes
set in the flight director and command a 10° nose up pitch attitude. Call for
flaps APPROACH (flaps 15 or flaps 7 if climb gradient is a factor) and press
the heading button on the flight director control panel.
If a GPS approach (or overlay) was programmed into the FMS and the missed
approach procedure is sequenced by use of the go-around button, the pilot fly-
ing may elect to press the NAV button on the flight director instead of the head-
ing button and follow the missed approach by way of the FMS.

MAP-32
MAXIMUM THRUST NORMAL CLIMB THRUST

CLIMB
FLAP RETRACTION 1. CLIMB AS REQUIRED
DECISION POINT 1. AT A PRE-DETERMINED ALTITUDE, 2. THROTTLES—MCT, OR
AS REQUIRED

SIMULTANEOUSLY: CONSIDERING TERRAIN, AND AT
3. AFTER TAKEOFF/CLIMB
1. SELECT GO-AROUND A MINIMUM AIRSPEED OF

VAPP + 10 KT—FLAPS UP CHECKLIST—COMPLETED
2. APPLY MAX POWER
3. ROTATE 10˚ NOSE UP (COMMAND BARS) 2. ACCELERATE TO NORMAL CLIMB
4. CHECK/SET FLAPS TO 15˚ SPEED
5. SELECT HDG OR NAV ON F/D
POSITIVE RATE
1. GEAR—UP
"GO-AROUND"
AIRPORT
Figure MAP-15. Missed Approach—Normal

MAXIMUM THRUST MAXIMUM CONTINUOUS
CLIMB
DECISION POINT FLAP RETRACTION 1. CLIMB AS REQUIRED AT VENR

2. THROTTLES—MCT, OR AS
1. AT VAPP + 10 KT (MINIMUM)—
SIMULTANEOUSLY: REQUIRED
FLAPS UP
1. SELECT GO-AROUND 3. AFTER TAKEOFF/CLIMB
2. ACCELERATE TO VENR
2. APPLY MAX POWER ON CHECKLIST—COMPLETED
GOOD ENGINE
3. ROTATE TO COMMAND BARS POSITIVE RATE
(10˚ NOSE UP ATTITUDE)
4. CHECK/SET FLAPS TO 15˚ 1. GEAR—UP
5. SELECT HDG OR NAV ON F/D 2. AIRSPEED—VAPP UNTIL
1,500' AGL OR CLEAR
OF OBSTACLES,
WHICHEVER IS HIGHER
1,500' AGL (MIN)

"GO-AROUND"
AIRPORT
MAP-33
Figure MAP-16. Missed Approach—Single Engine

As with the stall recovery procedures, as the engines accelerate, they will tend
to force the nose down. It will be necessary to increase the back pressure on
the yoke to maintain a pitch-up attitude. Once a positive rate of climb is es-
tablished, call for gear up and FLC mode on the flight director, which should
be accomplished by the pilot not flying.
Follow the published missed approach procedure or the procedure given by ATC.
If both engines are operating normally, adjust power and pitch as needed, and
climbing safely, maintain a reasonable speed and call for flaps up while ac-
celerating through V APP + 10 KIAS minimum.
If only one engine is available, maintain T/O thrust and adjust pitch as nec-
essary to maintain V APP while climbing to a safe altitude. Leave the flaps in
the APPROACH position until a safe altitude is achieved and accelerating
through V APP +10 KIAS.
The use of FLC is very beneficial to maintaining the best climb gradient. If
speed on the go-around is well above V APP , adjust the pitch to achieve V APP
and press the touch control steering (TCS) button to synchronize the com-
mand bars to the displayed airspeed (or use the pitch trim wheel to adjust FLC
to the desired V APP ).
Some airports may require a minimum missed approach climb gradient.To de-
termine the aircrafts single engine climb performance during missed ap-
proach, consult the “Approach Gross Climb” charts in the AFM.
LANDING PROCEDURES
Figure MAP-17 provides a guideline for a typical landing from a visual approach.
The actual touchdown is on the main gear with a slightly nose-high attitude. After
touchdown, extend the speedbrakes, and apply the wheel brakes as necessary.
NOTE
On single-engine approaches, do not lower the flaps
to LAND until the landing is assured.
After touchdown, extend the speedbrakes, ensure the throttles are at idle and
raise the thrust reverser levers to the deploy position after nosewheel con-
tact. When the DEPLOY light illuminates, the thrust reverser levers may be
raised to apply power to the engines. Do not exceed 75% of takeoff thrust
with the thrust reverser levers. Apply wheel brakes as necessary to stop the
airplane. Ensure the thrust reversers are in idle reverse by 60 KIAS during
the landing roll. When the thrust reversers are no longer needed, return the
thrust reverser levers to the stow position and ensure that all thrust reverser
annunciators extinguish.

ON FINAL
1. FLAPS—35˚ (NORMAL) OR
15˚ (SINGLE-ENGINE)
2. AIRSPEED (MIN)—VREF (FLAPS 35˚)
OR VAPP (FLAPS 15˚)

3. IF SINGLE-ENGINE, FLAPS 35˚ AND
VREF WHEN LANDING IS ASSURED
DOWNWIND LEG
(1,500' AGL)
1. AIRSPEED—160 - 180 KIAS
2. FLAPS—15˚
ABEAM TOUCHDOWN
1. GEAR—DOWN *
2. BEFORE LANDING
NOTE:
IN GUSTY WIND CONDITIONS, INCREASE VREF BY
1/2 OF THE GUST FACTOR IN EXCESS OF 5 KT.
* IF BEING RADAR-VECTORED TO A VISUAL PATTERN, EXTEND THE GEAR TURN TO FINAL

ON BASE LEG. IF BEING RADAR VECTORED FOR A STRAIGHT-IN APPROACH,
LOWER THE GEAR NOT LATER THAN THREE MILES FROM THE THRESHOLD. 1. BEGIN DESCENT
2. AIRSPEED (MIN)—MINIMUM
** MINIMUM MANEUVERING SPEED IS VREF + 10 KT (FLAPS 35˚) MANEUVERING SPEED **
MAP-35
OR VAPP + 20 KT (FLAPS 15˚).
Figure MAP-17. VFR Approach—Normal/Single Engine

NOTE
Use of thrust reversers is not permitted during touch-
and-go landings.
Due to possible FOD to the engine during taxi, keep

use of the thrust reversers to a minimum.
ADJUSTMENTS TO LANDING DISTANCE

• Antiskid inoperative ........................... Multiply landing distance by 1.6
• Reduced flap landing.......................... Multiply landing distance by 1.4
• Wet runway ............................................ Refer to advisory information,

Section VII, in the AFM.
• Icy runway.............................................. Refer to advisory information,

Section VII, in the AFM.
NOTE
Following excerpt from the Citation Excel/XLS
Operating Manual: Wheel Fusible Plug Considerations
—Brake application reduces the speed of an airplane
by means of friction between the brake stack compo-
nents. The friction generates heat, which increases the
temperature of the brake and wheel assembly, result-
ing in an increased tire pressure. Each main wheel in-
corporates fuse plugs, which melt at a predetermined
temperature, to prevent a possible tire explosion due
to excessively high tire pressure. Flight crews must take
precautions when conducting repetitive traffic cir-
cuits, including multiple landings and/or multiple re-
jected takeoffs, to prevent overheating the brakes,
which could melt the fuse plugs and cause loss of all
tire pressure and possible tire and wheel damage.
During such operations, available runway permit-
ting, minimize brake usage, and consider cooling the
brakes in flight with the landing gear extended.
Maximizing use of reverse thrust and extending speed
brakes will assist in bringing the airplane to a stop.
HYDROPLANING SPEEDS
The formula used to determine the speed at which a tire is likely to hydroplane
on a wet runway is stated as:
Hydroplane Speed = 7.7 Tire Pressure
From the above formula, the nose gear hydroplane speed is about 88 knots
and the main gear about 113 knots.

LANDING LIMITATIONS
The maximum landing weight is restricted by:
1. Maximum certified landing weight (structural).
2. Maximum landing weight permitted by climb requirements.
3. Maximum landing weight permitted by landing field length.
4. Maximum landing weight permitted by brake energy limits.
For high-pressure altitudes and temperatures, the approach climb configura-

tion may be more restrictive and require a lower landing weight than the
landing climb configuration. Therefore, the “Maximum Landing Weight
Permitted by Climb Requirements” chart, found in the AFM, depicts the land-
ing weight as limited by the approach climb (Table MAP-4).
Table MAP-4. LANDING LIMITATIONS
APPROACH CLIMB LANDING CLIMB

SPEED: VAPP (1.3S1) VLC (1.3 VSO)
(APPROACH CLIMB SPEED) (LANDING CLIMB SPEED)
THRUST SETTING: TAKEOFF (ONE ENGINE) TAKEOFF (TWO ENGINE)
FLAP POSITION: TAKEOFF LAND
GEAR POSITION: UP DOWN
REQUIRED CLIMB 2.1% GROSS 3.2% GROSS

GRADIENT
The AFM charts, “LANDING DISTANCE—FEET, Actual Distance,” provide

the horizontal distance necessary to land and come to a complete stop from
a point 50 feet over the runway threshold at V REF (130% of the stall speed in
the landing configuration). At that point, thrust is reduced to idle.

CROSSWIND LANDING
METHOD NO. 1:
The aircraft is flown down final approach with runway centerline alignment
maintained with normal drift correction. Approaching the threshold, lower
the upwind wing to maintain no drift and apply opposite rudder to maintain
alignment with runway centerline. Fly the airplane onto the runway. Do not
allow drift to develop. Keep full aileron deflection during the landing roll.
METHOD NO. 2:
The “crab” or wings-level method may be continued until just before touch-
down. Then, with wings level, apply rudder pressure to align the airplane with
the runway centerline at the moment of touchdown. Fly the airplane onto the
runway. Do not allow drift to develop. Keep full aileron deflection during the
landing roll.
FLAPS INOPERATIVE LANDING

(NOT IN LANDING POSITION)
When planning a reduced flap approach and landing (Figure MAP-18), the
landing weight of the airplane must be considered. An attempt should be
made to reduce this weight if possible, especially if runway length is mar-
ginal, due to the higher approach and landing speeds required for a reduced
flap configuration. Compute the normal V REF and add adjusted speeds.
Program the adjusted V REF for the new reduced flap V REF speed. Fly the final
approach at the adjusted V REF plus 10 knots maximum and reduce to the ad-
justed V REF prior to crossing the threshold.
NOTE
The reduced flap landing distance is 40% longer than
normal.
To preclude excessive float during landing, allow

the airplane to touch down in a slightly flatter atti-
tude than on a normal landing.
NOTE
Reduced flap adjusted V REF speeds:
• FLAPS 15°— V APP
• FLAPS 7°—V REF +10 KIAS
• FLAPS 0° or Unknown—V REF +15 KIAS

ON FINAL
1. SET UP A NORMAL SINK RATE/
VERTICAL PATH
2. PLAN TO REDUCE SPEED TO
ADJUSTED VREF NO LATER THAN
50' ABOVE THRESHOLD
3. TOUCHDOWN WITH MINIMUM

FLARE (APPROX. 300 -500 FPM)
DOWNWIND LEG (1,500' AGL)

1. COMPUTE AND SET ADJUSTED VREF
FOR A REDUCED FLAP LANDING
2. AIRSPEED—ADJUSTED VREF +10 KT
ABEAM TOUCHDOWN
1. GEAR—DOWN *
2. FLAPS INOPERATIVE APPROACH AND
LANDING CHECKLIST—COMPLETED
TURN TO FINAL
1. BEGIN DESCENT
* IF BEING RADAR VECTORED TO A VISUAL PATTERN, 2. MAXIMUM BANK ANGLE—30˚
EXTEND GEAR ON BASE LEG. IF BEING VECTORED 3. AIRSPEED (MIN)—ADJUSTED
FOR A STRAIGHT-IN APPROACH, LOWER GEAR NOT VREF + 10 KT
MAP-39
LATER THAN THREE MILES FROM THE THRESHOLD.
Figure MAP-18. Visual Approach and Landing with Flaps Inoperative

PRACTICAL TEST
The Flight Standards Service of the FAA has developed a Practical Test
Standards (PTS) book, which is used by all examiners in determining the pro-
ficiency of a pilot. The PTS is divided into two sections, “Preflight Preparation
and Preflight Procedures,” and “In-flight Maneuvers and Postflight Procedures.”
Within these sections are specific items that must be tested called “Areas of
Operation.” Within these areas are the tasks to be performed.
Listed below are the areas required by the PTS and a brief description of each.
PREFLIGHT PREPARATION
Task A—Equipment Examination
An oral examination regarding the systems of the aircraft including normal,
abnormal, and emergency operations.
Task B—Performance and Limitations

An evaluation of the performance and limitations of the aircraft using the ap-
propriate manuals and references to determine them.
PREFLIGHT PROCEDURES
Task A—Preflight Inspection
A thorough inspection of the aircraft interior and exterior looking for possi-
ble defects and corrective action, including manuals, quantities, and sur-
rounding area.
Task B—Powerplant Start

Proper procedure for starting and monitoring engines.
Task C—Taxiing
Proper taxi techniques and ground collision avoidance.
Task D—Pretakeoff Checks

Determining if the aircraft is safe for flight including proper airspeeds, en-
gine parameters, performance considerations, and clearances.

TAKEOFF AND DEPARTURE PHASE

Task A—Normal and Crosswind Takeoff
Performing a normal and crosswind takeoff using proper control movements
and power settings
Task B—Instrument Takeoff

Performing a takeoff into instrument meteorological conditions prior to reach-
ing 100 feet AGL.
Task C—Powerplant Failure During Takeoff

Performing a takeoff while experiencing an engine failure after V 1 but prior
to V R . Demonstrating proper control movements and directional control.
Task D—Rejected Takeoff

Performing an aborted takeoff after recognition of an engine or system failure.
Task E—Instrument Departure

Perform an instrument departure using appropriate charts or ATC clearances.
IN-FLIGHT MANEUVERS
Task A—Steep Turns
Perform a turn in IMC with a bank angle of 45° in two different directions.
Task B—Approaches to Stalls

Perform stalls in the clean landing and takeoff, or approach, configurations
in IMC using recommended recovery techniques.
Task C—Powerplant Failure

Demonstrates proper handling techniques during an engine failure, includ-
ing proper shutdown and restart procedures.

INSTRUMENT PROCEDURES
Task A—Instrument Arrival
Perform an instrument arrival to an aerodrome using appropriate charts or ATC
clearances.
Task B—Holding
Enter a published or assigned holding pattern at appropriate speeds and fol-
low ATC instructions.
Task C—Precision Instrument Approaches

Two precision approaches must be performed. One must be manually flown
with a powerplant failure using raw data or a flight director, at the discretion
of the examiner, and it must be completed to a missed approach or a landing.
Task D—Nonprecision Instrument Approaches

At least two nonprecision instrument approaches, one of which must include
a procedure turn, using two different navaids. One of these approaches must
be flown manually without receiving radar vectors.
Task E—Circling Approach

Perform a circling approach to a runway from an instrument approach with
no straight in minimums, or from an instrument approach to a runway other
than the intended runway of landing.
Task F—Missed Approach

At least two missed approaches must be completed. One must be from a pre-
cision approach, one must be a published missed approach procedure and one
must be with one engine inoperative.
LANDINGS AND APPROACHES TO LANDINGS

Task A—Normal and Crosswind Approaches and
Landings
Perform a normal and crosswind landing using proper control techniques, good
directional control, and stabilized airspeed.
Task B—Landing From a Precision Approach

One of the landings required must be from a precision instrument approach.

Task C—Approach and Landing With a Powerplant

Failure
One of the landings must be with an engine failure using proper handling tech-
niques and checklist procedures.
Task D—Landing From a Circling Approach

Perform a landing from a circling approach avoiding excessive bank angles
and rates of descent. Obstacle avoidance, aircraft maneuvering, and descent
from MDA are prime considerations.
Task E—Rejected Landing

Performs a rejected landing from an altitude approximately 50 feet above the
runway threshold, using proper procedures and techniques.
Task F—Landing From a No Flap or a Nonstandard Flap

Approach
Perform a landing without the use of flaps using proper checklist procedures
and airspeed control.
NORMAL AND ABNORMAL PROCEDURES

Demonstrate proper procedures for normal and abnormal system operations.
EMERGENCY PROCEDURES
Demonstrates proper emergency procedures appropriate for aircraft.
POSTFLIGHT PROCEDURES
Demonstrates proper procedures for after landing, taxiing, and ramping of
aircraft following checklist and ATC instructions.
PARKING AND SECURING

Demonstrates proper parking and securing techniques including aircraft
records.

PTS TOLERANCES
The PTS outlines tolerances allowed for each task listed under the “Areas of
Operation.” The tolerances are fairly standard.
Takeoff and Missed Approach

• Headings ±5°
• Airspeeds ±5 knots
• Altitudes ±100 feet
Basic Attitude: Enroute, Steep Turns, etc.

• Altitude ±100 feet
• Airspeed ±10 knots
• Heading ±10°
• Bank angle ±5°
Stalls
• Announces first indication of stall.
• Recovers with minimum loss of altitude.
Precision Approaches
• Needle deviation 1/2 dot
Nonprecision Approaches
• MDA +50, –0 feet
• 1/2 dot deviation or ±5° from course
Circling
• MDA +100, –0 feet
• Angle of Bank—Maximum of 30°
Landings
• Touchdown and stop in a safe manner.

WEIGHT AND BALANCE

CONTENTS
Page
WB-1
DEFINITIONS .................................................................................
WB-2
GENERAL .......................................................................................
WB-2
Weight .....................................................................................
WB-2
Balance....................................................................................
WB-2
Basic Formula .........................................................................
Weight Shift Formula.............................................................. WB-2
Weight Addition or Removal .................................................. WB-3
WB-3
FORMS.............................................................................................
Cessna Aircraft Company’s Computerized
Weight and Balance .............................................................. WB-15
FOR TRAINING PURPOSES ONLY WB-i

ILLUSTRATIONS
Figure Title Page
WB-1 Airplane Weighing Form ............................................ WB-4
WB-2 Weight and Balance Record ........................................ WB-5
WB-3 XLS Crew and Passenger Weight and Moment Table WB-6
WB-4 Excel Crew and Passengers Compartments Weight
and Moment Tables
(Standard Center Club Seat Arrangement) .................. WB-7
WB-5 XLS Baggage and Cabinet Compartments
Weight and Moment Tables ........................................ WB-8
WB-6 Excel Baggage and Cabinet Compartments
Standard Weight and Moment Tables .......................... WB-9
WB-7 Fuel Loading Weight and Moment Table .................. WB-10
WB-8 XLS Center-of-Gravity Limits Envelope Graph ...... WB-11
WB-9 Excel Center-of-Gravity Limits Envelope Graph ...... WB-12
WB-10 XLS Weight-and-Balance Worksheet ........................ WB-13
WB-11 Excel Weight-and-Balance Worksheet ...................... WB-14
WB-12 Weight and Balance Computation Form
(Identical for Excel and XLS) .................................. WB-16
FOR TRAINING PURPOSES ONLY WB-iii

WEIGHT AND BALANCE

DEFINITIONS
Manufacturer’s Empty Weight—Weight of structure, powerplants, furnish-
ings, systems, and other items of equipment that are an integral part of a par-
ticular configuration.
Standard Empty Weight—Manufacturer’s empty weight plus standard items.
Standard Items—Equipment and fluids not an integral part of a particular air-

plane and not a variation for the same type of airplane. These items may in-
clude, but are not limited to, the following:
• Unusable fuel
• Engine oil
• Toilet fluid
• Serviced fire extinguisher
• All hydraulic fluid
• Trapped fuel
Basic Empty Weight—Standard empty weight plus installed optional equipment.
Operational Takeoff Weight—Maximum authorized weight for takeoff. It is
subject to airport, operational, and related restrictions. This is the weight at
the start of the takeoff run and must not exceed maximum design takeoff weight.
Operational Landing Weight—Maximum authorized weight for landing. It is
subject to airport, operational, and related restrictions. It must not exceed max-
imum design landing weight.
Useful Load—Difference between maximum design takeoff weight and basic
empty weight. It includes payload, usable fuel, and other usable fluids not
included as operational items.
Usable Fuel—Fuel available for airplane propulsion.
Unusable Fuel—Fuel remaining after a fuel runout test has been completed
in accordance with governmental regulations. It includes draining unusable
fuel plus unusable portion of trapped fuel.
Trapped Fuel—Fuel remaining when the airplane is defueled by normal means
using the procedures and attitudes specified for draining the tanks.
Actual Zero Fuel Weight—Basic empty weight plus payload. It must not ex-
ceed maximum design zero fuel weight.
Payload—Maximum design zero fuel weight minus basic empty weight. This
is the weight available for crew, passenger baggage, and cargo.
MAC—Mean Aerodynamic Chord. The chord of an imaginary airfoil having
the same mathematical aerodynamic properties of the actual wing.
FOR TRAINING PURPOSES ONLY WB-1

GENERAL
WEIGHT
Airplane maximum weights are predicated on structural strength. It is nec-
essary to ensure the airplane is loaded within the various weight restrictions
to maintain structural integrity.
BALANCE
Balance, or the location of the center of gravity (CG), deals with airplane sta-
bility. The horizontal stabilizer must be capable of providing an equalizing
moment, which is produced by the remainder of the airplane. Since the amount
of lift produced by the horizontal stabilizer is limited, the range of movement
of the CG is restricted so proper airplane stability is maintained.
Stability increases as the CG moves forward. If the CG is out of the for-

ward limit, the airplane may become so “stable” the elevator cannot pro-
duce enough downward lift to be rotated at the proper speed or flared for
landing.
With the CG out of the aft CG limit, the stability decreases. Here the hori-
zontal stabilizer may not have enough nose down elevator travel to counter-
act a nose-up pitching movement. This could result in an unrecoverable stall
possibly ending in a spin.
BASIC FORMULA
Weight x Arm = Moment
This is the basic formula upon which all weight and balance calculations are based.
Remember the CG (arm) can be found by adapting the formula as follows:
Arm (CG) = Moment Weight
WEIGHT SHIFT FORMULA

The above formula can be utilized to shift weight if the CG is found to be out
of limits. Use of this formula avoids working the entire problem over again
by trial and error.
Shifted weight = Distance CG moved
Total weight Distance weight was moved
WB-2 FOR TRAINING PURPOSES ONLY

Example: Condiments weighing 100 pounds are moved from the tail compart-
ment to the refreshment center. Weight and balance previously calculated is
as follows:
Weight of aircraft.................................................................. 19,000 pounds
Current CG location............................................................... 326.16 inches
Weight of condiments................................................................ 100 pounds
Arm of luggage compartment.................................................. 431.0 inches
Arm of refreshment center..................................................... 172.09 inches
Inserting the values into the weight shift formula:

100 Distance CG moved
19,000 = 431.0 - 172.09
Cross multiplying gives the following result:
Distance CG moved = (100) X (431.0 – 172.09) / 19,000
Distance CG moved = 1.36
Since the weight was brought from the luggage compartment to the refreshment
center (weight moved forward, CG moved forward) the new CG would be:
New CG location = 326.16 – 1.36 = 324.8
WEIGHT ADDITION OR REMOVAL

If weight is to be added or removed after the weight and balance has been
computed, a simple formula can be used to figure the new CG.
Weight added (or removed) (X) Distance CG moved
New total weight = Distance between the weight
arm and the old CG arm
If it is desired to find the weight change needed to accomplish a particular

CG change, the formula can be adapted as follows:
Weight addition (or removal) (X) = Distance CG moved

Old total weight Distance between the weight
arm and the new CG arm
FORMS
The Weight and Balance forms are discussed in the following pages. Examples
of the forms are included in Figures WB-1 through WB-12. Forms WB-1
through WB-12 are in the AFM appropriate to the passenger seating and bag-
gage/cabinet configuration of each particular aircraft.

Figure WB-1
The airplane weight, CG arm, and moment (divided by 100) are all listed at
the bottom of this form as the airplane is delivered from the factory (Figure
WB-1). Ensure the basic empty weight figures listed are current and have not
been amended.
Figure WB-1. Airplane Weighing Form

Figure WB-2
The Weight and Balance Record amends the Airplane Weighing Form (Figure
WB-2). After delivery, if a service bulletin is applied to the airplane or if equip-
ment is removed or added that would affect the CG or basic empty weight, it
must be recorded on this form in the AFM. The crew must always have ac-
cess to the current airplane basic weight and moment in order to be able to
perform weight and balance computations.
Figure WB-2. Weight and Balance Record

Figure WB-3 (XLS)

Moment arms and calculated moments/100 are listed for each individual seat
for the standard seat arrangement (Figure WB-3). If an optional seating con-
figuration is installed in the aircraft, ensure the proper chart for that config-
uration is in the AFM.
MOMENT/100
SEAT 1 SEAT 3 SEAT 5 SEAT 7 SEAT 10 SEAT 10 LH

WEIGHT OR OR OR OR FWS AFT SFS
(POUNDS) SEAT 2 SEAT 4 SEAT 6 SEAT 8 ARM = FS ARM = FS ARM = FS
ARM = FS ARM = FS ARM = FS ARM = FS 181.24 IN. 205.60 IN. 357.99 IN.
F.S. 136.32
136.32 IN. 234.39 IN. 286.54 IN. 322.62 IN.
50 68.16 117.20 143.27 161.31 90.62 102.80 179.00

60 81.79 140.63 171.92 193.57 108.74 123.36 214.79
F.S. 181.24
70 95.42 164.07 200.58 225.83 126.87 143.92 250.59
80 109.06 187.51 229.23 258.10 144.99 164.48 286.39
90 122.69 210.95 257.89 290.36 163.12 185.04 322.19 F.S. 205.60
100 136.32 234.39 286.54 322.62 181.24 205.60 357.99

110 149.95 257.83 315.19 354.88 199.36 226.16 393.79
120 163.58 281.27 343.85 387.14 217.49 246.72 429.59 F.S. 234.39
130 177.22 304.71 372.50 419.41 235.61 267.28 465.39
140 190.85 328.15 401.16 451.67 253.74 287.84 501.19
150 204.48 351.59 429.81 483.93 271.86 308.40 536.99
160 218.11 375.02 458.46 516.19 289.98 328.96 572.78
F.S. 286.54
170 231.74 398.46 487.12 548.45 308.11 349.52 608.58
180 245.38 421.90 515.77 580.72 326.23 370.08 644.38
190 259.01 445.34 544.43 612.98 344.36 390.64 680.18
200 272.64 468.78 573.08 645.24 362.48 411.20 715.98
F.S. 322.62
210 286.27 492.22 601.73 677.50 380.60 431.76 751.78
220 299.90 515.66 630.39 709.76 398.73 452.32 787.58
230 313.54 539.10 659.04 742.03 416.85 472.88 823.38
240 327.17 562.54 687.70 774.29 434.98 493.44 859.18 F.S. 357.99
250 340.80 585.98 716.35 806.55 453.10 514.00 894.98
260 354.43 609.41 745.00 838.81 471.22 534.56 930.77
270 368.06 632.85 773.66 871.07 489.35 555.12 966.57
280 381.70 656.29 802.31 903.34 507.47 575.68 1002.37
290 395.33 679.73 830.97 935.60 525.60 596.24 1038.17
300 408.96 703.17 859.62 967.86 543.72 616.80 1073.97
310 422.59 726.61 888.27 1000.12 561.84 637.36 1109.77
320 436.22 750.05 916.93 1032.38 579.97 657.92 1145.57
330 449.86 773.49 945.58 1064.65 598.09 678.48 1181.37
340 463.49 796.93 974.24 1096.91 616.22 699.04 1217.17
Figure WB-3. XLS Crew and Passenger Weight and Moment Table

Figure WB-4 (Excel)

Moment arms and calculated moments/100 are listed for each individual seat
for the standard center club seat arrangement (Figure WB-4). If an optional
seating configuration is installed in the aircraft, ensure the proper chart for
that configuration is in the AFM.
Figure WB-4. Excel Crew and Passengers Compartments Weight

and Moment Tables (Standard Center Club Seat
Arrangement)

Figure WB-5 (XLS)

This form contains the arms and moments/100 for each compartment of the
standard configuration aircraft (Figure WB-5). The maximum weight listed
is the maximum placarded weight for each compartment. Remember this
limit is structural in nature. It is based on the maximum weight the flooring
in that area can support.
LH & RH AFT CLOSET

MOMENT/100 MOMENT/100
NAVIGATION AFT
CHART CASE CLOSET
WEIGHT ARM = WEIGHT ARM =
(POUNDS) FS 158.10 IN (POUNDS) FS 374.00 IN
5 7.91 5 18.70
10 15.81 10 37.40
15 23.72 15 56.10
20 74.80
25 93.50
30 112.20
CHART CASES 35 130.90
RH FORWARD 40 149.60
45 168.30
CLOSET 50 187.00 FS 158.10
MOMENT/100
55 205.70 FS 166.38
FORWARD
CLOSET
60 224.40 FS 173.20
WEIGHT ARM = 65 243.10
(POUNDS) FS 166.38 IN 68 254.32
5 8.32
10 16.64 BAGGAGE
15 24.96 COMPARTMENT
20 33.28 CONTENTS
25 41.60
MOMENT/100
30 49.91 TAIL CONE
35 58.23 COMPARTMENT
40 66.55 WEIGHT ARM =
(POUNDS) FS 431.00 IN
45 74.87
50 83.19 20 86.20
40 172.40
56 93.17
60 258.60
80 344.80
LH 100 431.00
REFRESHMENT 120
140
517.20
603.40
CENTER 160 689.60
MOMENT/100 180 775.80
WEIGHT REFRESHMENT 200 862.00
(POUNDS) CENTER 220 948.20
ARM = 240 1034.40
FS 173.20 in. 260 1120.60
10 17.32 280 1206.80
20 34.64 300 1293.00
30 51.96 320 1379.20
40 69.28 340 1465.40
50 86.60 360 1551.60
60 103.92 380 1637.80 FS 374.00
70 121.24 400 1724.00
80 138.56 420 1810.20
90 155.88 440 1896.40
100 173.20 460 1982.60
110 190.52 480 2068.80
120 207.84 500 2155.00
130 225.16 520 2241.20
141 244.21 540 2327.40
560 2413.60
FS 431.00
580 2499.80
600 2586.00
620 2672.20
640 2758.40
660 2844.60
680 2930.80
700 3017.00
Figure WB-5. XLS Baggage and Cabinet Compartments

Weight and Moment Tables

Figure WB-6 (Excel)

This form contains the arms and moments/100 for each compartment of the
standard configuration aircraft (Figure WB-6). The maximum weight listed
is the maximum placarded weight for each compartment. Remember this
limit is structural in nature. It is based on the maximum weight the flooring
in that area can support.
Figure WB-6. Excel Baggage and Cabinet Compartments

Standard Weight and Moment Tables

Figure WB-7 (Excel/XLS)

All of the weight and moment tables have arms listed for various locations
except the fuel table. Notice the arm varies depending on the quantity of use-
able fuel (Figure WB-7).
Figure WB-7. Fuel Loading Weight and Moment Table

Figure WB-8—Center-of-Gravity Envelope (XLS)

After summing all the weights and moments, and calculating the CG for the
flight, it is necessary to determine whether the CG is within allowable lim-
its (CG envelope) (Figure WB-8). To plot the location of the CG on the graph,
follow the horizontal weight line of the loaded aircraft to the corresponding
vertical line for the calculated CG. If the intersection falls in the CG enve-
lope, the aircraft is loaded within limits.
Figure WB-8. XLS Center-of-Gravity Limits Envelope Graph

Figure WB-9—Center-of-Gravity Envelope (Excel)

After summing all the weights and moments, and calculating the CG for the
flight, it is necessary to determine whether the CG is within allowable lim-
its (CG envelope) (Figure WB-9). To plot the location of the CG on the graph,
follow the horizontal weight line of the loaded aircraft to the corresponding
vertical line for the calculated CG. If the intersection falls in the CG enve-
lope, the aircraft is loaded within limits.
Figure WB-9. Excel Center-of-Gravity Limits Envelope Graph

Figure WB-10—Weight and Balance Worksheet (XLS)

A step-by-step precess is outlined for determining weight and CG limits by
this form (Figure WB-10). The payload computations are made in the left col-
umn, while the rest of the computations are done in the right column.
Figure WB-10. XLS Weight-and-Balance Worksheet

Figure WB-11—Weight-and-Balance Worksheet (Excel)

A step-by-step precess is outlined for determining weight and CG limits by
this form (Figure WB-11). The payload computations are made in the left col-
umn, while the rest of the computations are done in the right column.
Figure WB-11. Excel Weight-and-Balance Worksheet

CESSNA AIRCRAFT COMPANY’S COMPUTERIZED

WEIGHT AND BALANCE
Included with each new aircraft’s publication package, is a diskette contain-
ing individualized Weight and Balance information that allows the flight
crew to compute weight and balance from a PC. Diskette information is for-
matted in Microsoft Excel.
Operating Instructions
After loading the diskette into a PC:
1. Double click to open.
2. Click on “Enable Macros.”
3. A menu chart listing various seating options will appear over the Weight
and Balance Form (Figure WB-12).
NOTE
Only the Excel form is shown. XLS procedures are
identical.
• Select appropriate seat option for aircraft (Forms WB-4 through WB-6).
• Click, OK. Appropriate Weight and Balance form will display the
aircraft’s Basic Empty Weight and Moment in block 1 (right side) and
the selected seating option.
4. Complete left side of form with appropriate weights. Type in the weights
or use a weight chart by clicking the gray box adjacent to the arm in the
weight column.
5. Payload (subtotal) will automatically calculate as each weight is entered.

Concurrently, right side of form will display automatic calculation of
PAYLOAD WEIGHT and MOMENT and ZERO FUEL WEIGHT and
MOMENT in block 3.
6. The Center-of-Gravity envelope on the bottom of the form will continually

plot current CG locations in 400-pound increments.
7. Click on “COMPUTE” box at the top of the form to insert ramp fuel in
block 4, FUEL LOADING.
NOTE
If ZFW CG is out of the envelope a message will ap-
pear to, “please check your inputs and try again.” Fuel
loading cannot be inserted until ZFW CG is adjusted.

Figure WB-12. Weight and Balance Computation Form

(Identical for Excel and XLS)

8. After ramp fuel weight is inserted, the program will prompt to insert “fuel
reserves,” (included in the ramp fuel weight).
NOTE
If the ramp fuel weight inserted would cause the air-
craft weight to exceed Maximum Ramp Weight in
block 5, fuel loading in block 4 will automatically ad-
just not to exceed 20,200 (20,400 for XLS) pounds
in block 5.
9. Block 6, LESS FUEL FOR TAXIING, is protected and cannot be changed,

(200 pounds).
10. Block 7, TAKEOFF WEIGHT, will automatically compute after block 4,

FUEL LOADING, is inserted.
11. Block 8, LESS FUEL TO DESTINATION, is computed automatically by

subtracting reserve and taxi (200 pounds) fuel from ramp fuel inserted in
block 4.
12. Block 9, LANDING WEIGHT, is automatically calculated by adding

reserve fuel to ZFW (block 3).
13. Completed form will not allow CG out of the envelope (refer to CG plot
on Center-of-Gravity envelope on bottom of form).
14. The form may now be printed if desired.
15. If desired, saving flight crew weights and various cabinet compartment
weights (if they remain constant), will essentially save the form as Basic
Operating Weight (BOW). Calculating further trips may then be computed
by inserting only passenger weights, baggage compartment weights and
fuel.

PERFORMANCE
CONTENTS
Page
AIRPLANE FLIGHT MANUAL (AFM) PERFORMANCE
SPECIFICATIONS ......................................................................... PER-1
General .................................................................................. PER-1
Standard Performance Conditions......................................... PER-1
Variable Factors Affecting Performance ............................... PER-3
Definitions ............................................................................. PER-4
FLIGHT PLANNING—XLS ......................................................... PER-8
Specifications ........................................................................ PER-8
Takeoff Performance ........................................................... PER-10
Climb Performance ............................................................. PER-21
Cruise Performance............................................................. PER-22
Descent Performance .......................................................... PER-24
Reserve Fuel........................................................................ PER-25
Holding Performance .......................................................... PER-25
Landing Performance .......................................................... PER-26
Stall Speeds ......................................................................... PER-30
Mission Planning................................................................. PER-31
FLIGHT PLANNING—EXCEL.................................................. PER-35
Specifications ...................................................................... PER-35
Takeoff Performance ........................................................... PER-38
Climb Performance ............................................................. PER-49
Descent Performance .......................................................... PER-54
Fuel Reserves ...................................................................... PER-55
Holding Fuel ....................................................................... PER-55
Landing Performance .......................................................... PER-56
Mission Planning................................................................. PER-61
FOR TRAINING PURPOSES ONLY PER-i

SPECIAL PROCEDURES—XLS and EXCEL ........................... PER-65

Short Field Operation.......................................................... PER-65
Adverse Field Conditions.................................................... PER-66
Engine Anti-Ice ................................................................... PER-67
Passenger Comfort .............................................................. PER-68
Bird Ingestion Precautions .................................................. PER-69
Turbulent Air Penetration.................................................... PER-69
Cold Weather Operation...................................................... PER-69
Ground Deice/Anti-ice Operations ..................................... PER-71
SERVICING—XLS and EXCEL ................................................. PER-71
Fuel...................................................................................... PER-71
Oil........................................................................................ PER-73
Hydraulic............................................................................. PER-74
Oxygen ................................................................................ PER-74
Fire Bottles .......................................................................... PER-74
Landing Gear and Brakes Pneumatic System ..................... PER-75
Tires..................................................................................... PER-75
Toilet ................................................................................... PER-75
Airplane Cleaning and Care ................................................ PER-75
Deice Boots ......................................................................... PER-76
Engines................................................................................ PER-77
Interior Care ........................................................................ PER-77
Windows and Windshields .................................................. PER-78
Oxygen Masks..................................................................... PER-79
PER-ii FOR TRAINING PURPOSES ONLY

TABLES
Table Title Page
XLS
PER-1 Decision, Rotation and Takeoff Safety Speeds ........ PER-10
PER-2 Takeoff Field Length—15° Flaps ............................ PER-11
PER-3 Takeoff Field Length—7° Flaps .............................. PER-16
PER-4 250 KIAS/M 0.65 Climb ........................................ PER-21
PER-5 High-Speed Cruise .................................................. PER-22
PER-6 Long-Range Cruise .................................................. PER-23
PER-7 High Speed and Normal Descent ............................ PER-24
PER-8 Holding Speed and Fuel Flow.................................. PER-25
PER-9 Landing Distance—Actual ...................................... PER-26
PER-10 Stall Speeds.............................................................. PER-30
PER-11 Wind Correction Factors .......................................... PER-31
PER-12 Flight Time and Fuel Burn for
Selected Distances .................................................. PER-32
PER-13 Range/Payload Capability........................................ PER-34
PER-14 Decision, Rotation and Takeoff Safety Speeds ........ PER-38
EXCEL
PER-15 Takeoff Field Length—15° Flaps ............................ PER-39
PER-16 Takeoff Field Length—7° Flaps .............................. PER-44
PER-17 Climb Speeds .......................................................... PER-49
PER-18 Maximum Rate Climb.............................................. PER-50
PER-19 250 Knot/.62 Mach Cruise Climb............................ PER-51
PER-20 High-Speed Cruise .................................................. PER-52
PER-21 Long-Range Cruise .................................................. PER-53
PER-22 Normal and High Speed Descent ............................ PER-54
PER-23 Holding Speed and Fuel Flow.................................. PER-55
PER-24 Landing Distance .................................................... PER-56
PER-25 Stall Speed .............................................................. PER-60
PER-24 Landing Distance (Cont).......................................... PER-60
PER-26 Wind Correction Factors .......................................... PER-61
PER-27 Flight Time and Fuel Burn For
Selected Distances .................................................. PER-62
PER-28 Range/Payload Capability........................................ PER-64
FOR TRAINING PURPOSES ONLY PER-iii

PERFORMANCE
AIRPLANE FLIGHT MANUAL (AFM)
PERFORMANCE SPECIFICATIONS
GENERAL
Certification
The Model 560XL is certified under CFR Part 25, which governs the certifi-
cation of transport category airplanes. Part 25 performance requirements en-
sure specific single-engine climb capability throughout the flight.
Approved Airplane Flight Manual (AFM)

In accordance with Part 25, Airplane Flight Manual (AFM), Section IV,
Performance Section, contains only single-engine takeoff and climb data. All
takeoff data is based upon losing thrust on one engine at the worst possible
moment—near or right at V 1. The AFM contains no enroute cruise informa-
tion, but does contain landing data. This data is based upon the conditions,
factors and assumptions discussed below.
STANDARD PERFORMANCE CONDITIONS

All performance data in the AFM is based on flight test data and accessory losses.
1. Thrust ratings, including engine installation bleed air and accessory losses.
2. Full temperature accountability within the operational limits for which the
airplane is certified.
NOTE
Should ambient air temperature or altitude be below
the lowest temperature or altitude shown on the per-
formance charts, use the performance at the lowest
value shown.
Flap Handle Position Flap Deflection
a. Takeoff TO 7°
b. Takeoff TO/APPR 15°
c. Enroute UP 0°
d. Approach TO/APPR 15°
e. Landing LAND 35°
3. All takeoff and landing performance data is based on a paved, dry or wet
runway.
4. The takeoff performance data was obtained using the following

procedures and conditions.
FOR TRAINING PURPOSES ONLY PER-1

Single Engine Takeoff—Accelerate Go

a. Power was set static in the TO DETENT and verified to correspond to
Figure 4-8, AFM (Takeoff/Go Around Thrust Settings), and then the
brakes were released.
b. The pilot recognized engine failure at V1.
c. Positive rotation to +10° was made at VR and pitch was adjusted to
achieve V2 by 35 feet AGL dry runway.
d. The landing gear was retracted when a positive climb rate was
established.
e. V2 was maintained from the 35-foot point above the runway to 1,500
feet AGL.
f. The airplane was accelerated to V2 +10 KIAS at which time the flaps
were retracted and the acceleration continued to VENR. Power was
reduced to the climb detent and the climb was continued.
Takeoff—Accelerate Stop
Figure 4-8, AFM (Takeoff/Go-Around Thrust Settings), then brakes
were released.
b. The pilot recognized the necessity to stop because of engine failure or
other reasons just prior to V1.
c. Maximum pilot braking effort was initiated at V1 and continued until
the airplane came to a stop.
d. Both throttles were brought to idle immediately after brake
application.
e. Directional control was maintained through the rudder pedals and
differential braking as required.
f. Antiskid was ON during tests.
g. Speedbrakes were not used.
h. Thrust reversers were not used.
i. Wet runways only, for thrust reverser credit, the thrust reverser on the
operating engine was deployed immediately after the throttle reached
idle. Maximum reverse thrust was selected immediately after thrust
reverser deployed and was maintained to 60 KIAS, followed thereafter
by idle reverse thrust until the airplane came to a stop.
Multiengine Takeoff
Figure 4-8, AFM (Takeoff/Go-Around Thrust Settings) then brakes
were released.
PER-2 FOR TRAINING PURPOSES ONLY

b. Positive rotation to +10° was made at VR and pitch adjusted to achieve

V2 +10 by 35 feet AGL.
c. The landing gear was retracted when a positive climb rate was
established. Flaps were retracted at 400 feet.
5. Landing performance data was obtained using the following

procedures and conditions:
Landing
a. Landing preceded by a steady 3° angle approach down to the 50-foot
height point with airspeed at V REF in the landing configuration
(Flaps—LAND, Gear—Extended).
b. Two-engine thrust setting during approach was selected to maintain the

3° approach angle at VREF.
c. Idle thrust was established at the 50-foot height point and the throttles
remained at that setting until the airplane stopped.
d. A minimal rotation to a landing attitude was accomplished to ensure a

firm touchdown on the main gear.
e. Maximum wheel braking was applied immediately on nosewheel

contact and continued throughout the landing roll.
f. The antiskid system was ON during all tests.
g. Speedbrakes were not used.
h. Thrust reversers were not used.
VARIABLE FACTORS AFFECTING PERFORMANCE

Details of variables affecting performance are given with tables in the AFM
to which they apply. Assumptions which relate to all performance calcula-
tions, unless otherwise stated, are:
1. Cabin pressurization.
2. Anti-ice OFF.
3. Humidity corrections on thrust have been applied according to applicable
regulations.
4. Wind correction information is presented on the charts in the AFM. They
are taken as tower winds, 32.8 feet (10 meters) above runway surface.
Factors have been applied as prescribed in the applicable regulations. In
the tables, negative represents tailwind and positive represents headwind.
5. Gradient correction factors can be applied to gradients less than or equal to
2% downhill or 2% uphill. In the AFM tables, negative represents downhill
gradients and positive represents uphill gradients.

DEFINITIONS
Accelerate-Stop Distance—The distance required to accelerate to V 1 and
abort the takeoff and come to a complete stop with maximum braking applied
at V 1 .
Airport Barometric Altitude—Indicated altitude with altimeter set to air-

port altimeter setting while at airport elevation.
Altitude—All altitudes used in the AFM are pressure altitudes unless

otherwise stated.
Anti-ice Systems—The following systems comprise the anti-ice systems

which affect performance in the AFM:
1. Engine Anti-ice.
2. Wing Anti-ice.
Performance, when referred to ANTI-ICE ON, is based on all systems
being operated at the same time.
The pitot-static and angle-of-attack anti-ice system and horizontal tail

deice do not affect performance.
Calibrated Airspeed (KCAS)—Indicated airspeed (knots) corrected for po-

sition error and assumes zero instrument error.
Cat II—Category II operation. A straight-in ILS approach to the runway of

an airport under Category II ILS instrument approach procedure.
Climb Gradient—The ratio of the change in height during a portion of a climb

to the horizontal distance traversed in the same time interval (gradient = rise
over run).
Deice Systems—The horizontal stabilizer boots are the only deice system.
Demonstrated Crosswind—The demonstrated crosswind velocity of 24 knots

(measured at 10 meters above runway surface) is the velocity of the crosswind
component for which adequate control of the airplane during takeoff and land-
ing was actually demonstrated during certification tests. This is not limiting.
Engine Out Accelerate-Go Distance—The horizontal distance from brake

release to the point at which the airplane attains a height of 35 feet above the
runway surface “dry” or 15 feet “wet” (reference zero), on a takeoff during
which an engine is recognized to have failed at V 1 and the takeoff continued.
Gross Takeoff Flight Path—The takeoff flightpath that the airplane can
actually achieve under ideal conditions.
Gross Climb Gradient—The climb gradient that the airplane can actually
achieve with ideal ambient conditions (smooth air).

Indicated Airspeed (KIAS)—Airspeed indicator reading (knots). Zero in-

strument error is assumed.
Indicated Mach Number—The displayed Mach number value includes

position error.
ISA—International Standard Atmosphere; +15°C SL (standard), subtract 2°

per thousand feet altitude increase.
Landing Distance—The distance from a point 50 feet above the runway sur-
face to the point at which the airplane comes to a full stop on the runway.
Landing Field Length—Landing distance adjusted for operational factors.
Level Off Altitude—The barometric altitude at which second segment climb

ends.
Mach Number—The ratio of true airspeed to the speed of sound.
Net Climb Gradient—The gross climb gradient reduced by 0.8% during the
takeoff phase and 1.1% during enroute. This conservatism is required by spe-
cial clearance determinations to account for variables encountered in service.
Net Takeoff Flightpath—Takeoff flightpath used to determine obstacle clear-

ance. Uses net climb gradients to climb to a height of 1,500 feet above the run-
way surface.
OAT—Outside Air Temperature or Ambient Air Temperature. The free air static
temperature obtained either from ground meteorological sources or from in-flight
temperature indications, adjusted for instrument error and compressibility ef-
fects. Used interchangeably with Temperature (refer to Performance Tables, AFM).
Position Correction—A correction applied to indicated airspeed or altitude

to eliminate the effect of the location of the static pressure source on the in-
strument reading. No position corrections are required when using perform-
ance section charts in Section IV of the AFM, since all airspeeds and altitudes
in Section IV are presented as “indicated” values, except for stall speeds which
are presented as “calibrated” values.
RAT—Ram Air Temperature. Indicated outside air temperature as read from

the RAT display. This must be corrected for ram air temperature rise to ob-
tain true outside air temperature, (subtract ram air temperature rise from
RAT display to obtain true air temperature).
Reference Zero—The point in the takeoff flight path at which the airplane
is 35 feet (dry runway) or 15 feet (wet runway) above the takeoff surface and
at the end of the takeoff distance required.
Residual Ice—That ice which is not completely removed from the leading
edge stagnation areas of the wing and horizontal stabilizer by the surface anti-
ice/deice systems during operation in icing conditions. Refer to Section III
and IV of the AFM for applicable procedures.

Takeoff Climb Increment (TCI)—Altitude increment to be added to the

airport barometric altitude to obtain level-off altitude. This increment
includes corrections for nonstandard temperature.
Takeoff Field Length—The takeoff field length given for each combination
of gross weight, ambient temperature, altitude, wind, and runway gradients
is the greatest of the following:
1. 115% of the two-engine horizontal takeoff distance from start (static) to a

height of 35 feet above the runway surface.
2. Accelerate-stop distance, wet or dry runway, as appropriate.
3. The engine-out accelerate-go distance to 35 feet for dry runways and 15

feet for wet runways.
No specific identification is made on the charts (see AFM) concerning

which of these distances governs a specific case.
True Airspeed (KTAS)—The airspeed (knots) of an airplane relative to

undisturbed air.
True Mach Number—The displayed Mach with position error removed.
V 1 —Takeoff Decision Speed. The distance to continue the takeoff to 35 feet

(dry runway) or 15 feet (wet runway) will not exceed the scheduled takeoff
field length if recognition occurred at V 1 (accelerate-go). The distance to bring
the airplane to a full stop (accelerate-stop) will not exceed the scheduled take-
off field length provided that maximum brakes are applied at V 1 .
V 2—Takeoff Safety Speed. The climb speed is the actual speed at 35 feet above
the runway surface as demonstrated in flight during takeoff with one engine
inoperative.
V 35 —Actual speed at 35 feet above the runway surface as demonstrated in

flight during takeoff with both engines operating.
V A —Maximum Maneuvering Speed. The maximum speed at which applica-

tion of full available aerodynamic control will not overstress the airplane. V A
speed is a function of weight versus altitude.
V APP —Landing approach airspeed (1.3 V S1 ) with 15° flap position, landing
gear up.
V ENR —Single-engine enroute climb speed (V YSE ) or best rate-of-climb sin-

gle-engine. The Excel utilizes one reference speed, 160 KIAS at all weights.
V FE —Maximum Flap Extended Speed. The highest speed permissible with

wing flaps in a prescribed extended position.
V LE —Maximum Landing Gear Extended Speed. The maximum speed at

which an airplane can be safely flown with the landing gear extended.

V LO —(Extension). Maximum Landing Gear Extension Speed. The maxi-

mum speed at which the landing gear can be safely extended.
V LO—(Retraction). Maximum Landing Gear Retracting Speed. The maximum

speed at which the landing gear can be safely retracted.
V MCA —Minimum airspeed in the air in the takeoff configuration at which

directional control can be maintained when one engine is suddenly made in-
operative. V MCA is a function of engine thrust which varies with altitude and
temperature. The V MCA of 90 KIAS was determined at maximum takeoff
thrust and maximum takeoff weight.
V MCG —Minimum speed on the ground in the takeoff configuration at which

directional control can be maintained when one engine is suddenly made in-
operative, using only aerodynamic controls. V MCG is a function of both air-
plane weight and engine thrust which varies with altitude and temperature.
AC configuration airplanes, V MCG is 98 KIAS and was determined for max-
imum takeoff thrust. AB configuration airplanes, V MCG is 81 KIAS and was
determined for maximum takeoff thrust and the rudder bias system operational.
V MCL —Minimum airspeed in the air, in the landing configuration, at which

directional control can be maintained, when one engine is suddenly made in-
operative. V MCL is a function of engine thrust which varies with altitude and
temperature. V MCL of 92 KIAS was determined at maximum takeoff thrust
and maximum landing weight.
V MO/MMO —Maximum Operating Limit Speed.
V R —The speed at which rotation is initiated during takeoff to attain V 2 climb

speed at or before a height of 35 feet above the runway surface has been reached.
V REF —The airspeed equal to the landing 50-foot point speed (1.3 V SO ) with
full flaps and landing gear extended.
V SB —Maximum operating speed with speedbrakes in the extended position.
V SO —The stalling speed or the minimum steady flight speed in the landing
configuration.
V S1—The stalling speed or the minimum steady flight speed obtained in a spec-
ified configuration.
Visible Moisture—Visible moisture includes but is not limited to, the fol-
lowing conditions: fog with visibility less than one mile, wet snow and rain.
Wet Runway—A runway is considered wet when there is sufficient moisture

on the surface to appear reflective, but without significant areas of standing
water.
Wind—The wind velocities recorded as variables on the charts in the AFM

are to be understood as the headwind or tailwind components of the actual
winds at 32.8 feet (10 meters) above the runway surface (tower winds).

FLIGHT PLANNING—XLS
This Flight Planning guide is for the purpose of providing specific informa-
tion for evaluating the performance of the Cessna Citation XLS (Model 560XL).
This guide is developed from Flight Manual and Operating Manual data.
This document is not intended to be used in lieu of the FAA approved
Airplane Flight Manual (AFM) or Operating Manual. The data included
herein does not constitute an offer and is subject to change without notice.
SPECIFICATIONS

SPECIFICATIONS
Basic Performance
Takeoff Distance, Sea Level, ISA, MTOW 3,560 ft 1,085 m
Landing Distance, Sea Level, ISA, MLW 3,180 ft 969 m
Rate of Climb - 2 Engines 3,500 ft/min 1,067 m/min
Rate of Climb - 1 Engine 800 ft/min 244 m/min
Typical Cruise Speeds 415 - 435 KTAS
Airspeed Limitations
Maximum Operating Limit
MMO (26,515 ft / 8,082 m and above) M 0.75 Indicated
VMO (8,000 ft to 26,515 ft / 8,082 m) 305 KIAS 565 km/hr
VMO (Below 8,000 ft / 2,438 m) 260 KIAS 482 km/hr
Maximum Flap Speed (VFE)
Partial Flaps - 7° & 15° 200 KIAS 371 km/hr
Full Flaps - 35° 175 KIAS 324 km/hr
Max Landing Gear Oper - Extending (VLO) 250 KIAS 463 km/hr
Max Landing Gear Oper - Retracting (VLO) 200 KIAS 371 km/hr
Max Landing Gear Extended Speed (VLE) 250 KIAS 463 km/hr
Max. Speed Brake Operation Speed (VSB) No limit No limit
Minimum Control Speed, Air (VMCA) 90 KIAS 167 km/hr
Minimum Control Speed, Ground (VMCG) 81 KIAS 150 km/hr
Certified Weights
Maximum Ramp Weight 20,400 lb 9,253 kg
Maximum Takeoff Weight 20,200 lb 9,163 kg
Maximum Landing Weight 18,700 lb 8,482 kg
Maximum Zero Fuel Weight 15,100 lb 6,849 kg
Maximum Fuel Capacity (6.7 lb/gal) 6,740 lb 3,057 kg
Basic Operating Weight

Typically-Equipped Empty Weight 12,400 lb 5,625 kg
Two Crew & Furnishings 400 lb 181 kg
Basic Operating Weight 12,800 lb 5,806 kg
Payload
Useful Payload and Fuel 7,600 lb 3,447 kg
Maximum Payload 2,300 lb 1,043 kg
Payload at Full Fuel 860 lb 390 kg

TAKEOFF PERFORMANCE
14 CFR FAR 25 takeoff field lengths are shown on the following pages. FAR
25 defines takeoff distance as the greater of accelerate-stop, accelerate-go with
one engine inoperative, or 115% of the all engine takeoff distance to a point
35 feet above the runway. These factors are reflected in the takeoff field
lengths presented.
Second segment climb limitations are presented at the bottom of each take-
off field length table. Second segment climb refers to the ability of the air-
craft to meet certain climb rates after takeoff with one engine inoperative.
Second segment climb limitations are a function of temperature, elevation,
and aircraft weight.
Two flap settings are shown for the aircraft: 15° and 7°. A flap setting of 15°
is preferred to minimize runway length and runway speeds. In those situa-
tions where second segment climb requirements are too limiting for 15° of
flaps, a 7° flap setting is available. A 7° flap setting requires greater runway
length but provides greater second segment climb capability.
A paved, level, dry runway with zero wind is assumed. Runway lengths shown
are based on the aircraft anti-ice systems being off and the cabin bleed air on.
Table PER-1. Decision, Rotation and Takeoff Safety Speeds

Table PER-2. TAKEOFF FIELD LENGTH—15° FLAPS
TAKEOFF PERFORMANCE
TAKEOFF FIELD LENGTH - 15° FLAPS
(Over 35 Foot Screen Height)
Dry Runway, Zero Wind, Anti-Ice Off, Cabin Bleed Air On
Elevation = Sea Level

Ambient Temp ------------------------------------------- Takeoff Weight (lb) --------------------------------------------
°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 3,390 3,330 3,200 2,940 2,700 2,660 2,670 2,710
10 / 50 3,500 3,440 3,310 3,040 2,790 2,730 2,750 2,790
15 / 59 3,560 3,490 3,360 3,090 2,830 2,770 2,790 2,820
20 / 68 3,610 3,550 3,410 3,130 2,870 2,810 2,830 2,860
25 / 77 3,690 3,620 3,480 3,200 2,930 2,830 2,850 2,880
30 / 86 3,940 3,870 3,700 3,390 3,100 2,830 2,740 2,760
35 / 95 4,270 4,190 4,000 3,630 3,310 3,010 2,730 2,630
40 / 104 4,710 4,620 4,400 3,990 3,600 3,250 2,930 2,640
45 / 113 5,470 5,330 4,990 4,420 3,980 3,560 3,180 2,850
50 / 122 — — — 5,050 4,390 3,900 3,470 3,080
Climb Wght Temp
Limits °C/°F 45/113 45/113 47/117 51/124 54/129 54/129 54/129 54/129
Field Length at
Temp Limits (ft) 5,470 5,330 5,290 5,210 4,960 4,270 3,750 3,310
Elevation = 1,000 Feet

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 3,530 3,480 3,340 3,070 2,810 2,700 2,720 2,750
10 / 50 3,650 3,590 3,450 3,170 2,900 2,780 2,800 2,830
15 / 59 3,710 3,650 3,500 3,220 2,950 2,820 2,840 2,870
20 / 68 3,770 3,710 3,560 3,270 2,990 2,860 2,870 2,900
25 / 77 3,950 3,880 3,720 3,410 3,120 2,850 2,810 2,840
30 / 86 4,270 4,190 4,000 3,630 3,320 3,020 2,740 2,720
35 / 95 4,650 4,560 4,350 3,940 3,560 3,230 2,920 2,630
40 / 104 5,220 5,090 4,790 4,330 3,900 3,500 3,140 2,820
45 / 113 — — — 4,860 4,300 3,850 3,420 3,050
50 / 122 — — — — 4,910 4,240 3,770 3,330
Climb Wght Temp
Limits °C/°F 42/108 42/108 44/111 48/118 51/124 52/126 52/126 52/126
Field Length at
Temp Limits (ft) 5,560 5,420 5,400 5,320 5,080 4,510 3,930 3,460

Table PER-2. TAKEOFF FIELD LENGTH—15° FLAPS (Cont)
TAKEOFF PERFORMANCE
TAKEOFF FIELD LENGTH - 15°° FLAPS

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 3,690 3,630 3,480 3,200 2,930 2,750 2,770 2,800
10 / 50 3,800 3,750 3,590 3,300 3,020 2,830 2,850 2,870
15 / 59 3,860 3,810 3,650 3,350 3,070 2,870 2,880 2,910
20 / 68 3,980 3,910 3,750 3,440 3,150 2,880 2,890 2,910
25 / 77 4,280 4,200 4,010 3,660 3,340 3,040 2,780 2,800
30 / 86 4,630 4,540 4,330 3,930 3,560 3,230 2,920 2,690
35 / 95 5,070 4,970 4,720 4,280 3,850 3,460 3,130 2,810
40 / 104 — 5,690 5,330 4,700 4,220 3,780 3,370 3,020
45 / 113 — — — 5,430 4,710 4,160 3,700 3,270
50 / 122 — — — — — 4,770 4,110 3,620
Climb Wght Temp
Limits °C/°F 39/102 40/104 41/106 45/113 48/118 50/122 50/122 50/122
Field Length at
Temp Limits (ft) 5,660 5,690 5,500 5,430 5,190 4,770 4,110 3,620

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 3,720 3,660 3,510 3,230 2,950 2,720 2,730 2,760
0 / 32 3,840 3,780 3,630 3,330 3,050 2,810 2,820 2,840
10 / 50 3,970 3,900 3,750 3,440 3,150 2,890 2,900 2,920
15 / 59 4,040 3,970 3,820 3,500 3,200 2,930 2,940 2,960
20 / 68 4,310 4,230 4,040 3,690 3,370 3,070 2,850 2,870
25 / 77 4,650 4,560 4,340 3,940 3,580 3,250 2,940 2,760
30 / 86 5,040 4,940 4,700 4,260 3,840 3,470 3,130 2,810
35 / 95 5,620 5,480 5,140 4,630 4,170 3,730 3,350 3,000
40 / 104 — — — 5,180 4,570 4,080 3,630 3,220
45 / 113 — — — — 5,290 4,560 4,020 3,550
Climb Wght Temp
Limits °C/°F 36/97 37/99 39/102 42/108 45/113 48/118 48/118 48/118
Field Length at
Temp Limits (ft) 5,780 5,790 5,750 5,520 5,290 5,060 4,320 3,790

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 3,890 3,820 3,660 3,360 3,080 2,810 2,780 2,810
0 / 32 4,020 3,950 3,790 3,480 3,180 2,900 2,870 2,890
10 / 50 4,160 4,090 3,920 3,600 3,290 3,000 2,940 2,960
15 / 59 4,350 4,270 4,070 3,730 3,410 3,100 2,920 2,930
20 / 68 4,680 4,590 4,370 3,970 3,610 3,280 2,970 2,830
25 / 77 5,050 4,950 4,710 4,260 3,850 3,480 3,150 2,830
30 / 86 5,500 5,370 5,100 4,600 4,150 3,720 3,350 3,000
35 / 95 — — 5,690 5,000 4,490 4,020 3,580 3,200
40 / 104 — — — — 5,020 4,420 3,930 3,470
45 / 113 — — — — — 5,160 4,400 3,870
Climb Wght Temp
Limits °C/°F 33/91 34/93 36/97 39/102 42/108 45/113 46/115 46/115
Field Length at
Temp Limits (ft) 5,920 5,920 5,850 5,600 5,380 5,160 4,570 3,960

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,070 4,000 3,820 3,510 3,210 2,930 2,830 2,850
0 / 32 4,210 4,130 3,950 3,630 3,320 3,030 2,920 2,940
5 / 41 4,280 4,210 4,020 3,690 3,380 3,080 2,960 2,980
10 / 50 4,420 4,340 4,140 3,790 3,470 3,150 2,970 2,980
15 / 59 4,730 4,640 4,410 4,010 3,660 3,320 3,010 2,900
20 / 68 5,080 4,980 4,740 4,290 3,880 3,520 3,180 2,860
25 / 77 5,490 5,380 5,110 4,610 4,160 3,730 3,370 3,020
30 / 86 6,110 5,960 5,590 4,980 4,480 4,020 3,580 3,210
35 / 95 — — — 5,520 4,860 4,340 3,860 3,420
40 / 104 — — — — — 4,880 4,290 3,780
Climb Wght Temp
Limits °C/°F 30/86 31/88 33/91 36/97 39/102 43/109 44/111 44/111
Field Length at
Temp Limits (ft) 6,110 6,110 6,010 5,700 5,470 5,440 4,800 4,130

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,180 4,100 3,930 3,610 3,300 3,010 2,940 2,960
0 / 32 4,320 4,240 4,060 3,730 3,410 3,110 3,030 3,050
5 / 41 4,440 4,350 4,160 3,820 3,490 3,180 3,050 3,070
10 / 50 4,780 4,690 4,470 4,050 3,700 3,360 3,040 2,970
15 / 59 5,140 5,040 4,790 4,340 3,920 3,560 3,220 2,890
20 / 68 5,550 5,440 5,170 4,660 4,200 3,770 3,410 3,060
25 / 77 6,110 5,960 5,590 5,020 4,520 4,050 3,620 3,250
30 / 86 — — — 5,490 4,880 4,360 3,880 3,450
35 / 95 — — — — 5,440 4,760 4,220 3,730
40 / 104 — — — — — 5,500 4,690 4,110
Climb Wght Temp
Limits °C/°F 26/79 27/81 29/84 33/91 36/97 40/104 42/108 42/108
Field Length at
Temp Limits (ft) 6,250 6,240 6,130 5,900 5,620 5,500 5,020 4,300

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,300 4,220 4,050 3,720 3,400 3,100 3,040 3,060
0 / 32 4,490 4,410 4,210 3,870 3,530 3,220 3,120 3,130
5 / 41 4,810 4,720 4,490 4,080 3,730 3,390 3,070 3,040
10 / 50 5,190 5,090 4,840 4,380 3,960 3,600 3,250 2,940
15 / 59 5,600 5,490 5,220 4,700 4,240 3,810 3,450 3,100
20 / 68 6,130 5,980 5,650 5,080 4,570 4,090 3,660 3,280
25 / 77 — — 6,280 5,490 4,920 4,400 3,920 3,490
30 / 86 — — — — 5,360 4,750 4,220 3,730
35 / 95 — — — — — 5,290 4,600 4,060
40 / 104 — — — — — — 5,250 4,470
Climb Wght Temp
Limits °C/°F 22/72 22/72 25/77 29/84 33/91 37/99 40/104 40/104
Field Length at
Temp Limits (ft) 6,420 6,260 6,280 6,030 5,800 5,620 5,250 4,470


TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-20 / -4 4,310 4,230 4,050 3,720 3,400 3,110 3,030 3,050
-10 / 14 4,460 4,380 4,190 3,850 3,520 3,210 3,130 3,150
0/ 32 4,850 4,760 4,530 4,120 3,770 3,430 3,110 3,110
5/ 41 5,240 5,130 4,880 4,420 4,000 3,640 3,290 3,010
10 / 50 5,660 5,550 5,280 4,750 4,290 3,860 3,490 3,140
15 / 59 6,140 6,000 5,700 5,120 4,610 4,130 3,700 3,320
20 / 68 — — 6,300 5,550 4,970 4,440 3,960 3,530
25 / 77 — — — 6,170 5,360 4,780 4,250 3,760
30 / 86 — — — — 6,040 5,200 4,600 4,050
35 / 95 — — — — — — 5,060 4,410
Climb Wght Temp
Limits °C/°F 17/63 18/64 20/68 25/77 30/86 34/93 38/100 38/100
Field Length at
Temp Limits (ft) 6,430 6,420 6,300 6,170 6,040 5,780 5,510 4,650

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-20 / -4 4,470 4,390 4,200 3,860 3,530 3,210 3,120 3,140
-10 / 14 4,650 4,570 4,360 4,000 3,660 3,330 3,210 3,230
0/ 32 5,280 5,180 4,930 4,460 4,040 3,680 3,320 3,080
5/ 41 5,710 5,600 5,320 4,800 4,320 3,910 3,530 3,170
10 / 50 6,180 6,050 5,750 5,170 4,650 4,170 3,740 3,360
15 / 59 — — 6,300 5,590 5,010 4,480 3,990 3,560
20 / 68 — — — 6,180 5,410 4,830 4,290 3,800
25 / 77 — — — — 6,030 5,220 4,630 4,090
30 / 86 — — — — — 5,850 5,010 4,410
35 / 95 — — — — — 6,680 5,640 4,780
Climb Wght Temp
Limits °C/°F 13/55 14/57 16/61 21/70 26/79 30/86 35/95 36/97
Field Length at
Temp Limits (ft) 6,580 6,560 6,450 6,320 6,180 5,850 5,640 4,880

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 3,850 3,780 3,620 3,310 3,010 2,730 2,580 2,590
10 / 50 3,970 3,900 3,740 3,420 3,110 2,820 2,650 2,670
15 / 59 4,040 3,970 3,800 3,470 3,160 2,870 2,690 2,700
20 / 68 4,100 4,030 3,860 3,520 3,210 2,910 2,730 2,740
25 / 77 4,210 4,120 3,950 3,610 3,280 2,970 2,750 2,760
30 / 86 4,570 4,460 4,230 3,850 3,500 3,170 2,850 2,660
35 / 95 5,120 5,000 4,700 4,170 3,760 3,400 3,050 2,730
40 / 104 5,930 5,780 5,420 4,760 4,180 3,700 3,320 2,960
45 / 113 7,080 6,880 6,410 5,570 4,840 4,200 3,640 3,230
50 / 122 — — 7,590 6,510 5,590 4,810 4,130 3,530
Climb Wght Temp
Limits °C/°F 48/118 48/118 50/122 53/127 54/129 54/129 54/129 54/129
Field Length at
Temp Limits (ft) 7,850 7,610 7,590 7,260 6,400 5,440 4,630 3,940

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 4,020 3,950 3,780 3,450 3,140 2,850 2,630 2,640
10 / 50 4,150 4,080 3,900 3,560 3,240 2,940 2,700 2,710
15 / 59 4,220 4,140 3,970 3,620 3,300 2,990 2,740 2,750
20 / 68 4,300 4,220 4,030 3,680 3,350 3,040 2,780 2,790
25 / 77 4,540 4,450 4,240 3,870 3,510 3,180 2,860 2,730
30 / 86 5,050 4,940 4,650 4,150 3,770 3,400 3,060 2,740
35 / 95 5,730 5,590 5,250 4,630 4,070 3,670 3,290 2,940
40 / 104 6,700 6,520 6,100 5,330 4,650 4,050 3,580 3,180
45 / 113 8,020 7,790 7,230 6,230 5,380 4,650 4,000 3,470
50 / 122 — — — 7,430 6,320 5,390 4,600 3,920
Climb Wght Temp
Limits °C/°F 45/113 45/113 47/117 50/122 52/126 52/126 52/126 52/126
Field Length at
Temp Limits (ft) 8,020 7,790 7,760 7,430 6,810 5,770 4,890 4,150


TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
0 / 32 4,200 4,120 3,940 3,600 3,280 2,970 2,680 2,690
10 / 50 4,340 4,270 4,080 3,720 3,390 3,070 2,770 2,760
15 / 59 4,420 4,340 4,140 3,780 3,440 3,120 2,810 2,800
20 / 68 4,570 4,480 4,280 3,900 3,540 3,210 2,890 2,800
25 / 77 5,030 4,910 4,630 4,170 3,780 3,420 3,070 2,750
30 / 86 5,630 5,490 5,160 4,560 4,060 3,660 3,290 2,940
35 / 95 6,450 6,280 5,880 5,160 4,520 3,960 3,540 3,160
40 / 104 7,600 7,380 6,870 5,960 5,170 4,480 3,870 3,420
45 / 113 — — — 7,040 6,030 5,170 4,430 3,790
50 / 122 — — — — 7,260 6,130 5,180 4,380
Climb Wght Temp
Limits °C/°F 42/108 42/108 44/111 47/117 50/122 50/122 50/122 50/122
Field Length at
Temp Limits (ft) 8,180 7,940 7,920 7,590 7,260 6,130 5,180 4,380

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,240 4,170 3,990 3,640 3,310 3,000 2,710 2,650
0 / 32 4,390 4,310 4,120 3,760 3,420 3,100 2,800 2,730
10 / 50 4,550 4,470 4,260 3,890 3,540 3,200 2,890 2,810
15 / 59 4,640 4,560 4,350 3,960 3,600 3,260 2,940 2,850
20 / 68 5,020 4,910 4,630 4,200 3,810 3,440 3,100 2,780
25 / 77 5,590 5,450 5,130 4,540 4,080 3,680 3,300 2,950
30 / 86 6,290 6,140 5,760 5,060 4,440 3,950 3,540 3,160
35 / 95 7,250 7,050 6,590 5,740 5,000 4,350 3,810 3,390
40 / 104 — 8,350 7,740 6,660 5,740 4,950 4,260 3,680
45 / 113 — — — — 6,840 5,820 4,950 4,210
Climb Wght Temp
Limits °C/°F 39/102 40/104 41/106 44/111 48/118 48/118 48/118 48/118
Field Length at
Temp Limits (ft) 8,280 8,350 8,040 7,740 7,770 6,530 5,500 4,640

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,440 4,360 4,170 3,800 3,460 3,130 2,820 2,700
0 / 32 4,600 4,520 4,310 3,930 3,570 3,240 2,920 2,780
10 / 50 4,780 4,690 4,480 4,070 3,700 3,350 3,020 2,860
15 / 59 5,050 4,930 4,680 4,240 3,850 3,480 3,130 2,830
20 / 68 5,580 5,450 5,130 4,540 4,110 3,710 3,330 2,980
25 / 77 6,230 6,080 5,710 5,030 4,420 3,960 3,550 3,170
30 / 86 7,060 6,870 6,430 5,620 4,910 4,280 3,800 3,390
35 / 95 8,150 7,910 7,360 6,380 5,520 4,780 4,120 3,630
40 / 104 — — — 7,540 6,440 5,520 4,720 4,030
45 / 113 — — — — 7,930 6,660 5,610 4,730
Climb Wght Temp
Limits °C/°F 36/97 37/99 39/102 42/108 45/113 46/115 46/115 46/115
Field Length at
Temp Limits (ft) 8,410 8,440 8,470 8,200 7,930 6,960 5,840 4,910

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,660 4,570 4,360 3,980 3,610 3,270 2,950 2,750
0 / 32 4,830 4,740 4,520 4,110 3,740 3,380 3,050 2,830
5 / 41 4,920 4,820 4,600 4,180 3,800 3,440 3,100 2,870
10 / 50 5,110 5,000 4,750 4,310 3,910 3,540 3,180 2,880
15 / 59 5,610 5,480 5,160 4,590 4,150 3,750 3,370 3,010
20 / 68 6,220 6,060 5,700 5,020 4,430 3,990 3,580 3,200
25 / 77 6,970 6,790 6,360 5,570 4,880 4,270 3,820 3,400
30 / 86 7,920 7,700 7,180 6,240 5,430 4,710 4,090 3,640
35 / 95 — — 8,260 7,110 6,130 5,280 4,540 3,910
40 / 104 — — — — 7,360 6,240 5,300 4,500
Climb Wght Temp
Limits °C/°F 33/91 34/93 36/97 39/102 42/108 44/111 44/111 44/111
Field Length at
Temp Limits (ft) 8,630 8,630 8,590 8,310 8,030 7,370 6,160 5,160

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,790 4,700 4,490 4,080 3,710 3,360 3,030 2,850
0 / 32 4,960 4,870 4,650 4,220 3,840 3,470 3,130 2,940
5 / 41 5,110 5,010 4,780 4,330 3,930 3,560 3,200 2,960
10 / 50 5,660 5,530 5,210 4,640 4,200 3,790 3,410 3,040
15 / 59 6,250 6,100 5,730 5,060 4,480 4,040 3,620 3,230
20 / 68 6,980 6,800 6,370 5,590 4,900 4,310 3,860 3,440
25 / 77 7,890 7,670 7,160 6,240 5,440 4,730 4,130 3,670
30 / 86 9,040 8,780 8,150 7,040 6,080 5,250 4,530 3,930
35 / 95 — — — 8,200 7,000 5,980 5,110 4,350
40 / 104 — — — — — 7,130 5,990 5,050
Climb Wght Temp
Limits °C/°F 30/86 31/88 33/91 36/97 39/102 42/108 42/108 42/108
Field Length at
Temp Limits (ft) 9,040 9,040 8,900 8,530 8,140 7,740 6,450 5,400

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-10 / 14 4,940 4,850 4,620 4,200 3,820 3,460 3,120 2,950
0 / 32 5,170 5,070 4,840 4,390 3,980 3,600 3,240 3,020
5 / 41 5,640 5,510 5,190 4,670 4,230 3,820 3,430 3,070
10 / 50 6,270 6,120 5,750 5,080 4,520 4,080 3,660 3,270
15 / 59 7,000 6,820 6,390 5,620 4,930 4,360 3,900 3,480
20 / 68 7,900 7,690 7,180 6,270 5,460 4,750 4,170 3,710
25 / 77 9,000 8,740 8,130 7,040 6,090 5,260 4,540 3,970
30 / 86 — — — 8,000 6,860 5,890 5,040 4,310
35 / 95 — — — — 8,030 6,800 5,760 4,870
40 / 104 — — — — — 8,140 6,780 5,650
Climb Wght Temp
Limits °C/°F 26/79 27/81 29/84 33/91 36/97 40/104 40/104 40/104
Field Length at
Temp Limits (ft) 9,250 9,240 9,080 8,820 8,330 8,140 6,780 5,650

TAKEOFF PERFORMANCE

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-20 / -4 4,950 4,860 4,630 4,210 3,830 3,460 3,120 2,940
-10 / 14 5,130 5,040 4,800 4,360 3,960 3,580 3,230 3,040
0/ 32 5,650 5,520 5,210 4,720 4,270 3,850 3,460 3,100
5/ 41 6,280 6,130 5,770 5,100 4,570 4,120 3,700 3,300
10 / 50 7,020 6,840 6,420 5,650 4,960 4,410 3,950 3,520
15 / 59 7,890 7,680 7,180 6,280 5,480 4,770 4,220 3,750
20 / 68 9,000 8,740 8,140 7,050 6,110 5,290 4,570 4,010
25 / 77 — — 9,300 7,980 6,850 5,890 5,050 4,320
30 / 86 — — — — 7,830 6,660 5,660 4,810
35 / 95 — — — — — 7,730 6,490 5,450
Climb Wght Temp
Limits °C/°F 21/70 22/72 25/77 29/84 33/91 37/99 38/100 38/100
Field Length at
Temp Limits (ft) 9,240 9,220 9,300 8,930 8,620 8,290 7,130 5,930

°C / °F 20,200 20,000 19,500 18,500 17,500 16,500 15,500 14,500
-20 / -4 5,140 5,050 4,810 4,370 3,970 3,590 3,230 3,030
-10 / 14 5,360 5,260 5,010 4,540 4,120 3,730 3,350 3,120
0/ 32 6,290 6,140 5,780 5,120 4,610 4,160 3,730 3,330
5/ 41 7,030 6,850 6,430 5,660 4,970 4,450 3,990 3,550
10 / 50 7,900 7,690 7,190 6,290 5,500 4,790 4,260 3,790
15 / 59 8,960 8,710 8,120 7,050 6,120 5,300 4,580 4,050
20 / 68 — — 9,270 7,970 6,860 5,900 5,070 4,340
25 / 77 — — — 9,150 7,790 6,640 5,660 4,810
30 / 86 — — — — — 7,570 6,380 5,380
35 / 95 — — — — — — 7,300 6,080
Climb Wght Temp
Limits °C/°F 17/63 18/64 20/68 25/77 29/84 34/93 36/97 36/97
Field Length at
Temp Limits (ft) 9,480 9,460 9,270 9,150 8,730 8,520 7,520 6,240

CLIMB PERFORMANCE
Table PER-4. 250 KIAS/M 0.65 CLIMB
CLIMB PERFORMANCE
250 KIAS / M 0.65 CLIMB
ISA, Zero Wind, Anti-Ice Off
Time, Fuel, and Distance To Climb
Pressure ------------------------------------ Takeoff Weight (lb) ------------------------------------
Altitude (ft) 20,200 19,000 18,000 16,000 14,000
15,000 Min 4 4 4 3 3
Lb 219 203 191 167 144
NM 20 19 18 15 13
25,000 Min 8 8 7 6 5
Lb 374 347 325 282 243
NM 42 39 37 32 27
29,000 Min 11 10 9 8 7
Lb 454 419 391 339 290
NM 57 52 49 42 36
31,000 Min 12 11 10 9 7
Lb 489 451 420 363 310
NM 64 58 54 47 40
33,000 Min 13 12 11 9 8
Lb 523 481 448 386 329
NM 71 65 60 51 44
35,000 Min 14 13 12 10 9
Lb 559 513 476 410 349
NM 79 72 66 57 48
37,000 Min 15 14 13 11 9
Lb 597 546 506 434 368
NM 87 79 73 62 52
39,000 Min 17 16 14 12 10
Lb 641 583 539 460 389
NM 98 89 82 69 58
41,000 Min 20 17 16 13 11
Lb 695 627 577 488 411
NM 113 101 92 76 63
43,000 Min 23 20 18 15 12
Lb 766 681 621 520 435
NM 134 116 104 86 70
45,000 Min 29 24 21 17 14
Lb 888 758 679 557 461
NM 172 141 123 97 79

CRUISE PERFORMANCE
Table PER-5. HIGH-SPEED CRUISE
CRUISE PERFORMANCE
HIGH SPEED CRUISE
ISA, Anti-Ice Off
Cruise Speed & Fuel Flow

Pressure ------------------------------------- Cruise Weight (lb) -------------------------------------
Altitude (ft) 20,000 19,000 18,000 16,000 14,000
5,000 KTAS 280 280 280 280 280
Lb/Hr 1,635 1,625 1,615 1,598 1,583
10,000 KTAS 353 353 353 353 353
Lb/Hr 2,027 2,019 2,010 1,995 1,982
15,000 KTAS 379 379 379 379 379
Lb/Hr 2,005 1,994 1,984 1,966 1,950
21,000 KTAS 414 414 414 414 414
Lb/Hr 2,063 2,049 2,036 2,012 1,991
23,000 KTAS 427 427 427 427 427
Lb/Hr 2,087 2,072 2,058 2,032 2,010
25,000 KTAS 432 434 435 437 439
Lb/Hr 2,019 2,023 2,024 2,024 2,025
27,000 KTAS 436 437 439 441 443
Lb/Hr 1,945 1,947 1,948 1,949 1,949
29,000 KTAS 437 438 439 442 444
Lb/Hr 1,852 1,847 1,849 1,852 1,855
31,000 KTAS 438 439 441 441 441
Lb/Hr 1,763 1,764 1,764 1,738 1,705
33,000 KTAS 437 437 437 437 437
Lb/Hr 1,662 1,643 1,625 1,591 1,559
35,000 KTAS 433 433 433 433 433
Lb/Hr 1,533 1,512 1,493 1,458 1,425
37,000 KTAS 431 431 431 431 431
Lb/Hr 1,426 1,400 1,377 1,339 1,305
39,000 KTAS 431 431 431 431 431
Lb/Hr 1,359 1,327 1,297 1,251 1,214
41,000 KTAS 428 431 431 431 431
Lb/Hr 1,278 1,272 1,235 1,175 1,133
43,000 KTAS 417 424 430 431 431
Lb/Hr 1,172 1,176 1,179 1,116 1,062
45,000 KTAS 399 413 422 431 431
Lb/Hr 1,082 1,097 1,094 1,072 1,003

Table PER-6. LONG-RANGE CRUISE
CRUISE PERFORMANCE
LONG RANGE CRUISE
ISA, Anti-Ice Off
Cruise Speed & Fuel Flow

Pressure ------------------------------------- Cruise Weight (lb) -------------------------------------
Altitude (ft) 20,000 19,000 18,000 16,000 14,000
5,000 KTAS 271 271 270 268 240
Lb/Hr 1,572 1,558 1,544 1,507 1,318
10,000 KTAS 275 273 272 270 260
Lb/Hr 1,374 1,355 1,337 1,298 1,224
15,000 KTAS 283 280 275 274 273
Lb/Hr 1,243 1,206 1,167 1,135 1,105
21,000 KTAS 296 292 288 280 268
Lb/Hr 1,115 1,078 1,048 980 900
23,000 KTAS 305 298 293 284 274
Lb/Hr 1,092 1,047 1,012 941 870
25,000 KTAS 314 307 301 288 277
Lb/Hr 1,073 1,030 987 905 834
27,000 KTAS 319 314 309 290 283
Lb/Hr 1,038 1,001 964 865 804
29,000 KTAS 321 319 315 302 292
Lb/Hr 996 969 937 858 792
31,000 KTAS 324 321 320 308 292
Lb/Hr 966 936 909 832 747
33,000 KTAS 325 323 321 312 303
Lb/Hr 936 905 872 802 734
35,000 KTAS 329 326 323 318 310
Lb/Hr 917 881 846 784 719
37,000 KTAS 333 330 326 323 315
Lb/Hr 899 860 823 763 695
39,000 KTAS 342 338 334 327 320
Lb/Hr 905 862 820 747 677
41,000 KTAS 353 348 342 332 324
Lb/Hr 923 875 824 736 662
43,000 KTAS 367 358 353 340 330
Lb/Hr 956 891 838 738 656
45,000 KTAS 374 369 367 351 339
Lb/Hr 986 915 869 751 658

DESCENT PERFORMANCE
Table PER-7. HIGH SPEED AND NORMAL DESCENT
DESCENT PERFORMANCE
HIGH SPEED & NORMAL DESCENT
ISA, Zero Wind, Anti-Ice Off,
Speed Brakes Retracted, Gear & Flaps Up
Time, Fuel, and Distance To Descend

High Speed — 3,000 FPM Normal — 2,000 FPM
Pressure ------- End of Cruise Weight (lb) ------- ------- End of Cruise Weight (lb) -------
Altitude (ft) 18,000 15,000 12,000 18,000 15,000 12,000
15,000 Min 6 5 5 8 8 8
Lb 62 64 78 127 138 150
NM 28 26 25 38 38 37
25,000 Min 9 9 9 13 13 13
Lb 124 133 153 245 261 278
NM 49 48 47 71 71 70
31,000 Min 11 11 11 16 16 16
Lb 164 175 198 317 335 355
NM 64 62 62 93 92 92
33,000 Min 12 11 11 17 17 17
Lb 175 187 211 338 356 376
NM 68 67 66 100 99 99
35,000 Min 12 12 12 18 18 18
Lb 185 197 221 356 374 395
NM 73 71 71 107 106 106
37,000 Min 13 13 13 19 19 19
Lb 193 206 231 371 390 411
NM 78 76 76 113 113 113
39,000 Min 14 13 13 20 20 20
Lb 200 213 239 385 404 425
NM 82 81 80 120 120 120
41,000 Min 14 14 14 21 21 21
Lb 206 220 246 398 417 438
NM 87 85 85 127 127 126
43,000 Min 15 15 15 22 22 22
Lb 214 227 253 410 428 450
NM 93 90 90 134 134 133
45,000 Min 16 16 15 23 23 23
Lb 222 235 260 421 439 461
NM 98 96 95 141 141 140

RESERVE FUEL
Reserve Fuel Allowances
Based on four passengers, ISA, zero wind.
VFR Fuel Reserves (at 15,000 feet)
Day (30 minutes) ............................................................................... 554 lb
Night (45 minutes) ............................................................................. 834 lb
IFR Fuel Reserves (Alternate plus 45 minutes at 15,000 feet))
100 Nautical Mile Alternate............................................................ 1,324 lb
NBAA Fuel Reserves*
* NBAA IFR Reserves are defined as the amount of fuel for the following
profile:
• A 5-minute approach at sea level
• Climb to 5,000 feet
• A 5-minute hold at 5,000 feet
• Climb to cruise altitude for the diversion to the alternate airport
• Cruise at long range cruise power
• Descend to sea level
• Land with 30 minutes of holding fuel at 5,000 feet
HOLDING PERFORMANCE
Table PER-8. HOLDING SPEED AND FUEL FLOW
ISA, Anti-Ice Off, Speed Brakes Retracted, Gear & Flaps Up
Holding Speed & Fuel Flow

------------------------------------ Pressure Altitude (ft) ------------------------------------
Weight (lb) KIAS S.L. 5,000 10,000 15,000 20,000 25,000 30,000
17,000 190 1,256 1,160 1,069 995 932 890 855
16,000 185 1,215 1,118 1,031 960 892 847 814
15,000 180 1,176 1,076 993 926 851 807 774
14,000 175 1,138 1,034 957 892 812 768 735
13,000 170 1,085 985 920 852 775 729 696

LANDING PERFORMANCE
Table PER-9. LANDING DISTANCE—ACTUAL
LANDING PERFORMANCE
LANDING DISTANCE - ACTUAL
(Distance from 50 Feet Above the Runway)
Flaps 35°, Dry Runway, Zero Wind, Anti-Ice On or Off

Ambient Temp ------------------------------------------- Landing Weight (lb) -------------------------------------------
°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
0 / 32 3,060 3,040 2,980 2,860 2,740 2,620 2,490 2,370
10 / 50 3,140 3,120 3,060 2,940 2,820 2,680 2,550 2,430
15 / 59 3,180 3,160 3,100 2,970 2,850 2,720 2,590 2,460
20 / 68 3,230 3,200 3,140 3,010 2,890 2,750 2,620 2,490
25 / 77 3,270 3,240 3,180 3,050 2,920 2,790 2,650 2,520
30 / 86 3,310 3,280 3,220 3,090 2,960 2,820 2,680 2,550
35 / 95 3,350 3,320 3,260 3,120 2,990 2,850 2,710 2,580
40 / 104 3,390 3,360 3,300 3,160 3,030 2,890 2,740 2,610
45 / 113 3,430 3,400 3,330 3,200 3,060 2,920 2,770 2,630
50 / 122 3,470 3,440 3,370 3,230 3,100 2,950 2,800 2,660
Lndg Wght Temp
Limits °C/°F 52/126 54/129 54/129 54/129 54/129 54/129 54/129 54/129
VREF (KIAS) 117 117 115 112 109 106 102 99

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
0 / 32 3,150 3,120 3,060 2,940 2,820 2,690 2,560 2,430
10 / 50 3,230 3,200 3,140 3,010 2,890 2,760 2,620 2,490
15 / 59 3,270 3,250 3,180 3,050 2,930 2,790 2,650 2,520
20 / 68 3,320 3,290 3,220 3,090 2,970 2,830 2,690 2,550
25 / 77 3,360 3,330 3,270 3,130 3,000 2,860 2,720 2,580
30 / 86 3,400 3,370 3,310 3,170 3,040 2,900 2,750 2,610
35 / 95 3,440 3,410 3,350 3,210 3,080 2,930 2,790 2,640
40 / 104 3,480 3,460 3,390 3,250 3,110 2,960 2,820 2,670
45 / 113 3,520 3,500 3,430 3,290 3,150 3,000 2,850 2,700
50 / 122 — 3,540 3,470 3,320 3,180 3,030 2,880 2,730
Lndg Wght Temp
Limits °C/°F 49/120 50/122 51/124 52/126 52/126 52/126 52/126 52/126
VREF (KIAS) 117 117 115 112 109 106 102 99

Table PER-9. LANDING DISTANCE—ACTUAL (Cont)

LANDING DISTANCE - ACTUAL

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
0 / 32 3,230 3,210 3,140 3,020 2,890 2,760 2,620 2,490
10 / 50 3,320 3,290 3,230 3,100 2,970 2,830 2,690 2,560
15 / 59 3,360 3,340 3,270 3,140 3,010 2,870 2,730 2,590
20 / 68 3,410 3,380 3,310 3,180 3,050 2,900 2,760 2,620
25 / 77 3,450 3,420 3,360 3,220 3,080 2,940 2,790 2,650
30 / 86 3,500 3,470 3,400 3,260 3,120 2,980 2,830 2,680
35 / 95 3,540 3,510 3,440 3,300 3,160 3,010 2,860 2,720
40 / 104 3,580 3,550 3,480 3,340 3,200 3,050 2,890 2,750
45 / 113 3,620 3,590 3,520 3,380 3,240 3,080 2,930 2,780
50 / 122 — — — 3,420 3,270 3,120 2,960 2,810
Lndg Wght Temp
Limits °C/°F 46/115 47/117 48/118 50/122 50/122 50/122 50/122 50/122
VREF (KIAS) 117 117 115 112 109 106 102 99

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-10 / 14 3,240 3,220 3,160 3,030 2,900 2,770 2,640 2,510
0 / 32 3,330 3,300 3,240 3,110 2,980 2,840 2,700 2,570
10 / 50 3,420 3,390 3,330 3,190 3,060 2,910 2,770 2,630
15 / 59 3,460 3,440 3,370 3,230 3,100 2,950 2,810 2,670
20 / 68 3,510 3,480 3,410 3,270 3,140 2,990 2,840 2,700
25 / 77 3,550 3,530 3,460 3,310 3,180 3,030 2,880 2,730
30 / 86 3,600 3,570 3,500 3,360 3,220 3,060 2,910 2,760
35 / 95 3,640 3,620 3,540 3,400 3,260 3,100 2,950 2,800
40 / 104 3,690 3,660 3,590 3,440 3,300 3,140 2,980 2,830
45 / 113 — — 3,630 3,480 3,330 3,170 3,020 2,860
Lndg Wght Temp
Limits °C/°F 43/109 44/111 45/113 48/118 48/118 48/118 48/118 48/118
VREF (KIAS) 117 117 115 112 109 106 102 99



°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-10 / 14 3,340 3,320 3,250 3,120 3,000 2,860 2,720 2,590
0 / 32 3,430 3,410 3,340 3,200 3,070 2,930 2,790 2,650
10 / 50 3,520 3,500 3,430 3,290 3,150 3,000 2,860 2,710
15 / 59 3,570 3,540 3,470 3,330 3,190 3,040 2,890 2,750
20 / 68 3,620 3,590 3,520 3,370 3,230 3,080 2,930 2,780
25 / 77 3,660 3,630 3,560 3,410 3,270 3,120 2,960 2,810
30 / 86 3,710 3,680 3,610 3,460 3,310 3,160 3,000 2,850
35 / 95 3,760 3,730 3,650 3,500 3,350 3,190 3,030 2,880
40 / 104 3,800 3,770 3,700 3,540 3,390 3,230 3,070 2,910
45 / 113 — — — 3,590 3,430 3,270 3,110 2,950
Lndg Wght Temp
Limits °C/°F 40/104 41/106 42/108 45/113 46/115 46/115 46/115 46/115
VREF (KIAS) 117 117 115 112 109 106 102 99

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-10 / 14 3,450 3,420 3,360 3,220 3,090 2,950 2,800 2,670
0 / 32 3,540 3,510 3,450 3,310 3,170 3,020 2,870 2,730
5 / 41 3,590 3,560 3,490 3,350 3,210 3,060 2,910 2,770
10 / 50 3,630 3,600 3,530 3,390 3,250 3,100 2,940 2,800
15 / 59 3,680 3,650 3,580 3,430 3,290 3,140 2,980 2,830
20 / 68 3,730 3,700 3,620 3,480 3,330 3,170 3,020 2,860
25 / 77 3,770 3,740 3,670 3,520 3,370 3,210 3,050 2,900
30 / 86 3,820 3,790 3,720 3,560 3,410 3,250 3,090 2,930
35 / 95 3,870 3,840 3,760 3,610 3,460 3,290 3,130 2,970
40 / 104 — — 3,810 3,650 3,500 3,330 3,160 3,000
Lndg Wght Temp
Limits °C/°F 37/99 38/100 40/104 43/109 44/111 44/111 44/111 44/111
VREF (KIAS) 117 117 115 112 109 106 102 99


Elevation = 6000 Feet

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-10 / 14 3,560 3,530 3,460 3,320 3,190 3,040 2,890 2,750
0 / 32 3,650 3,630 3,560 3,410 3,270 3,120 2,970 2,820
5 / 41 3,700 3,670 3,600 3,460 3,310 3,160 3,000 2,850
10 / 50 3,750 3,720 3,650 3,500 3,350 3,200 3,040 2,890
15 / 59 3,800 3,770 3,690 3,540 3,400 3,240 3,080 2,920
20 / 68 3,850 3,820 3,740 3,590 3,440 3,280 3,110 2,960
25 / 77 3,890 3,860 3,790 3,630 3,480 3,310 3,150 2,990
30 / 86 3,940 3,910 3,830 3,680 3,520 3,350 3,190 3,020
35 / 95 — 3,960 3,880 3,720 3,560 3,390 3,220 3,060
40 / 104 — — — 3,760 3,610 3,430 3,260 3,090
Lndg Wght Temp
Limits °C/°F 34/93 35/95 37/99 40/104 42/108 42/108 42/108 42/108
VREF (KIAS) 117 117 115 112 109 106 102 99
Elevation = 7000 Feet

°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-10 / 14 3,670 3,640 3,570 3,430 3,290 3,140 2,990 2,840
0 / 32 3,770 3,740 3,670 3,520 3,380 3,220 3,060 2,910
5 / 41 3,820 3,790 3,720 3,570 3,420 3,260 3,100 2,950
10 / 50 3,870 3,840 3,770 3,610 3,460 3,300 3,140 2,980
15 / 59 3,920 3,890 3,810 3,660 3,510 3,340 3,180 3,020
20 / 68 3,970 3,940 3,860 3,700 3,550 3,380 3,210 3,050
25 / 77 4,020 3,990 3,910 3,750 3,590 3,420 3,250 3,090
30 / 86 4,070 4,040 3,960 3,790 3,640 3,460 3,290 3,120
35 / 95 — — — 3,840 3,680 3,500 3,330 3,160
40 / 104 — — — — 3,720 3,540 3,370 3,190
Lndg Wght Temp
Limits °C/°F 31/88 32/90 33/91 37/99 40/104 40/104 40/104 40/104
VREF (KIAS) 117 117 115 112 109 106 102 99



°C / °F 18,700 18,500 18,000 17,000 16,000 15,000 14,000 13,000
-20 / -4 3,680 3,650 3,580 3,440 3,300 3,150 3,000 2,850
-10 / 14 3,790 3,760 3,690 3,540 3,390 3,240 3,080 2,930
0/ 32 3,890 3,860 3,790 3,630 3,480 3,320 3,160 3,000
5/ 41 3,940 3,910 3,840 3,680 3,530 3,360 3,200 3,040
10 / 50 3,990 3,960 3,890 3,730 3,570 3,410 3,240 3,080
15 / 59 4,050 4,010 3,940 3,770 3,620 3,450 3,280 3,110
20 / 68 4,100 4,070 3,990 3,820 3,660 3,490 3,320 3,150
25 / 77 4,150 4,120 4,040 3,870 3,710 3,530 3,360 3,190
30 / 86 — — 4,090 3,920 3,750 3,570 3,400 3,220
35 / 95 — — — — 3,800 3,620 3,440 3,260
Lndg Wght Temp
Limits °C/°F 27/81 28/82 30/86 34/93 38/100 38/100 38/100 38/100
VREF (KIAS) 117 117 115 112 109 106 102 99
STALL SPEEDS
Table PER-10. STALL SPEEDS
Zero Angle of Bank, Landing Gear Up or Down, KCAS
Stall Speeds
----------------------------------- Flap Position ------------------------------------
Weight (lb) Land 15° 7° Up
20,200 94 99 102 106
20,000 93 98 102 105
19,000 91 96 99 103
18,000 89 94 97 100
17,000 86 91 94 97
16,000 84 88 92 95
15,000 81 83 86 89
14,000 79 83 86 89

MISSION PLANNING
Criteria
The mission planning table (Table PER-12) provides flight time and fuel
burn statistics for selected distances and altitudes.
Flight time represents the time for the climb, cruise and descent portion of
the mission. No allowance has been added for taxi, takeoff, approach, or ATC
procedures. Fuel burn represents the total amount of fuel consumed for taxi,
climb, cruise, and descent. There is a taxi and takeoff allowance of 135
pounds of fuel included in all fuel burn figures. NBAA IFR fuel reserves (100
NM) are considered in each case, but are not included in the fuel burn figure.
The mission planning table reflects the cruise climb schedule of 250 knots/.65
Mach, high-speed cruise, and high-speed descent schedules. Standard day con-
ditions are assumed with zero wind enroute. The effects of wind can be de-
termined from the wind correction factors shown in Table PER-11. Apply the
wind correction factor to the zero wind flight time and fuel burn to estimate
the impact of wind.
Typical cruise altitudes for various distances are:
Distance (nm) Typical Cruise Altitude (ft)

0 - 100 6,000 - 18,000
101 - 200 17,000 - 31,000
201 - 300 29,000 - 39,000
301 - 500 37,000 - 41,000
501 - 900 39,000 - 43,000
901 + 39,000 - 45,000
Table PER-11. WIND CORRECTION FACTORS
Wind Correction Factors *

True Airspeed -------------- Headwinds (kt) -------------- --------------- Tailwinds (kt) ---------------
(kt) 100 75 50 25 0 25 50 75 100
320 1.45 1.31 1.18 1.08 1.00 0.93 0.86 0.81 0.76
340 1.42 1.28 1.17 1.08 1.00 0.93 0.87 0.82 0.77
360 1.38 1.26 1.16 1.07 1.00 0.93 0.88 0.83 0.78
380 1.36 1.25 1.15 1.07 1.00 0.94 0.88 0.84 0.79
400 1.33 1.23 1.14 1.06 1.00 0.94 0.89 0.84 0.80
420 1.31 1.22 1.13 1.06 1.00 0.94 0.89 0.85 0.81
440 1.29 1.21 1.13 1.06 1.00 0.95 0.90 0.85 0.81
* Wind Correction Factor is calculated as KTAS divided by the sum of KTAS ± wind component

Table PER-12. FLIGHT TIME AND FUEL BURN FOR

SELECTED DISTANCES
MISSION PLANNING
FLIGHT TIME & FUEL BURN
------------------------------------------------- Cruise Altitude (ft) -------------------------------------------------

15,000 25,000 31,000 33,000 35,000
Dist Time Fuel Time Fuel Time Fuel Time Fuel Time Fuel
(nm) (min) (lb) (min) (lb) (min) (lb) (min) (lb) (min) (lb)
200 0:33 1,189 0:31 1,105 0:32 1,026 0:32 1,001 0:32 978
300 0:49 1,710 0:45 1,572 0:46 1,424 0:45 1,369 0:46 1,318
400 1:05 2,232 0:59 2,040 0:59 1,824 0:59 1,737 1:00 1,660
500 1:21 2,755 1:13 2,508 1:13 2,225 1:13 2,108 1:14 2,003
600 1:37 3,279 1:27 2,976 1:27 2,628 1:27 2,479 1:28 2,347
700 1:53 3,806 1:41 3,445 1:41 3,032 1:41 2,852 1:42 2,693
800 2:09 4,334 1:55 3,915 1:54 3,437 1:54 3,226 1:56 3,040
900 2:24 4,866 2:09 4,388 2:08 3,844 2:08 3,602 2:10 3,388
1,000 2:23 4,861 2:22 4,253 2:22 3,979 2:24 3,737

1,100 2:35 4,662 2:36 4,358 2:38 4,088
1,200 2:49 5,070 2:50 4,740 2:52 4,441

1,300 3:03 5,478 3:04 5,122 3:05 4,795
1,400 3:17 5,506 3:19 5,151

1,500 Assumptions:
• 250 KIAS / M 0.65 climb
• High-speed cruise
1,600 • High-speed descent
• ISA, zero winds enroute
1,700 • Flight time includes climb, cruise and descent
• Fuel burn includes 135 pounds for taxi and takeoff
• Four passengers @ 200 pounds each, two crew
1,800 • NBAA IFR Reserves - 100 nm (1,210 lb) Reserves are not included in the fuel burn
figures
1,900

Table PER-12. FLIGHT TIME AND FUEL BURN FOR

SELECTED DISTANCES (Cont)
FOR SELECTED DISTANCES
------------------------------------------------- Cruise Altitude (ft) -------------------------------------------------

37,000 39,000 41,000 43,000 45,000
Time Fuel Time Fuel Time Fuel Time Fuel Time Fuel Dist
(min) (lb) (min) (lb) (min) (lb) (min) (lb) (min) (lb) (nm)
0:33 961 0:32 946 0:32 941 200
0:47 1,274 0:46 1,239 0:46 1,216 0:47 1,196 0:48 1,184 300
1:00 1,590 1:00 1,533 1:00 1,493 1:01 1,458 1:02 1,435 400
1:15 1,906 1:14 1,829 1:14 1,771 1:15 1,722 1:16 1,689 500
1:28 2,224 1:28 2,126 1:28 2,050 1:29 1,988 1:30 1,944 600
1:43 2,543 1:42 2,425 1:42 2,330 1:43 2,256 1:44 2,201 700
1:57 2,863 1:57 2,724 1:57 2,612 1:57 2,525 1:58 2,461 800
2:11 3,185 2:11 3,025 2:11 2,896 2:11 2,796 2:12 2,722 900
2:25 3,509 2:25 3,328 2:25 3,183 2:25 3,070 2:26 2,985 1,000
2:39 3,834 2:38 3,633 2:39 3,471 2:39 3,346 2:40 3,252 1,100
2:53 4,160 2:52 3,939 2:52 3,763 2:53 3,624 2:54 3,522 1,200
3:06 4,488 3:06 4,249 3:06 4,055 3:07 3,904 3:08 3,795 1,300
3:20 4,818 3:20 4,561 3:20 4,350 3:21 4,189 3:22 4,069 1,400
3:34 5,149 3:34 4,874 3:34 4,648 3:35 4,476 3:36 4,336 1,500
3:48 5,482 3:48 5,190 3:49 4,948 3:49 4,767 3:50 4,605 1,600
4:02 5,507 4:03 5,252 4:03 5,059 4:05 4,876 1,700
4:17 5,351 4:19 5,148 1,800

4:34 5,423 1,900

Table PER-13. RANGE/PAYLOAD CAPABILITY
MISSION PLANNING
RANGE / PAYLOAD CAPABILITY

NBAA IFR Reserves (100 nm), ISA,
Zero Wind, High-Speed Cruise
12
10
Number of Passengers
0
0 200 400 600 800 1000 1200 1400 1600 1800 2000
NBAA IFR Range (nautical miles)
Assumptions:
2 crew, passengers at 200 pounds each
Cruise at FL 450

FLIGHT PLANNING—EXCEL
This Flight Planning guide is for the purpose of providing specific informa-
tion for evaluating the performance of the Cessna Citation Excel (Model
560XL).
This guide is developed from Flight Manual and Operating Manual data. This
document is not intended to be used in lieu of the FAA approved Airplane Flight
Manual (AFM) or Operating Manual. The data included herein does not con-
stitute an offer and is subject to change without notice.
SPECIFICATIONS


INTENTIONALLY LEFT BLANK

TAKEOFF PERFORMANCE
Table PER-14 shows decision, rotation and takeoff speeds for aircraft with
rudder bias system installed.
FAR Part 25 takeoff field lengths are shown in Tables PER-15 and PER-16.
Part 25 defines takeoff distance as the greater of accelerate-stop, accelerate-
go with one engine inoperative, or 115% of the all engine takeoff distance to
a point 35 feet above the runway. These factors are reflected in the takeoff
distances presented.
Second segment climb limitations are presented at the bottom of each take-
off chart for reference. Second segment climb refers to the ability of the air-
craft to meet certain climb rates after takeoff with one engine inoperative.
Second segment climb limitations are a function of temperature, elevation and
aircraft weight.
Two flap settings are shown for the aircraft: 15° and 7°. A flap setting of 15°
is preferred to minimize runway length and runway speeds. In those situa-
tions where second segment climb requirements are two limiting for 15° of
flaps, a 7° flap setting is available. A 7° flap setting requires greater runway
length but provides greater second segment climb capability.
A paved, level, dry runway with zero wind is assumed. Runway lengths shown
are based on the aircraft’s anti-ice system being off.
Table PER-14. DECISION, ROTATION AND TAKEOFF SAFETY SPEEDS











CLIMB PERFORMANCE
Two climb schedules are shown on the following pages: Maximum Rate
Climb and Cruise Climb.
Table PER-17 shows the indicated airspeeds at various altitudes for the var-
ious climb schedules.
The Maximum Rate Climb schedule results in the minimal amount of time to
reach a selected altitude (Table PER-18).
The Cruise Climb schedule provides a balance between forward speed and
rate of climb (Table PER-19).
Each climb schedule is based on the climb starting at sea level. Weights rep-
resent the weight of the aircraft at the start of the climb.
Table PER-17. CLIMB SPEEDS

Table PER-18. MAXIMUM RATE CLIMB

Table PER-19. 250 KNOT/.62 MACH CRUISE CLIMB

CRUISE PERFORMANCE
The High-Speed Cruise schedule is shown in Table PER-20.
Table PER-20. HIGH-SPEED CRUISE

CRUISE PERFORMANCE
The Long-Range Cruise schedule is shown in Table PER-21.
Table PER-21. LONG-RANGE CRUISE

DESCENT PERFORMANCE
The Normal and High-Speed Descent schedule is shown in Table PER-22. The
time distance and fuel used from a given altitude is based on descending to
sea level.
Table PER-22. NORMAL AND HIGH SPEED DESCENT

FUEL RESERVES
HOLDING FUEL
The Holding Speed and Fuel Flow schedule is shown in Table PER-23.
Table PER-23. HOLDING SPEED AND FUEL FLOW

LANDING PERFORMANCE
Landing Distance schedules are shown in Table PER-24.
Stall Speed is shown in Table PER-25.
Table PER-24. LANDING DISTANCE

Table PER-24. LANDING DISTANCE (Cont)



Table PER-25. STALL SPEED

MISSION PLANNING
Criteria
Wind correction factors are shown in Table PER-26. The factors are calcu-
lated as KTAS divided by the sum of KTAS ± wind component.
The mission planning table (Table PER-27) provides flight time and fuel
burn statistics for selected distances and altitudes.
Flight time represents the time for the climb, cruise and descent portion of
the mission. No allowance has been added for taxi, takeoff or approach. Fuel
burn represents the total amount of fuel consumed for taxi, climb, cruise, and
descent. There is a taxi allowance of 125 pounds of fuel included in all fuel
burn figures. IFR fuel reserves are considered in each case, but are not in-
cluded in the fuel burn figure.
The mission planning table reflects a climb using the cruise climb schedule
of 250 knots/.62 Mach, cruise at high speed cruise and descent using the high
speed descent schedule. Standard day conditions are assumed with zero wind
enroute. The effects of wind can be determined from the wind correction fac-
tors table below. Apply the wind correction factor to the zero wind flight time
and fuel burn to estimate the impact of wind.
Range and payload capability is shown in Table PER-28.
Typical cruise altitudes for various distances are:
Table PER-26. WIND CORRECTION FACTORS

Table PER-27. FLIGHT TIME AND FUEL BURN FOR SELECTED

DISTANCES

Table PER-27. FLIGHT TIME AND FUEL BURN FOR SELECTED

DISTANCES (Cont)

Table PER-28. RANGE/PAYLOAD CAPABILITY

SPECIAL PROCEDURES—XLS AND EXCEL

SHORT FIELD OPERATION
For takeoff, taxi into position as close to the approach end as possible and
apply takeoff thrust while holding the brakes. Airplane Flight Manual take-
off field length data assumes a static run-up and use of all available runway.
When specified thrust is set, release the brakes. Rotate smoothly right at V R
as a delay will result in degradation of takeoff performance. Retract the gear
when positively climbing and climb at V 2 (V 2 + 10 KIAS multiengine) with
T.O. (7°) or T.O. & APPR. (15°) flaps until clear of any obstacles.
Landing field length data in the FAA approved Airplane Flight Manual as-
sumes a steady 3° approach angle and a threshold crossing speed of V REF at
an altitude of 50 feet, with thrust reduced to idle at that point. In practice, it
is suggested that for minimum field operations the threshold be crossed at a
comfortable obstacle clearance altitude allowing some deceleration to take
place approaching the runway. Touchdown should occur with maximum avail-
able runway remaining at minimum safe speed.
The energy to be dissipated during rollout is directly related to airplane

weight and velocity at touchdown. Although weight is normally dictated by
cabin loading and reserves required, flight planning into short fields should
include avoiding carrying excess weight in stored fuel. This consideration of-
fers the side benefit of improved enroute performance. Velocity is something
that can be controlled in nearly every case. Precise speed control is impor-
tant in the short field environment. A 1% increase in speed will require ap-
proximately 2% more rollout distance. Excessive speed and late throttle
reduction will also increase “float” prior to touchdown.
In general, short field landings are accomplished the same as normal land-
ings except for heavier braking and closer attention to touchdown point and
speed. A stabilized approach at V REF provides the best possible starting point
because any corrections necessary will be small. Establish a glide angle that
will safely clear any obstacles and result in touchdown as comfortably close
to the approach end as feasible.
Avoid a very flat approach as they generally result in excessive power being
required in close and the vertical gust protection margin is reduced. At ap-
proximately 50 feet AGL, power reduction is normally begun to cross the thresh-
old at a speed not in excess of V REF . Check the throttles at idle and avoid an
excessive flare that may cause the airplane to float. Deceleration will take place
much more rapidly on the runway than it will airborne.
If thrust reversers are not used, extend the speed brakes while lowering the
nose and commence braking with steady maximum pressure. Once braking
has begun, back pressure on the yoke will create elevator drag without affect-
ing weight on the gear provided the nosewheel is not lifted off the runway.
For landings utilizing thrust reversers, after touchdown on the mains, lower
the nose, extend speed brakes, and deploy the thrust reversers. Forward pres-
sure on the yoke should be applied during reverser deployment. Check illu-
mination of the ARM, UNLOCK and DEPLOY lights. Once the thrust reversers

are deployed, apply maximum reverse thrust power. Once braking has begun
and maximum reverse power is reached, back pressure on the yoke will pro-
vide additional weight on the main gear provided the nose is not raised. At
60 KIAS return the thrust reverser levers to the idle reverse detent position.
Leave the thrust reversers deployed for aerodynamic drag and idle reverse
power.
ADVERSE FIELD CONDITIONS

The Airplane Flight Manual presents takeoff field length data for dry, wet,
and hard surface runways. The AFM landing data assumes a dry, hard surface
runway. Precipitation-covered runway conditions will degrade braking effec-
tiveness and will require significantly greater actual takeoff and landing field
lengths.
Considerations for landing on a precipitation-covered runway are similar to

those for short field operations where speed is minimized and maximum roll-
out distance is made available. Runway composition, condition and con-
struction, the amount of precipitation and the depth of main landing gear tire
tread remaining affect the magnitude of braking degradation, so it is impos-
sible to apply a fixed factor to cover all conditions. Again, maximizing roll-
out runway available and touching down at minimum safe speed will provide
the greatest possible margin.
Use of the thrust reversers on precipitation-covered runways is the same as

that for a landing on a normal or dry runway. Cockpit visibility is not ham-
pered by blowing rain, snow, or ice thrown forward by the thrust reversers ex-
cept at low speed with idle reverse. Single-engine reversing during crosswind
landings on precipitation-covered runways should be used with discretion.
Precipitation-covered and icy runways present particular hazards which must be

understood in order to achieve effective braking. Under normal braking condi-
tions the antiskid system is very effective in preventing skids and in producing
minimum stopping distances, with the pilot applying and maintaining steady max-
imum pressure. However, on a precipitation- or ice-covered runway, the phenom-
enon of dynamic hydroplaning may greatly reduce the antiskid effectiveness,
because the wheels either do not spin up equally or do not spin up to the antiskid
threshold speed. It is important to maintain properly inflated tires with good tread
depth, and because groundspeed is critical, to avoid tailwinds when operating
in these conditions. When braking on precipitation-covered runways, ensure the
wheels are down and tracking prior to applying brakes. This will give the wheels
time to spin up. Ensure maximum weight is on the wheels, i.e., deploy speed brakes
and retract flaps. If runway permits, utilize maximum aerodynamic braking and
thrust reversers to slow the airplane prior to braking. When braking is commenced,
gradually apply steady pressure until antiskid cycling begins. As long as the an-
tiskid is cycling, maintain the pressure. If long antiskid pressure dumps occur
due to hydroplaning, release the brakes to allow the wheels to spin up again and
then gradually reapply pressure until antiskid cycling resumes.
After landing on ice or slush, a complete check of the airplane, including over-
board vents and controls surfaces, should be conducted.

ENGINE ANTI-ICE
The importance of proper system use cannot be overemphasized as serious
engine damage can result from ice ingestion. Its function is preventative in
nature and flight into visible moisture, with an outside air temperature below
+10°C indicated RAT should be anticipated, so the system is on and operat-
ing when icing conditions are encountered. Turning it on, after ice has accu-
mulated, could result in ice from the inlet being freed and ingested by the engine.
Bleed air anti-icing of the engine inlet alone is available at idle power and
above; however, approximately 70 % N 2 rpm is required to maintain the ENG
ANTI-ICE annunciator extinguished when operated in conjunction with
WING ANTI-ICE. In descent, it should be turned on well before entering an
icing environment to ensure sufficient time is available for all system param-
eters to be met.
Engine icing may occur before ice formation is observed on the wings, there-
fore, surface icing should not be used to verify possible engine icing. The EN-
GINE ANTI-ICE system must be operated any time the airplane is operated
in visible moisture below +10°C indicated ram air temperature (RAT) or
when airframe icing is occurring. Refer to Section II of the Airplane Operating
Manual and/or Chapter 10 of the FSI Pilot Training Manual (PTM), Vol. 2,
for an explanation of the ice protection systems.
NOTE
If ambient temperature is approximately 15°C or
warmer, the ENG ANTI-ICE L/R annunciators may
not illuminate when anti-ice is selected ON. To en-
sure that bleed air is flowing to the engine inlet, the
crew should observe a momentary small decrease in
N 2 when ENGINE ON is selected.
During sustained ground operations in freezing pre-

cipitation the engines should be operated for 15 sec-
onds out of every 4 minutes at 60% N 2 or above to
preclude ice forming on engine probes or internal
components.
CAUTION
During sustained ground operations in freezing con-

ditions, if the engines are operated at idle, ice may
form on engine probes and internal components. This
may cause engine vibration and erroneous RAT in-
dications. By increasing the engine speed to 60% N 2
or higher, the engine vibration will be eliminated
and the RAT indication will read correctly. The pilot
should accomplish this procedure prior to reading
RAT to compute takeoff N 1 settings.

PASSENGER COMFORT
Passenger comfort can be broadly delineated into two categories of environ-
mental/pressurization and pilot technique. Some pointers are as follows:
• When parked in daylight in hot weather, it is suggested the cabin

window shades be closed to reduce solar heat transfer. An optional
exterior windshield cover performs the same function for the cockpit,
and is very effective.
• The interior temperature can be controlled on the ground by use of the

vapor-cycle air conditioning (operating from generator or GPU power)
and/or air-cycle machine. In flight, the vapor-cycle air-conditioning
system can be used up to 18,000 feet to augment ACM cooling
capabilities (if installed). Refer to Section II of the Airplane Operating
Manual, Environmental and Temperature Control and/or Chapter 11 of
the FSI PTM, Vol. 2, for a complete description and operation of
system components.
• To warm the interior, a combination of hot engine bleed air is mixed

with cold air from the air-cycle machine. This temperature can be
controlled through a wide range of settings. Refer to Section II of the
Airplane Operating Manual, Environmental and Temperature Control,
and/or Chapter 11 of the FSI PTM, Vol. 2, for a complete description
and operation of the system components.
• Increasing or decreasing engine bleed air extraction can cause a slight

momentary bump in cabin pressure. Always check power stabilized at
idle when changing the PRESS SOURCE SELECT on the ground.
• The abbreviated checklist is designed to enable the cockpit crew to

perform all prestart functions in advance. This permits items such as
the warning test to be completed before cabin crew and passenger
boarding, and accelerates the ramp departure without compromising
safety or thoroughness.
• Leaving the chocks, brake checks can be done lightly and smoothly. If
heavy braking is required on landing roll, using up elevator to create
drag also counters the nose down pitching moment, so that deceleration
feel in the cabin is less abrupt. Do not apply excessive back pressure,
as weight may be lifted from the main wheels, decreasing braking
effectiveness and increasing the possibility of a blown tire.
• Utilizing proper pressurization system procedures, coupled with a

thorough understanding of the automatic controller and indicators
greatly simplifies operation. Optimum system performance, in terms of
passenger comfort, is best achieved by proper selection of landing field
elevation and by not making power changes simultaneously.
• Although it is not mandatory, use of the yaw damper is recommended

when hand flying the airplane. It reduces pilot rudder input required
and the airplane rides better in rough air. The yaw damper must be off
for takeoff and landing.

• Power management has an impact on cabin comfort and changes

should be made smoothly and symmetrically. An approximate estimate
of synchronization can be made by observing the rpm gages, and exact
adjustments made audibly or with the engine synchronizer. Although
the higher pitched turbine sound is generally more noticeable in the
cockpit, the lower, fan out-of-synchronization sound is usually more
pronounced in the area of the rear seats.
• Good crew coordination and smooth operation of controls and systems

serves the best interests of safety, economy, and passenger comfort.
BIRD INGESTION PRECAUTIONS

Studies have indicated that bird strikes are more likely to occur from the sur-
face to approximately 4,000 feet AGL. As a precaution against engine flame-
out due to bird ingestion, it is recommended that the engine ignitors be ON
when flying at or below 4,000 feet AGL, or anytime the crew has reason to
suspect that the potential for a bird strike exists.
TURBULENT AIR PENETRATION

Flight through severe turbulence should be avoided if possible. The follow-
ing procedures are recommended for flight in severe turbulence.
1. Ignition................................................................................................... ON
2. Airspeed ....................... Approximately 180 KIAS (do not chase airspeed)
3. Maintain a constant attitude without chasing the altitude. Avoid sudden

large control movements.
4. Operation of autopilot is recommended using basic pitch and lateral mode

only.
COLD WEATHER OPERATION

Operation of the airplane has been demonstrated after prolonged exposure to
ground ambient temperature of –40°C (–40°F). This was the minimum tem-
perature achieved in cold weather testing. The operational procedures in this
section are recommended for operations where prolonged exposure to tem-
peratures below –10°C (+14°F) is anticipated or has occurred.
1. If the aircraft has been cold soaked at temperatures below –10°C (+14°F)
it is recommended the battery and crew oxygen masks be removed and
stored at a temperature above –10°C (+14°F). If the battery has been cold
soaked at temperatures below –10°C (+14°F), battery warmup to at least
–10°C (+14°F) is required. This temperature may be checked with the
battery temperature gage. Proper battery warmup may require extended
application of heat to the battery.

2. The use of engine preheat should not be required at temperatures down to

–40°C (–40°F). However, it should be verified after engine start and
before flight there are no visible oil leaks.
3. The avionics may require warmup after cold soak. This may require as
long as 30 minutes. All avionics must be operating properly before flight
as indicated by the following:
a. RAT indication stable and correct.
b. Standby Flight Display aligned and indicating correctly.
c. PFDs and MFD including air data displays indicating correctly.
d. FMS CDUs and Radio Management Units (RMUs) indicating and

operating correctly with no visible waviness or distortion.
e. Audio reception is available on all applicable avionics.
4. After 2 hours or longer exposure to ambient temperature of –10°C (+14°F)

or colder, cabin temperature must be held at or above +10°C (+50°F) for a
minimum of 15 minutes prior to takeoff. This temperature ensures proper
deployment and operation of the passenger oxygen masks. Cabin
temperature can be determined by the CAB TEMP indicator or a handheld
thermometer. This can be accomplished by taxiing the airplane to a
suitable area and increasing power above idle (approximately ≥65% N2) to
obtain duct supply temperatures of approximately 200°F.
Engine preheating is best accomplished by installing the engine covers and

directing hot air through the oil filler access door. A heater hose can be placed
in the tail cone with the door propped as far closed as possible to minimize
heat loss. With sufficient hose length, the cabin and cockpit area can be
warmed through the pilot’s side window.
The W/S TEMP annunciator may not test after cold soak at extremely cold
temperatures. If this occurs, repeat the test after the cabin has warmed up.
The test must be completed prior to flight.
If a start is attempted and the starter will not motor to 8% N 2 minimum, ter-
minate the start sequence. Advancing the throttle to idle below 8% N 2 can be
damaging to the engine and battery. Battery voltage below 11 volts after the
start button is pressed indicates a potential for an unsuccessful start.
Do not set the parking brake if the anticipated cold soak temperature is -15°C
(5°F) or below.
Maximum heat from the air-conditioning system is obtained with the right en-
gine operating and the PRESS SOURCE SELECT in NORM. Switching the tem-
perature control selector to MANUAL, and selecting MANUAL HOT for 10
seconds, ensures the temperature mixing valve is in the full hot position. Turning
on the CKPT RECIRC fan to HI will increase air circulation in the cockpit.
Operating the right engine above idle rpm increases temperature and airflow.

Utilizing APU bleed air, if equipped, will heat the interior much quicker than
engine bleed on the ground.
Because the airplane utilizes two separate controls for the cockpit and the cabin,
comfortable temperature ranges can be obtained at both locations. Separate
zone sensors for both the cockpit and cabin ensure accurate readings through-
out the comfort range.
Use of MANUAL mode of the AUTO TEMP SELECT should be restricted to

below 31,000 feet altitude in order to prevent possible overheating of the air
cycle machine, which would result in automatic actuation of the emergency
pressurization system.
Operating in extremely cold temperatures reduces the solubility and super cools
any water particles in the fuel, increasing the possibility of fuel system icing.
The five tank, and one fuel filter drains under each wing should be drained fre-
quently and thoroughly. It is possible for water to settle in the sump and freeze,
blocking the drain, in which case heat should be applied until fuel flows freely.
Maintain heat after flow begins to ensure all particles have melted and collect
the drainage in a clear, clean container to inspect for water globules.
GROUND DEICE/ANTI-ICE OPERATIONS

Ground deicing/anti-icing procedures are contained in the Airplane Operating
Manual, Section IV, OPERATING INFORMATION, or the Airplane Flight
Manual (AFM), Section VII, ADVISORY INFORMATION.
SERVICING—XLS AND EXCEL

FUEL
A variety of fuels can be used in the airplane. Commercial kerosene Jet-A,
Jet A-1, JET-B, JET-3, JP-4, JP-5 and JP-8 are approved fuels.
Ethylene glycol monomethyl ether (EGME) and diethylene glycol monomethyl

ether (DIEGME) are approved for use but are not required.
Procedure For Adding Ethylene Glycol Monomethyl Ether

(EGME) Fuel Additive
Use the following procedure to blend anti-icing additive as the airplane is being
refueled through the wing filler caps:
1. Attach MIL-I-27686 additive to refuel nozzle, making sure blender tube

discharges in the refueling stream.
2. Start refueling while simultaneously fully depressing and slipping ring

over trigger of blender.

WARNING
Anti-icing additives containing ethylene glycol
monomethyl ether (EGME) are harmful if inhaled,
swallowed, or absorbed through the skin, and will
cause eye irritation. Also, it is combustible. Before
using this material, refer to all safety information on
the container.
CAUTION
Assure the additive is directed into the flowing fuel
stream and the additive flow is started after the fuel
flow starts and is stopped before fuel flow stops. Do
not allow concentrated additive to contact coated in-
terior of fuel tank or airplane painted surface.
Use not less than 20 fluid ounces of additive per 156

gallons of fuel or more than 20 fluid ounces of addi-
tive per 104 gallons of fuel.
Procedure For Adding DIethylene Glycol Monomethyl

Ether (DIEGME) Fuel Additive
NOTE
Service experience has shown that DIEGME has pro-
vided acceptable protection from bacterial growth in
fuel systems.
Use the following procedure to blend anti-icing additive as the airplane is being
refueled through the wing filler caps:
1. Attach MIL-I-85470 additive to refuel nozzle, making sure blender tube

discharges in the refueling stream.
2. Start refueling while simultaneously fully depressing and slipping ring

over trigger of blender.
CAUTION
Diethylene glycol monomethyl ether (DIEGME) is

slightly toxic if swallowed and may cause eye red-
ness, swelling and irritation. Also, it is combustible.
Before using this material, refer to all safety infor-
mation on the container. Assure the additive is directed
into the flowing fuel stream with the additive flow
started after the fuel flow starts and stopped before
fuel flow stops. Do not allow concentrated additive
to contact coated interior of fuel tank or airplane
painted surface.

Use not less than 20 fluid ounces of additive per 156

gallons of fuel or more than 20 fluid ounces of addi-
tive per 104 gallons of fuel.
Procedure For Checking Fuel Additives

1. Prolonged storage of the airplane will result in a water buildup in the fuel
which “leaches out” the additive. An indication of this is when an
excessive amount of water accumulates in the fuel tank sumps. The
concentration can be checked using an anti-icing additive concentration
test kit available from Cessna Aircraft Company, Citation Marketing
Division, Wichita, KS 67277. It is imperative that the instructions for the
test kit be followed explicitly when checking the additive concentration.
The concentrations by volume for the EGME/DIEGME shall be 0.10
percent minimum and 0.15 percent maximum, either individually or mixed
in a common tank. Fuel, when added to the tank, should have a minimum
concentration of 0.10 percent by volume.
OIL
Each engine oil tank has an oil filler neck with cap assembly and sight indi-
cator. Oil is added to each engine directly through the filler neck and quan-
tity is measured at the sight indicator in U.S. quarts. An accurate check of oil
quantity can only be made when the engine is hot, and should be accomplished
10 minutes after engine shutdown.
CAUTION
Persons who handle engine oil are advised to mini-
mize skin contact with used oil, and promptly re-
move any used oil from their skin. A laboratory study,
while not conclusive, found substances which may
cause cancer in humans. Thoroughly wash used oil
off skin as soon as possible with soap and water. Do
not use kerosene, thinners or solvents to remove used
engine oil. If waterless hand cleaner is used, always
apply skin cream after using.
BRITISH PETROLEUM 2380, CASTROL 5000, AEROSHELL TURBINE

OIL 500, AEROSHELL TURBINE OIL 560, ROYCO TURBINE OIL 500,
MOBIL JET OIL 254 and MOBIL JET OIL II are all approved oils. Normally
different brands of oil should not be mixed; however, if oil replenishment is
required, and oil of the same brand as tank contents is not available, follow
procedures set forth in Section I of the Operating Manual under APPROVED
OILS. The type of oil used in each airplane is noted in the engine logbook,
as well as on a placard inside the filler access door.
The latest revision of Pratt and Whitney Canada, Inc. Service Bulletin 7001
may also be consulted for approved oils.

CAUTION
When changing from an existing lubricant formula-
tion to a “third generation” lubricant formulation
(Aeroshell Turbine Oil 560 or Mobil Jet 254), the en-
gine manufacturer strongly recommends that such a
change should only be made when an engine is new
or freshly overhauled. For additional information on
use of third generation oils, refer to the engine man-
ufacturer’s pertinent oil service bulletins.
HYDRAULIC
Servicing the main hydraulic reservoir requires equipment capable of deliv-
ering hydraulic fluid under pressure and is normally performed by mainte-
nance personnel. The reservoir should be serviced with one of the approved
fluids: SKYDROL 500 B-4, or LD-4, LD-5; or Hyjet IVA Plus only.
The hydraulic brake reservoir can be serviced by removing the left nose com-
partment lower liner to allow access to the brake reservoir. The filler plug can
then be removed and the reservoir filled to within one-half inch of the open-
ing. The brake reservoir should be serviced with one of the approved fluids,
SKYDROL 500B or equivalent.
OXYGEN
The oxygen filler valve is located just inside the access door in the right for-
ward avionics compartment, near the aft end of the compartment. Oxygen serv-
icing should be done by maintenance personnel using breathing oxygen
conforming to MIL-O-27210, Type 1. Refer to the cockpit gage while serv-
icing to prevent overfill.
Oxygen pressure will vary with ambient temperature. In very cold ambient
temperatures, the oxygen pressure indication may appear low, but may, in ac-
tuality, be appropriate for the temperature condition.
NOTE
Refer to Chapter 12 of the Airplane Maintenance
Manual, Oxygen Service Requirements, Pressure
Variations Chart.
FIRE BOTTLES
Under-serviced fire bottles must be exchanged by authorized maintenance
facilities.

LANDING GEAR AND BRAKES PNEUMATIC SYSTEM

The emergency gear and brake bottle should be serviced when the pressure
gage reads below 1,800 psi. Maintenance personnel should perform the serv-
icing with high pressure nitrogen and refill the bottle to 2,050 psi. Servicing
is accomplished through a charging valve on the bottle pressure gage.
TIRES
Main gear tire pressures should be maintained at 210 psi and nose tire at 130
psi. Since tire pressure will decrease as the temperature drops, a slight over
inflation can be used to compensate for cold weather. Main tires inflated at
21°C should be overinflated 1.5 psi for each 6°C drop in temperature antic-
ipated at the coldest airport of operation. Nose tires at 21°C should be over-
inflated only 0.5 PSI for each 6°C anticipated drop in temperature.
Worn tires and underinflated tires both contribute to lowering the speed at
which hydroplaning occurs on precipitation covered runways. Refer to Adverse
Field Conditions in this section for a discussion of hydroplaning.
TOILET
The airplane may be equipped with either a carry out flush toilet or an exter-
nally serviceable flush toilet. Both types require servicing when the liquid
level becomes too low or when the liquid appears to have incorrect chemical
balance. Instructions for servicing the toilets are found in Chapter 12 of the
Airplane Maintenance Manual.
AIRPLANE CLEANING AND CARE

Painted Surfaces
The exterior of a new airplane is painted with a polyurethane two-compo-
nent topcoat which, unlike early coatings, does not require exposure to air
for complete cure to occur. The care required by the finish will not change
as the paint ages.
The finish should be cleaned only by washing with clean water and mild
soap, followed by rinse water and drying with a soft cloth or chamois.
Minimize flying through rain, hail or sleet for a few weeks to protect the new
paint.
To help prevent development of corrosion, particularly filiform corrosion, the

airplane should be spray-washed at least every two or three weeks (especially
in warm, damp, and salty environments) and waxed with a good grade of water
repellent wax to help keep water from accumulating in skin joints and around
countersinks. A heavier coating of wax on the leading edge, on the vertical tail
and on the engine nose cones helps reduce abrasions encountered in these areas.

Polyurethane topcoats are designed with UV inhibitors to slow the degrada-

tion caused by exposure. The inhibitors concentrate near the surface of the
coating during the initial stages of cure. Care must be taken during any buff-
ing, polishing, or power waxing so this surface layer is disturbed only to the
smallest extent necessary. However, with special care, buffing, polishing, and
power waxing is acceptable. Wax products containing silicones should be
avoided as they contribute to buildup of P-static, especially if the surface is
well buffed to produce a shine.
DEICE BOOTS
The deice boots on the horizontal stabilizer leading edges have a special elec-
trically-conductive coating to bleed off static charges which cause radio in-
terference and may perforate the boots. Servicing operations should be done
carefully, to avoid damaging this conductive coating or tearing the boots.
To prolong the life of surface deice boots, they should be washed and serv-
iced on a regular basis. Keep the boots clean and free from oil, grease and
other solvents which cause rubber to swell and deteriorate. Clean the boots
with mild soap and water, then rinse thoroughly with clean water. Outlined
below are the recommended cleaning and servicing procedures.
CAUTION
Use only the following instructions when clean-
ing boots. Disregard instructions which recom-
mend Petroleum-based liquids (methyl-
ethylketone, nonleaded gasoline, etc.) which can
harm the boot material.
NOTE
Isopropyl alcohol can be used to remove grime which
cannot be removed using soap. If isopropyl alcohol
is used for cleaning, wash area with mild soap and
water, then rinse thoroughly with clean water.
To improve the service life of the boots and to reduce the adhesion of ice, it
is recommended that the deice boots be treated with AGE MASTER No. 1
or ICEX.
AGE MASTER No. 1, used to protect the rubber against deterioration from
ozone, sunlight, weathering, oxidation and pollution, and ICEX, used to help
retard ice adhesion and for keeping deice boots looking new longer, are both
products of, and recommended by, B.F. Goodrich.
The application of both AGE MASTER No. 1 and ICEX should be in accor-
dance with the manufacturer’s recommended directions as outlined on the
containers.

CAUTION
Protect adjacent areas, clothing, and use plastic or rub-
ber gloves during applications, as Age Master No. 1
stains and ICEX contains silicone which makes paint
touchup almost impossible.
Ensure the manufacturer’s warnings and cautions are

adhered to when using Age Master No. 1 and ICEX.
If a high gloss finish is desired on the deice boots, AKROSEAL coating

(available from Huber Janitorial Supplies, 114 North St. Francis Street,
Wichita, Ks, 67202) may be used in lieu of AGE MASTER NO. 1 and/or ICEX.
Preparation for application of ACROSEAL is the same as required for AGE
MASTER NO. 1 and ICEX. Apply a thin layer of ACROSEAL on the clean
and dry surface of the deice boot with a cloth swab. Let dry thoroughly and
hand buff with a soft cloth.
Small tears and abrasions can be repaired temporarily without removing the
boots and the conductive coating can be renewed.
ENGINES
The engine compartments should be cleaned using a suitable solvent. Most
efficient cleaning is done using a spray-type cleaner. Before spray cleaning,
ensure protection is afforded for other components which may be adversely
affected by the solvent. Refer to the Airplane Maintenance Manual for proper
lubrication of components after engine cleaning.
INTERIOR CARE
To remove dust and loose dirt from the upholstery, headliner and carpet,
clean the interior regularly with a vacuum cleaner.
Blot any spilled liquid promptly with cleansing tissue or rags. Do not pat the
spot; press the blotting material firmly and hold it for several seconds.
Continue blotting until no more liquid is absorbed. Scrape off sticky materi-
als with a dull knife, then spot clean the area.
Oily spots may be cleaned with household spot removers, used sparingly. Before
using any solvent, read the instructions on the container and test it on an ob-
scure place on the fabric to be cleaned. Never saturate the fabric with a
volatile solvent; it may damage the padding and backing material.

WARNING
Use all cleaning agents in accordance with the man-
ufacturer’s recommendations. The use of toxic or
flammable cleaning agents is discouraged. If these
cleaning agents are used, ensure adequate ventilation
is provided to prevent harm to the user and/or dam-
age to the airplane.
Soiled upholstery and carpet may be cleaned with foam-type detergent, used
according to the manufacturer’s instructions. To minimize wetting the fab-
ric, keep the foam as dry as possible and remove it with a vacuum cleaner.
The plastic trim, instrument panel and control knobs need only be wiped
with a damp cloth. Oil and grease on the control wheel and control knobs can
be removed with a cloth moistened with kerosene. Volatile solvents, such as
mentioned in paragraphs on care of the windshield, must never be used since
they soften and craze the plastic.
WINDOWS AND WINDSHIELDS

The glass windshields and forward (fixed) cockpit side windows, and the acrylic
aft (openable) cockpit windows, and the cabin windows should be kept clean
at all times. Recommended products and materials for washing and protect-
ing the windows and windshields are listed in Chapter 12 of the Airplane
Maintenance Manual. The acrylic windows should be kept clean and waxed
at all times. To prevent scratches and crazing, wash them carefully with
plenty of soap and water, using the palm of the hand to feel and dislodge dirt
and mud. A soft cloth, chamois or sponge may be used, but only to carry water
to the surface. Rinse thoroughly, then dry with a clean, moist chamois.
Rubbing the surface of the plastic with a dry cloth builds up an electrostatic
charge which attracts dust particles in the air. Wiping with a moist chamois
will remove both the dust and this charge.
Remove oil and grease with a cloth moistened with kerosene. Never use gaso-
line, benzine, acetone, carbon tetrachloride, fire extinguisher fluid, lacquer
thinner or glass cleaner. These materials will soften the acrylic and may
cause it to craze.
After removing dirt and grease, if the surface is not badly scratched, it should
be waxed with a good grade of commercial wax. The wax will fill in minor
scratches and help prevent further scratching. Apply a thin, even coat of wax
and bring it to a high polish by rubbing lightly with a clean, dry soft flannel
cloth. If the surface is badly scratched, refer to the Airplane Maintenance
Manual for approved repairs.
Do not use a canvas cover on the windshield, unless freezing rain or sleet is
anticipated. Canvas covers may scratch the acrylic surface.

OXYGEN MASKS
The crew masks are permanent-type masks which contain a microphone for
radio transmissions. The passenger masks are oro-nasal type which forms
around the mouth and nose area. All masks can be cleaned with alcohol. Do
not allow solution to enter microphone or electrical connections. Apply tal-
cum powder to external surfaces of passenger mask rubber face-piece.


CONTENTS
Page
CREW RESOURCE MANAGEMENT......................................... CRM-1
CREW CONCEPT BRIEFING GUIDE ........................................ CRM-5
Introduction .......................................................................... CRM-5
Common Terms .................................................................... CRM-5
Pretakeoff Briefing (IFR/VFR) ............................................ CRM-5
Crew Coordination Approach Sequence .............................. CRM-6
FOR TRAINING PURPOSES ONLY CRM-i

ILLUSTRATIONS
Figure Title Page
CRM-1 Situational Awareness in the Cockpit ........................ CRM-1
CRM-2 Command and Leadership ........................................ CRM-1
CRM-3 Communication Process............................................ CRM-2
CRM-4 Decision Making Process .......................................... CRM-2
CRM-5 Crew Performance Standards .................................... CRM-3
FOR TRAINING PURPOSES ONLY CRM-iii

CAPTAIN COPILOT REMEMBER

INDIVIDUAL INDIVIDUAL
S/A S/A 2+2=2
— OR —
2+2=5
GROUP (SYNERGY)
S/A
IT'S UP TO YOU!
CLUES TO IDENTIFYING:
• Loss of Situational Awareness
• Links in the Error Chain
OPERATIONAL
1. FAILURE TO MEET TARGETS

2. UNDOCUMENTED PROCEDURE
3. DEPARTURE FROM SOP
4. VIOLATING MINIMUMS OR LIMITATIONS
5. NO ONE FLYING AIRPLANE
6. NO ONE LOOKING OUT WINDOW
7. COMMUNICATIONS
HUMAN
8. AMBIGUITY
9. UNRESOLVED DISCREPANCIES
10. PREOCCUPATION OR DISTRACTION
11. CONFUSION OR EMPTY FEELING
12.
Figure CRM-1. Situational Awareness in the Cockpit
LEADERSHIP STYLES
LAISSEZ-
AUTOCRACTIC AUTHORITARIAN DEMOCRATIC
FAIRE
STYLE LEADERSHIP LEADERSHIP
STYLE
(EXTREME) STYLE STYLE
(EXTREME)
PARTICIPATION
LOW HIGH
COMMAND — Designated by Organization
— Cannot be Shared
LEADERSHIP — Shared Among Crewmembers
— Focuses on "What's Right," not "Who's Right"
Figure CRM-2. Command and Leadership
FOR TRAINING PURPOSES ONLY CRM-1

INTERNAL EXTERNAL INTERNAL

BARRIERS BARRIERS BARRIERS
THINK:
RECEIVE
OPERATIONAL • Solicit and give
NEED SEND GOAL feedback
• Listen carefully
• Focus on behavior,
not people
• Maintain focus on
the goal
• Verify operational
outcome is achieved
FEEDBACK
ADVOCACY: To increase other's S/A INQUIRY: To increase your own S/A

• State Position • Decide What, Whom, How to ask
• Suggest Solutions • Ask Clear, Concise Questions
• Be Persistent and Focused • Relate Concerns Accurately
• Listen Carefully • Draw Conclusions from Valid Information
• Keep an Open Mind
— REMEMBER —
Questions enhance communication flow. Do not give in to the temptation to ask questions
when advocacy is required. Use of advocacy or inquiry should raise a "red flag."
Figure CRM-3. Communication Process
HINTS:
• Identify the problem:
• Communicate it
EVALUATE RECOGNIZE • Achieve agreement
RESULT NEED • Obtain commitment
IDENTIFY
AND
• Consider appropriate SOPs
DEFINE
IMPLEMENT PROBLEM • Think beyond the obvious
RESPONSE alternatives
COLLECT • Make decisions as a result of
FACTS
the process
SELECT A IDENTIFY
RESPONSE ALTERNATIVES • Resist the temptation to make
WEIGH IMPACT
OF ALTERNATIVES an immediate decision and
then support it with facts.
Figure CRM-4. Decision Making Process
CRM-2 FOR TRAINING PURPOSES ONLY

SITUATIONAL AWARENESS
a. Accomplishes appropriate preflight planning.
b. Sets and monitors targets.
c. Stays ahead of the aircraft by preparing for expected or contingency
situations.
d. Monitors weather, aircraft systems, instruments, and ATC communications.
e. Shares relevant information with the rest of the crew.
f. Uses advocacy/inquiry to maintain/regain situational awareness.
g. Recognizes error chain clues and takes actions to break links in the chain.
h. Communicates objectives and gains agreement when appropriate.
i. Uses effective listening techniques to maintain/regain situational awareness.
STRESS
a. Recognizes symptoms of stress in self and others.
b. Maintains composure, calmness, and rational decision making under stress.
c. Adaptable to stressful situations/personalities.
d. Uses stress management techniques to reduce effects of stress.
e. Maintains open, clear lines of communications when under stress.
COMMUNICATION
a. Establishes open environment for interactive communication.
b. Conducts adequate briefings to convey required information.
c. Recognizes and works to overcome barriers to communications.
d. Operational decisions are clearly stated to other crewmembers and
acknowledged.
e. Crewmembers are encouraged to state their own ideas, opinions, and
recommendations.
f. Crewmembers are encouraged to ask questions regarding crew actions.
g. Assignments of blame is avoided. Focuses on WHAT is right, and not WHO is
right.
h. Keeps feedback loop active until operational goal/decision is achieved.
i. Conducts debriefings to correct substandard/inappropriate performance and to
reinforce desired performance.
Figure CRM-5. Crew Performance Standards (Sheet 1 of 2)

SYNERGY AND CREW CONCEPT

a. Ensures that group climate is appropriate to operational situation.
b. Coordinates flight crew activities to achieve optimum performance.
c. Uses effective team building techniques.
d. Demonstrates effective leadership and motivation techniques.
e. Uses all available resources.
f. Adapt leadership style to meet operational and human requirements.
WORKLOAD MANAGEMENT
a. Communicates crew duties and receives acknowledgement.
b. Sets priorities for crew activities.
c. Recognizes and reports overloads in self and in others.
d. Eliminates distractions in high workload situations.
e. Maintains receptive attitude during high workload situations.
f. Uses other crewmember.
g. Avoids being a "one man show."
DECISION MAKING
a. Anticipates problems in advance.
b. Uses SOPs in decision making process.
c. Seeks information from all available resources when appropriate.
d. Avoids biasing source of information.
e. Considers and weighs impact of alternatives.
f. Selects appropriate courses of action in a timely manner.
g. Evaluates outcome and adjusts/reprioritizes.
h. Recognizes stress factors when making decisions and adjusts accordingly.
i. Avoids making a decision and then going in search of facts that support it.
ADVANCED/AUTOMATED COCKPITS
a. Follows automation related SOPs.
b. Specifies pilot and copilot duties and responsibilities with regard to
automation.
c. Verbalizes and acknowledges entries and changes in flight operation.
d. Verifies status and programming of automation.
e. Selects appropriate levels of automation.
f. Programs automation well in advance of maneuvers.
g. Recognizes automation failure/invalid output indications.
Figure CRM-5 Crew Performance Standards (Sheet 2 of 2)

CREW CONCEPT BRIEFING GUIDE

INTRODUCTION
Experience has shown that adherence to SOPs helps to enhance individual and
crew cockpit situational awareness and will allow a higher performance level
to be attained. Our objective is for standards to be agreed upon prior to flight
and then adhered to, such that maximum crew performance is achieved. These
procedures are not intended to supersede any individual company SOP, but
rather are examples of good operating practices.
See Maneuvers and Procedures chapter for call-outs.
COMMON TERMS
PIC Pilot in Command
Designated by the company for flights requiring more than one pilot.
Responsible for conduct and safety of the flight. Designates pilot
flying and pilot not flying duties.
PF Pilot Flying
Controls the airplane with respect to assigned runway, course, alti-

tude, airspeed, etc., during normal and emergency conditions.
Accomplishes other tasks as directed by the PIC.
PNF Pilot Not Flying
Maintains ATC communications, copies clearances, accomplishes

checklists and other tasks as directed by the PIC.
B Both
PRETAKEOFF BRIEFING (IFR/VFR)

NOTE
The following briefing is to be completed during
item 1 of the Pretakeoff checklist. The PF will accom-
plish the briefing.
Pilot flying will review at a minimum:
1. Runway (direction and condition)
2. SID/DP/STAR/FMSP/IAP
3. Power settings, speeds
4. Abnormal/emergency procedures prior to and after decision speed

5. Emergency return intentions
6. Instrument approach procedure: FAF, FAF altitude, initial rate of descent,

DA/DH/MDA, time to missed approach point (if required), missed
approach procedure
7. Crewmember responsibilities during takeoff/DP and approach/landing
8. Ask for and clarify questions regarding the briefing
CREW COORDINATION APPROACH SEQUENCE

During the Approach Checklist, the PF transfers aircraft control and briefs
the crew coordination approach sequence. This should be completed as early
as possible, prior to initiating an IFR approach (prior to the FAF).
See Maneuvers and Procedures section for call-outs.

SYSTEMS REVIEW—EXCEL
CONTENTS
Page
SQUAT SWITCH INPUTS............................................................. SRE-1
EMERGENCY BUS CONDITION................................................ SRE-2
LIGHTING ..................................................................................... SRE-3
Cockpit Panel Lights ............................................................. SRE-3
Cockpit Overhead Lights....................................................... SRE-3
Cabin Lighting....................................................................... SRE-4
Emergency Lighting (EMER LTS)........................................ SRE-6
Exterior Lights....................................................................... SRE-7
Tail Cone Compartment lights .............................................. SRE-8
Pulselite system (Optional) ................................................... SRE-8
ELECTRICAL SYSTEM ............................................................... SRE-8
POWERPLANT............................................................................ SRE-16
Ignition ................................................................................ SRE-18
FIRE PROTECTION .................................................................... SRE-19
Sensing Loops and Control Units ....................................... SRE-19
Operation............................................................................. SRE-20
FUEL ............................................................................................ SRE-21
HYDRAULICS............................................................................. SRE-25
POWER BRAKES AND ANTISKID........................................... SRE-34
EMERGENCY BRAKES............................................................. SRE-36
FLIGHT CONTROLS .................................................................. SRE-36
ICE AND RAIN PROTECTION .................................................. SRE-44
PNEUMATICS/AIR CONDITIONING....................................... SRE-54
PRESSURIZATION ..................................................................... SRE-60
SERVICE AIR .............................................................................. SRE-65
OXYGEN...................................................................................... SRE-66
AUXILIARY POWER UNIT ....................................................... SRE-68
Electronic Control Unit (ECU) ........................................... SRE-68
FOR TRAINING PURPOSES ONLY SRE-i

Fuel System......................................................................... SRE-69

Oil System........................................................................... SRE-69
Pneumatic System ............................................................... SRE-71
Electrical System................................................................. SRE-73
Fire Protection..................................................................... SRE-73
Exterior Preflight................................................................. SRE-75
APU Control Panel and Annunciator Functions ................. SRE-75
APU Start Sequence............................................................ SRE-78
Start Power Logic................................................................ SRE-80
APU Operating Limitations ................................................ SRE-84
Battery and APU Starter Cycle Limitations ........................ SRE-86
AVIONICS.................................................................................... SRE-86
ANTENNA AND DRAIN TUBE................................................. SRE-98
SRE-ii FOR TRAINING PURPOSES ONLY

ILLUSTRATIONS
Figure Title Page
SRE-1 Cabin/Entry Lights Panel .......................................... SRE-5
SRE-2 DC Power Distribution .............................................. SRE-9
SRE-3 Pilot Circuit-Breaker Panel ...................................... SRE-10
SRE-4 Copilot Circuit-Breaker Panel.................................. SRE-11
SRE-5 PW545A Cross-Section .......................................... SRE-17
SRE-6 Engine Fire Extinguishing System .......................... SRE-20
SRE-7 Engine Fire Detection System ................................ SRE-21
SRE-8 Fuel System—Normal Operation ............................ SRE-22
SRE-9 Fuel System—Crossfeed (R to L)............................ SRE-24
SRE-10 Hydraulic System—Open Center ............................ SRE-26
SRE-11 Speedbrake System—Normal Operation
(Extended)................................................................ SRE-27
SRE-12 Gear System—Normal Retraction .......................... SRE-28
SRE-13 Gear System—Normal Extension............................ SRE-29
SRE-14 Gear System—Emergency Extension...................... SRE-30
SRE-15 Thrust Reversers—Stowed ...................................... SRE-32
SRE-16 Thrust Reversers—Deployed .................................. SRE-33
SRE-17 Power Brake/Antiskid System ................................ SRE-35
SRE-18 Flight Controls ........................................................ SRE-37
SRE-19 Rudder Bias System ................................................ SRE-39
SRE-20 Rudder Bias System—Engine Failure .................... SRE-39
SRE-21 Two-Position Horizontal Stabilizer.......................... SRE-43
SRE-22 Pitot-Static System .................................................. SRE-45
SRE-23 Windshield Anti-Ice System .................................... SRE-47
SRE-24 Wing/Engine Anti-Ice System ................................ SRE-49
SRE-25 Wing Leading Edge Cross Section .......................... SRE-51
SRE-26 Tail Deice System .................................................... SRE-53
SRE-27 Bleed-Air Precooler ................................................ SRE-55
SRE-28 Air Conditioning System with APU ........................ SRE-57
SRE-29 Vapor Cycle Air Conditioning System .................... SRE-59
FOR TRAINING PURPOSES ONLY SRE-iii

SRE-30 Pressurization Control Panel.................................... SRE-61

SRE-31 Pressurization System .............................................. SRE-62
SRE-32 Autoschedule Boundary .......................................... SRE-63
SRE-33 High Altitude Landing Graph .................................. SRE-64
SRE-34 High Altitude Departure Graph................................ SRE-64
SRE-35 Service Air System .................................................. SRE-65
SRE-36 Oxygen System ........................................................ SRE-67
SRE-37 APU Annunciators, Copilot Panel .......................... SRE-70
SRE-38 APU Control Panel .................................................. SRE-72
SRE-39 First Engine Start (R)—APU
Generator On Line .................................................. SRE-74
SRE-40 APU Engine Start On Ground (Engines OFF) ........ SRE-79
SRE-41 APU Start—On Ground, Using GPU ...................... SRE-81
SRE-42 Second Engine Start (L)—APU
Generator On Line .................................................. SRE-82
SRE-43 APU Start On Ground (Generator Assist)................ SRE-83
SRE-44 APU Start—In Flight (Battery Only) ...................... SRE-85
SRE-45 Primus 1000 System Block Diagram ...................... SRE-87
SRE-46 Standby Flight Display—Meggitt............................ SRE-93
SRE-47 Standby HSI ............................................................ SRE-95
SRE-48 Excel Antenna and Drain Tube Locations .............. SRE-99
TABLES
Table Title Page
SRE-1 APU Operating Limits ............................................ SRE-84
SRE-iv FOR TRAINING PURPOSES ONLY

SYSTEMS REVIEW—EXCEL
SQUAT SWITCH INPUTS
Left main squat switch only
In flight, it enables:
1. Flight hour meter
2. Digital clock to record elapsed flight time
3. Opening of emergency pressurization valve
4. Landing gear handle locking solenoid to be energized
5. TAS probe heater (Rosemount)
6. Enables flight idle (with EECs operating)
7. Normal (auto) control of pressurization
8. Enables ram air door modulation for precoolers
On the ground, it enables:

1. Pressurization controller opens outflow valves (<85% N2)
2. Prepressurization during takeoff
3. Generator-assisted starts
4. Engine bypass valve for precooler operation
5. Ground idle (with EECs operating)
6. O v e r r i d e s o p t i o n a l P u l s e l i t e S y s t e m t o s t e a d y i l l u m i n a t i o n
(without the optional pulselite switch)
The safe-flight angle-of-attack indexer lights on the glareshield are enabled

any time aircraft nose gear is down and locked and aircraft is not on the
ground (enables indexer 20 seconds after takeoff with nose gear down).
Left and right squat switches in parallel:
In flight, they enable:

1. Stick-shaker operation
On ground, they enable:

1. Thrust reverser deployment (either squat switch)
2. Stick-shaker test
3. Antiskid locked wheel protection (both squat switches)
FOR TRAINING PURPOSES ONLY SRE-1

EMERGENCY BUS CONDITION

The following components and systems are operative when only the battery
bus and emergency bus(es) are powered (GENs off line, BATT in EMER):
1. Cockpit flood and glareshield lights

2. L and R fan speed/ N 1 indicators (tapes and digits)
3. Standby HSI
4. Standby pitot/static heater
5. Flap control
6. Gear control
7. Hydraulic control valve
8. Gear warning
9. L and R ignition (secondary only)
10. Stabilizer control
11. 1 and 2 audio panels
12. COMM 1
13. NAV 1 (standby HSI)
14. AHRS 2 (standby HSI)
15. RMU 1
16. Standby radio control head
17. Voltmeter (in BATT or EMER only)
18. Emergency exit lights (interior and exterior)
19. ELT
20. Secondary flight display system
21. Auxiliary gear control (manual)
22. Emergency brakes (manual)
23. Manual pressurization control “cherry picker”
24. Cabin altitude/differential pressure indicator
25. Oxygen pressure gauge
26. Passenger oxygen valve (OFF and ON only)
27. Magnetic compass
28. Engine and wing anti-ice bleed air only (no cross-flow)
29. Backlighting for the N 1 gauge, and standby HSI
SRE-2 FOR TRAINING PURPOSES ONLY

LIGHTING
COCKPIT PANEL LIGHTS
Panel lights are controlled by the master panel ON–OFF toggle switch,
(DAY–NIGHT), on the pilot lower instrument panel (PANEL LIGHT).
With master switch ON, the following rheostats control light intensity:
• LEFT DIM: LH digital clock, AOA indicator, voltmeter, LH and RH

ammeters, LH PFD display controller, LH PFD bezel.
• CENTER DIM: Wet compass, standby HSI, pressurization controller,

differential pressure gauge, engine instruments, fuel temperature
gauge, aileron and rudder trim tab indicators, RAT indicator, radar con-
trol panel, MFD controller, MFD bezel, FMS panels.
• RIGHT DIM: Oxygen gauge, RH digital clock, hour meter, battery tem-
perature indicator, RH PFD display controller, RH PFD bezel, cock-
pit voice recorder, ECU controller.
• EL DIM: Throttle quadrant, pilot and copilot switch panels, light

switch panels, tilt panel, thrust reverser switch panels, landing gear
control panel, LH and RH CB panels, LH and RH side wall subpan-
els. (Electrical power to the EL panels is supplied by a small 40–60
VAC inverter in the nose avionics compartment.)
NOTE
Placing the master panel switch ON dims the annun-
ciator panel and ignition lights, and illuminates two
red windshield ice detection post lights. Following
rheostats are powered directly from the emergency
bus (not connected to the master DAY–NIGHT
switch).
• GLARESHIELD AUXILIARY LIGHTS: Rheostat on the pilot side-

wall subpanel.
• FLOOD DIM: Overhead floods on the pilot lower switch panel.
COCKPIT OVERHEAD LIGHTS

Two map/chart lights controlled by LH and RH rheostats on the pilot and copi-
lot sidewall subpanels.
Two sets of paired emergency DC powered floodlights, one set in the over-
head and one set in the bottom of the annunciator panel assembly, illuminate
the forward cockpit area and the engine instruments respectively. They all il-
luminate or extinguish simultaneously by rotating the FLOOD rheostat on the
pilot lower instrument panel.

CABIN LIGHTING
Cabin lighting consists of overhead reading lamps, overhead indirect fluo-
r e s c e n t l i g h t s , a f t va n i t y l i g h t s , f o r wa r d a n d a f t d iv i d e r l i g h t s , N O
SMOKING/FASTEN SEAT BELT and EXIT lights, dropped aisle footwell
lighting, and forward work station lights.
Reading Lights
Directionally adjustable reading lamps are located above each seat including the
aft vanity seat and controlled by switches adjacent to the outboard arm rests.
Indirect Fluorescent Lights

Cabin overhead indirect lighting is controlled by a switch on a cabin light switch
panel on the forward cabin entry door frame. Initially pushing the switch il-
luminates the lights bright and after a few seconds, pushing the switch again
dims the lights. The next push extinguishes the lights. The lights may also be
controlled by a CABIN LIGHT switch on the galley light panel.
Dropped Aisle Lighting

Strip lights on both sides of the footwell aisle are normally controlled by the
AISLE lighting switch on the galley light panel. A portion of each strip light
also illuminates when the emergency exit lights are activated, refer to EMER-
GENCY EXIT LIGHTS below.
Divider, Aft Closet and Work Station Lights

Divider and work station lights are controlled by switches on the galley light
panel. The aft closet light is activated when the closet door is opened and ex-
tinguishes as the closet door is closed.
Cabin/Entry Lights Panel

Cabin entry lights are powered directly from the battery bus and are activated
ON and OFF with ENTRY LIGHT switches on the cabin entry door forward
frame light panel and the galley light panel (Figure SRE-1). Entry lights con-
sist of four reading lamps in the passenger compartment (one above the for-
ward and aft seats on each side) and a light above the emergency exit door in
the aft vanity area. Door entry lights on either side of the lower door frame
and six lights in the door steps are also included.

CABIN
LIGHT
ENTRY
LIGHT
LEFT UPPER DOOR RIGHT UPPER

DOOR PIN DOOR PIN
LOCKS
LEFT LOWER RIGHT LOWER

DOOR PIN DOOR PIN
DOOR
SEAL
CABIN
DOOR
VENT DOOR DOOR HANDLE
Figure SRE-1. Cabin/Entry Lights Panel
Passenger Information Signs

The PASS SAFETY switch on the tilt panel controls passenger information
signs in the cabin as follows:
• PASS SAFETY—ON: Sounds an audible chime and illuminates the

SEATBELT ON/NO SMOKING signs. Illuminates all emergency exit
lights if the EMER LTS switch is ARMED.
• SEAT BELT—ON: Sounds an audible chime and illuminates the
SEATBELT ON only sign.

EMERGENCY LIGHTING (EMER LTS)

Emergency exit lights consist of cabin lights, emergency exit illumination and
identification, emergency egress and ground illumination for emergency
evacuation. Emergency lighting is controlled by the EMER LTS ARM–ON–OFF
switch on the pilot instrument panel. Emergency lighting may be powered by
main DC power or power supplied from two emergency nicad battery packs
(one behind the pilot left side console and one aft in the vanity area). Placing
the EMER LTS switch ON illuminates all emergency exit lights, both inte-
rior and exterior. Placing the switch to ARM illuminates the lights if a 5
gravity (G) force is experienced, a loss of main DC power occurs (i.e., dual
generator failure with BATT switch in EMER, or PASS SAFETY switch is
ON). An amber light adjacent to the switch illuminates with the switch OFF
with main DC power (BATT switch ON) to alert the crew to ARM the system
prior to flight.
Interior Lights
• Three reading lamps in the passenger compartment (one forward right
side, one forward and one aft on left side).
• One reading lamp above emergency exit door in vanity area.
• All EXIT sign lighting.
• RH and LH footwell (dropped aisle) strip lights illuminate to direct

occupants to the exit doors (partial illumination only).
Exterior Lights
• Two lights in the right side fuselage that illuminates top of the right
wing.
• One light in the right side fuselage forward of the wing root that il-
luminates the ground in front of the wing.
NOTE
The forward emergency nicad battery pack provides
emer power to illuminate the exit sign over the cabin
door, the reading light opposite the cabin door, the
reading light just aft of the cabin door and the dropped
aisle strip lighting on left side. The aft emergency
nicad battery pack provides emergency power to il-
luminate the exit signs on the aft divider and above
the emergency exit door, an overhead light above the
emergency exit door, the reading lamp on the left
rear side of the passenger compartment forward of the
aft divider, right side dropped aisle strip lighting, and
the three exterior emergency lights.

EXTERIOR LIGHTS
Navigation
Navigation lights are wing tip lights (red–left, green–right) and a white light
on the tail stinger, controlled by the NAV ON–OFF switch on the tilt panel.
Anticollision
Anticollision lights are high intensity white pulsating strobes mounted on the
extreme outboard of each wingtip, controlled by the GND REC/ANTICOLL
ON–OFF switch on the tilt panel.
Ground Recognition
The ground recognition light is a red beacon light on the top of the rudder. It
is controlled by the GND REC/ANTI COLL ON–OFF switch on the tilt panel.
Wing Inspection
Wing inspection lights are mounted in both sides of the fuselage forward of
the wing leading edges. When activated, they illuminate the entire leading
edges of both wings. The lights are controlled by the WING INSP ON–OFF
switch on the tilt panel within the ANTI ICE/DEICE switch section.
Landing/Recognition/Taxi Lights
Landing and recognition lights are mounted side by side on each forward
wingtip light assembly. The landing lights are installed outboard and are
canted downward slightly. The inboard recognition lights illuminate directly
ahead.
Two fixed position seal beam taxi lights are mounted in the belly fairings on
each side of the fuselage. The lights also supplement the landing lights.
Control
Landing, recognition and taxi lights are all controlled by individual ON–OFF
switches on the center pedestal. The following switch positions function as
follows:
• REC/TAXI–ON (down)—“BELLY” TAXI LIGHTS and RECOGNI-

TION lights ON.
• OFF—All LANDING, TAXI and RECOGNITION lights OFF.
• LANDING LIGHTS— ON—All LANDING, TAXI and RECOGNI-

TION lights ON.

TAIL CONE COMPARTMENT LIGHTS

Tail Cone Maintenance Compartment Light
The tail cone maintenance compartment light is powered from the battery bus
and controlled by an ON–OFF switch in the right forward side of the com-
partment (adjacent to the electrical J-box). If the light is inadvertently left
on, closing the access door extinguishes the light.
Baggage Compartment Light

The baggage compartment is illuminated by three lights, two in the overhead
ceiling and one in the sidewall. The lights are powered from the battery bus and
controlled by a switch in the forward access door closeout assembly. If the light
is inadvertently left on, closing the door extinguishes the light.
PULSELITE SYSTEM (OPTIONAL)

The Precise Flight, Inc. Automatic Pulselite System allows the taxi (belly) and
recognition lights to pulse. The taxi and recognition lights automatically pulse
when both REC/TAXI switches are positioned ON (down) and the airplane is
airborne. Positioning either or both switches to LANDING LIGHTS–ON de-
activates the system and all LANDING, RECOGNITION and TAXI lights re-
vert to steady illumination. An optional pulselight switch may be installed next
to the LANDING/RECOGNITION/TAXI light switch which overrides the
squat switch to allow pulsing of the TAXI and RECOGNITION lights on the
ground. The switch must be ON in addition to having both REC/TAXI LIGHTS
selected ON for the lights to pulse, airborne or on the ground. Refer to SUP-
PLEMENT 2, PRECISE FLIGHT–AUTOMATIC PULSELITE SYSTEM in the
Airplane Flight Manual (AFM) for detailed operating procedures.
ELECTRICAL SYSTEM
Electrical system schematics are shown after the ELECTRICAL SYSTEM text
(Figures SRE-2, SRE-3 and SRE-4).
1. Battery switch (power distribution switch), controls the battery isolation

relay):
• 0FF—Battery isolation relay is deenergized open and the relay between
the crossfeed and emergency buses remain deenergized closed. The
relay between the battery bus and the emergency buses is deenergized
open. The voltmeter is not connected to the battery bus, so normal
system voltage cannot be read.
• EMER—Battery isolation relay is deenergized open and the relay

between the crossfeed and emergency buses is energized open. The
relay between the battery bus and the emergency buses is energized
closed. The battery supplies power to the battery bus and emergency
buses. If the generators are on line, the generators supply power to the
feed and crossfeed buses. This switch position isolates the battery from
generator charging power. The voltmeter indicates only battery voltage,
unless the VOLTAGE SEL switch is positioned to L–GEN or R–GEN.

NORMAL OPERATIONS (GROUND OR INFLIGHT, APU GENERATOR OFF LINE, AVIONICS ON)
LEFT CIRCUIT BREAKER PANEL RIGHT CIRCUIT BREAKER PANEL
ENGINE START
L DISENGAGE R
EMER SYS AFT GEN EMER AVN
J-BOX OFF
START
SYS SYS LMT CB DISG L R AVN AVN
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER
L FEED BUS CROSSFEED BUS R FEED BUS RELAY
EMER PWR RELAY AVN EMER RELAY 25A

A A
L GEN BATTERY EMER AVN
NOTE: SWITCH SWITCH E
BATT
RED BORDERS
A L GEN ISOLATION M R GEN
ON 25A
DENOTE EMERGENCY P ON RELAY RELAY E RELAY ON I
U OFF R N
POWERED BUSES OFF OFF T
28.5 EMER 28.5 E
RESET 28.5 RESET R
B
LEGEND U
START START I
L GEN RELAY RELAY R GEN O
L GENERATOR S BUS BATTERY BUS BUS R
GCU GCU
APU 1
R GENERATOR 7
RELAY APU RELAY
L BATTERY
R 5
ENGAGED A
APU GENERATOR STARTER STARTER
GPU
FIELD GEN BATT RELAY OVER (32.5 VDC) GEN FIELD
BATTERY RELAY DISCONNECT VOLTAGE RELAY
RELAY
GPU INTERIOR POWER
INPUT
SRE-9
Figure SRE-2. DC Power Distribution

SRE-10

LEGEND
LH MAIN DC BUS
RH MAIN DC BUS
EMERGENCY BUS
Figure SRE-3. Pilot Circuit-Breaker Panel

LEGEND
LH MAIN DC BUS
RH MAIN DC BUS
EMERGENCY BUS
SRE-11
Figure SRE-4. Copilot Circuit-Breaker Panel

• BATT—Battery isolation relay is energized closed and the relay

between the crossfeed bus and emergency buses is deenergized closed.
The relay between the battery bus and the emergency buses is
deenergized open. All power is normally supplied by the generators or,
if off line, by the battery. The voltmeter indicates system voltage.
2. Generator switches (left and right)
• ON—Activates the GCU, which may close the power relay.
• OFF—Disables the GCU, opens the power relay, not the field relay.
• RESET—Momentarily resets the field relay.
3. Generator control units (digital GCUs with BITE lights):
• The GCUs regulate 30-VDC generators to 28.5 VDC.
• The GCUs provide protection for the generators and the electrical
system.
• The GCUs parallel the generators to share the system load; the
generators must be within 0.3 volts and approximately 10% of system
load, not to exceed a 30-amp split.
• Each GCU mounted in the tail cone has four red BITE lights (fault
lights). The GCU fault lights may indicate a GCU fault, overvoltage, a
ground fault, or a system problem. The LEDs self-test at power-up.
Flashing LEDs can be extinguished by resetting the generator switch
three times within three seconds if no fault exists.
4. Voltmeter select switch:
• BATT—Voltage is read from the hot battery bus when the battery
switch is in the BATT or EMER position only; the switch is spring
loaded to the BATT position.
• LH or RH GEN—Voltage can be read from the respective generator

output to its power relay. Accurate generator voltage is normally
obtained with the selected generator off-line.
5. Emergency bus items
LH CB panel:
• Cockpit flood and glareshield lights (AUX panel)
• L and R fan speed (N1)
• Standby HSI
• Gear warning (lights and horn)
• Standby P/S heater

• L and R ignition (SEC)
• Flap control
• Stabilizer control
• Gear control
• Hydraulic control valve
RH CB panel:
• Audio 1 and 2
• COMM 1
• NAV 1
• AHRS 2
• Standby radio control (COMM 1 CB)
• RMU 1 (COMM 1 CB)
Placing the battery switch to either ON or OFF causes the emergency relays
to relax, connecting the emergency buses to the crossfeed bus. Placing the
battery switch to EMER, energizes the emergency relay connecting the emer-
gency buses to the battery bus. With the generators on line, placing the bat-
tery switch to OFF does not cause loss of power to the emergency buses
(isolates battery from generators). Loss of both generators require that the
battery switch be placed in the EMER position, which load sheds the main
DC feed buses and allows the battery to power only the battery bus and emer-
gency buses. This extends the battery life to approximately 30 minutes.
6. Battery bus items:
• Tail cone baggage compartment light
• Tail cone access compartment lights
• Cabin entrance lights
• Voltmeter (BATT switch in ON or EMER only)
• Emergency buses (BATT switch in EMER only)
• APU preflight panel (tail cone)
• LAV service lights
• Engine chip detector lights (if installed)
7. Crossfeed bus items:

• Left and right ignition (NORM and ON)
• Left and right taxi lights (Belly)
• Cockpit WEMAC fan
• Forward and aft air conditioner evaporator fans
• Optional APU
8. LH or RH GEN OFF annunciator:
• Indicates the respective power relay is open.
• If voltage indicates near zero, the field relay is tripped; reset is

possible.
• If voltage indicates normal, the power relay is open and the field relay
is not tripped; reset is not probable.
9. Current limiters (225 amp):
• Should open due to system malfunction only (AFT J-BOX LMT).
• If failed, prevents generators from being paralleled.
• If failed prior to start, the engine on the side of the failed current limiter
cannot be started. The MASTER WARNING on the opposite side
illuminates steady.
10. Standby flight display—Normally powered through the STBY PWR CB

on the pilot CB panel. If power is lost at the circuit breaker, provided the
STBY PWR switch is ON, then the secondary flight display battery pack
(nose compartment) automatically provides power to the following (amber
light adjacent to switch on):
• Tube
• Back lighting for the standby HSI, N1 gauges
11. Auxiliary avionics battery pack (nose compartment):
• Emergency power supply for the AHRS 1 and 2
• Emergency power supply for the flight management system (on ground
only, through FMS GND PWR CB on RH CB panel)
• For any of the above to occur, the STBY PWR switch must be ON
12. The emergency lighting system consists of a EMER LTS control switch,
two 18-cell rechargeable nicad battery packs, and various cabin interior
and exterior lights and signs. When activated, the emergency lighting
system will use main DC or battery power for system operation. If main

DC or battery power is unavailable or otherwise becomes depleted, the

system operates from its two battery packs for a minimum of 10 minutes.
The battery packs are constantly trickle-charged to capacity so they can be
available as a final source of emergency lighting power. Cabin emergency
lighting activates for the following reasons:
• ARM (automatic activation due to):
• Loss of main DC power with battery switch in EMER (uses

battery pack power)
• 5 g impact (uses main DC, battery, or battery pack power)
• Passenger safety switch selection (requires main DC or battery

power only)
• ON (manual activation) (uses main DC, battery, or battery pack power)
• OFF (system deactivation)
NOTE
The amber light to the left of the EMER LTS switch
illuminates when the airplane battery switch is placed
ON and the EMER LTS switch is OFF. The amber light
can be extinguished by placing the EMER LTS switch
to ARM or ON.
13. AC alternators:
• Two engine-driven 3.0 KVA alternator, one mounted on each engine.
• Provides 115/200 volt, three-phase, 200 to 400 Hz power to electrically

heat the windshields.
• Controlled by WINDSHIELD ANTI-ICE switches on the cockpit tilt

panel (main DC).

POWERPLANT
Pratt and Whitney PW545A (Figure SRE-5)
3,804 pounds thrust
Bypass ratio 4.0:1
Electronic engine control (EEC) in AUTO mode provides:
• Detented throttle positions for automatic takeoff, climb, and cruise

thrust settings (N 1 governing)
• Automatic thrust setting compensation for anti-ice ON
• Automatic engine idle governing for flight idle (57–62% N 2 ) and

ground idle (48–51% N 2 )
NOTE
Ground idle rpm is achieved 8 seconds after landing
(WOW).
• Acceleration and deceleration limiting
• N 1 and N 2 speed limiting
• N 2 underspeed limiting (at climb detent and above)
• Overspeed protection (N 2 )
• N 1 or N 2 synchronization
• Engine diagnostic system (EDS) functions
• Closed loop bleed valve (BOV) control
• Independent overspeed protection (N 2 ) (~3% above redline)
NOTE
In AUTO mode, it is still the responsibility of the pilot
to monitor N 1 and N 2 limits to prevent an overspeed
condition and ITT limits to prevent an overtemper-
ature condition.

LEGEND
INDUCTION AIR
EXHAUST AIR
COMBUSTION
CHAMBER
CENTRIFICAL
COMPRESSION AIR
AXIAL
COMPRESSOR AIR
TURBINE AIR
Figure SRE-5. PW545A Cross-Section

Electronic engine control (EEC) in manual mode:
The fuel control unit (FCU) takes over full control of the engine speed in re-
sponse to the throttle position. In MANUAL mode, the throttle directly con-
trols the FCU by means of a mechanical linkage. MANUAL mode provides
the following functions:
• Pilot adjustable power setting (N 2 governing above minimum fuel

flow N 2 )
• Idle governing (N 2 governing) at flight idle and anti-ice idle
• Acceleration and deceleration limiting (ratio unit control)
• N 2 speed limiting (~3% above redline)
• Limited engine diagnostic system functions (EDS)
Bleed-off valve control (BOV):
• Bleeds off excessive P 2.8 compressed air, preventing compressor

surges and stalls.
• Electrically controlled operation by the EEC.
• BOV protection fails to a pneumatic backup mode with the loss of main
DC power.
IGNITION
A single, dual-channel exciter box with two ignitor plugs per engine. Burst mode
type ignition that produces 6–7 sparks per second for the first 30 seconds, then
one spark per second, thereafter. Green ignition light verifies that DC power
is available to the exciter box. If one ignitor plug fails during engine starts, the
engine starts normally and the ignition light remains illuminated.
Ignition Switch:
• NORM—Auto ignition for start, and for engine or wing/engine anti-
ice on (powered by the crossfeed bus).
• ON—Pilot select ignition—used for takeoffs, landings, turbulence,

stalls, precipitation and emergency descents. Powered by the cross-
feed bus with main DC power.
• SEC—Placing the ignition switch to the secondary (SEC) position pow-

ers the ignition system from the emergency bus.

Oil
Maximum consumption is 0.2 pph, measured over a 10-hour period or one
quart in 10 hours. Check oil level 10 minutes after shutdown.
Do not mix nonapproved oil brands or oils of different viscosities.
The switch to “third generation” oils should not be accomplished unless an

engine is new or freshly overhauled.
Oil pressure fluctuations are normal. Oil pressure indicator measures differ-
ential oil pressure.
Fuel
Engine-driven fuel pump—A two-stage pump in the fuel control unit.
Hydromechanical fuel control unit (FCU)—Normally controlled by EEC.
Flow divider valve—Regulates fuel flow to the primary and secondary fuel
manifolds. Secondary manifold activates at 26–28% N 2 .
Emergency shutoff valve—Shuts down engine to protect airframe.
FIRE PROTECTION
The engine fire protection system is composed of sensing loops, two control
units (one for each engine) in the tail cone, one ENG FIRE warning switch-
light for each engine, one FIRE DET SYS L–R annunciator for each engine,
two fire extinguisher bottles which are activated from the cockpit, a FIRE EXT
BOTL LOW annunciator, and a fire detection circuit test (Figure SRE-6).
Detection and extinguishing system electrical power is supplied from normal
DC power.
SENSING LOOPS AND CONTROL UNITS

Each engine nacelle contains heat-sensing cables, or loops. Each loop is
mounted around the lower engine accessory section and surrounds the engine
combustion section, outside the bypass duct. The loops are connected to con-
trol units that monitor the electrical resistance. As the loop is heated, its
electrical resistance decreases until, at a temperature of approximately 500°F,
a circuit is completed to the control unit to illuminate the applicable red ENG
FIRE switchlight. The fire detection loops are continuously monitored. Should
one of the loops fail, the associated FIRE DET SYS L–R annunciator would
illuminate to indicate that fire detection is no longer available on that side.

LH RH
ENGINE ENGINE
FIRE FIRE
BOTTLE 1 BOTTLE 2
ARMED ARMED
PUSH PUSH
FIRE BOTTLE 1
FIRE BOTTLE 2
FIRE LOOP FIRE LOOP
RUDDER FIRE
BIAS DET SYS
FIRE EXT
BOTL LOW L R
LEGEND
FIRE BOTTLE #1 DISCHARGE
Figure SRE-6. Engine Fire Extinguishing System
OPERATION
An engine fire light or overheat condition is indicated by illumination of the ap-
plicable ENG FIRE switchlight on the glareshield (Figure SRE-7). Depressing
the illuminated ENG FIRE switchlight causes both white BOTTLE ARMED
switchlights to illuminate, arming the circuits to the bottles for operation. In ad-
dition, the generator field relay opens (GEN OFF annunciator illuminates) and
provides a ground to power the fuel and hydraulic firewall shutoff valves closed
(causing the respective LO FUEL PRESS, LO HYD FLOW, F/W SHUTOFF an-
nunciators to illuminate). The circuit to the thrust reverser isolation valve is dis-
abled, preventing deployment of the thrust reverser on that engine.
Depressing either illuminated BOTTLE ARMED switchlight fires the explo-
sive cartridge on the selected bottle, releasing its contents into the engine na-
celle. The BOTTLE ARMED switchlight extinguishes.
Depressing the ENG FIRE switchlight a second time opens the fuel and hy-
draulic firewall shutoff valves, and disarms the extinguishing system.
Due to the location of the fire bottles, the bottle pressures cannot be checked
on preflight. If either or both fire extinguisher bottle pressure is low, the amber
FIRE EXT BOTL LOW annunciator illuminates to alert the crew.

LH RH
ENGINE ENGINE
FIRE FIRE
FIRE
DET SYS
L R
FIRE LOOP FIRE LOOP
Figure SRE-7. Engine Fire Detection System
FUEL
Refer to Figure SRE-8 for a schematic indicating normal operation of the fuel
system.
Total capacity is 1,006 U.S. gallons (approximately 6,790 pounds).
Boost pump switches:
• ON—DC boost pumps are continuously energized.
• OFF—DC boost pumps are deenergized and does not operate.
• NORM—DC boost pumps are energized automatically for start, cross-

feed, and low fuel pressure.
Tank-to-tank transfer rate during crossfeed is approximately 700–900 pph (en-

gines shutdown).
Low fuel pressure light illuminates at a decreasing pressure of 5 psi and ex-
tinguishes at an increasing pressure of 7 psi.

SRE-22
PRE-CHECK
DE-FUEL SELECT LEVERS
LEVERS
SPPR ADAPTOR
(360 +/- 20 LBS.)
(163 KG)
SCAVENGE
LO FUEL EJECTOR
LEVEL
SURGE TANK
L R

FUEL
x x BOOST
FUEL TRANSFER TUBES L R

R FUEL
HOPPER
PRIMARY EJECTOR
LEGEND
(300-1200 PSIG)
NEG / POS PRESSURE RELIEF VALVE LO FUEL
MOTIVE PRESS
FUEL TANK VENT FLOW
L R
MOTIVE FLOW SHUTOFF
SPR FUEL SHUTOFF VALVE SOLENOID VALVE (NO)
P P
HIGH LEVEL PILOT VALVE FOHE
FUEL
FLTR BP FUEL FILTER
LOW LEVEL PILOT VALVE VENT F/W
L R
ENG DRIVEN SHUTOFF
FUEL BOOST PUMP FUEL PUMP
FUEL INSIDE THE TANK
FCU L R
CHECK VALVE REFUEL PRESSURE FUEL FLOW
FCU FLOW DIVIDE CROSS FEED VALVE
EJECTOR PUMP
MOTIVE FLOW PRESSURE EMERGENCY FUELSHUTOFF VALVE
(MOTORIZED) APU
(MECHANICAL)
F/W SHUTOFF VALVE (MOTORIZED) APU FUEL
JET PUMP PRESSURE SHUTOFF
P LOW FUEL PRESSURE SWITCH (5.3 PSIG)
FUEL VALVE
BOOST PUMP PRESSURE
x LOW LEVEL FLOAT SWITCH (360+/-20 LBS - 163 KG) XFEED
TEMPERATURE SENSOR (-60oC - 70oC) SCAVENGE PUMP PRESSURE
Figure SRE-8. Fuel System—Normal Operation

Low fuel level light illuminates at approximately 360 plus or minus 20 pounds
of usable fuel remaining in the respective tank; input is from a float switch.
Illumination of the FUEL GAUGE annunciator indicates a fault has been de-
tected in the respective fuel gauging system. Do not shut down DC power (BATT
switch to OFF), after engine shut down, until checking and recording the fuel
conditioner BITE lights.
The fuel filter is on the engine, downstream of the fuel-oil heat exchanger
(FOHE), eliminating the need for fuel anti-ice additives. It is still recommended
to use Prist or other approved fuel additives on a regular basis for the anti-
fungal properties of the additive.
NOTE
Av-gas is not an approved fuel
Crossfeed operation—Select the tank desired to supply fuel to both engines

and transfer to the other tank. Selecting crossfeed initiates the following se-
quence of events (Figure SRE-9):
1. DC boost pump on the side supplying fuel activates (respective FUEL

BOOST illuminates steady).
2. Crossfeed valve opens. (FUEL XFEED advisory light illuminates when

the valve is fully open.)
3. Motive flow valve on the receive side is energized closed three seconds
after crossfeed is selected. (Transfer rate depends on operating engine(s)
requirements.)
Selecting the crossfeed switch to OFF, reverses the above process. Should the
crossfeed valve fail to close, the FUEL XFEED advisory light illuminates flash-
ing and activates the MASTER CAUTION lights steady.
It is important to allow the crossfeed system to complete the above process

before reselecting a different crossfeed switch position.
If the opposing boost pump activates (on the receiving side), it would indi-
cate a timing problem with the crossfeed valve. To rectify the problem, reset
the opposing boost pump (turn the opposing side FUEL BOOST switch to ON
then back to NORM). Do not turn a FUEL BOOST switch OFF and leave it
there; OFF is OFF.

FUEL BOOST LEGEND

ON FUEL HOPPER
O
F
F OPERATING BOOST
NORM PUMP
CROSSFEED CROOSFEED
VALVE OPEN
L R MOTIVE FLOW
TANK OFF TANK SHUTOFF VALVE
FUEL INSIDE THE TANK

MOTIVE FLOW PRESSURE
L R
ENG ENG
JET PUMP PRESSURE
BOOST PUMP PRESSURE
SCAVENGE PUMP PRESSURE
FUEL
BOOST
FUEL L R
XFEED
Figure SRE-9. Fuel System—Crossfeed (R to L)

HYDRAULICS
1. Reservoir Quantities:
• Overfull.................................................................................. 360 cu. in.
• Full ....................................................................................... 215 cu. in.
• Refill ...................................................................................... 175 cu. in.
• LO HYD LEVEL annunciator................................................. 74 cu. in.
2. Engine driven pump flow rate is 3.25 gpm/12.29 liters-per-minute

maximum.
3. Open center system (Figure SRE-10):

• System control valve open (normal) ............................................. 60 psi
• System control valve closed (system operational) ................... 1,500 psi
4. Speedbrakes (Figure SRE-11):
• Held closed mechanically; held open by trapped hydraulic fluid.
• Retract normally with switch in RETRACT; Automatically retracts if
either throttle is advanced beyond 80 to 85% N2 PLA (power lever
angle).
• Requires normal DC power to remain extended. With loss of normal
DC power, the safety valve relaxes open to allow trapped fluid to
escape and the speedbrakes blow to a trail position.
• System pressure is required for extension and normal retraction.
• The NO TAKEOFF annunciator illuminates if the speedbrakes are
extended on the ground.
5. Landing gear (Figures SRE-12, SRE-13 and SRE-14):

• DC power is required for normal hydraulic extension and retraction of
the landing gear. Control provided through the GEAR CONTROL
circuit breaker on the left CB panel and powered by the emergency DC
bus system.
• Gear is held in the up position by uplocks; held in the down position by

downlocks. Operational hydraulic pressure is used to raise and lower
the gear.
• Freefall/blowdown backup system. Manual retraction of the uplocks

allows the gear to freefall, then pneumatics provide positive
downlock; pneumatic system retracts the uplocks in the event the
manual release does not retract the uplocks, also providing positive
downlock of the gear.
• Aural warning, unsafe gear down: 1) both throttles below 70% N2,
flaps beyond 15°; 2) both throttles below 70% N2, radio altitude less
than 500 feet; 3) both throttles below 70% N 2 , radio altimeter
inoperative, airspeed less than 150 KIAS.

SRE-26
SUBSYSTEM CONTROL VALVES
MAXIMUM SYSTEM OPERATION—1,500 PSI MAX

F F (.55 GPM)

LANDING
OPEN CENTER OPERATION—60 PSI

GEAR LO HYD

FLOW
SPEEDBRAKES
L R

RETURN LINES
WING FLAPS
HORIZONTAL HYD CONTROL

STABILIZER VALVE (N/O)
PRESSURE
THRUST SWITCH (LOADING VALVE)
REVERSERS
P
RELIEF VALVE
OPENS AT 1,350 PSI
FILTER
F/W SHUTOFF
MOTORIZED
VALVE
EXEL R ENGINE
F/W
PUMP
LEGEND SHUTOFF
(74 CU)
SUPPLY SUCTION
LO HYD L R
RETURN PRESSURE LEVEL
HYD
#1 SYS HIGH PRESSURE (MAIN)
PRESS HYDRAULIC RESERVOIR (TAIL CONE)
Figure SRE-10. Hydraulic System—Open Center

FLAPS
UP 0°
TRIM TO
CLB T.O.
NOSE 200 KIAS 7°
DOWN
CRU
STAB
T T
O H T.O. &
R APPR 15°
O 200 KIAS
T
NOSE
UP
T
L
E
IDLE
MIS COMP
LAND 35°
SPEED
BRAKE CUT
175 KIAS
SPD BRK

OFF
ENGINE SYNC
EXTEND
LH RH MUST BE OFF
FAN OFF TURB FOR TAKEOFF
RETRACT & LANDING
EXTEND
SPEED BRAKE
SWITCH
SPEED BRAKE SPEED BRAKE

ACTUATOR ACTUATOR
LO HYD BYPASS
LEVEL
SAFETY
HYD
PRESS
CONTROL VALVE
VALVE
1500 PSI PRESSURE SYSTEM LOADING VALVE

RELIEF VALVE
CHECK CHECK
VALVE VALVE
LEGEND HYDRAULIC HYDRAULIC
PUMP PUMP
SUPPLY SUCTION
RETURN
LOW FULL OVER FULL
SUCTION
HYDRAULIC RESERVOIR
RETURN PRESSURE
LO HYD
#1 SYS HIGH PRESSURE (MAIN) LEVEL
HYD
PRESS
SRE-27
Figure SRE-11. Speedbrake System—Normal Operation (Extended)

SRE-28
PRESSURE FROM HYDRAULIC PUMP TO HYDRAULIC RESERVOIR

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE
SHUTTLE VALVE

LANDING GEAR LANDING GEAR

ACTUATOR UPLOCK UPLOCK ACTUATOR
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
L R UPLOCK LANDING GEAR
H H ACTUATOR
LEGEND
UP ANTI-
RETRACT PRESSURE
SKID
LANDING ON
GEAR RETURN PRESSURE
DOWN NITROGEN
BLOW DOWN EMERGENCY NITROGEN
OFF BOTTLE
Figure SRE-12. Gear System—Normal Retraction

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE

SHUTTLE VALVE
ACTUATOR ACTUATOR
UPLOCK UPLOCK
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
H H LEGEND
ACTUATOR
EXTEND PRESSURE
UP ANTI-
SKID
LANDING ON RETURN PRESSURE
GEAR
DOWN NITROGEN EMERGENCY NITROGEN

BLOW DOWN
OFF BOTTLE
SRE-29
Figure SRE-13. Gear System—Normal Extension

SRE-30
FROM HYDRAULIC PUMP TO HYDRAULIC RESERVOIR

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE

SHUTTLE VALVE
ACTUATOR ACTUATOR
UPLOCK UPLOCK
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
L R UPLOCK LANDING GEAR LEGEND
H H ACTUATOR
#1 SYS HIGH PRESSURE (MAIN)
UP ANTI- RETURN PRESSURE

SKID
LANDING ON
GEAR
EMERGENCY NITROGEN
DOWN NITROGEN
BLOW DOWN
OFF BOTTLE
Figure SRE-14. Gear System—Emergency Extension

6. Thrust reversers (Figures SRE-15 and SRE-16):

• Normal DC power is required.
• Only one squat switch is required (left, right or both) to allow the
control valve to energize to the deploy position when commanded.
• Each emergency stow switch is powered by the opposite side thrust

reverser circuit breaker.
• Illumination of either the ARM or UNLOCK light while in flight

triggers the MASTER WARNING flasher.
• If the thrust reversers do not operationally check (including emergency

stow), flight must not be attempted.
• The thrust reversers should be in idle power at and below 60 knots.
• The use of thrust reversers to back the aircraft is prohibited.
7. Flaps:
• DC power is required for flap actuation. Flap control is through the flap
handle on the throttle pedestal. DC power is supplied by the emergency
bus, through the FLAP CONTROL circuit breaker on the left CB
panel.
• Detented flap positions are provided at the 7° and 15° positions. The
flaps can be selected at any position between zero and 35°. Flap
position is shown via a pointer to the left of the flap lever.
• In the event of electrical failure, the flap solenoid valve remains in the
neutral position, and flap position cannot be changed.
• Mechanical interconnect prevents asymmetrical flap condition.
• Flaps may be selected from full up (0°) to full down (35°).
• Detented flap positions: 7° (T.O.) and 15° (T.O. and APPR).
• If hydraulic system failure occurs with the flaps retracted, they cannot be
extended. If the flaps are extended and hydraulic system failure occurs,
they remain in the selected position, unless the flap handle is moved.
Movement of the flap handle energizes the solenoid valve, and the flaps
blow to a trail position as determined by air loads present.
• The NO TAKEOFF annunciator illuminates when the flaps are set less
than 7° or greater than 15° on the ground.

SRE-32
ISOLATION VALVES

THRUST REVERSER
THRUST REVERSER
STOW
CONTROL VALVE CONTROL VALVE STOW
EMER ARM
ARM EMER (SQUAT SWITCH & (SQUAT SWITCH &
UNLOCK
THROTTLE LEVERS) THROTTLE LEVERS) UNLOCK
NORM DEPLOY
DEPLOY NORM
LO HYD FLOW VALVE PRESSURE FLOW VALVE LO HYD
FLOW (LO HYD FLOW) SWITCH (LO HYD FLOW) FLOW
(ARM LIGHT)
THRUST REVERSER L R L R
LEVERS
PRESSURE
LO HYD
LEVEL SWITCH HYD CONTROL
FLAPS
UP
T.O.
0°
7°
HYD VALVE (LOADING
PITCH
TRIM
T.
T
H
R
O
200 KIAS
T.O. &
APPR 15°
PRESS VALVE)
T 200 KIAS
O.
T
L LAND
NOSE E 35°
UP 173 KIAS
OFF
N1
ENGINE SYNC
OFF
N2
MUST BE OFF
FOR TAKEOFF
& LANDING
PRESSURE
RELIEF VALVE
OPENS @
1350 PSI
HYDRAULIC
LEGEND LOW LEVEL SWITCH HYDRAULIC PUMP
HYDRAULIC (LO HYD LEVEL)
STATIC FLOW PUMP RESERVOIR
LO HYD
#1 SYS LOW LEVEL
PRESSURE (MAIN)
HYD
SUPPLY SUCTION PRESS
Figure SRE-15. Thrust Reversers—Stowed

ISOLATION VALVES

THRUST REVERSER
THRUST REVERSER
STOW
EMER ARM
ARM EMER (SQUAT SWITCH & (SQUAT SWITCH &
UNLOCK
NORM DEPLOY
DEPLOY NORM
(ARM LIGHT)
LEVERS
PRESSURE
LO HYD
FLAPS
UP
T.O.
0°
7°
HYD VALVE (LOADING
PITCH
TRIM
T.
T
H
R
O
200 KIAS
T.O. &
APPR 15°
PRESS VALVE)
T 200 KIAS
O.
T
L LAND
NOSE E 35°
UP 173 KIAS
OFF
N1
ENGINE SYNC
OFF
N2
MUST BE OFF
FOR TAKEOFF
& LANDING
PRESSURE
RELIEF VALVE
OPENS @
1350 PSI
HYDRAULIC
LOW LEVEL SWITCH PUMP
(LO HYD LEVEL) RESERVOIR
#1 SYS HIGH PUMP
PRESSURE (MAIN) LO HYD
LEVEL
SUPPLY SUCTION
HYD
SRE-33
RETURN PRESSURE PRESS
Figure SRE-16. Thrust Reversers—Deployed

8. Horizontal stabilizer:
• The two-position horizontal stabilizer is hydraulically actuated and

electronically controlled.
• Flap handle position and airspeed automatically determines the

position of the stabilizer.
• See “Flight Controls” later in this chapter for description.
POWER BRAKES AND ANTISKID

• Separate from the main aircraft hydraulic system, with the reservoir
in the nose (Figure SRE-17).
• Normal DC power required to operate the accumulator pump.
• Antiskid protection is available only when power brakes are operational.
CAUTION
Do not pull the PWRBRKS circuit breaker to prevent
the power brake pump from cycling. With the circuit
breaker disengaged, the power brake system is inop-
erative and the toe pedals are disabled. Braking is then
available only by use of the emergency brake system.
• Pneumatic brakes are a backup for the power brakes; no differential brak-
ing and no antiskid protection available with pneumatic braking.
• Antiskid protection drops out at 10 knots and below.
• Touchdown protection prevents power brake and antiskid operation until:
1. Either or both squat switches in the ground-on-ground (GOG)

mode.
2. Wheel speed greater than 40 knots, locked wheel protection pro-

vides a pressure dump command to the slow wheel when velocity
is 50% less than the fast wheel.
• Digital antiskid provides continuous monitoring of the system with

the gear up or down.
• If a fault is detected, the ANTISKD INOP annunciator illuminates.

Analog antiskid BITE (fault) indicators are in the left nose avionics
compartment.

TEST
BRAKE SYSTEMHYDRAULIC SPARE OFF FIRE
LO BRK RESERVOIR (SKYDROL/HYJET) WRN
N
AVN LDG
PRESS GEAR
L S R ANNU BATT
H H ANTI-SKID ANTI TEMP
INOP SKID STICK
UP ANTI-SKID SHAKER
ON OVER
SPEED T / REV

W/S TEMP
DOWN OFF ACCUMULATOR 28 VDC

HYDRAULIC
PUMP
PRESSURE
BIT FAULT INDICATORS SWITCHS
P P 900 PSI LO BRK
POWER BRAKE PRESS
--SQUAT DISAGREE
--CNTL UNIT
--VALVE
--R XDCR
--L XDCR
VALVE 1230-1500 PSI ANTI-SKID

INOP
ANTI-SKID
Servo Valve
Left & Right Digital Anti-Skid VENT
Squat Switch Control Unit PARKING
EMERGENCY
BRAKE
BRAKE HANDLE
CHECK VALES
NITROGEN
<
<
L/R WHEEL TRANSDUCER INPUTS BLOW DOWN

LEGEND SHUTTLE VALVES BOTTLE XL
<
<
SUPPLY
RETURN PRESSURE
SYS HIGH PRESSURE (MAIN)

EMERGENCY NITROGEN/
SRE-35
CONTROL PRESSURE (MSTR CYL)
Figure SRE-17. Power Brake/Antiskid System

EMERGENCY BRAKES
• A pneumatic brake system is available in the event the hydraulic brake
system fails (see Figure SRE-17).
• Uses air pressure from the pneumatic bottle. Bottle pressure is ade-
quate for stopping the aircraft, even if the landing gear has been pneu-
matically extended.
• Pulling the red EMER BRAKE PULL lever mechanically actuates the
emergency brake valve. Air pressure to the brakes is metered in di-
rect proportion to the amount of lever movement.
• Differential braking is not possible, since air pressure is applied to

both brakes simultaneously. Releasing the handle vents pneumatic pres-
sure from the brakes.
• Do not depress the brake pedals while applying emergency air brakes.
FLIGHT CONTROLS
All primary flight controls (ailerons, elevators, and rudder) are manually ac-
tuated with cables and pulleys and are dual interconnected (Figure SRE-18).
Secondary flight controls consist of trim tabs, speedbrakes, flaps, and a two-
position horizontal stabilizer (Figure SRE-18).
1. Ailerons:
• Maximum aileron travel is 19° up and 15° down.
• Trim tab on the left aileron only has a maximum travel is 20° up and
down.
• Trim is mechanically controlled with a wheel at the rear of the center

pedestal.
2. Elevators:
• Maximum elevator travel is 19° up and 15° down.
• Left and right trim tab travel is 5° up and 15° down.
• Trim is electrical (pitch) or mechanical.
• Left electrical pitch trim overrides right.
• Electrical pitch trim must be tested before flight.
• Electrical trim can be interrupted with the red AP/TRIM DISC button
on either yoke.

RUDDER
ELEVATOR TRIM TAB
RUDDER TRIM TAB
ELEVATOR TRIM TAB

AILERON ELEVATOR
FLAPS
SPEED BRAKES
AILERON TRIM TAB
SRE-37
Figure SRE-18. Flight Controls

3. Rudder:
• Maximum rudder travel is 28.5° either side of centerline.
• Trim tab (servo tab) travel is 14° either side of centerline when rudder
is centered.
• Full rudder pedal deflection on ground deflects the nosewheel 20°

either side of centerline.
• Flight must not be attempted if the nosewheel steering is inoperative.
4. Rudder bias:
• The rudder bias system is comprised of a bleed-air shutoff valve and a

dual actuating pneumatic cylinder that drives a closed loop cable
connected to the rudder (Figures SRE-19 and SRE-20).
• With engines operating, bleed air is continuously available to the

pneumatic cylinder through the shutoff valve any time main DC or
battery power is available and the rudder bias circuit breaker is in. With
approximately equal thrust available from the engines, rudder bias is
balanced and does not affect rudder position. During periods of
unequal thrust (i.e., engine failure), the rudder bias system
automatically deflects toward the operating engine. This assists the
pilot to compensate for adverse yaw.
• With power available, the rudder bias shutoff valve energizes open and
ports engine bleed air to its respective side in the cylinder. If the shutoff
valve does not move to its full open position, the amber RUDDER
BIAS annunciator illuminates indicating the system has malfunctioned.
• The rudder bias circuit breaker must be pulled to deactivate the system.
• Rudder bias is deactivated during thrust reversers deployment or

EMER STOW use. This is accomplished by closing the bias shutoff
valve; however, the amber rudder bias annunciator does not illuminate.
• Rudder bias does not operate in the emergency bus condition.
5. Rudder bias heat:
• A flexible dual electric element heater blanket is wrapped around the

rudder bias cylinder to prevent freezing. Each blanket heats
automatically between 4–16°C with main DC or battery power
available. The blankets each have its own thermostat (RTD—resistive
thermal device) for redundancy.
• Main DC power for the system is supplied from a circuit breaker in the
aft J-box.

RUDDER
BIAS
FIRE EXT
BOTL LOW
HEATER
BLANKET
BIAS
ACTUATOR
SHUTOFF
VALVE
RUDDER
BIAS HTR
BIAS
HEATER
FAIL
LEGEND
BLEED AIR
Figure SRE-19. Rudder Bias System
RUDDER
BIAS
FIRE EXT
BOTL LOW
HEATER
BLANKET
BIAS
VALVE
RUDDER
BIAS HTR
BIAS
HEATER
FAIL
LEGEND
STATIC FLOW
BLEED AIR
Figure SRE-20. Rudder Bias System—Engine Failure

• Upon aircraft power-up, the heating system PCB in the aft J-box
conducts a test of the two blanket thermostats. If either blanket fails the
test, the BIAS HEATER FAIL annunciator on the cockpit panel flashes
until pressed. The annunciator then illuminates steady until
maintenance is performed.
• Upon aircraft power-up and after self-test, the heating system heats the
cylinder, if required, to 16°C. While heating is in progress, the BIAS
HEATER FAIL annunciator illuminates steady. Refer to the “Master
Warning” section for further description.
6. Flaps:
• Flaps are electrically controlled and hydraulically actuated through the

flap handled on the center pedestal. Electrical control power is
provided from the emergency bus through the FLAP CONTROL
circuit breaker on the left CB panel.
• With the loss of electrical power (circuit breaker out), the flaps remain
in the last position. The flaps cannot be moved.
• With loss of hydraulic power, the flaps remain in the last position
unless the flap handle is moved, after which the flaps blow to a “trail”
position dependent upon air-load forces.
• Normally, during flap actuation the HYD PRESS annunciator

illuminates until the flaps reach the commanded position.
• Flap positions ranging from 0–35° can be selected with the flap handle.
Although a wide range of positions can be selected, 0° is used for
cruise only, 7° and 15° are approved for takeoff, and 35° is used for
landing. Flap handle detents and speed placards are available at the flap
handle.
• On ground, if the flaps are not set to 7° or 15° (TO or APPR) position,
the NO TAKEOFF annunciator illuminates.
• Flaps are held extended with trapped hydraulic fluid and held retracted
mechanically.
• Mechanical interconnects between flap segments prevent asymmetrical

flap conditions.
• Flap handle movements to certain positions affect horizontal stabilizer

movement.

INTENTIONALLY LEFT BLANK

7. Horizontal stabilizer (Figure SRE-21):

• The two-position stabilizer is a secondary flight control used to achieve
optimum angle of attack for takeoff and flight configurations.
• The two-position horizontal stabilizer is electrically controlled and
hydraulically actuated.
• The stabilizer is electrically (emergency bus) controlled by the flap
handle and moves in concert with flap movement. Flaps 0° commands
a stabilizer full up (+1°) position used for climb, cruise, and descent.
Flap handle positions from 1° to 35° commands the stabilizer to its full
down (–2°) position used during takeoff, landing, and approach.
Stabilizer positions between full up and full down only occur due to
malfunction.
• Horizontal stabilizer movement in either direction requires 25 seconds.
The HYD PRESS annunciator illuminates during stabilizer movement
until the system achieves its commanded position.
• If the horizontal stabilizer position is not correct according to the flap
handle position command, an amber stabilizer miscompare
annunciation (STAB MISCOMP) illuminates to alert the crew. On
ground, this miscompare causes a NO TAKEOFF annunciation. Refer
to “Stab Miscomp” description in the “Master Warning” section.
• The horizontal stabilizer system incorporates an airspeed sensor, which
receives pitot-static input from the standby pitot-static system.
Stabilizer movement is restricted with airspeeds greater than 215 ± 10
KIAS.
• Horizontal stabilizer movement restriction due to excessive speed is
accomplished by an overspeed signal from the airspeed sensor to the
system printed circuit board (PCB), which restricts proper positioning of
the arming valve and stabilizer control valve.
• The horizontal stabilizer does not move when commanded by the flap
handle for the following reasons:
1. Hydraulic pressure is not available.
2. Malfunction.
3. IAS in excess of 215 ± 10 KIAS.
4. Landing gear is first selected and is in transit (landing gear has
priority over the horizontal stabilizer only through the stabilizer
PCB).
WARNING
Do not retract flaps above 200 KIAS. Associated sta-
bilizer movement can cause a significant nose down
pitch upset if the movement is not prevented.
• The horizontal stabilizer receives power from the emergency bus.

HORIZONTAL STABILIZER
CRUISE POSITION
SPEED ABOVE 200 KIAS
FLAP
CONTROL

VALVE
(EMER BUS)
P HYD PRESS HYD

SWITCH PRESS
HYDRAULIC CONTROL
VALVE (LOADING VALVE)(NO)
STAB
MIS COMP
FLAP HANDLE HYDRAULIC
POSITION PUMP
HYDRAULIC
RESERVOIR STABILIZER
FLAPS POSITION
UP 0°
TRIM TO
CLB T.O.
NOSE 200 KIAS 7°
DOWN
CRU +1
T
O
T
H T.O. & SPEED –2
R
O
APPR
200 KIAS
15°
PCB
T
T
(UP) SENSOR
L
E
NOSE
UP IDLE
215 (+/– 10) KIAS
LAND
SPEED
175 KIAS
35°
(DN) (EMER BUS)
BRAKE CUT
OFF
ENGINE SYNC HORZ STAB
RETRACT
LH RH
FAN
MUST BE OFF
OFF TURB FOR TAKEOFF
& LANDING CONTROL
LEGEND
VALVE STBY PITOT/STATIC RETURN PRESSURE
EXTEND
(EMER BUS) HYDROMECHANICAL INPUT
ARMING VALVE SUPPLY SUCTION
ACTUATOR
#1 SYS LOW PRESSURE (MAIN)
STATIC FLOW
SRE-43
Figure SRE-21. Two-Position Horizontal Stabilizer

8. Control lock:
• Secures the three primary flight controls in the neutral position and
secures the throttles in cut-off position.
• Towing with the control lock engaged is not recommended on ground

because of the on-ground nosewheel-rudder interconnect. Nosewheel
steering can be damaged if the nosewheel is turned too far. The
maximum range available for towing is 60° left or right of center.
9. Stall warning stick shaker:

• System includes a stick shaker on each control column and an angle-
of-attack system. Stick shaker power is provided through the AOA
circuit breaker on the left CB panel. When the stall warning system
senses an approaching stall, both shakers vibrate until the condition is
corrected.
• Stick shaker vibration activates a .79–.88 AOA (8–10% above stall) and
above depending on flap setting.
• The system is required to be tested and operate for dispatch. Refer to

“Master Warning—Rotary Test” for indications.
• Additional stall warning is achieved with a stall strip on each wing root
by producing buffets.
ICE AND RAIN PROTECTION

1. Ice detection:
• Two red ice detection barrel lights mounted on the top of the
instrument panel glareshield reflect a glow to warn the crew if ice
accumulates on the windshields at the extreme inboard area.
• Wing inspection lights on each side of the fuselage illuminate the wing
leading edges.
2. Pitot-static heat—Two minute limit on ground operation (Figure SRE-

22):
• Three pitot tubes (L, R, and STBY)
• Six static ports (L, R, and STBY)
• AOA vane
• Separate annunciators for AOA, L, R, and STBY pitot-static
systems.
• Standby pitot-static heat is powered from the emergency bus.

P/S
HTR
LEFT PITOT SYSTEM L R RIGHT PITOT SYSTEM

410 00
410 00
MADC MADC 300
FD FAIL
S
G
ATT 1
2
20 20
41500
FD FAIL 41500
IAC IAC
ATT 1 10 10
300 S 20 20 2 00
G 215 410 00
2 00
10 10
200 10 10
2 00
215 410 00
00
200 10 10 100
40500
360 DH
729 M 29 92 IN
CRS HDG 1 WPT VOR 1
DATA DATA
100
40500 360 +I0
360 DH 3
729 M 29 92 IN
CRS HDG 1 WPT VOR 1 2

360 +I0
3
CRS 1
2 0
CRS 1 VOR 1
1
0 ADF 1 2
GSPD
VOR 1
1
HDG ------ KTS 3
310
ADF 1 2
GSPD
HDG ------ KTS 3
310
STD
HONEYWELL
STD
HONEYWELL
TAS
PROBE
LH STATIC RH STATIC
PORTS PORTS
LEGEND
LH STATIC
RH STATIC STBY
10
3
2
15
4 5 6
20
PSI 7
9
25
30
P/S HTR
5 1 9 35
STANDBY STATIC
0 10
DIFF 40
1013 MB PRESS 50
M 5M
PITOT &
0
CABIN ALT
X1000 FT
WINDSHIELD 450 200
STATIC 100
ON 400 10 10
STANDBY PITOT L R 390 240 00

O'RIDE 300
10 10
900
ON
200
20 20
29.92
800
IN
STANDBY PITOT
LH PITOT APR ATT
BARO
AIRSPEED
OFF OFF
RH PITOT SENSOR
SRE-45
(HORIZONTAL STABILIZER)
Figure SRE-22. Pitot-Static System

3. Windshield anti-ice (Figure SRE-23):
• The windshield is heated with AC current to provide anti-ice and defog

capability.
• Two 3-phase 115 VAC alternators are on each engine accessory gear
box (N2 rpm) to provide current for heating of the windshields forward
and side panels. The rear side windows are not electrically heated.
• The left and right alternator bearings are monitored for wear with the
white AC BEARING annunciator. When illuminated, the alternator has
approximately 20 hours of operations remaining.
• One 3-gang AC circuit breaker for each alternator is in the baggage

compartment.
• Left and right DC controllers regulate the respective alternators and

control temperature via embedded windshield sensors. The left and
right controller circuit breakers are on the cockpit left CB panel.
• Two left temperature sensors and two right temperature sensors are
used by the controller to regulate windshield temperature at 110°F
(43°C). If the system malfunctions and windshield temperature reaches
140°F (60°C), system overtemperature circuitry deactives AC power to
the entire left or right system. This condition causes the amber W/S
O’HEAT and the W/S FAULT annunciators to flash.
After temperature cool down of the affected side to 115°F (46°C), the
system can automatically reset and again apply AC power for heating.
If the side reaches the overtemperature limit again, the system will shut
down. This on/off condition is called “cycling”. AFM “Abnormal
Procedures” must be followed.
• Left or right system malfunctions other than overheat conditions

activate the amber W/S FAULT annunciator only. AFM “Abnormal
• Windshield sensors are tested with the rotary test knob—WS TEMP
per AFM “Normal Procedures.” See “Rotary Test” section for
procedures.
NOTE
The W/S FAULT annunciator may not test after cold
soak at extremely cold temperatures. If this occurs,
repeat the test after the cabin has warmed up. The test
must be completed prior to flight.
I f t h e w i n d s h i e l d i s h e a t s o a ke d a b ove + 5 6 ° C
(+134°F), the test results in a W/S FAULT annunci-
ator illuminating.

SECONDARY
140°F/60°C SENSOR
S S
W/S
O'HEAT PRIMARY
L R P P SENSOR

110°F/43°C
(NORM TEMP)
DC DC
CONTROLLER W/S CONTROLLER
FAULT
L R
WINDSHIELD
LH RH
ALTERNATOR L R ALTERNATOR
O'RIDE
ON
LEGEND AC
BEARING OFF
LH ALTERNATOR
L R
SRE-47
RH ALTERNATOR
Figure SRE-23. Windshield Anti-Ice System

• The left and right WINDSHIELD switches on the tilt panel have three
positions:
• O’RIDE—Allows the controllers to heat the windshield to 110°F
at a faster heating rate than normal. The controller continues to
regulate windshield temperature at 110°F.
• ON (normal operating position)—Allows the controller to heat the

windshield to 110°F at a normal “ramp-up” heating rate. The
controller continues to regulate windshield temperature at 110°F.
• OFF—Deactivates the controller and the respective alternator.
• The left and right rear unheated windshield panels are individually
defogged via cockpit underfloor air. The amount of air is controlled
with the WINDOW VENT knobs on the left and right cockpit side
panels.
4. Engine anti-ice (Figure SRE-24):
• Engine anti-ice heats the fan nose cone, T1 and T0 probes, the nacelle
lip, and stator vanes.
• The following are heated anytime the engine is operating:
• Fan nose cone—P2.8 bleed air
• T1 probe—P3 bleed air
• With the engine or engine/wing anti-ice switch ON, the following are
heated:
• T0 probe—Electric
• Nacelle lip—P3 bleed air
• First two sets of stator vanes—P3 bleed air
• With an engine or engine/wing anti-ice switch ON, the nacelle and

stator anti-ice bleed-air valves (PRSOVs) deenergize open and the T0
probe receives electricity. The amber ENGINE ANTI-ICE annunciator
illuminates steady until the nacelle temperature rises above 60°F
(15°C). Above 60°F, the annunciator extinguishes indicating normal
operation.
NOTE
If ambient temperature is approximately 59°F (15°C)
or warmer, the ENG ANTI-ICE L–R annunciators
may not illuminate when anti-ice is selected ON. To
ensure that bleed air is flowing to the engine inlet,
the crew should observe a momentary small decrease
in N 2 when ENGINE ON is selected.

WING WING WNG XFLOW WING/ENGINE
ON L R
ANTI-ICE XFLOW O'HEAT ON
VALVE (106°C)
220° 220°
L R (N/C) L R OFF
ON
OFF ENGINE
160° 160°

160° 160°
(71°C) EMER
L WING PRESS
ANTI-ICE VALVE
PRSOV (N/C)
(N/O)
L PRECOOLER
ENG
ANTI-ICE 60° (15°C)
60° R NACELLE
L R ANTI-ICE PRSOV (N/O)
P3 P3
560° 560°
R STATOR ANTI-ICE PRSOV (N/O)
LEGEND
BLD AIR
PURGE AIR O'HEAT
L R
P3 BLEED AIR
RAM AIR
WING AND ENGINE ANTI-ICE ON
WING BLEED-AIR SHUTOFF CAPABILITY
SRE-49
DUE TO AN O'HEAT CONDITION
Figure SRE-24. Wing/Engine Anti-Ice System

• In flight with the engine or engine/wing anti-ice switch ON, if the

nacelle temperature does not exceed 60°F or the stator valve does not
open within 4 minutes 45 seconds, the steady amber engine anti-ice
annunciator begins to flash indicating a malfunction. AFM “Abnormal
Procedures” are required.
• During ground operations, the annunciator illuminates steady only
during warmup or for malfunction. The engine T 1 probe is not
monitored for electrical current or heat.
• For flights into icing conditions, the anti-ice system requires a preflight
test per AFM “Normal Procedures.”
• With engine or engine/wing anti-ice switch ON, engine ignition
activates continuously.
• With a loss of main DC power (emergency bus condition), the nacelle
and stator anti-ice valves fail-safe open. The T0 probe is not heated.
• The T1-FCU sensor (flush-mounted sensor) is heated indirectly with
nacelle heat.
• Engine anti-ice must be selected when temperature is +10°C or less
with visible moisture (ground or in flight).
5. Wing anti-ice (Figure SRE-24 and SRE-25):

• Precooled engine bleed air (P3) is used to heat the wing leading edge.
The bleed-air pressure is regulated at 16 psi.
• With the engine/wing anti-ice switch ON, the wing anti-ice bleed-air
valve (PROSV) deenergizes open and the amber WING ANTI-ICE
annunciator illuminates steady. When the wing leading edge
temperature exceeds 230°F (110°C), the annunciator extinguishes
indicating normal operation.
• In flight, the WING ANTI-ICE annunciator begins flashing if the wing
does reach operating temperature within 4 minutes and 45 seconds.
AFM “Abnormal Procedures” are required.
• The wing undertemperature switches are enabled with engine/wing
anti-ice switch ON.
• On ground, the WING ANTI-ICE annunciator illuminates steady only.
• With a loss of main DC power (emergency bus condition), the wing
anti-ice valves fail-safe open.
• Wing anti-ice is not recommended for use during ground operations.

WING
O'HEAT
L R

160°F SWITCH FUEL BOUNDARY
HEA
T SH
IELD
PUR
GE P
ASS
AIR A
FLO GE
W
BLE
DEFLECTOR SHIELD ED
AIR
SRE-51
Figure SRE-25. Wing Leading Edge Cross Section

NOTE
The wing anti-ice valve is held closed as a result of
a bleed-air overheat condition on the respective side.
This automatic action protects the wing leading edge
from excessive heat.
Wing crossflow (see Figure SRE-24):

• Wing crossflow is selected ON with the WING XFLOW switch. When
selected ON, the crossflow valve energizes open and connects the left
and right wing bleed air systems.
• Wing crossflow is used per AFM abnormal conditions.
• The crossflow valve fail-safes closed.
• There is no annunciation during cross-flow use.
Wing overheat protection (see Figure SRE-25):
• Wing overheat protection is provided by two 160°F (71°C) spar sensors

in the cooling purge passage airflow between the heated leading edge
and the fuel tank.
• On ground or in flight, if an overheat condition is sensed by either

switch, the amber WING O’HEAT L–R annunciator illuminates
flashing. This condition automatically closes the respective side wing
anti-ice valve (PRSOV) causing the wing to become cold, or not heat
up if wing anti-ice is selected.
As the overheat condition cools below the 160°F value, the wing anti-
ice valve automatically reactivates if its switch is ON. This OFF/ON
activation is called “cycling.” AFM “Abnormal Procedures” must be
consulted.
• The wing over-temp sensors are activated with or without wing anti-ice
selected ON.
6. Deice boots (Figure SRE-26):
• The tail deice system for the horizontal stabilizer is a pneumatic boot
system.
• Engine bleed air, service air (23 psi), is used to inflate and deflate the
boots.
• The system consists of a control switch in the cockpit, timer/logic PC

boards, two control valves, two pressure switches, two rubber deice
boots, and deice annunciators.
• The TAIL AUTO–OFF–MANUAL switch is on the ANTI ICE/DEICE

switch panel. The system is powered by main DC power through the
TAIL DEICE circuit breaker on the left CB panel.

NOTE:
XL USES A SINGLE LOGIC BOARD
XL/XLS
LOGIC
AUTO MODE–ONE
18 SECOND CYCLE BOARD
EVERY 3 MINUTES PRECOOLED
SERVICE BLEED
AIR PRESSURE
TAIL (ENG OR APU)
TAIL

DEICE
AUTO
5 OFF 23 PSI
PRESSURE
REGULATOR
MANUAL
LEGEND
VACUUM
RIGHT GENERATOR BELOW
16 PSI
16 PSI
PRESSURE
VACUUM PRESSURE SWITCH
SERVICE AIR
P P
16 PSI & ABOVE
COMBINATION VACUUM
EJECTOR/SOLENOID VALVES (NC)
L BOOT R BOOT
SRE-53
Figure SRE-26. Tail Deice System

• Selecting AUTO starts the automatic 18-second inflation cycle. The left
boot inflates during the first 6 seconds (white TL DEICE PRESS–L
advisory light illuminates), and then returns to the vacuum position,
extinguishing the annunciator light. After a 6-second pause, the right
boot inflates (white TL DEICE PRESS–R advisory light illuminates)
during the last 6 seconds, then extinguishes. Approximately 3 minutes
later, the cycle repeats itself.
• Placing the control switch to MANUAL, bypasses the timer logic and
simultaneously inflates both deice boots. The boots remain inflated as
long as the switch is held in the MANUAL position. Recommended
inflation time is 6 to 8 seconds and should be repeated at 3- to 5-minute
intervals as long as icing conditions are encountered. Both white
advisory TL DEICE PRESS L–R lights illuminate simultaneously as
both boots inflate.
• The amber TL DEICE FAIL L–R annunciator illuminates flashing if the

control switch is in AUTO and boot inflation pressure does not reach 16
psi, or the boots do not cycle properly due to timer or control valve
failure. The annunciator also illuminates flashing if the control switch has
been turned OFF and the timer and/or control valve(s) are still energized.
• Loss of normal DC power prevents tail deice function.
• The tail deice boots should not be activated at indicated RAT

temperatures below –35°C/–31°F at airspeeds at or above 150 KIAS or
–40°C/F at airspeeds below 150 KIAS. Boot cracking may result.
PNEUMATICS/AIR CONDITIONING
• Hot, P3 engine bleed air is used for environmental/pressurization, wing
anti-ice, and service air. This bleed air must be reduced in temperature by
use of a cross-flow type heat exchanger or “precooler.” Engine anti-ice
does not use precooled air; it uses raw bleed air off the side of the engine.
• During ground operations, the Excel aircraft uses cool engine fan
bypass air to flow into the precooler and extract heat from the engine
hot bleed air as it flows through into the bleed air manifolds. The
targeted temperature after the bleed air exits the precooler is 405°F
(207°C) (Figure SRE-27).
• After takeoff and in flight (WOW), ram air is used instead of engine
fan bypass air for precooler use. Ram air inters the precooler through a
NACA vent door under each engine pylon.
• The amber BLEED AIR O’HEAT L–R annunciator illuminates if the

bleed-air temperature exiting the precooler reaches 560°F (293°C). This
overheat condition automatically causes the respective side wing anti-ice
shutoff valve to close if open or not open if later selected for use.
• Excessively hot bleed air exiting the precooler can be shut off by
selecting the opposite side source with the source selector knob.

XL PRECOOLER OPERATION ON GROUND ONLY

LEGEND
GND GROUND TEMP SENSOR 405°C
AIR IN-FLIGHT TEMP SENSOR 475°C
560° OVERTEMP SENSOR
FAN AIR
VALVE
P3 ENGINE
PRECOOLER FAN AIR
BLEED AIR
CROSS-FLOW
MIXER
EXHAUST
VENT
GND
DC
AIR PRECOOLER
CONTROL
560°
BLD AIR
PRECOOLER AIR O'HEAT
TO SYSTEMS L R
XL PRECOOLER OPERATION IN FLIGHT

LEGEND
ELECTRICALLY
ACTUATED DOOR
P3 ENGINE
PRECOOLER BLEED AIR
CROSS-FLOW
MIXER
EXHAUST
VENT
GND
DC
RAM
AIR PRECOOLER
CONTROL
I N -FLIGHT AIR
560°
BLD AIR
TO SYSTEMS L R
WEIGHT OFF WHEELS
Figure SRE-27. Bleed-Air Precooler

• When operating at or above maximum cruise thrust (CLB detent), and

RAT is approximately 0°C or warmer, selection of WING/ENGINE
ANTI-ICE ON may cause the BLEED AIR O’HEAT L–R annunciator
to illuminate. The situation is not hazardous and corrects itself within a
few seconds.
The pressurization source selector has the following positions:
• OFF—All valves are closed; bleed air is still available for service air
and anti-ice/deice.
• LH—The left flow control valve is relaxed open; the right flow control
valve is energized closed. The ACM receives air from the left engine
only (6 ppm airflow).
• NORMAL—The left and right flow control valves are relaxed open
(this is the fail-safe condition of the system), providing normal airflow
from both engines to the ACM (12 ppm total airflow).
• RH—The right flow control valve is relaxed open; the left flow control
valve is energized closed. The ACM receives air from the right engine
• EMER—The emergency pressure flow control valve is energized open;
the left and right flow control valves are energized closed. Airflow to the
ACM is stopped and control of temperature is with the left throttle.
Emergency pressurization is not available on the ground. Emergency
pressurization is provided by the left engine only and airflow is diverted to
the forward portion of the dropped aisle ducts on the left side of the cabin.
Temperature control:
• Normal DC power is required for automatic and manual modes.

• Temperature control is provided separately for both cockpit and cabin.
• Temperature is controlled in the cockpit and cabin by mixing constant
temperature cool air from the ACM system with unconditioned warm
bleed air. This moderated warm air is directed to the warm air
distribution network for the cockpit and cabin areas.
• Cold air for the overhead WEMACS is supplied directly from the ACM
system.
• Components which are common to the cockpit and cabin temperature
control systems include temperature control valves (TCVs), mixing
muffs, temperature sensors, zone sensors, duct overheat switches, and
the temperature controller.
• AIR DUCT O’HEAT CKPT/CAB annunciators illuminate if the aft
end of the respective air supply duct (cabin or cockpit) temperature has
overheated (300°F/149°C).
• Additional sources of cold or hot air are provided by the optional
auxiliary power unit (APU) or vapor cycle air-conditioning unit (cold
air only) (Figure SRE-28).

CKPT TEMP SEL CABIN TEMP SEL
AIRDUCT
AUTO 7 1 AUTO O'HEAT ENG P3
BLEED AIR
CKPT CAB
COLD HOT COLD HOT CKPT CAB
SEL SEL
PRECOOLER
SUPPLY SUPPLY
MANUAL MANUAL

T
OZONE
CONVERTER
T T T APU
COCKPIT ARM REST
ZONE Z
SENSOR FLOOR
R FLOW
CONTROL
TCV (16 PSI)
(NO)
FOOT WARMERS
COCKPIT AREA
WEMACS WEMAC
WATER SEPARATOR APU BAV
BOOST
TCV
T
EMER
ACM
PRESS
T
WEMACS TCV ACM
O'HEAT
CABIN ZONE
SENSOR Z
AISLE
MIXING (NO)
FLOOR
MUFF
ARM REST BLD AIR
T
O'HEAT
475°F
LEGEND T
EMER L R
EMER (PRSOV) (NC) 560°F
PRECOOLED BLEED AIR PRESS
ANTI-SKID
COLD ACM AIR INOP
CABIN/COCKPIT UNDER-FLOOR DUCTING
STATIC FLOW
SRE-57
Figure SRE-28. Air Conditioning System with APU

Supplemental air conditioning—Vapor cycle system (Figure SRE-29)
• Additional supplemental cooling of the cabin and cockpit is provided

by a vapor-cycle air-conditioning unit in the tail cone. It is a
conventional compressor/condenser R134a system that is powered by
the main DC electrical system. It may be operated by itself, or in
conjunction with the ECU/ACM.
• The system consists of a belt-driven compressor, condenser, two cabin

evaporator/fans (one forward and one aft), and a cockpit-mounted
control unit on the copilot lower instrument panel.
• A barometric pressure switch is incorporated to shut down the

compressor drive motor above 18,000 feet.
• Supplemental cooling air is provided through the upper WEMACS by

the aft evaporator system, and through a floor outlet grill near the
cockpit by the forward evaporator system.
NOTE
The vapor cycle system is removed if an optional
APU is installed.

AC - FANS AHRS 2
CKPT AC ON OFF
HDG REV ATT REV ADC REV
RECIRC
LO LO DG L SLEW
HI
AC WEMAC
BOOST
TEST R SLEW
LO HIGH HIGH

AIR
EXHAUST
INTAKE
WEMACS
BAROMETRIC
FWD SWITCH
EVAPORATOR (18,000 ft)
FAN
UNIT AFT
FLOOR
GRILL EVAPORATOR
FAN UNIT
VAPOR CYCLE
WEMACS MACHINE
LEGEND
VENT AIR
AIR INTAKE
EXHAUST
COMPRESSOR DISCHARGE
SRE-59
COLD AIR
Figure SRE-29. Vapor Cycle Air Conditioning System

PRESSURIZATION
• The pressurization controls are on the center pedestal tilt panel (Figure
SRE-30)
• Normal DC power and 23 psi (service air) air/vacuum are required for
both AUTO and ISOBARIC MODE operation. AUTO mode also
requires input from the No. 1 ADC (Figure SRE-31).
• ISOBARIC MODE is the result of loss of No. 1 ADC input. It is

indicated by an amber LED on the face of the cabin pressure controller.
Control pressurization using FL and CA mode in the altitude select
window.
• MANUAL MODE can be selected at any time. It requires no DC

power source nor normal vacuum to operate. “cherry picker” uses
cabin pressure for closing the outflow valves and nose wheel well low
pressure static air for vacuum to open the outflow valves. Will not
override the 14,500 ±500 cabin altitude limiter valves, nor does it
override the MAX cabin differential protection of 9.5 ±0.1 psid.
• Provides a sea level cabin to 25,230 feet, with a 9.3 ± 0.1 psid.
Provides a 6,800 feet cabin altitude at 45,000 feet.
• Normal DC power and vacuum is required to operate the cabin dump

valve. Cabin dump does not override the cabin altitude limit valves.
• Airplane is depressurized on the ground (left squat switch) with less

than 85% throttle position angle.
• Above 85% throttle position angle on the ground, the pressurization

controller goes into the takeoff/prepressurization mode. This increases
cabin pressure to preclude pressure bumps upon takeoff. Normal
prepressurization descends the cabin to approximately 50 feet below
takeoff field elevation during the takeoff run.
• The pressurization controller (Figure SRE-31) provides both normal

(8,000 feet and below) and high altitude (8,100 to 14,000 feet)
autoschedule pressurization control, depending upon what set landing
altitude (SLA) is selected in the SET ALT window of the controller
(Figure SRE-32). High altitude mode prevents the CAB ALT
annunciator from illuminating at a cabin pressure of 10,000 feet while
descending or climbing below FL250 (Figures SRE-33 and SRE-34).
However, it illuminates if cabin altitude should reach 14,500 feet.
• The controller is programmed to limit cabin climb and descent rates to

+600/–500 fpm respectively.
• High altitude mode climb and descent rates are limited to a maximum
of +2,500/–1,500 fpm respectively.

ANTI-ICE / DEICE PRESSURIZATION
PRESS SYSTEM SELECT

PITOT & WINDSHIELD WINDSHIELD
STATIC AIR WING INSP UP EMER DUMP 15 20
18
ON L R ON ON MANUAL M ON
O'RIDE 10 25
A SET ALT 4 5 6
3 PSI 7
ON N FL EXER 8 30
2
U 1 9
A 5 0 10 35
OFF L DIFF 40
OFF OFF OFF AUTO
DOWN
NORM 0 0
PRESS 50
CABIN ALT
WNG XFLOW WING/ENGINE RATE X1000 FT
TAIL
ON L ON R AUTO DEPRESSURIZE CABIN BEFORE LANDING
PRESS SOURCE
OFF OFF NORM
ON CKPT TEMP SEL CABIN TEMP SEL
LH RH
OFF ENGINE MANUAL
LIGHTS
PASS GND REC/ AUTO AUTO
SAFETY NAV TAIL FLOOD ANTI-COL OFF EMER
ON ON ON ON
GND CKPT CAB
OFF REC COLD HOT COLD HOT
ON SEL SEL
SEAT BELT OFF OFF OFF SUPPLY SUPPLY
ON MANUAL MANUAL
SRE-61
Figure SRE-30. Pressurization Control Panel

SRE-62
PRESS SYSTEM SELECT

UP EMER DUMP 15 20
MANUAL M ON
10 25
A SET ALT 4 5 6
3 PSI 7
N FL EXER 8 30
2
U 1 9
A 5 0 10 35
AUTO L DIFF 40
NORM PRESS 50
DOWN 0
CABIN ALT
RATE X1000 FT
DEPRESSURIZE CABIN BEFORE LANDING

OUTSIDE
STATIC
28 VDC SOURCE
#1 AIR CABIN AIR

DATA
COMPUTER
NOSE WHEEL
WELL STATIC HIGH
SOURCE ALT
SIGNAL CABIN AIR
VACUUM
EJECTOR
> 1.5 PSID
CABIN AIR
CABIN AIR SHUTTLE

VALVE CABIN AIR
1.5 PSI
ORIFICE CABIN AIR
OUTSIDE
STATIC
SOURCE VACUUM
FLAPS
CAB ALT 23 PSI
LEGEND
UP 0°
BLEED AIR
TO
TRIM
NOSE
CLB T.O.
200 KIAS 7°
STATIC PRESSURE
DOWN
CRU
T T
O H T.O. &
R
O
T
APPR
200 KIAS
15°
SERVICE AIR
T
L
E
NOSE
UP IDLE
LAND
175 KIAS
35° CABIN AIR
SPEED
BRAKE CUT
OFF
ENGINE SYNC
LH RH MUST BE OFF
RETRACT & LANDING
VACUUM
EXTEND
Figure SRE-31. Pressurization System

Max Delta P Limit Autoschedule Boundary
45000
40000 Cruise @ FL410

35000
30000
Aircraft Descent to SLA
Altitude 25000
(FT) Climb to FL410
20000
Cabin @ SLA
15000 1500 ft above SLA
Takeoff
10000
from 1000 FT
5000
Negative Delta P Limit
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)

SRE-63
Figure SRE-32. Autoschedule Boundary

45000
Aircraft climbs to
35000
Cabin Holds @ 78000 ft until
Cabin Climbs Acft descends below FL 245
30000
Aircraft to and maintains
7800 ft. at 600 FPM
Altitude 25000 Cabin Climbs to
Landing Field
(FT)
20000 (NLT 1500 AGL)
15000 Takeoff from

3000 ft
10000
5000
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)
Figure SRE-33. High Altitude Landing Graph
45000
35000
Cabin will reach 8000 ft with
30000 Acft at approx. FL 250
Aircraft
Altitude 25000
(FT) Descent to
20000
SLA
Climb to Takeoff from
15000
FL 450 14000 ft
10000
5000
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)
Figure SRE-34. High Altitude Departure Graph

SERVICE AIR
• Bleed air supplied by the engines or an optional APU.
• Regulated at 23 psi.
• Used for (Figure SRE-35):
• Horizontal stabilizer deice boots, inflation pressure.
• Pressurization outflow valve operation.
• Cabin entrance primary door seal and acoustic door seals.
• Throttle detents, EECs AUTO mode.
FLAPS
UP 0°
THROTTLE TRIM
NOSE
DOWN
TO
CLB T.O.
200 KIAS 7°
DETENTS
CRU
T T
O H T.O. &
R APPR 15°
O 200 KIAS
T
T
L
E
NOSE
UP IDLE
LAND 35°
175 KIAS
SPEED
BRAKE CUT
OFF
ENGINE SYNC
LH RH MUST BE OFF
RETRACT & LANDING
EXTEND
DOOR SEALS
VACUUM EJECTOR
FOR OUTFLOW VALVES
23 PSI
PRECOOLER REGULATOR
PRECOOLER
L FLOW ACM
CONTROL
VALVE P3 ENG
BLEED AIR
LEGEND APU
BAV
SERVICE AIR
VACUUM APU
BLEED AIR
BLEED AIR TO DEICE
SYSTEM
Figure SRE-35. Service Air System

OXYGEN
• 50-cubic-foot bottle is standard with an option of a 76-cubic-foot
bottle in the right side of the lower nose compartment (Figure SRE-
36).
• The bottle pressurization green arc is marked from 1,600 to 1,800 psi.
This does not ensure oxygen availability to the crew or cabin.
• Quick-donning EROS crew masks are stowed in a retainer below the

crewmember side windows. The masks have an integral microphone
and pressure regulator. Three positions are afforded: EMER (for
pressure breathing), 100%, and diluter demand. Masks must be stowed
properly to qualify as quick-donning masks.
• Passenger masks are stowed in overhead containers. Passenger oxygen

selector on the pilot console has three positions: OFF (crew only);
AUTO (masks automatically drop if cabin pressure exceeds
approximately 14,500 feet, with normal DC power available); ON
(manual drop).
• With the OXYGEN selector in AUTO; if cabin altitude exceeds 14,500

feet, passenger masks drop automatically. If cabin pressure is restored
to normal values, the solenoid valve is deenergized at approximately
12,000 feet cabin altitude, shutting off oxygen flow to the passenger
masks.
• Oxygen cylinder is serviced through a service port in the lower aft sill
of the right nose compartment (aviator breathing oxygen only!).
• A green overboard discharge indicator (disc) is below the aft edge of

the nose compartment door. Missing or ruptured disc indicates that the
oxygen cylinder has overpressurized and maintenance must be
performed before flight.

14,500 +/- 500
FILLER VALVE &
PROTECTIVE CAP
Cabin Altitude
OVERBOARD
COPILOTS
DISCHARGE FACE MASK
INDICATOR CYLINDER
PRESSURE GAUGE 5 AMP
OXYGEN
CB

28 - VOLT
DC
SHUTOFF
VALVE ALTITUDE
PRESSURE
SWITCH
PRESSURE
REGULATOR OVERHEAD
OXYGEN CHECK
CYLINDER VALVE DROP BOX
SOLENOID
PILOTS FACE
MASK
LEGEND
PASS OXY
OXYGEN SUPPLY (HI PRESS) ON OFF AUTO
OFF ON
OXYGEN CYLINDER
PASS OXY
OXYGEN SUPPLY (REG MED PRESS) AUTO
STATIC FLOW OXYGEN CONTROL VALVE

Oxygen System
SRE-67
Automatic Deploy
Figure SRE-36. Oxygen System
AUXILIARY POWER UNIT

The Allied Signal Model RE100-XL is a fully automatic, constant speed gas
turbine engine mounted in a titanium steel fireproof enclosure in the tail
cone. It utilizes a single-stage centrifugal impeller and a single-stage turbine.
The APU requires main DC power, fuel from the right tank, and control sig-
nals from the aircraft for operation. Service is conducted through the APU
panel on the right-rear fuselage above the engine pylon. The following lists
some general APU highlights:
• The APU is optional equipment. The unit is installed in place of the

standard vapor cycle air conditioner.
• APU installation increases aircraft weight by approximately 100

pounds plus any required nose ballast.
• The APU generator provides 28 VDC power and bleed air for ground
and in-flight use.
• Maximum altitude for in-flight start is 20,000 feet.
• Maximum altitude for in-flight operation is 30,000 feet.
• The APU produces no thrust.
• The APU is not certified for unattended use.
ELECTRONIC CONTROL UNIT (ECU)

The ECU is responsible for automatic APU operations. The ECU supplies the
following functions:
• The ECU is powered with the APU MASTER switch ON.
• Built-in test during power-up.
• Automatic start control.
• Speed control.
• ECU autorelight function applies ignition at 94% rpm to prevent

flameout.
• Protective shutdown capability—Excessive EGT, rpm overspeed, low

oil pressure (LOP), high oil temperature.
• Start inhibit capability.
• Fault code storage.
• Fault reporting to the field service monitor (FSM). The FSM provides
download capability.

FUEL SYSTEM
• Fuel is normally supplied from the right wing fuel tank except during
left to right crossfeed operations.
• Right boost pump operates continuously during APU start and APU
operation. If crossfeeding from left to right, the left boost pump
supplies fuel for APU operations (the right boost pump deenergizes).
• When the right boost pump is operating for APU operations only, the
amber FUEL BOOST R annunciator does not illuminate.
• Fuel flow is 110 pph during loaded operation (generator online and
bleed valve open)
• Fuel flow indications are available in the FMS.
• APU fuel valve opens during the start sequence and closes for
normal/abnormal shutdown including APU fire.
OIL SYSTEM
• Oil reservoir is in the accessory gearbox. Oil quantity is approximately
1.5 US quarts.
• APU normally uses the same oil as the engines.
• Oil service is through the small door on the outside access panel.
• The oil reservoir is cooled with compressor intake.
• APU oil level is preflight checked 5 minutes or longer following
shutdown.
• The APU service panel in the tail cone is used to check oil level
electrically. Following a successful panel LAMP TEST, select PRE
FLT position:
• No illuminating lights indicates full oil.
• Amber illumination indicates 300 cc low of oil. APU operation is
permitted. Service at next opportunity.
• Red and amber illumination indicates 550 cc low of oil. APU
operation is prohibited. Oil service is required.
• APU service panel is battery bus powered.
• Low oil pressure switch (LOP) signals the ECU to initiate a protective
shutdown. The amber APU FAIL annunciator illuminates on the far
right cockpit panel (Figure SRE-37).
• High oil temperature signals the ECU to initiate a protective shutdown.
The amber APU FAIL annunciator illuminates.
• Magnetic chip collector is inspected by maintenance only.

Figure SRE-37. APU Annunciators, Copilot Panel

PNEUMATIC SYSTEM
• A main duty of the APU is to provide supplemental bleed air to the
aircraft environmental/pressurization and all service air systems.
• The ACM, TCVs and underfloor ducting, and deice boots are major
APU bleed air users.
• The APU cannot supply bleed air to the anti-ice systems.
• Bleed air from the APU is supplied through a bleed-air valve (BAV)
(see Figures SRE-28 and SRE-35).
• The BAV is controlled by the ECU and the BLEED AIR ON–OFF
switch located on the APU control panel.
• After start, with the BLEED AIR switch ON, the ECU opens the BAV
and supplies regulated bleed air to the aircraft bleed-air manifold.
When the BAV is open (or other than closed), the white BLEED VAL
OPEN annunciator illuminates.
• If an ACM O’HEAT condition occurs, the ECU commands the BAV to

close until the condition is cleared.
• Bleed air is regulated by the ECU according to EGT and inlet ambient
temperatures. As EGT increases, bleed air is reduced to maintain a safe
EGT. If EGT reaches 690°C, the ECU provides a protective shutdown
(Figure SRE-38).

MAP LIGHT
QUARTZ
0001248
TOTAL HOURS
DIM
OFF
APU SYSTEM
BLEED AIR GENERATOR
ON ON
BLEED VAL OPEN
O
F
READY TO LOAD
F
OFF RESET
APU RPM %
MAX RPM
108%
APU EGT 0 C
MAX EGT
6900
DC VOLTAGE
APU MASTER
START TEST ON
N
O
R
M
STOP PUSH OFF
Figure SRE-38. APU Control Panel

ELECTRICAL SYSTEM
• One 28 VDC, 300 constant ampere starter-generator is on the gearbox.
APU generator load has priority over bleed-air load. The ECU will
reduce bleed air as required to maintain 100% shaft rpm for generator
operation.
• The generator is controlled via its GCU and generator switch on the
APU control panel (Figure SRE-38). The GCU and three-position
generator switch operate identically to engine generator switches.
• When selected online, generator current is supplied to the system at the
crossfeed bus (Figure SRE-39). The APU generator and the engine
generators can all simultaneously supply power to the aircraft’s bus
system.
• Generator load is indicated with an amperage gage located on the far
right side of the cockpit control panel (see Figure SRE-37).
• Maximum generator loads (red lines) are 200A on ground and 230A in
flight up to 30,000 feet.
• The APU and engine generators are not interchangeable.
FIRE PROTECTION
• Fire detection—Uses a gas-filled fire detection loop inside the fireproof
APU enclosure. As heat increases, the gas expands and causes a
pressure sensor to activate the red APU FIRE switchlight on the far
right cockpit control panel (see Figure SRE-37). Upon fire detection
the following occur:
1. ECU automatically initiates an automatic shutdown.
2. APU generator goes off line (field relay opens).
3. APU fuel shutoff valve closes.
4. Fuel boost deenergizes.
5. The APU fire bottle is armed.
6. An APU fire protective shutdown is stored in the ECU memory.
• Fire extinguishing—One dedicated fire bottle is above the baggage
compartment ceiling.
• Fire bottle arming by the ECU is indicated with the illumination of the
red APU FIRE switchlight. Pressing the red switchlight fires the
contents of the bottle into the APU compartment.
• If the red switchlight is not pressed by the crew, the ECU fires the
bottle 8 seconds after the light illuminates.
• The fire detection loop and bottle is continuously monitored by the
ECU. If the loop malfunctions or bottle becomes low or discharged, the
ECU automatically shuts down the APU or inhibits its start. The amber
APU FAIL annunciation illuminates for either malfunction.

SRE-74
FIRST ENGINE START (R) USING APU GEN & BATTERY - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
SYS SYS DISG AVN AVN
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
P ON RELAY RELAY E RELAY ON I
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
U L GEN RELAY RELAY R GEN O
S BUS BATTERY BUS BUS R
GCU GCU
APU 1
RELAY 7
L APU RELAY R 5
LEGEND BATTERY ENGAGED A
STARTER STARTER
GPU
APU GENERATOR FIELD GEN BATT RELAY OVER (32.5 VDC) GEN FIELD
RELAY DISCONNECT VOLTAGE RELAY
RELAY
BATTERY GPU INTERIOR POWER
INPUT
Figure SRE-39. First Engine Start (R)—APU Generator On Line

EXTERIOR PREFLIGHT
• Check APU air inlets on the upper right rear fuselage—Check CLEAR
(compressor inlet, cooling inlet for the starter-generator).
• APU exhaust—Check CLEAR.
• Tail cone ram air inlet on the right rear fuselage below the pylon—
Check CLEAR.
• APU drains on the bottom of the rear fuselage .
• Check oil quantity lights on the service panel in the tail cone.
APU CONTROL PANEL AND ANNUNCIATOR FUNCTIONS

Prior to APU starting, the aircraft battery switch must be ON (see Figure
SRE-38).
NOTE
APU starts on the ground may be aircraft battery
starts only, EPU starts only (battery disconnect relay
opens during start), or aircraft generator(s) assisted
battery starts.
In-flight APU starts are battery only starts (squat
switch logic prevents generator-assisted APU starts).
In-flight starts are prohibited above 20,000 feet.
In-flight APU starts are prohibited after dual gener-
ator failure.
APU MASTER switch—The MASTER switch is placed ON to provide elec-

trical power to the ECU. The ECU performs APU power-up tests. After the
power-up tests are completed, the ECU accomplishes the prestart built-in test
equipment (BITE) test to ensure no faults exist that would inhibit a start. If
a fault is detected, the APU FAIL light illuminates.
APU FAIL light—Illuminates for an APU fault of low fire bottle pressure. APU
start attempt is prohibited when the APU FAIL light is illuminated.
APU TEST button—Performs a lamp test of annunciators (FIRE WARNING,

APU FAIL, APU RELAY ENGAGED, BLEED VALVE OPEN, READY TO
LOAD), digital indicators (RPM-50, EGT-500, DC VOLTS-00.0) and in-
tegrity of the APU fire system.
APU GENERATOR switch—ON position allows generator power to connect

to the airplane crossfeed bus after the READY TO LOAD light illuminates.
OFF position disconnects the APU generator from the crossfeed bus. RESET
position allows a possible reset of an APU generator tripped field relay.

APU BLEED AIR VALVE switch—ON position may be selected following

READY TO LOAD illumination. The ON position opens the BAV approxi-
mately half way, after a short delay the valve opens to full position. The
white BLEED VAL OPEN annunciator illuminates on the APU control panel
when the valve is other than closed. The OFF switch position closes the BAV
and the annunciator extinguishes when the valve is fully closed. AFM requires
the switch be OFF for start and shutdown.
NOTE
Any time the APU is operating, the service air sys-
tem is pressurized whether or not the bleed-air valve
is open or closed.
APU START/STOP switch—The ECU provides automatic starting after plac-

ing the MASTER switch “ON” followed by momentarily placing the APU
START switch to “START.” The ECU controls ignition and fuel automatically
during start as required for ambient conditions.
The aircraft right boost pump activates (FUEL BOOST–R annunciator remains
extinguished; LO FUEL PRESS–R goes out).
If the APU start is an engine generator(s) assisted start (ground only), the en-
gine start relay(s) close (engine start button(s) illuminates), and the APU start
logic commands the battery isolation relay open to protect the 225-amp cur-
rent limiters.
At 5% rpm, the ECU powers the ignition unit, fuel torque motor, and the APU
fuel solenoid valve (open). During start, the ECU controls fuel scheduling,
and continually monitors engine speed and EGT limits as determined by am-
bient conditions (T 2 ). If scheduled limits are exceeded, the ECU executes a
precautionary shutdown (APU FAIL illuminates). The fault code is stored in
memory for ease of maintenance during troubleshooting.
The STOP position initiates a simulated overspeed signal to the ECU to ini-
tiate an immediate shutdown. After commanding shutdown using the APU
START–STOP switch, the ECU remains powered until the APU MASTER
switch is placed OFF.
Following an APU shutdown for any reason, a restart must not be attempted
until 30 seconds after the rpm indicator displays 0%
APU RELAY ENGAGED light—Illuminates then extinguishes prior to the
READY TO LOAD light illuminating. At 50% speed, the speed sensor signals
the GCU to deenergize the start relay and the APU RELAY ENGAGED extin-
guishes. If the speed sensor fails and/or the GCU fails to open the start relay
at 50%, the ECU backs up the GCU and opens the start relay at 60% rpm.
READY TO LOAD light—At 95% rpm the start counter records the start.
At 95% rpm plus 4 seconds, the ECU shifts to onspeed control. The READY
TO LOAD illuminates (start is complete). The APU may now be loaded elec-
trically and pneumatically.

At 99% rpm, the ignition unit is deenergized.

At 100% rpm, the APU is considered onspeed. At 100% rpm, the ECU main-
tains constant rotor speed rpm at 100% plus or minus 1.0% (70,200 rpm), EGT
within limits and the DC VOLTAGE indicator should display 28.5 VDC.
If APU speed drops below 94%, the ignition unit automatically reenergizes,
unless the APU is in a protective or normal shutdown mode.
The programmed ECU onspeed EGT and overspeed shutdown limits are es-
tablished at 690°C (1275°F) and 108% respectively.
APU GENERATOR—After the READY TO LOAD illuminates, the APU gen-
erator may be placed online. Placing the APU generator switch ON, energizes
the APU generator power relay to connect the APU generator output to the
airplane crossfeed bus. The APU ammeter on the copilot instrument panel
should reflect an amperage load.
APU GEN OFF light—Indicates the APU generator relay is open with the APU
running onspeed.
APU AMMETER—200 amp maximum on ground, 230 amp maximum in flight.
HOBBS METERS—At the bottom of the APU panel. Begins recording APU
operation when normal oil pressure is sensed by the ECU. This meter is used
for generator maintenance.
APU FIRE light/button—Alerts the crew of an APU fire in the APU enclo-
sure. APU immediately shuts down. Pressing the button activates the extin-
guisher. Extinguisher automatically activates 8 seconds after the light
illuminates if the button is not pressed.

APU START SEQUENCE

• APU start is initiated via the spring-loaded three-position APU START
switch (see Figure SRE-38):
• APU START—Signals the ECU to automatically initiate the start

sequence.
1. The APU relay closes supplying power to the APU starter. The
APU RELAY ENGAGED annunciator illuminates (Figure SRE
40).
2. Right fuel boost activates.
3. APU fuel shutoff valve opens.
4. Ignition is applied at 5% rpm.
5. Starter cutout is at 50% rpm, at 60% rpm backup. APU

RELAY ENGAGED annunciator extinguishes.
6. ECU provides acceleration fuel scheduling
7. READY TO LOAD annunciation at 95% rpm + 4 seconds (see

Figure SRE-38).
8. Ignition terminates at 99% rpm
9. ECU maintains 100% normal rpm.
• NORM—Normal spring-loaded position for other than START or

STOP.
• STOP—ECU initiates a simulated overspeed signal that stops the

APU. The ECU remains powered for future starts.
NOTE
Following APU shutdown for any reason, an APU
restart must not be attempted until 30 seconds after
the rpm indicator reads zero.
WARNING
The airplane battery must be installed and the battery
switch in BATT position or the airplane generator(s)
must be on and operating prior to and during all APU
operation to assure fire protection system power.
• Start counter retains total APU starts (in the tail cone or in the cockpit
at the bottom of the APU control panel).

APU START - USING BATTERY ONLY - ENGINE GENERATORS NOT AVAILABLE - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
GEN

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B
START START I
GCU GCU
APU 1
RELAY 7
L APU RELAY R 5
STARTER STARTER
GPU
BATTERY FIELD GEN BATT RELAY OVER (32.5 VDC) GEN FIELD
RELAY
GPU INTERIOR POWER
INPUT
SRE-79
Figure SRE-40. APU Engine Start On Ground (Engines OFF)

START POWER LOGIC

• APU BATTERY START ON GROUND—GPU or engine generators
are not available (see Figure SRE-40).
• Battery supplies current through the APU relay for starting.
• APU GPU START ON GROUND – GPU is connected and available

(Figure SRE-41).
• GPU supplies current for start through the APU relay. The battery
disconnect relay opens and the battery does not supply current.
• FIRST ENGINE START, APU GENERATOR ON LINE—APU

generator and battery available (SNs 5253 and on or SB 560XL 49 10
incorporated) (see Figure SRE-39).
• APU generator current is supplied for engine start through the

APU relay and the isolation relay.
• Battery current is supplied through the APU relay for engine start.
• SECOND ENGINE START ON GROUND, APU AND ENGINE

GENERATORS ON LINE—Both generators and the battery are
available (Figure SRE-42):

APU relay.
• Engine generator current is supplied for cross start through the left
and right start relays.
• Battery isolation relay opens to prevent current travel.
• The battery supplies engine start current through the left and right
start relays.
• APU START ON GROUND, ENGINE GENERATORS ON LINE—

Both engine generators and the battery supply current for APU start
(Figure SRE-43).
• Engine generators supply current for APU start through the

respective start relays and the APU relay.
• The battery supplies engine start current through the APU relay.

APU START - USING GPU - ENGINE GENERATORS NOT AVAILABLE - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A

GEN
OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B
START START I
GCU GCU
APU 1
LEGEND RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
BATTERY STARTER STARTER
GPU
EXTERNAL DC RELAY DISCONNECT VOLTAGE RELAY
RELAY
INTERIOR POWER
SRE-81
GPU
Figure SRE-41. APU Start—On Ground, Using GPU

SRE-82
SECOND ENGINE START (L) USING R ENG GEN, APU GEN, & BATTERY - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
LEGEND GCU GCU
APU 1
RELAY 7
BATTERY L APU RELAY R 5
BATTERY ENGAGED A
STARTER STARTER
APU GENERATOR GPU
R GENERATOR RELAY
GPU INTERIOR POWER
INPUT
Figure SRE-42. Second Engine Start (L)—APU Generator On Line

APU START USING L & R ENGINE GENERATORS & BATTERY - GROUND ONLY - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
LEGEND GCU GCU
APU 1
R GENERATOR RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
STARTER STARTER
L GENERATOR GPU
BATTERY RELAY
GPU INTERIOR POWER
INPUT
SRE-83
APU GENERATOR
Figure SRE-43. APU Start On Ground (Generator Assist)

• APU START IN FLIGHT—Battery current only is available during

flight (Figure SRE-44):
• Battery supplies APU start current through the APU relay.
• Battery isolation relay opens to prevent engine generators from

participating.
APU OPERATING LIMITATIONS

• APU operation is prohibited until a satisfactory APU test has been
accomplished as contained in the “Normal Procedures” section of
Supplement 16 of the Excel Airplane Flight Manual (AFM).
• Starting the APU is prohibited whenever the APU FAIL light is

illuminated.
• APU start attempt is prohibited after a dual generator failure.
• Following shutdown for any reason, APU restart must not be attempted
until 30 seconds after the RPM indicator reads 0%.
• Applying deice (anti-ice fluid of any type) is prohibited with the APU
operating.
• Deployment of the thrust reversers for more than 30 seconds with the
APU operating is prohibited.
• The APU is not approved for unattended operation.
• The limits in Table SRE-1 apply to APU starting and operation
Table SRE-1. APU OPERATING LIMITS

OPERATING MAX MAX N1% FUEL MAX GEN AMBIENT
CONDITION: ALT FT EGT OC TEMP OC LOAD AMPS TEMP OC
(NOTE3) (NOTE2)
STARTING 20,000 690 -- Refer to basic -- -54 to 54

AFM fuel limits
RUNNING 30,000 690 108 Refer to basic 200 GND

AFM fuel limits 230 FLT -54 to 54
(NOTE 1)
NOTES
1. Transient current greater than 200 amperes is approved for APU cross
generator start of main engines.
2. APU Ammeter Instrument Markings:
a. Red Triangle = 200 amperes.
b. Red Line = 230 amperes.
3. APU automatically shuts down if EGT or rpm limits are exceeded.

APU START USING BATTERY ONLY - IN-FLIGHT - AVIONICS ON
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
GEN

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B START START I
GCU GCU
APU 1
LEGEND RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
L GENERATOR STARTER STARTER
GPU
R GENERATOR RELAY DISCONNECT VOLTAGE RELAY
RELAY
GPU INTERIOR POWER
INPUT
SRE-85
BATTERY
Figure SRE-44. APU Start—In Flight (Battery Only)

BATTERY AND APU STARTER CYCLE LIMITATIONS

• Battery and APU starter cycle limitations
• Starter limitation—Three APU start cycles per 30 minutes. Three
cycles of operation with 90 second rest period between start
cycles is permitted.
• Battery limitation—Nine APU start cycles per hour. (An APU
battery start counts as 1/3 if a normal engine battery start.)
NOTE
1. On the ground, no battery cycle is counted when
starting the main engines using a cross genera-
tor start from the APU generator or from a ground
power unit.
2. Use of an external power source with voltage in
excess of 28 VDC or current in excess of 1,000
amps may damage the starter. Minimum 800 amps
for start.
3. If battery limitation is exceeded, a deep cycle in-
cluding a capacity check must be accomplished to
detect possible cell damage. Refer to Chapter 24
of the Excel Maintenance Manual for procedure.
AVIONICS
All primary avionics systems and components are DC-powered Primus 1000
EFIS system. Sensor inputs include (Figure SRE-45):
• Dual Litef LCR-93 attitude and heading reference system (AHRS).
• Dual micro air data computers (MADC).
MADCs are powered by ADC 1 and 2 circuit breakers on RH CB panel and

provide the following data to the High level data link control bus (HLDC):
• Pitot pressure, total and static air temperature for TAS/CAS to the IC-
600s/615s for PFD airspeed tapes, MACH and VMO/MMO indications
and warning horn.
• Static pressure, pressure altitude, and baro-corrected altitude (inches or
hPa) for the PFD altitude tapes.
• Altitude change rate for the PFD vertical velocity indicators.
• TAS data for the FMS and MFD.
• Pressure altitude to the optional TCAS.
• Altitude information to the Kollsman pressurization controller (ADC
No. 1 only).
• Also outputs data for the transponder, flight data recorder, flight
director, and autopilot.

AHRS #1 AHRS #2
A #1 #2
A
H ATT MICRO AIR DATA ATT H
R COMPUTERS R
U HDG HDG U

FLUX FLUX
VALVE VALVE
AUTOPILOT
SERVOS DIGITAL DATA BUS
IAC IAC
#1 FD/AP #2
PFD 1
FD/AP
PITCH PFD 2
IC 600 OR 615 IC 600 OR 615

SENSOR INTERFACE SENSOR INTERFACE
FD COMPUTER FD COMPUTER
SYMBOL GENERATOR SYMBL GENERATOR
ROLL AUTOPILOT COMPUTER HSI TCAS WX BARO PRE NAV FMS
MAP
PLAN
TCAS
WX
TERR
NORM EMER HSI TCAS WX
TERR
BARO
RAD
PRE
VIEW
NAV FMS
TERR RAD VIEW
INC
ET2 PUSH TO ENTER RCL SKP NAV ADF NAV ADF
NAV ADF NAV ADF R
N OFF FMS OFF FMS
OFF FMS OFF FMS PUSH
PUSH G PUSH STD
PUSH STD ST1 ST2 PAG ESC OFF TO TEST
OFF TO TEST DEC OFF DATA
SET BARO
BARO MFD DIM Honeywell BRG PFD DIM MINIMUMS SET BRG
BRG PFD DIM MINIMUMS SET BRG
VAPP LNAV VNAV AL-VN 2000 FMS LNAV VASEL VGP

PFD MFD
242 YD OFF AP OFF CAT2 1400 160 AP ENG 9000
920 HDG FMS1 RWO1L
280
329
349 23.0 NM
12 MIN 160 E 20 20
FMS 1000
20 20 6 6
4 N 39 10 10 4
260 33
2 140 10000 2
YAW
10 10 1500 1
3 3 6 1
KHUT
242 13 80 125 98 20
30
240 WPT
60 120 00
1 1 002 4 2 10 10 1
10 10 2 2
6
220 4 WPT R 20 20 4
6 001 100 9500 6
W
20 20 OM I 1
–950
25.0 30
RAD BARO
200 MIN MIN
.750 M 2500 29.92 IN .261 M 200 STD
AOA DME H ICT AOA DME
HDG FMS1 RW0IL HDG FMS1 RW0IL
349 VOR1 VOR1
329 023 13 23.0 NM RW01L
9.9L
329 349 023 13 23.0 NM
.70 12 MIN DME1 TCAS DME2 .50 12 MIN
ICT ICT
33 N 157 KTS 13 NM 13 NM 33 N 157 KTS
3 FMS STATUS TCAS TEMP KHUT FMS STATUS
BRG PTR
RAT +6°C BRG PTR 3
30
MSG +13 +13 WPT MSG

FMS1 FMS1
30
ABOVE SAT –1°C 002
ADF APPR RELATIVE ADF APPR
6
40 SPEED
W
ET DR TCAS AUTO TAS 162 ET DR

-04
0:00:00 CLOCK GSPD 157 0:00:00 WPT
WIND 001 WIND
E
CLOCK 19:37:47 DEST CLOCK

-04
24
19:39:07 39
WEATHER RW01L 19:39:07 5 39
WEATHER WX/R/T 12 MIN WEATHER
12
WX/R/T 21 TAWS T4.5° A TAWS WX/R/T TAWS

T4.5° A S 15 STAB TGT TERRAIN T4.5° A
STAB TGT TERRAIN LX/ON INHIBIT STAB TGT RA 9.8NM+13 TERRAIN
LX/ON INHIBIT LX/ON TA 4.5NM-04 INHIBIT
SRE-87
Honeywell Honeywell Honeywell
PFD MFD PFD
Figure SRE-45. Primus 1000 System Block Diagram

The true airspeed (TAS) temperature probe (Rosemount) provides tempera-

ture data to the MADCs only. It is electrically anti-ice protected any time the
aircraft is weight-off-wheels and the avionics master power switch is on.
The RAT gauge source temperature is provided by normal DC from the EEC temp
sensor (T.0. probe) in the right engine inlet. If the right T.0. probe fails, No. 2
MADC automatically provides temperature information to the RAT indicator.
Integrated avionics computers (IAC):
• Dual IC-600 or IC-615 computers provide data processing for the pilot
and copilot EFIS system. Normally, IAC No. 1 powers the pilot PFD
and MFD; the No. 2 IAC powers the copilot PFD.
• Both IACs contain a sensor interface, flight director computer, and
symbol generator. Only the No. 1 IAC contains the autopilot computer.
• HDG, ATT, and ADC REV buttons enable the respective IAC to utilize
the other IACs AHRS or MADC data in the event of failure, thereby
providing redundancy.
• The SG1/MFD/SG2 selector on the MFD controller allows either IAC

to power all three displays in the event of IAC or symbol generator
failure.
Primus 880 weather radar:
• X-band alphanumeric digital radar with a display designed to provide

weather location and analysis, as well as ground mapping.
• Can be operated in conjunction with the EFIS and MFD equipment to

provide radar video displays.
• The radar transmitter is normally disabled on the ground. However, by

rapidly depressing the STAB switch four times within three seconds,
the transmitter radiates on the ground.
Primus II radio system:
• Dual remote radio management units (RMUs). RMU 1 is powered by

the emergency DC bus and RMU 2 is powered by main DC power.
• COM 1, NAV 1, ADF 1, etc., are controlled by the left RMU. COM 2,
NAV 2, ADF 2, etc., are controlled by the right RMU.
• VHF COMM is provided by the RZC-850 integrated communications

unit. Operates in the frequency range of 118.00 to 136.97 MHz, and
can be strapped to extend the upper frequency range to 152 MHz. It is
8.33 khz spacing capable.
• VHF NAV is provided by the RZC-850 integrated navigation unit.

Operates in the frequency range of 108.00 to 117.95 MHz. The system
encompasses the functions of VHF NAV, localizer and glide slope

receiver, and marker beacon receiver, as well as ADF and DME

functions.
• ADF NAV is provided by the DF-850 ADF receiver module, a

component of the RNZ-850 integrated navigation unit. Operates in the
frequency range of 100.00 to 1799.00 kHz in 0.5 increments.
• ATC TRANSPONDER function is provided by the XS-850

transponder module, a subunit of the RCZ-850 integrated
communication unit. It functions as a 4096 code mode A transponder,
as well as providing mode C (altitude) and mode S (collision
avoidance) information. Altitude information is provided by the
respective (1 or 2) AZ-850 micro air data computer in the pilot or
copilot Primus 1000 system.
• DME NAV function is provided by the optional Primus ll DME system

module. Each module is comprised of an RNZ-850 integrated
navigation unit, an NV-850 VHF NAV receiver and a DME-850
distance measuring module. The DME transmitter works in the L
frequency band and the receiver frequency range is from 962 to 1213
MHz. Normal DME function follows the VHF NAV receiver. However,
a hold function allows the tuning of military TACAN channels in order
to receive the DME portion of the TACAN signals. DME data is
displayed on two DI-850 indicators; one on the pilot and one on the
copilot instrument panels. DME data can also be displayed on the pilot
and copilot EHSIs.
• The STANDBY RADIO CONTROL (SRC) is normally on the center

instrument panel, to the right of the engine gauges. It contains normal
and emergency modes. The SRC is powered from the emergency DC
bus through the NAV1 circuit breaker. It acts as an additional tuning
source for the radio system (COM1 and NAV1).
Radio altimeter:
• The Collins ALT-55B radio altimeter displays radio altitude up to an

absolute altitude of 2,500 feet. Altitude is displayed on the bottom
center of the attitude sphere of the EADIs. Between 200 and 2,500
feet, the display is in 10-foot increments. Below 200 feet, it is in 5-
foot increments.
• Decision height (DH) selection is displayed digitally in the lower right

side of the EADI display. The decision height range is from 0 to 990
feet in 10-foot increments. The DH display can be removed with full
counterclockwise rotation of the DH/TST knob on the DC-550 display
controller. A decision height warning horn sounds when the airplane
reaches the decision height set on the pilot EADI.

Autopilot (AP):
• The autopilot and yaw damper are engaged by depressing the

AP–ENGAGE switchlight. With the flight director OFF, pitch and roll
are manually controlled with the turn knob and pitch wheel.
• Flight director (FD) mode(s) selected, the FD controls the autopilot.

• The autopilot may be switched to the pilot FD/PFD 1 or copilot
FD/PFD 2 by means of an illuminated selector switch (FD/AP–PFD1,
FD/AP–PFD2) on the center instrument panel.
• If a lateral mode is engaged, the autopilot holds the heading existing at

the moment of lateral mode disengagement.
• The autopilot/flight control system contains pitch, roll and yaw servos
that control the aircraft in accordance with manual or FD guidance to
the autopilot.
• The Primus 1000, IAC No. 1 contains the autopilot module for autopilot
control, consequently, if IAC No. 1 fails, the autopilot is inoperative.
• The autopilot may be temporarily disengaged by the touch control

steering (TCS) button on the yoke(s) but the yaw damper remains
engaged.
• The autopilot is normally disengaged one of three ways: 1) depressing

the AP/TRIM DISC red switch on either yoke; 2) electrically trimming
the elevator (yaw damper remains engaged); 3) depressing the go-
around button on either throttle.
• LOW bank limit may be selected manually by depressing the BANK

LIMIT–LOW switch on the controller (limits bank angle to 14°). Low
bank limit automatically engages climbing through 34,000 feet and
automatically disengages descending through 33,750 feet.
Flight director (FD):

• The P-1000 system incorporates one flight director in each IAC. A
single FD control panel on the center instrument panel is used to
control the FDs. Either crew member may control the aircraft through
the control panel by switching control with the FD/AP–PFD 1 or 2
selector as discussed above. If the FD/AP control is switched from one
pilot to the other, the AP reverts to basic mode. The FD must be
reprogrammed.
NOTE
When the FD/AP is coupled to the VOR, another lat-
eral mode must be selected prior to switching VOR
NAV frequencies. HDG mode may be used after syn-
chronizing HDG bug to the current airplane heading.
Basic ROLL may also be used.

Stall warning and AOA system:
• The angle-of-attack system is powered by 28 VDC from the left main

DC bus and incorporates an angle-of-attack sensor, a signal summing
unit, a vane heater monitor, an angle-of-attack indicator, a stick shaker,
and an indexer.
• The full-range-type indicator is calibrated from 0.1 to 1.0 and marked
with red, yellow, and white arcs. Lift being produced is displayed as a
percentage and, with flap position information, is valid for all airplane
configurations and weights. The area at the lower part of the scale (0.57
to 0.1) represents the normal operating range, except for approach and
landing. The narrow white arc (0.57 to 0.63) covers the approach and
landing range, and the middle of the white arc (0.6) represents the
optimum landing approach (VAPP or VREF). The yellow range (0.63 to
0.85) represents a caution area where the airplane is approaching a
critical angle-of-attack. The red arc (0.85 to 1.0) is a warning zone that
represents the area just prior to stick shaker activation and continuing
to full stall. At an indication of approximately 0.79 to 0.88 (depending
on flap setting and rate of deceleration) in the warning range, the stick
shakers activate.
• If the angle-of-attack system loses power or becomes inoperative for

other reasons, the needle deflects to the top of the scale and stows at a
1.0 indication.
NOTE
The airplane must not be flown if the stick shaker is
found to be inoperative on the preflight check or if
the angle-of-attack system is otherwise inoperative.
• Stick shakers are installed on the pilot and copilot control columns and
provide tactile warning of impending stall. The angle-of-attack
transmitter causes the stick shakers to be powered when the proper
threshold is reached.
WARNING
If the angle-of attack vane heater fails and the vane
becomes iced, the stick shaker may not operate or may
activate at normal approach speeds.

• The approach indexer, mounted on the pilot glareshield, provides a

heads-up display of deviation from the approach reference. The display
is in the form of three illuminated symbols used to indicate the airplane
angle-of-attack:
• When the airplane speed is on reference, the green center circle is

illuminated.
• As the speed decreases from reference (.6), the circle illumination

dims and the top red chevron illumination increases until the top
chevron is illuminated and the circle is extinguished. The top red
chevron points down, indicating that the angle-of-attack must be
decreased to eliminate the deviation.
• When the airplane is accelerating from the on-speed reference, the

illumination of the green circle dims and illumination of the
bottom yellow chevron increases until the circle is extinguished
and only the bottom chevron is illuminated. The bottom yellow
chevron points up to indicate that the angle-of-attack must be
increased to eliminate the deviation.
• The indexer is active any time the nose gear is down and locked
and the airplane is not on the ground. There is a 20-second delay
after takeoff before the indexer activates.
• Stall strips on the leading edge of each wing create turbulent

airflow at high angles of attack, causing a buffet to warn of
approaching stall conditions. This system is considered a backup
to the angle-of-attack stick shaker system in case of malfunctions
and electrical power failure.
Emergency instruments:
• Secondary flight display (SFD) (GH-3000)—see “Standby

Instruments” in the XLS Systems Review chapter in this manual.
• Secondary flight display (SFD). (Meggitt)
• The tube is a DC-powered cathode ray tube indicator combining

standby attitude indicator, altimeter, and airspeed/Mach
indications into one composite instrument (Figure SRE-46).
• The secondary flight display (SFD) contains solid state inertial

sensors for the measurement and presentation of aircraft pitch and
bank attitudes. The attitude display has an instantaneous display
range of 360° of bank and 50° of pitch. A moving tape on the right
side of the display includes a rolling digit display of altitude
calibrated in 100-foot increments. Baro data is set in the altitude
display by a knob on the bottom right side of the bezel. The setting is
displayed millibars at the top right of the display and in inches of
mercury at the bottom right of the display. On the left side of the
display, is a moving tape showing airspeed. Airspeed becomes active
at 40 knots.

MACH DISPLAY APR DISPLAY
VMO TAPE
BARO
DISPLAY IN
M. 457 ILS 1013 HP HECTO

280 PASCALS
500
260 10 10
AIRSPEED ALTITUDE
TAPE 250 70 00 TAPE
COURSE 240 10 10
GLIDESLOPE
INDICATORS
20 20
220
500 BARO
29.92 IN DISPLAY IN
INCHES Hg
BARO
APR ATT
BAROMETRIC SETTING
SRE-93
Figure SRE-46. Standby Flight Display—Meggitt

• The Mach number is displayed in the upper left corner of the

display and has a range of 0.35 to 0.999 Mach.
• Failure flag indications for airspeed and altitude are red crosses
covering the appropriate tape box, with all indications removed
from within the box. The failure flags for the Mach indication and
baro setting are a series of four red dashes in the appropriate
display area.
• The navigation display is selected by the APR button on the

bottom of the display bezel. Pressing the button once displays ILS
localizer information and glide slope flight director information
on the tube, provided the NAV 1 receiver is tuned to an ILS.
Pressing the button a second time displays back course localizer
information on the tube, provided the NAV 1 receiver is tuned to a
localizer back course frequency. Pressing the button a third time
removes all navigation information from the tube. VOR tracking
information is not available on the tube, but is available on the
standby HSI.
• Power to the standby flight display system is controlled by a

switch marked STBY PWR–ON/OFF/TEST on the pilot lower
instrument panel. The SFD has an emergency source of power
from an emergency battery pack in the right nose avionics
compartment. This battery pack also provides emergency
instrument lighting for the dual fan (N1) tachometers and the
standby horizontal situation indicator (HSI). The battery pack is
constantly charged by the airplane electrical system. The standby
instrument power switch must be ON for automatic transfer to its
emergency battery pack. This battery pack supplies a minimum of
30 minutes of power to the SFD. An amber STBY PWR ON light
next to the STBY PWR switch illuminates when the emergency
battery pack is powering the SFD.
• Standby horizontal situation indicator (HSI):
• The standby HSI is a three-inch instrument on the pilot instrument

panel, directly below the tube. It provides navigational guidance
in case of PFD/flight director failure, and is powered by the
emergency bus (Figure SRE-47).
• The standby HSI displays compass heading (No. 2 AHRS), and

navigation inputs from NAV 1, (i.e., glide slope, localizer
deviation, and airplane position relative to VOR radials). The
compass card is graduated in 5° increments, and a lubber line is
fixed at the fore and aft positions. A fixed reference airplane is in
the center of the HSI, aligned with the lubber line markings. In
addition, there is a course deviation bar and course cursor, as well
as a blue ADF needle, which displays ADF 1 bearings and rotates
around the outer portion of the dial (not available with loss of
normal DC power).

33 N
3
N
30
A
V
6
W
E
24
V
E
R
12
T
21 15
S
CRS ADF
Figure SRE-47. Standby HSI
• Magnetic compass:
• A standard liquid-filled magnetic compass is mounted above the

glare shield. Directly above the compass are the seating height
indicator balls.
Digital clock (Davtron):
• Two Davtron model M877 clocks on the pilot and copilot upper
instrument panels can display four functions: local time, GMT, flight
time, and elapsed time. Two versions of elapsed time may be selected:
count up or count down.
• The clock has two control buttons: SEL (select) and CTL (control). The
SEL button is used to select the desired function and the CTL button is
used to start and reset the selected mode.
• The flight time mode of the clock is enabled by a landing gear squat
switch, which causes the clock to operate any time the airplane weight
is off the landing gear. The flight time may be reset by the pilots.

Static wicks:
• A static electrical charge, commonly referred to as P (precipitation)

static, builds up on the surfaces of the airplane in flight and causes
interference in radio and avionics equipment operation. The static
wicks are installed on the wing and empennage trailing edges, and
dissipate static electricity in flight.
• There are a total of 20 static wicks:
• One on each wingtip.

• Four on each wing trailing edge outboard of the aileron.
• One on the trailing edge of each aileron.
• Two on the trailing edge of each elevator.
• Two on the upper trailing edge of the rudder.
• One on the top of the rudder.
• One on the tail stinger.
• One or more missing static wicks cause radio P-static.
• Some static wicks may be missing for dispatch. Refer to AFM
“Normal Procedures” or MEL for conditions.
TCAS ll (optional):
• TCAS ll detects and tracks aircraft in the vicinity of your own airplane.
It interrogates the transponders of other aircraft and analyzes the
signals to range and bearing, and relative altitude if it is being reported.
It then issues visual and aural advisories so that the crew may perform
appropriate vertical avoidance maneuvers. TCAS control is provided
through the RMUs.
EGPWS (optional):
• The optional Allied signal enhanced ground proximity warning system

(EGPWS) provides visual and aural warnings of terrain in the
following basic GPWS modes:
1. Excessive rate-of-descent with respect to terrain (Mode 1).
2. Excessive closure rates to terrain (Mode 2).
3. Negative climb before acquiring a predetermined terrain clearance

after takeoff or a missed approach (Mode 3).
4. Insufficient terrain clearance based on flap configuration (Mode 4).

5. Inadvertent descent below glide slope (Mode 5).
6. Minimums callout upon reaching DH (Mode 6).
7. SMART 500 callout—Altitude callout at 500 AGL (Mode 6).
8. Excessive bank angle alerting (Mode 6).
9. Windshear warning and windshear caution alerts (Mode 7).
• In addition, the enhanced ground proximity warning system provides

the following terrain map enhanced modes:
1. Terrain clearance floor exceedance.
2. Look-ahead cautionary terrain alerting and warning awareness.
3. Terrain awareness display. EGPWS provides display of

approximate terrain and obstacles. The terrain display is color and
intensity-coded (by density) to provide visual indication of the
relative vertical distance between the airplane and the terrain.
Area Navigation:
• Universal avionics systems UNS-1 Csp flight management system

(FMS) is a centralized control and master computer system, designed
to consolidate and optimize the acquisition, processing, interpretation
and display of certain aircraft navigation and performance data. The
UNS-1 Csp FMS system may be installed as GPS only or multisensor
system. Digital air data information (including baro-corrected altitude
and true airspeed) and heading input is required of all installations.
• Each individual navigational sensor is specifically designed for primary

navigation. The FMS system takes advantage of a particular sensor
good properties while minimizing its liabilities. The system processes
multiple range information from the DME, True airspeed data from the
air data computer, velocity and position information from the long
range navigation sensors, and aircraft heading, in order to derive one
best computed position (BCP).
• The FMS contains a memory capacity of up to 100,000 waypoints. The

stored Jeppesen data base provides the capacity for complete coverage for
SIDs, STARs, approaches, high/low airways, Navaids, IFR intersections
and airports with runways longer than 4,000 feet with IFR approaches in
the worldwide data base. It provides the capability for:
1. Pilot data storage
2. Company route data
3. Off line flight planning

4. Fuel management monitoring
5. Frequency management
6. Lateral guidance and steering
7. Vertical guidance—VNAV (with FD/AP coupling authority)
• The Honeywell FMZ is optional.
Locator beacon:
• The ELT 110-4 emergency locator transmitter (ELT) provides a

modulated omnidirectional signal, transmitted simultaneously on
emergency frequencies 121.50 and 243.00 MHz. The system is
activated by an impact of 5.0 +2/–0 g, or manually by a remote
ON–OFF switch forward of the pilot CB panel.
ANTENNA AND DRAIN TUBE

Antenna and drain tube locations are shown in Figure SRE-48.

NAV 1 & 2
FLUX VALVE
ADF 2 ACM AIR INLET

LOCATOR BCN
(OPTIONAL) COMM 1 APU EXHAUST
DIVERSITY TRANSPONDER 1 (LH SIDE) RH SIDE
DIVERSITY TRANSPONDER 2 (RH SIDE) LIGHTNING

ADF 1 DETECT APU AIR INLET
TCAS II UPPER RH SIDE
(OPTIONAL) SATCOM HF
GPS 1
GPS 2
(OPTIONAL)
RADAR
12 INCH
STORMSCOPE
(OPTIONAL)
ACM EXHAUST RH SIDE
APU FUEL DRAIN
TAILCONE FRESH AIR INLET RH SIDE
GLIDESLOPE ENGINE DRAIN
DME2 AFIS
DME1 BATTERY VENT
RADAR ALTIMETER HYDRAULIC RESERVOIR DRAIN
MARKER BEACON MAGNASTAR
TRANSPONDER 1 FWD LAVATORY
DRAIN REAR LAV / CONDENSER DRAIN
TCAS II LOWER
RADAR ALTIMETER
COM2
TRANSPONDER 2
GEAR BLOWDOWN
VENT
SRE-99
Figure SRE-48. Excel Antenna and Drain Tube Locations

SYSTEMS REVIEW—XLS
CONTENTS
Page
SQUAT SWITCH INPUTS............................................................. SRX-1
Left Main Squat Switch Only................................................ SRX-1
Left and Right Squat Switches in Parallel............................. SRX-2
COMPLETE GENERATOR FAILURE CONDITION .................. SRX-2
LIGHTING ..................................................................................... SRX-4
Cockpit Panel Lights ............................................................. SRX-4
Cockpit Overhead Lights....................................................... SRX-4
Cabin Lighting....................................................................... SRX-5
Cabin Emergency Lighting (EMER LTS) ............................. SRX-7
Exterior Lights....................................................................... SRX-8
Tail Cone Compartment Lights ............................................. SRX-8
Pulselite System .................................................................... SRX-9
ELECTRICAL SYSTEM ............................................................. SRX-10
General ................................................................................ SRX-10
POWERPLANT............................................................................ SRX-18
General ................................................................................ SRX-18
Ignition ................................................................................ SRX-20
FIRE PROTECTION .................................................................... SRX-21
Sensing Loops and Control Units ....................................... SRX-21
Operation............................................................................. SRX-22
FUEL ............................................................................................ SRX-24
HYDRAULICS............................................................................. SRX-28
POWER BRAKES AND ANTISKID........................................... SRX-39
EMERGENCY BRAKES............................................................. SRX-41
FLIGHT CONTROLS .................................................................. SRX-41
ICE AND RAIN PROTECTION .................................................. SRX-48
PNEUMATICS/AIR CONDITIONING....................................... SRX-58
PRESSURIZATION ..................................................................... SRX-62
SERVICE AIR .............................................................................. SRX-68
FOR TRAINING PURPOSES ONLY SRX-i

OXYGEN...................................................................................... SRX-70
AUXILIARY POWER UNIT (APU) ........................................... SRX-72
Electronic Control Unit (ECU) ........................................... SRX-72
Fuel System......................................................................... SRX-73
Oil System........................................................................... SRX-73
Pneumatic System ............................................................... SRX-74
Electrical System................................................................. SRX-75
Fire Protection..................................................................... SRX-78
Exterior Preflight................................................................. SRX-78
APU Control Panel and Annunciator Functions ................. SRX-79
APU OPERATING LIMITATIONS ............................................. SRX-88
Battery and APU Starter Cycle Limitations ........................ SRX-89
AVIONICS.................................................................................... SRX-90
Integrated Avionics Computers (IAC):................................ SRX-90
Comparison Monitor Annunciators (9) ............................... SRX-92
Primus 880 Weather Radar.................................................. SRX-93
Primus II Radio System ...................................................... SRX-93
Radio Altimeter ................................................................... SRX-94
Autopilot (AP)..................................................................... SRX-94
Flight Director (FD): ........................................................... SRX-95
Stall Warning and AOA System: ......................................... SRX-95
Standby Instruments............................................................ SRX-97
Clocks.................................................................................. SRX-99
TCAS II ........................................................................................ SRX-99
Terrain Awareness and Warning System (TAWS):.............. SRX-99
Area Navigation ................................................................ SRX-100
Locator Beacon ................................................................. SRX-101
Static Wicks ...................................................................... SRX-101
ANTENNA AND DRAIN TUBE .............................................. SRX-102
ILLUSTRATIONS
Figures Title Page
SRX-1 Cabin/Entry Lights Panel .......................................... SRX-6
SRX-ii FOR TRAINING PURPOSES ONLY

SRX-2 DC Power Distribution ............................................ SRX-11

SRX-3 Left CB Panel .......................................................... SRX-13
SRX-4 Right CB Panel ........................................................ SRX-14
SRX-5 PW545B Cross-Section .......................................... SRX-19
SRX-6 Engine Fire Extinguishing System .......................... SRX-22
SRX-7 Engine Fire Detection System ................................ SRX-23
SRX-8 Fuel System—Normal Operation ............................ SRX-25
SRX-9 Fuel System—Crossfeed (R to L)............................ SRX-27
SRX-10 Hydraulic System-Open Center .............................. SRX-29
SRX-11 Speedbrake System-Normal Operation (Extended) SRX-31
SRX-12 Gear System-Normal Retraction.............................. SRX-33
SRX-13 Gear System-Normal Extension .............................. SRX-34
SRX-14 Gear System-Emergency Extension ........................ SRX-35
SRX-15 Thrust Reversers-Stowed .......................................... SRX-37
SRX-16 Thrust Reversers-Deployed...................................... SRX-38
SRX-17 Power Brake/Antiskid System ................................ SRX-40
SRX-18 Flight Controls ........................................................ SRX-42
SRX-19 Rudder Bias System ................................................ SRX-44
SRX-20 Rudder Bias System—Engine Failure .................... SRX-44
SRX-21 Two-Position Horizontal Stabilizer.......................... SRX-46
SRX-22 Pilot-Static System .................................................. SRX-49
SRX-23 Windshield Anti-Ice System .................................... SRX-51
SRX-24 Wing/Engine Anti-Ice System ................................ SRX-53
SRX-25 Wing Leading Edge Cross Section .......................... SRX-55
SRX-26 Tail Deice System .................................................... SRX-57
SRX-27 Engine Bleed-Air Precooler .................................... SRX-59
SRX-28 Air Conditioning System ........................................ SRX-61
SRX-29 Pressurization System .............................................. SRX-63
SRX-30 Pressurization Control Panel .................................... SRX-64
SRX-31 Auto Schedule Boundary ........................................ SRX-65
SRX-32 High Altitude Landing Graph .................................. SRX-66
SRX-33 High Altitude Departure Graph................................ SRX-67
FOR TRAINING PURPOSES ONLY SRX-iii

SRX-34 Service Air System .................................................. SRX-69

SRX-35 Oxygen System ........................................................ SRX-71
SRX-36 APU Annunciators—Copilot Panel ........................ SRX-74
SRX-37 APU Control Panel .................................................. SRX-76
SRX-38 First Engine Start (R)—APU Generator On Line.... SRX-77
SRX-39 APU Start On Ground (Engines OFF) .................... SRX-82
SRX-40 APU Start On-Ground—GPU (EPU) ...................... SRX-83
SRX-41 Second Engine Start (L)—APU Generator On Line SRX-84
SRX-42 APU Start On Ground (Generator Assist)................ SRX-86
SRX-43 APU Start—In Flight (Battery Only) ...................... SRX-87
SRX-44 XLS Primus 1000 CDS System Block Diagram...... SRX-91
SRX-45 Standby Flight Display—GH 3000.......................... SRX-97
SRX-46 Standby HSI ............................................................ SRX-98
SRX-47 Excel Antenna and Drain Tube Locations ............ SRX-102
TABLES
Tables Title Page
SRX-1 APU Operating Limits ............................................ SRX-88
SRX-2 Comparison Monitor Annunciators (9).................... SRX-92
SRX-iv FOR TRAINING PURPOSES ONLY

SYSTEMS REVIEW—XLS
SQUAT SWITCH INPUTS
LEFT MAIN SQUAT SWITCH ONLY
In flight, it enables:
1. Flight hour meter
2. Digital clock to record elapsed flight time
3. Opening of emergency pressurization valve
4. Landing gear handle locking solenoid to be energized
5. TAS probe heater (Rosemount)
6. Enables flight idle (EECs operating)
7. Normal (auto) control of pressurization
On the ground, it enables:
1. Pressurization controller opens outflow valves (<85% N2)
2. Prepressurization during takeoff
3. Generator-assisted starts
4. Engine bypass valve for precooler operation
5. Ground idle (with EECs operating)
6. Overrides the pulselite system to steady illumination
7. Noncancelable master caution for LO BRK PRESS annunciator
The safe-flight angle-of-attack indexer lights on the glareshield are enabled

any time the airplane nose gear is down and locked and the airplane is not on
the ground (enables indexer 20 seconds after takeoff with nose gear down).
FOR TRAINING PURPOSES ONLY SRX-1

LEFT AND RIGHT SQUAT SWITCHES IN PARALLEL

In flight, they enable:
1. Stick-shaker operation
On ground, they enable:
1. Thrust reverser deployment (either squat switch)
2. Stick-shaker test
3. Antiskid locked wheel protection (both squat switches)
4. Stabilizer/flap switch miscompare
COMPLETE GENERATOR FAILURE CONDITION

The following components and systems are operative when only the battery
bus, emergency buses, and standby instruments are powered (GENs off line,
BATT switch EMER, STBY PWR switch ON):
1. Cockpit auxiliary panel and overhead floodlights
2. Interior entry lights
3. Cabin emergency lights (forward and aft emergency battery packs power
emergency lights when airplane battery voltage drops below emergency
pack voltage)
4. LH engine display indicator (AMLCD)
a. N1 (digits and tapes)
b. N2 (digits)
c. ITT (tapes)
d. Fuel flow (digits)
e. Oil temperature (digits)
f. RAT (digits)
5. Standby pitot and static heaters
6. Standby flight display (standby emergency battery pack)
7. Backlighting for standby HSI (standby emergency battery pack)
8. Standby HSI
SRX-2 FOR TRAINING PURPOSES ONLY

9. AHRS 2
10. Hydraulic control valve
11. Flap control
12. Landing gear control, indication, and warning
13. Auxiliary gear control (manual)
14. Emergency brakes (pneumatic)
15. Two-position stabilizer control
16. Pilot and copilot audio panel
17. Standby radio control head
18. RMU 1
19. NAV 1
20. COMM 1
21. LH and RH start PCBs
22. LH and RH secondary ignition
23. Voltmeter (BATT switch in ON or EMER position)
24. Manual pressurization control (“Cherry Picker”)
25. Cabin altitude/differential pressure indicator
26. Passenger oxygen valve (OFF and ON only)
27. Oxygen pressure gauge
28. Engine and wing anti-ice (bleed air with no cross-flow capability)
29. ELT

LIGHTING
COCKPIT PANEL LIGHTS
Panel lights are controlled by the master panel ON–OFF toggle switch
(DAY–NIGHT) on the pilot lower instrument panel (PANEL LIGHT).
With the master switch ON, the following rheostats control light intensity:
• LEFT DIM—Voltmeter, left and right ammeters, left PFD display

controller, audio panel 1, left FD mode selector.
• CENTER DIM—Standby HSI, DAVTRON clock, pressurization

controller, differential pressure gauge, engine instruments (AMLCD),
aileron and rudder trim tab indicators, radar control panel, MFD
controller, FMS panels, standby radio control, RMU 1 and 2 controls,
ECS digits, center console controllers.
• RIGHT DIM—Oxygen gauge, hourmeter, battery temperature

indicator, right PFD display controller, cockpit voice recorder, APU
ammeter, APU right panel, right FD mode selector, audio panel 2.
• EL DIM—Throttle quadrant, pilot and copilot switch panels, light

switch panels, tilt panel, thrust reverser switch panels, landing gear
control panel, left and right CB panels, left and right sidewall
subpanels, (electrical power to the EL panels is supplied by a small
40–60 VAC inverter in the nose avionics compartment).
NOTE
Placing the day/night switch ON, dims the annunci-
ator panel, stand alone annunciators/switches, and il-
luminates two red windshield ice detection post lights.
The following rheostats are powered directly from the
emergency bus (not connected to the master
DAY/NIGHT switch):
• G L A R E S H I E L D AU X I L I A RY L I G H T S —
Rheostat on the pilot sidewall subpanel.
• FLOODLIGHTS—Rheostat on the pilot lower

switch panel.
COCKPIT OVERHEAD LIGHTS

Two map/chart lights controlled by left and right rheostats are on the pilot
and copilot sidewall subpanels.
Three emergency DC powered floodlights (two in the overhead and one on

the bottom of the annunciator panel assembly) illuminate the forward cock-
pit area and the engine instruments respectively. They all illuminate or ex-
tinguish simultaneously by rotating the FLOOD rheostat on the pilot lower
instrument panel.

CABIN LIGHTING
Cabin lighting consists of overhead reading lamps, overhead indirect lights,
aft vanity lights, forward and aft divider lights, NO SMOKING/FASTEN
SEAT BELT and EXIT lights, dropped aisle footwell lighting, and forward
work station lights.
Reading Lights
Directionally adjustable reading lamps are above each seat, including the aft
vanity seat. Thy are controlled by switches adjacent to the outboard arm rests.
Indirect Lights
Cabin overhead indirect lighting is controlled by a switch on a cabin light switch
panel on the forward cabin entry door frame. Initially pushing the switch il-
luminates the lights “bright.” After a few seconds, pushing the switch again
dims the lights. The next push extinguishes the lights. The lights may also be
controlled by a CABIN LIGHT switch on the galley light panel.
Dropped Aisle Lighting

Strip lights on both sides of the footwell aisle are normally controlled by the
AISLE lighting switch on the galley light panel. A portion of each strip light
illuminates when the emergency exit lights are activated (refer to EMER-
GENCY EXIT LIGHTS later in this chapter).
Divider, Aft Closet, and Work Station Lights

Divider and work station lights are controlled by switches on the galley light
panel. The aft closet light activates when the closet door is opened and ex-
tinguishes as the closet door is closed.

Cabin/Entry Lights Panel

Cabin entry lights are powered directly from the battery bus and activate ON
and OFF with ENTRY LIGHT switches on the cabin entry door forward frame
light panel and the galley light panel (Figure SRX-1). Entry lights consist of
four reading lamps in the passenger compartment (one above the forward and
aft seats on each side) and a light above the emergency exit door in the aft
vanity area. Door entry lights on either side of the lower door frame and six
lights in the door steps are also included.
CABIN
LIGHT
ENTRY
LIGHT
LEFT UPPER DOOR RIGHT UPPER

DOOR PIN LOCKS DOOR PIN
LEFT LOWER RIGHT LOWER

DOOR PIN DOOR PIN
DOOR
SEAL
CABIN
DOOR
VENT DOOR DOOR HANDLE
Figure SRX-1. Cabin/Entry Lights Panel
Passenger Information Signs

The PASS SAFETY switch on the tilt panel controls passenger information
signs in the cabin as follows:
• PASS SAFETY–ON: Sounds an audible chime and illuminates the

SEATBELT ON/NO SMOKING signs. Illuminates all emergency exit
lights if the EMER LTS switch is ARMED.
• SEAT BELT–ON: Sounds an audible chime and illuminates only the

SEATBELT ON signs.

EMERGENCY LIGHTING (EMER LTS)

Emergency exit lights consist of cabin lights, emergency exit illumination and
identification, emergency egress, and ground illumination for emergency
evacuation. Emergency lighting is controlled by the EMER LTS ARM–ON–OFF
switch on the pilot instrument panel. Emergency lighting may be powered by
main DC power or power supplied from two emergency nicad battery packs
(one behind the pilot left side console and one aft in the vanity area). Placing
the EMER LTS switch ON illuminates all emergency exit lights, both inte-
rior and exterior. Placing the switch to ARM illuminates the lights if a 5
gravity (g) force is experienced, or a loss of main DC power occurs (i.e., dual
generator failure with BATT switch in EMER). Emergency lighting also il-
luminates when the PASS SAFETY switch is ON. An amber light adjacent to
the switch illuminates with the switch OFF and main DC power available (BATT
switch ON) to alert the crew to ARM the system prior to flight.
Interior Lights
• Three reading lamps in the passenger compartment (one forward right
side, one forward and one aft on left side)
• One reading lamp above the emergency exit door in the vanity area
• All EXIT sign lighting
• Right and left footwell (dropped aisle) strip lights illuminate to direct
occupants to the exit doors (partial illumination only)
Exterior Lights
• Two lights in the right fuselage illuminate at the top of the right wing.
• One light in the right fuselage forward of the wing root illuminates the
ground in front of the wing.
NOTE
The forward emergency nicad battery pack provides
emergency power to illuminate the exit sign over the
cabin door, the reading light opposite the cabin door,
the reading light just aft of the cabin door, and the
dropped aisle strip lighting on left side. The aft emer-
gency nicad battery pack provides emergency power
to illuminate the exit signs on the aft divider and
above the emergency exit door, an overhead light
above the emergency exit door, the reading lamp on
the left rear of the passenger compartment forward
of the aft divider, right dropped aisle strip lighting,
and the three exterior emergency lights.

EXTERIOR LIGHTS
Navigation
Navigation lights are wing tip lights (red–left, green–right) and a white light
on the tail stinger, controlled by the NAV ON–OFF switch on the tilt panel.
Anticollision
Anticollision lights are high intensity white pulsating strobes on each wingtip,
controlled by the GND REC/ANTI COLL ON–OFF switch on the tilt panel.
Ground Recognition
The ground recognition light is a red beacon light on the top of the rudder. It
is controlled by the GND REC/ANTI COLL ON–OFF switch on the tilt panel.
Wing Inspection
Wing inspection lights are on both sides of the fuselage forward of the wing
leading edges. When activated, they illuminate the entire leading edges of both
wings. The lights are controlled by the WING INSP ON–OFF switch on the
tilt panel within the ANTI ICE/DEICE switch section.
Landing/Recognition/Taxi Lights
Landing and recognition lights are side by side on each forward wingtip light
assembly. The landing lights are installed outboard and are canted downward
slightly. The inboard recognition lights illuminate directly ahead.
Two fixed position seal beam taxi lights are in the belly fairings on each side
of the fuselage. The lights also supplement the landing lights.
Control
Landing, recognition, and taxi lights are all controlled by individual ON–OFF
switches on the center pedestal. The switch positions function as follows:
• REC/TAXI–ON (down): “Belly” taxi lights and recognition lights ON.
• OFF: All landing, taxi, and recognition lights OFF.
• LANDING LIGHTS—ON: All landing, taxi, and recognition lights ON.

TAIL CONE COMPARTMENT LIGHTS

Tail Cone Maintenance Compartment Light
The tail cone maintenance compartment light is powered from the battery bus
and controlled by an ON–OFF switch in the right forward side of the com-
partment (adjacent to the electrical J-box). If the light is inadvertently left
on, closing the access door extinguishes the light.
Baggage Compartment Light

The baggage compartment is illuminated by three lights: two in the overhead
ceiling and one in the sidewall. The lights are powered from the battery bus
and controlled by a switch in the forward access door closeout assembly. If
the light is inadvertently left on, closing the door extinguishes the light.
PULSELITE SYSTEM
The Precise Flight, Inc. Automatic Pulselite System allows the taxi (belly)
and recognition lights to pulse. The taxi and recognition lights automatically
pulse when both REC/TAXI switches are in the ON position (down) and the
airplane is airborne. Positioning either or both switches to LANDING
LIGHTS–ON deactivates the system and all landing, recognition, and taxi
LANDING, RECOGNITION and TAXI lights revert to steady illumination.
The pulselight switch next to the LANDING/RECOGNITION/TAXI light
switch overrides the squat switch to allow pulsing of the TAXI and RECOG-
NITION lights on the ground. The switch must be ON and both REC/TAXI
LIGHTS must be selected ON for the lights to pulse, airborne or on the
ground. Refer to “Supplement 2, Precise Flight—Automatic Pulselite System”
in the Airplane Flight Manual (AFM) for detailed operating procedures.

ELECTRICAL SYSTEM
GENERAL
DC power distribution is shown in Figure SRX-2.
1. Battery switch (power distribution switch)—Controls the battery isolation

relay:
• OFF—Battery isolation relay is deenergized open and the relay

between the crossfeed and emergency buses remain deenergized
closed. The relay between the battery bus and the emergency buses is
deenergized open. The voltmeter is not connected to the battery bus, so
normal system voltage cannot be read.
• EMER—Battery isolation relay is deenergized open and the relay

between the crossfeed and emergency buses is energized open. The
relay between the battery bus and the emergency buses is energized
closed. The battery supplies power to the battery bus and emergency
buses. If the generators are online, they supply power to the feed and
crossfeed buses. The EMER switch position isolates the battery from
the generator charging power. The voltmeter indicates only battery
voltage, unless the VOLTAGE SEL switch is in the L–GEN or R–GEN
position.
• BATT—Battery isolation relay is energized closed and the relay

between the crossfeed bus and emergency buses is deenergized closed.
The relay between the battery bus and the emergency buses is
deenergized open. All power is normally supplied by the generators or,
if off line, by the battery. The voltmeter indicates system voltage.
2. Generator switches (left and right)—300 amp hr

• ON—Activates the GCU, which may close the power relay
• OFF—Disables the GCU, opens the power relay, not the field relay
• RESET—Momentarily resets the field relay
3. Generator control units (digital GCUs with BITE lights):

• The GCUs regulate 30 VDC generators to 28.5 VDC.
• The GCUs provide protection for generators and the electrical system.
• The GCUs parallel the generators to share the system load; the
generators must be within 0.3 volts and approximately 10% of system
load, not to exceed a 30-amp split.
• Each GCU in the tail cone has four red BITE lights (fault lights). The
GCU fault lights may indicate a GCU fault, overvoltage, a ground fault,
or a system problem. For a detailed fault, hold the generator switch in
RESET and match the arrays of LEDs to the list of fault IDs. The LEDs
self-test at power-up. Flashing LEDs can be extinguished by resetting the
generator switch three times within 3 seconds if no fault exists.

NORMAL OPERATIONS (GROUND OR INFLIGHT, APU GENERATOR OFF LINE, AVIONICS ON)
LEFT CIRCUIT BREAKER PANEL RIGHT CIRCUIT BREAKER PANEL
ENGINE START
L DISENGAGE R
EMER SYS AFT GEN EMER AVN
J-BOX OFF
START
SYS SYS LMT CB DISG L R AVN AVN
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER

A A
NOTE: SWITCH SWITCH E
BATT
RED BORDERS
ON 25A
DENOTE EMERGENCY P ON RELAY RELAY E RELAY ON I
U OFF R N
POWERED BUSES OFF OFF T
28.5 EMER 28.5 E
RESET 28.5 RESET R
B
LEGEND U
START START I
L GEN RELAY RELAY R GEN O
L GENERATOR S BUS BATTERY BUS BUS R
GCU GCU
APU 1
R GENERATOR 7
RELAY APU RELAY
L BATTERY
R 5
ENGAGED A
APU GENERATOR STARTER STARTER
GPU
BATTERY RELAY DISCONNECT VOLTAGE RELAY
RELAY
GPU INTERIOR POWER
INPUT
SRX-11
Figure SRX-2. DC Power Distribution

4. Voltmeter select switch:
• BATT—Voltage is read from the hot battery bus when the battery
switch is in the BATT or EMER position only; the switch is spring
loaded to the BATT position.
• LH or RH GEN—Voltage can be read from the respective generator

output to its power relay. Accurate generator voltage is normally
obtained with the selected generator off-line.
5. Emergency bus items:
Left CB panel (Figure SRX-3)
• Cockpit auxiliary panel lights
• Cockpit (overhead) floodlights
• LH engine display indicator (AMLCD)
• Standby pitot and static heaters
• Hydraulic control
• Flap control
• Landing gear control
• Landing gear warning
• Stab control
• LH secondary ignition
• RH secondary ignition
Right CB panel (Figure SRX-4)
• Standby HSI
• AHRS 2
• Audio panel 1 (switches to AVN EMER bus with BATT in EMER)
• Audio panel 2
• NAV 1
• COMM 1

ENVIRONMENTAL LIGHTS ENGINE INSTRUMENTS FLIGHT INSTRUMENTS
NORM MANUAL ANTI- WING GND COCKPIT L ENG L OIL L FUEL STBY STBY
PRESS TEMP NAV COLL INSP REC FLOOD PULSE DISPLAY PRESS QTY PWR ATT
5 3 5 71/2 5 5 5 1 5 2 2 15 5
EMER AUTO EL L CENTER R AUX R ENG R OIL R FUEL CENTER FLT HR

PRESS TEMP PANEL PANEL PANEL PANEL PANEL DISPLAY PRESS QTY CLOCK METER
5 3 3 5 5 5 3 5 2 2 2 2

ANTI-ICE ENGINE SYSTEMS

L L PITOT TAS AOA L STBY P/S L L F/W L FIRE L FUEL FUEL L
ENG/WING STATIC HEATER HEATER W/S HTR IGNITION L EEC SHUTOFF DETECT BOOST CONTROL START
71/2 71/2 10 10 5 10 71/2 71/2 71/2 2 15 5 2
R R PITOT TAIL R R ENG R F/W R FIRE R FUEL R

ENG/WING STATIC DEICE W/S IGNITION R EEC VIB MON SHUTOFF DETECT BOOST START
71/2 71/2 5 5 71/2 71/2 5 71/2 2 15 2

SYSTEMS
CABIN DOOR WARNING SMOKE L THRUST W/S FLAP GEAR PITCH SPEED BATT
MONITOR LTS 1 DETECTOR REVERSER AIR AOA CONTROL CONTROL TRIM BRAKE CVR TEMP
1/
5 5 7 2 15 5 5 5 5 5 5 71/2
DC POWER
LH FEED RH FEED NOSE WARNING RUDDER R THRUST GLRSHLD HYD STAB GEAR A/P AHRS AUX PWR SKID
1 2 3 1 2 3 WHL RPM LTS 2 BIAS REVERSER FANS CONTROL CONTROL WARN SERVO BATT CHG BRKS CONTROL
1/ 1/ 1/
50 50 50 50 50 50 5 5 7 2 7 2 3 5 5 7 2 15 15 5
LEGEND
LH MAIN DC BUS
RH MAIN DC BUS
EMERGENCY BUS
SRX-13
Figure SRX-3. Left CB Panel

SRX-14
AVIONICS
AUDIO WARN COMM NAV AHRS AHRS 1 XPDR DME ADF ADC IC PFD PFD MFD
1 AUDIO 1 1 1 1 AUX 1 1 1 WARN 1 1 CONT 1 1 1 RADAR
5 5 71/2 5 5 5 5 5 5 5 5 71 / 2 5 15 15 71/2
AUDIO WARN COMM NAV AHRS AHRS 2 XPDR DME ADF ADC IC PFD PFD MFD RADAR

2 AUDIO 2 2 2 2 AUX 2 2 2 PHONE 2 2 CONT 2 2 CONT CONT
1/ 1/
5 5 7 2 5 5 5 5 5 5 5 5 7 2 5 15 5 5
STBY GPS CABIN CABIN FMS FMS HF RAD LIGHTNING

HSI 3 BRIEFER DISPLAY 1 GND PWR DATALINK SATCOM CHECKLIST RESET HF TCAS FDR ALT DETECT
1/
5 5 5 10 2 5 5 5 5
SECURITY TAS/SAT FMS DATALINK SATCOM MFD

SYSTEM SELCAL IND TAWS 2 COMM HP A 2
5 5
DSPTCH SATCOM SATCOM

GND PWR ANTENNA R/T
APU AC POWER AVIONICS DC POWER

FIRE LIGHTNING DSPTCH GND LH FEED RH FEED
LEGEND MASTER DETECT ECU FDR DETECT AHRS PWR CTRL INVERTER 1 2 1 2
LH MAIN DC BUS 5 5 10 50 50 50 50
RH MAIN DC BUS
EMERGENCY BUS
Figure SRX-4. Right CB Panel

Placing the battery switch to either BATT or OFF causes the emergency
relays to relax, connecting the emergency buses to the crossfeed bus.
Placing the battery switch to EMER, energizes the emergency relay
connecting the emergency buses to the battery bus. With the generators
online, placing the battery switch to OFF does not cause loss of power to
the emergency buses (isolates battery from generators). Loss of both
generators requires the battery switch be placed in the EMER position,
which load sheds the main DC feed buses and allows the battery to power
only the battery bus and emergency buses. This extends the battery life to
approximately 30 minutes.
6. Battery bus items:
• Tail cone baggage compartment light
• Tail cone access compartment lights
• Cabin entry lights
• Voltmeter (BATT switch in ON or EMER only)
• Emergency buses (BATT switch in EMER only)
• LAV service lights
• Engine chip detector lights (if installed)
7. Crossfeed bus items (not inclusive):
• Cockpit recirculation fan
• LH and RH ignitor source
• LH landing lights
• LH landing/recognition lights

8. LH or RH GEN OFF annunciator light:
• Indicates the respective power relay is open.
• If voltage indicates near 0, the field relay is tripped; reset may be

possible.
• If voltage indicates normal, the power relay is open and the field relay
is not tripped; reset is not probable.
9. Current limiters (225 amp):
• Should open if respective generator is supplying in excess of 225 amps

through the current limiter (AFT J-BOX LMT).
• If failed, only the opposite generator can power the crossfeed bus.
• If failed prior to start, the engine on the side of the failed current limiter
cannot be started. The MASTER WARNING on the opposite side will
illuminate steady.
10. Goodrich standby flight display unit—Normally powered through the

STBY PWR CB on the left CB panel. If power is lost at the circuit
breaker, provided the STBY PWR switch is ON, then the standby flight
display battery pack (nose compartment) automatically provides power to
the following (amber light adjacent to switch on):
• Goodrich tube
• Back lighting for the standby HSI and left display (AMLCD)
11. Auxiliary avionics battery pack (right nose compartment):
• Auxiliary power supply for the AHRS 1 and 2
• Auxiliary power supply for the flight management system (on ground
only, through FMS GND PWR CB on right CB panel)
• For any of the above to occur, the STBY PWR switch must be ON

12. The emergency lighting system consists of a EMER LTS control switch,
two 18-cell rechargeable nicad battery packs, and various cabin interior
and exterior lights and signs. When activated, the emergency lighting
system will use main DC or battery power for system operation. If main
DC or battery power is unavailable or otherwise becomes depleted, the
system operates from its two battery packs for a minimum of 10 minutes.
The battery packs are constantly trickle-charged to capacity so they can be
available as a final source of emergency lighting power. Cabin emergency
lighting activates for the following reasons:
• ARM (automatic activation due to):
• Loss of main DC power with battery switch in EMER (uses

battery pack power)
• 5 g impact (uses main DC, battery, or battery pack power)
• Passenger safety switch selection (requires main DC or battery

power only)
• ON (manual activation) (uses main DC, battery, or battery pack power)
• OFF (system deactivation)
NOTE
The amber light to the left of the EMER LTS switch
illuminates when the airplane battery switch is placed
ON and the EMER LTS switch is OFF. The amber light
can be extinguished by placing the EMER LTS switch
to ARM or ON.
13. AC Alternators:
• One engine-driven 3.0 KVA alternator on each engine.
• Provides 115/200 volt, three-phase, 200 to 400 Hz power to electrically

heat the windshields.
• Controlled by WINDSHIELD ANTI-ICE switches on the cockpit tilt

panel (main DC) and the power control init (PCU) in the tail
compartment.

POWERPLANT
GENERAL
• Pratt and Whitney PW545B (Figure SRX-5)
• 4,095 pounds thrust at sea level (25°C/77°F)
• Bypass ratio 3.8:1
Electronic engine control (EEC) in AUTO mode provides:
• Detented throttle positions for automatic takeoff, climb, and cruise

thrust settings (N1 governing)
• Automatic thrust setting compensation for anti-ice ON
• Automatic engine idle governing for flight idle (57–62% N2) and
ground idle (48–51% N2)
NOTE
Ground idle rpm is achieved 8 seconds after landing
(WOW).
• Acceleration and deceleration limiting
• N1 and N2 speed limiting
• N2 underspeed limiting (at climb detent and above)
• Overspeed protection (N2)
• N1 or N2 synchronization
• Engine diagnostic system (EDS) functions
• Independent overspeed protection (N2) (~3% above redline)
NOTE
In AUTO mode, it is still the responsibility of the pilot
to monitor N 1 and N 2 limits to prevent an overspeed
condition and ITT limits to prevent an overtemper-
ature condition.

LEGEND
INDUCTION AIR
EXHAUST AIR
COMBUSTION
CHAMBER
CENTRIFICAL
COMPRESSION AIR
AXIAL
COMPRESSOR AIR
TURBINE AIR
Figure SRX-5. PW545B Cross-Section

Electronic engine control (EEC) in manual mode:
The fuel control unit (FCU) takes over full control of the engine speed in re-
sponse to the throttle position. In MANUAL mode, the throttle directly con-
trols the FCU by means of a mechanical linkage. MANUAL mode provides
the following functions:
• Pilot adjustable power setting (N2 governing above minimum fuel flow
N2)
• Idle governing (N2 governing) at flight idle and anti-ice idle
• Acceleration and deceleration limiting (ratio unit control)
• N2 speed limiting (~3% above redline)
• Limited engine diagnostic system functions (EDS)
Bleed-off valve control (BOV):
• Bleeds off excessive P2.8 compressed air, preventing compressor surges

and stalls.
• Electrically controlled operation by the EEC.
• BOV protection fails to a pneumatic backup mode with the loss of

main DC power.
IGNITION
A single, dual-channel exciter box with two ignitor plugs per engine. Burst
mode type ignition that produces 6–7 sparks per second for the first 30 sec-
onds, then one spark per second, thereafter. Green IGN annunciation on the
AMLCD engine indication verifies that DC power is available to the exciter
box. If one ignitor plug fails during engine start, the engine starts normally
and the ignition light remains illuminated during the start sequence and ter-
minates at the normal time.
Ignition Switch:
• NORM—Autoignition for start, and for engine or wing/engine anti-ice
on (powered by the crossfeed bus).
• ON—Pilot select ignition (used for takeoffs, landings, turbulence,

stalls, precipitation, and emergency descents). Powered by the
crossfeed bus with main DC power.
• SEC—Placing the ignition switch to the secondary (SEC) position

powers the ignition system from the emergency bus.

Oil
Maximum consumption is 0.2 pph, measured over a 10-hour period or one quart
in 10 hours. Check oil level 10 minutes after shutdown.
Do not mix nonapproved oil brands or oils of different viscosities.
Do not switch to “third generation” oils unless an engine is new or freshly

overhauled.
Oil pressure fluctuations are normal. Oil pressure indicator measures differ-
ential oil pressure.
Fuel
Engine-driven fuel pump—A two-stage pump in the fuel control unit.
Hydromechanical fuel control unit (FCU)—Fuel metering is normally con-

trolled by EEC.
Flow divider valve—Regulates fuel flow to the primary and secondary fuel
manifolds.
Secondary manifold activates at 26–28% N 2 .
Emergency shutoff valve—Shuts down engine to protect airframe.
FIRE PROTECTION
The engine fire protection system is composed of sensing loops, two control
units (one for each engine) in the tail cone, one ENG FIRE warning switch-
light for each engine, one FIRE DET SYS L–R annunciator for each engine,
two fire extinguisher bottles that activate from the cockpit (Figure SRX-6),
a FIRE EXT BOTL LOW annunciator, and a fire detection circuit test (Figure
SRX-7). Detection and extinguishing system electrical power is supplied
from normal DC power.
SENSING LOOPS AND CONTROL UNITS

Each engine nacelle contains one heat-sensing cable (or loop). The loop sur-
rounds the engine accessory section and the outer bypass duct near the com-
bustion section. It connects to a control unit that monitors electrical resistance.
As the loop is heated, its electrical resistance decreases until, at a tempera-
ture of approximately 450°F, a circuit is completed to the control unit to il-
luminate the applicable red ENG FIRE switchlight. The fire detection loops
are continuously monitored. Should the loop fail, the associated FIRE DET
SYS L/R annunciator illuminates to indicate fire detection is no longer avail-
able on that side.

LH RH
ENGINE ENGINE
FIRE FIRE
BOTTLE 1 BOTTLE 2
ARMED ARMED
PUSH PUSH
FIRE BOTTLE 1
FIRE BOTTLE 2
FIRE LOOP FIRE LOOP
RUDDER FIRE
BIAS DET SYS
FIRE EXT
BOTL LOW L R
LEGEND
Figure SRX-6. Engine Fire Extinguishing System
OPERATION
An engine fire light or overheat condition is indicated by illumination of the
applicable ENG FIRE switchlight on the glareshield (Figure SRX-7). Depressing
the illuminated ENG FIRE switchlight causes both white BOTTLE ARMED
switchlights to illuminate, arming the circuits to the bottles for operation. In
addition, the generator field relay opens (GEN OFF annunciator illuminates)
and provides a ground to power the fuel and hydraulic firewall shutoff valves
closed (causing the respective LO FUEL PRESS, LO HYD FLOW, F/W SHUT-
OFF annunciators to illuminate). The circuit to the thrust reverser isolation
valve is disabled, preventing deployment of the thrust reverser on that engine.

LH RH
ENGINE ENGINE
FIRE FIRE
FIRE
DET SYS
L R
FIRE LOOP FIRE LOOP
Figure SRX-7. Engine Fire Detection System
Depressing either illuminated BOTTLE ARMED switchlight fires the explo-

sive cartridge on the selected bottle, releasing its contents into the engine na-
celle. The BOTTLE ARMED switchlight extinguishes.
Depressing the ENG FIRE switchlight a second time opens the fuel and hy-
draulic firewall shutoff valves, and disarms the extinguishing system.
Due to the location of the fire bottles, bottle pressures cannot be checked on
preflight. If pressure is low on either (or both) fire extinguisher bottles, the
amber FIRE EXT BOTL LOW annunciator illuminates to alert the crew.

FUEL
Refer to Figure SRX-8 for Fuel System—Normal Operation.
Total useable fuel capacity is 1,006 U.S. gallons (approximately 6,790

pounds).
Boost pump switches:
• ON—DC boost pumps are continuously energized, FUEL BOOST

L–R annunciated.
• OFF—DC boost pumps are deenergized and will not operate.
• NORM—DC boost pumps are energized automatically for start,

crossfeed, and low fuel pressure with annunciation of FUEL BOOST
L–R. The right boost pump is energized automatically for and during
APU operation, but annunciation (FUEL BOOST–R) only occurs
during start, crossfeed, and low fuel pressure.
Tank-to-tank transfer rate during crossfeed is approximately 700–900 pph (en-

gines shutdown).
Low fuel pressure light illuminates at a decreasing pressure of 5 psi and ex-
tinguishes at an increasing pressure of 7 psi.
Low fuel level light illuminates at approximately 360 ± 20 pounds of usable

fuel remaining in the respective tank; input is from a float switch.
Illumination of the FUEL GAUGE annunciator indicates a fault has been de-
tected in the respective fuel gauging system. Do not shut down DC power (BATT
switch to OFF) after engine shut down until checking and recording the fuel
conditioner BITE lights.
The fuel filter is on the engine downstream of the fuel-oil heat exchanger
(FOHE), eliminating the need for fuel anti-ice additives. It is still recommended
to use Prist or other approved fuel additives on a regular basis for the anti-
fungal properties of the additive.
NOTE
Av-gas is not an approved fuel.

PRE-CHECK
DE-FUEL SELECT LEVERS
LEVERS
SPPR ADAPTOR
(360 +/- 20 LBS.)
(163 KG)
SCAVENGE
LO FUEL EJECTOR
LEVEL
SURGE TANK
L R
FUEL

x x BOOST
FUEL TRANSFER TUBES L R

R FUEL
HOPPER
PRIMARY EJECTOR
LEGEND
(300-1200 PSIG)
LO FUEL
NEG / POS PRESSURE RELIEF VALVE MOTIVE PRESS
FLOW
FUEL TANK VENT L R
MOTIVE FLOW SHUTOFF
SPR FUEL SHUTOFF VALVE SOLENOID VALVE (NO)
P P
HIGH LEVEL PILOT VALVE FUEL FOHE
FLTR BP FUEL FILTER
LOW LEVEL PILOT VALVE VENT F/W
L R ENG DRIVEN SHUTOFF
FUEL BOOST PUMP FUEL INSIDE THE TANK FUEL PUMP
FCU L R
CHECK VALVE REFUEL PRESSURE FUEL FLOW
FCU FLOW DIVIDE CROSS FEED VALVE
EJECTOR PUMP MOTIVE FLOW PRESSURE EMERGENCY FUELSHUTOFF VALVE
(MOTORIZED) APU
(MECHANICAL)
F/W SHUTOFF VALVE (MOTORIZED) APU FUEL
JET PUMP PRESSURE
SHUTOFF
P LOW FUEL PRESSURE SWITCH (5.3 PSIG)
BOOST PUMP PRESSURE FUEL VALVE
x LOW LEVEL FLOAT SWITCH (360+/-20 LBS - 163 KG) XFEED
SRX-25

TEMPERATURE SENSOR (-60oC - 70oC)
Figure SRX-8. Fuel System—Normal Operation

Crossfeed Operation—Select the tank desired to supply fuel to both engines

and transfer to the other tank. With the boost pump switches in NORM, se-
lecting crossfeed initiates the following sequence of events (Figure SRX-9):
1. DC boost pump on the side supplying fuel activates (respective FUEL

BOOST illuminates steady).
2. Crossfeed valve opens (FUEL XFEED advisory light illuminates when the
valve is fully open).
3. Motive flow valve on the receive side is energized closed 3 seconds

after crossfeed is selected (transfer rate depends on operating engine(s)
requirements).
NOTE
If the boost pump on the receiving side is ON or the
boost pump on the crossfeeding side is OFF, no cross-
feed will take place.
Selecting the crossfeed switch to OFF, reverses the above process. Should the
crossfeed valve fail to close, the FUEL XFEED advisory light illuminates flash-
ing and activates the MASTER CAUTION lights steady.
It is important to allow the crossfeed system to complete the above process

before reselecting a different crossfeed switch position.
If the opposing boost pump activates (on the receiving side), it indicates a
timing problem with the crossfeed valve. To rectify the problem, reset the op-
posing boost pump (turn the opposing side FUEL BOOST switch to ON then
back to NORM). Do not turn a FUEL BOOST switch OFF and leave it there;
OFF is OFF.

LEGEND
FUEL HOPPER FUEL INSIDE THE TANK
OPERATING BOOST MOTIVE FLOW PRESSURE
PUMP
CROOSFEED JET PUMP PRESSURE
VALVE OPEN
MOTIVE FLOW BOOST PUMP PRESSURE
SHUTOFF VALVE
FUEL BOOST
ON
O
F
F
NORM
CROSSFEED
L R
TANK OFF TANK
L R
ENG ENG
FUEL
BOOST
FUEL
XFEED L R
Figure SRX-9. Fuel System—Crossfeed (R to L)

HYDRAULICS
1. Reservoir Quantities:
• Overfull.................................................................................... 360 cu in
• Full ......................................................................................... 215 cu in
• Refill ........................................................................................ 175 cu in
• LO HYD LEVEL annunciator................................................... 74 cu in
• Empty........................................................................................... 5 cu in
2. Engine-driven pump flow rate is 3.25 gpm/1229 lpm maximum.
3. Open center system (Figure SRX-10):
• System control valve open (normal) ............................................. 60 psi
• System control valve closed (system operational) ................... 1,500 psi

SUBSYSTEM CONTROL VALVES

F F (.55 GPM)

LANDING

GEAR LO HYD

FLOW
SPEEDBRAKES
L R

RETURN LINES
WING FLAPS
HORIZONTAL HYD CONTROL

STABILIZER VALVE (N/O)
PRESSURE
THRUST SWITCH (LOADING VALVE)
REVERSERS
P
RELIEF VALVE
OPENS AT 1,350 PSI
FILTER
F/W SHUTOFF
MOTORIZED
VALVE
XLS R ENGINE
F/W
PUMP
LEGEND SHUTOFF
(74 CU)
SUPPLY SUCTION LO HYD L R
LEVEL
RETURN PRESSURE
HYD
#1 SYS HIGH PRESSURE (MAIN) PRESS HYDRAULIC RESERVOIR (TAIL CONE)
SRX-29
Figure SRX-10. Hydraulic System-Open Center

4. Speedbrakes (Figure SRX-11):
• Held closed mechanically; held open by trapped hydraulic fluid.
• Retract normally with switch in RETRACT; Automatically retracts if

either throttle is advanced beyond 80 to 85% N2 PLA (power lever
angle).
• Requires normal DC power to remain extended. With loss of normal

DC power, the safety valve relaxes open to allow trapped fluid to
escape and the speedbrakes blow to a trail position.
• System pressure is required for extension and normal retraction.
• The white SPD BRK EXTEND annunciator illuminates when both

speedbrake segments are extended. Speedbrakes fully extend in
approximately 1.5 seconds.
• The NO TAKEOFF annunciator illuminates if the speedbrakes are

extended on the ground.

FLAPS
UP 0°
TRIM TO
CLB T.O.
NOSE 200 KIAS 7°
DOWN
CRU
STAB
T T
O H T.O. &
R APPR 15°
O 200 KIAS
T
NOSE
UP
T
L
E
IDLE
MIS COMP
LAND 35°
SPEED
BRAKE CUT
OFF
175 KIAS
ENGINE SYNC
SPD BRK
EXTEND
LH RH MUST BE OFF
RETRACT & LANDING

EXTEND
SPEED BRAKE
SWITCH
SPEED BRAKE SPEED BRAKE

ACTUATOR ACTUATOR
LO HYD BYPASS
LEVEL
SAFETY
HYD
PRESS
CONTROL VALVE
VALVE
1500 PSI PRESSURE SYSTEM LOADING VALVE

RELIEF VALVE
CHECK CHECK
VALVE VALVE
PUMP PUMP
SUPPLY SUCTION
RETURN
LOW FULL OVER FULL
SUCTION
HYDRAULIC RESERVOIR
RETURN PRESSURE
LO HYD
#1 SYS HIGH PRESSURE (MAIN) LEVEL
HYD
PRESS
SRX-31
Figure SRX-11. Speedbrake System-Normal Operation (Extended)

5. Landing gear (Figures SRX-12, SRX-13 and SRX-14):
• DC power is required for normal hydraulic extension and retraction of

the landing gear. Control provided through the GEAR CONTROL
circuit breaker on the left CB panel and powered by the emergency DC
bus system.
• Gear is held in the up position by uplocks; held in the down position by

hydraulic actuators. Operational hydraulic pressure is used to raise and
lower the gear.
• Freefall/blowdown backup system. Manual retraction of the uplocks

allows the gear to freefall, then pneumatics provide positive downlock
(pneumatic system retracts the uplocks in the event the manual release
does not).
• Aural warning, unsafe gear down: 1) both throttles below 70% N2,
flaps beyond 15°; 2) both throttles below 70% N2, radio altitude less
than 500 feet; 3) both throttles below 70% N 2 , radio altimeter
inoperative, airspeed less than 150 KIAS.

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE
SHUTTLE VALVE

ACTUATOR UPLOCK UPLOCK ACTUATOR
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
H H ACTUATOR LEGEND
RETRACT PRESSURE
UP ANTI-
SKID
GEAR
DOWN NITROGEN
EMERGENCY NITROGEN
BLOW DOWN
OFF BOTTLE
SRX-33
Figure SRX-12. Gear System-Normal Retraction

SRX-34

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE
SHUTTLE VALVE

ACTUATOR ACTUATOR
UPLOCK UPLOCK
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
H H ACTUATOR LEGEND
EXTEND PRESSURE
UP ANTI-
SKID
GEAR
DOWN NITROGEN EMERGENCY NITROGEN

BLOW DOWN
OFF BOTTLE
Figure SRX-13. Gear System-Normal Extension

LO HYD
LEVEL
EMERGENCY - FLUID
HYD RETURN VALVE
PRESS CONTROL VALVE
SHUTTLE VALVE

ACTUATOR ACTUATOR
UPLOCK UPLOCK
SHUTTLE VALVE
LO BRK
PRESS
UNLOCK ANTISKD
INOP
N T-HANDLE
O
H H ACTUATOR
UP ANTI-
SKID
LANDING ON
GEAR
DOWN NITROGEN
BLOW DOWN
OFF BOTTLE
SRX-35
Figure SRX-14. Gear System-Emergency Extension

6. Thrust Reversers (Figures SRX-15 and SRX-16):
• Normal DC power is required.
• Only one squat switch is required (left, right, or both) to allow the
control valve to energize to the deploy position when commanded.
• Each emergency stow switch is powered by the opposite side thrust

reverser circuit breaker.
• Illumination of either the ARM or UNLOCK light while in flight

triggers the MASTER WARNING flasher.
• If the thrust reversers do not operationally check (including emergency

stow), flight must not be attempted.
• The thrust reversers should be in idle power at and below 60 knots.
• The use of thrust reversers to back the airplane is prohibited
• Normal thrust reverser use or a reverser selected to EMER STOW

deactivates rudder bias.
7. Flaps:
• Flaps are hydraulically actuated and electronically controlled.
• See FLIGHT CONTROLS—Flaps for description.
8. Two-position horizontal stabilizer:
• Horizontal stabilizer is hydraulically actuated and electronically

controlled
• See FLIGHT CONTROLS—Horizontal Stabilizer for description

ISOLATION VALVES
THRUST REVERSER
THRUST REVERSER

STOW
EMER ARM
(SQUAT SWITCH & (SQUAT SWITCH &
ARM EMER
UNLOCK
NORM DEPLOY
DEPLOY NORM
(ARM LIGHT)
LEVERS
PRESSURE
LO HYD
FLAPS
UP
T.O.
0°
7°
HYD VALVE (LOADING
PITCH
TRIM
T.
T
H
R
O
200 KIAS
T.O. &
APPR 15°
PRESS VALVE)
T 200 KIAS
O.
T
L LAND
NOSE E 35°
UP 173 KIAS
OFF
N1
ENGINE SYNC
OFF
N2
MUST BE OFF
FOR TAKEOFF
& LANDING
PRESSURE
RELIEF VALVE
OPENS @
1350 PSI
HYDRAULIC
STATIC FLOW PUMP RESERVOIR
LO HYD
#1 SYS LOW LEVEL
PRESSURE (MAIN)
HYD
SRX-37
SUPPLY SUCTION PRESS
Figure SRX-15. Thrust Reversers-Stowed

SRX-38
ISOLATION VALVES
THRUST REVERSER
THRUST REVERSER

STOW
EMER ARM
(SQUAT SWITCH & (SQUAT SWITCH &
ARM EMER
UNLOCK
NORM DEPLOY
DEPLOY NORM
(ARM LIGHT)
LEVERS
PRESSURE
LO HYD
FLAPS
UP
T.O.
0°
7°
HYD VALVE (LOADING
PITCH
TRIM
T.
T
H
R
O
200 KIAS
T.O. &
APPR 15°
PRESS VALVE)
T 200 KIAS
O.
T
L LAND
NOSE E 35°
UP 173 KIAS
OFF
N1
ENGINE SYNC
OFF
N2
MUST BE OFF
FOR TAKEOFF
& LANDING
PRESSURE
RELIEF VALVE
OPENS @
1350 PSI
HYDRAULIC
#1 SYS HIGH PUMP RESERVOIR
PRESSURE (MAIN) LO HYD
SUPPLY SUCTION LEVEL
HYD
RETURN PRESSURE PRESS
Figure SRX-16. Thrust Reversers-Deployed

POWER BRAKES AND ANTISKID

• Separate from the main airplane hydraulic system, with the reservoir in
the aft fairing (Figure SRX-17).
• Normal DC power required to operate the accumulator pump. Power is

available to the pump motor whenever the landing gear handle is
selected “down.”
• Antiskid protection is available only when power brakes are

operational.
CAUTION
Do not pull the PWRBRKS circuit breaker to prevent
the power brake pump from cycling. With the circuit
breaker disengaged, the power brake system is inop-
erative and the toe pedals are disabled. Braking is then
available only by use of the emergency brake system.
• LO BRK PRESS annunciator with noncancelable MASTER

CAUTION illuminates if low brake pressure exists on the ground.
• Pneumatic brakes are a backup for the power brakes; no differential

braking and no antiskid protection available with pneumatic braking.
• Antiskid protection ceases below 10 knots and is available at speeds

between 10 and 175 knots.
• Touchdown protection prevents power brake and antiskid operation

until:
1. Wheel speed reaches 59 knots or 3 seconds after a squat switch is

in the ground-on-ground (GOG) mode.
2. Locked wheel protection provides a pressure dump command to

the slow wheel when its velocity is 30% less than the fast wheel.
• Digital antiskid provides continuous monitoring of the system.
• If any fault except squat switch disagree is detected, the ANTISKD

INOP annunciator illuminates. Analog antiskid BITE(fault) indicators
are in the left nose avionics compartment.

SRX-40
CABIN PRESSURE
TEST
OFF FIRE
SPARE WARN
AVN LDG
GEAR FLUID RESERVOIR
BATT
ANNU TEMP
ANTI STICK
SKID SHAKER
OVER T/REV
SPEED W/S TEMP

POWER BRAKE
CABLES
PUMP MOTOR
PEDAL
BIT FAULT INDICATOR
--CNTL UNIT
--VALVE
--L XDCR
--R XDCR
--SQUAT DISAGREE
1230-1500 PSI
ACCUMULATOR
P
PRESSURE
LO BRK
PRESS
LINE
UNLOCK
ANTISKD
N LO BRK P INOP
L
O
R
PRESS POWER
H H BRAKE 900 PSI
ANTISKD
VALVE
INOP
UP ANTI-
SKID
RETURN
LANDING
GEAR
ON LINES NITROGEN
ANTI-SKID BLOW DOWN
DOWN BOTTLE
SERVO
OFF VALVE PARKING BRAKE
LEGEND 28 VDC MAIN DIGITAL ANTI-SKID VENT

RETURN PRESSURE WOW CONTROL UNIT WOW
#1 SYS HIGH PRESSIRE (MAIN)
METERED BRAKE PRESSURE

WHEEL TRANSDUCER INPUTS
EMERGENCY NITROGEN PNEUMATIC LINE
Figure SRX-17. Power Brake/Antiskid System

EMERGENCY BRAKES
• A pneumatic brake system is available in the event the hydraulic brake
system fails (Figure SRX-17).
• Uses air pressure from the pneumatic bottle. Bottle pressure is

adequate for stopping the airplane, even if the landing gear has been
pneumatically extended.
• Pulling the red EMER BRAKE PULL lever mechanically actuates the
emergency brake valve. Air pressure to the brakes is metered in direct
proportion to the amount of lever movement.
• Differential braking is not possible, since air pressure is applied to both

brakes simultaneously. Releasing the handle vents pneumatic pressure
from the brakes. Do not excessively cycle the handle due to rapid
depletion of the pneumatic bottle.
• Do not depress the brake pedals while applying emergency air brakes.
FLIGHT CONTROLS
All primary flight controls (ailerons, elevators, and rudder) are manually ac-
tuated with cables and pulleys and are dual interconnected. Secondary flight
controls consist of trim tabs, speedbrakes, flaps, and a two-position horizon-
tal stabilizer (Figure SRX-18).
1. Ailerons:
• Maximum aileron travel is 19° up and 15° down.
• Trim tab on the left aileron only has a maximum travel is 20° up and
down.
• Trim is mechanically controlled with a wheel at the rear of the center

pedestal.
2. Elevators:
• Maximum elevator travel is 19° up and 15° down.
• Left and right trim tab travel is 5° up and 15° down.
• Trim is electrical (pitch) or mechanical.
• Left electrical pitch trim will override right.
• Electrical pitch trim must be tested before flight.
• Electrical trim can be interrupted with the red AP/TRIM DISC button
on either yoke.

SRX-42
RUDDER
ELEVATOR TRIM TAB
RUDDER TRIM TAB
ELEVATOR TRIM TAB

AILERON ELEVATOR
FLAPS
SPEED BRAKES
AILERON TRIM TAB
Figure SRX-18. Flight Controls

3. Rudder:
• Maximum rudder travel is 28.5° either side of centerline.
• Trim tab (servo tab) travel is 14° either side of centerline when rudder
is centered.
• Full rudder pedal deflection on ground deflects the nosewheel 20°

either side of centerline with brake assist the maximum nosewheel
deflection is approximately 90°.
• Flight must not be attempted if the nosewheel steering is inoperative.
4. Rudder bias:
• The rudder bias system is comprised of a bleed-air shutoff valve and a

dual actuating pneumatic cylinder that drives a closed loop cable
connected to the rudder (Figures SRX-19 and SRX-20).
• With engines operating, bleed air is continuously available to the

pneumatic cylinder through the shutoff valve any time main DC or
battery power is available and the rudder bias circuit breaker is in. With
approximately equal thrust available from the engines, rudder bias is
balanced and does not affect rudder position. During periods of
unequal thrust (i.e., engine failure), the rudder bias system
automatically deflects toward the operating engine. This assists the
pilot to compensate for adverse yaw.
• With power available, the rudder bias shutoff valve energizes open and
ports engine bleed air to its respective side in the cylinder. If the shutoff
valve does not move to its full open position, the amber RUDDER BIAS
annunciator illuminates indicating the system has malfunctioned.
• The rudder bias circuit breaker must be pulled to deactivate the system.
• Rudder bias is deactivated during thrust reversers deployment or EMER

STOW use. This is accomplished by closing the bias shutoff valve;
however, the amber rudder bias annunciator does not illuminate.
• Rudder bias does not operate in the emergency bus condition.
5. Rudder bias heat:
• A flexible dual electric element heater blanket is wrapped around the

rudder bias cylinder to prevent freezing. Each blanket heats
automatically between 4–16°C with main DC or battery power
available. The blankets each have its own thermostat (RTD—resistive
thermal device) for redundancy.
• Main DC power for the system is supplied from a circuit breaker in the
aft J-box.

RUDDER
BIAS
FIRE EXT
BOTL LOW
HEATER
BLANKET
BIAS
ACTUATOR
SHUTOFF
VALVE
RUDDER
BIAS HTR
BIAS
HEATER
FAIL
LEGEND
BLEED AIR
Figure SRX-19. Rudder Bias System
RUDDER
BIAS
FIRE EXT
BOTL LOW
HEATER
BLANKET
BIAS
VALVE
RUDDER
BIAS HTR
BIAS
HEATER
FAIL
LEGEND
STATIC FLOW
BLEED AIR
Figure SRX-20. Rudder Bias System—Engine Failure

• Upon aircraft power-up, the heating system PCB in the aft J-box
conducts a test of the two blanket thermostats. If either blanket fails the
test, the BIAS HEATER FAIL annunciator on the cockpit panel flashes
until pressed. The annunciator then illuminates steady until
maintenance is performed.
• Upon aircraft power-up and after self-test, the heating system heats the
cylinder, if required, to 16°C. While heating is in progress, the BIAS
HEATER FAIL annunciator illuminates steady. Refer to the “Master
Warning” section for further description.
6. Flaps:
• Flaps are electrically controlled and hydraulically actuated through the

flap handled on the center pedestal. Electrical control power is
provided from the emergency bus through the FLAP CONTROL
circuit breaker on the left CB panel.
• With the loss of electrical power (circuit breaker out), the flaps remain
in the last position. The flaps cannot be moved.
• With loss of hydraulic power, the flaps remain in the last position
unless the flap handle is moved, after which the flaps blow to a “trail”
position dependent upon air-load forces.
• Normally, during flap actuation the HYD PRESS annunciator

illuminates until the flaps reach the commanded position.
• Flap positions ranging from 0–35° can be selected with the flap handle.
Although a wide range of positions can be selected, 0° is used for
cruise only, 7° and 15° are approved for takeoff, and 35° is used for
landing. Flap handle detents and speed placards are available at the flap
handle.
• On ground, if the flaps are not set at or between 7° or 15° (TO or

APPR) position, the NO TAKEOFF annunciator illuminates.
• Flaps are held extended with trapped hydraulic fluid and held retracted
mechanically.
• Mechanical interconnects between flap segments prevent asymmetrical

flap conditions.
• Flap handle movements to certain positions affect horizontal stabilizer

movement.
7. Horizontal stabilizer (Figure SRX-21):
• The two-position stabilizer is a secondary flight control used to achieve

optimum angle of attack for takeoff and flight configurations.
• The two-position horizontal stabilizer is electrically controlled and

hydraulically actuated.

SRX-46
HORIZONTAL STABILIZER
CRUISE POSITION
SPEED ABOVE 200 KIAS
FLAP
CONTROL
VALVE

(EMER BUS)
P HYD PRESS HYD

SWITCH PRESS
HYDRAULIC CONTROL
VALVE (LOADING VALVE)(NO)
STAB
MIS COMP
FLAP HANDLE HYDRAULIC
POSITION PUMP
HYDRAULIC
RESERVOIR STABILIZER
FLAPS POSITION
UP 0°
TRIM TO
CLB T.O.
NOSE 200 KIAS 7°
DOWN
CRU +1
T
O
T
H T.O. & SPEED –2
R
O
APPR
200 KIAS
15°
PCB
T
T
(UP) SENSOR
L
E
NOSE
UP IDLE
215 (+/– 10) KIAS
LAND
SPEED
175 KIAS
35°
(DN) (EMER BUS)
BRAKE CUT
LH
OFF
RH
ENGINE SYNC
MUST BE OFF
HORZ STAB
OFF TURB FOR TAKEOFF
RETRACT
FAN
& LANDING CONTROL
VALVE STBY PITOT/STATIC LEGEND
EXTEND
(EMER BUS) HYDROMECHANICAL INPUT RETURN PRESSURE
ARMING VALVE
ACTUATOR
SUPPLY SUCTION
#1 SYS LOW PRESSURE (MAIN)
STATIC FLOW
Figure SRX-21. Two-Position Horizontal Stabilizer

• The stabilizer is electrically (emergency bus) controlled by the flap

handle and moves in concert with flap movement. Flaps 0° commands
a stabilizer full up (+1°) position used for climb, cruise, and descent.
Flap handle positions from 1° to 35° commands the stabilizer to its full
down (–2°) position used during takeoff, landing, and approach.
Stabilizer positions between full up and full down only occur due to
malfunction.
• Horizontal stabilizer movement in either direction requires 25 seconds.

The HYD PRESS annunciator illuminates during stabilizer movement
until the system achieves its commanded position.
• If the horizontal stabilizer position is not correct according to the flap

handle position command, an amber stabilizer miscompare
annunciation (STAB MISCOMP) illuminates to alert the crew. On
ground, this miscompare causes a NO TAKEOFF annunciation. Refer
to “Stab Miscomp” description in the “Master Warning” section.
• The horizontal stabilizer system incorporates an airspeed sensor, which

receives pitot-static input from the standby pitot-static system.
Stabilizer movement is restricted with airspeeds greater than 215 ± 10
KIAS.
• Horizontal stabilizer movement restriction due to excessive speed is

accomplished by an overspeed signal from the airspeed sensor to the
system PCB (printed circuit board), which restricts proper positioning of
the arming valve and stabilizer control valve.
• The horizontal stabilizer will not move when commanded by the flap
handle for the following reasons:
1. Hydraulic pressure is not available.
2. Malfunction.
3. IAS in excess of 215 ± 10 KIAS.
4. Landing gear is first selected and is in transit (landing gear has

priority over the horizontal stabilizer only through the stabilizer
PCB).
WARNING
Do not retract flaps above 200 KIAS. Associated sta-
bilizer movement can cause a significant nose down
pitch upset if the movement is not prevented.
• The horizontal stabilizer receives power from the emergency bus.

8. Control Lock:
• Secures the three primary flight controls in the neutral position and
secures the throttles in cut-off position.
• Towing with the control lock engaged is not recommended on ground

because of the on-ground nosewheel-rudder interconnect. Nosewheel
steering can be damaged if the nosewheel is turned too far. The
maximum range available for towing is 60° left or right of center.
9. Stall warning stick shaker:
• System includes a stick shaker on each control column and an angle-

of-attack system. Stick shaker power is provided through the AOA
circuit breaker on the left CB panel. When the stall warning system
senses an approaching stall, both shakers vibrate until the condition is
corrected.
• Stick shaker vibration activates a .79–.88 AOA (8–10% above stall) and
above depending on flap setting.
• The system is required to be tested and operate for dispatch. Refer to

“Master Warning—Rotary Test” for indications.
• Additional stall warning is achieved with a stall strip on each wing root
by producing buffets.
ICE AND RAIN PROTECTION

1. Ice detection:
• Two red ice detection barrel lights on the top of the instrument panel
glareshield reflect a glow to warn the crew if ice accumulates on the
windshields at the extreme inboard area. These are activated by the
light panel switch in the ON position.
• Wing inspection lights on each side of the fuselage illuminate the wing
leading edges.
2. Pitot-static heat—Two minute limit on ground operation (Figure SRX-22):
• Three pitot tubes (L, R, and STBY)
• Six static ports (L, R, and STBY)
• AOA vane
• Separate annunciators for AOA, L, R, and STBY pitot-static systems.
• Standby pitot-static heat is powered from the emergency bus.

P/S
HTR
LEFT PITOT SYSTEM L R RIGHT PITOT SYSTEM

MADC MADC
LNAV VNAV 2000 LNAV VASEL
VAPP AL-VN FMS VGP
280
242 YD OFF AP OFF CAT2 1400
920 IAC IAC 160
160 E 20
AP ENG
20
9000
FMS 1000
20 20 6 6
260 4 10 10 4
2 140 10000 2
10 10 1500 1
3 80
6 20
1
242
240 13 60 125 98 00
120
1 1 4 2 10 10 1
220
10
20
10
20 OM I
2
4
6 DATA DATA 100
R
1
20
30
20
2
4
9500 6
–950
RAD BARO
200 MIN MIN
.750 M 2500 29.92 IN .261 M 200 STD
AOA DME AOA DME
HDG FMS1 RW0IL HDG FMS1 RW0IL
349 023 VOR1 349 023 VOR1
329 13 23.0 NM 329 13 23.0 NM
.70 12 MIN .50 12 MIN
33 N 157 KTS
33 N 157 KTS
3 FMS STATUS KHUT FMS STATUS
BRG PTR BRG PTR 3
30
MSG +13 WPT MSG

FMS1 FMS1
30
002
ADF APPR ADF APPR
6
W
ET DR ET DR
0:00:00 0:00:00 WPT
WIND 001 WIND
E
CLOCK CLOCK
-04
24
19:39:07 39 19:39:07 5 39
WEATHER WEATHER
12
WX/R/T 21 TAWS WX/R/T TAWS

T4.5° A S 15 T4.5° A
STAB TGT TERRAIN STAB TGT RA 9.8NM+13 TERRAIN
INHIBIT INHIBIT
TAS
LX/ON LX/ON TA 4.5NM-04
PROBE
Honeywell Honeywell
LEGEND LH STATIC RH STATIC

PORTS PORTS
LH STATIC
RH STATIC
STANDBY STATIC
STBY
STANDBY PITOT 10
15
4 5 6
20
25 P/S HTR
3 PSI 7
9 30
2
5 1 9 35
80 30.15 in 0 10
DIFF 40
PRESS 50
PITOT &
0
CABIN ALT
LH PITOT WINDSHIELD
60 00
1500
X1000 FT
STATIC 40
1
30
10 10
13 20
00
ON L R 9
10 10
RH PITOT O'RIDE 1000
ON
33
M
N
STANDBY PITOT
AIRSPEED
OFF OFF
SENSOR
SRX-49
(HORIZONTAL STABILIZER)
Figure SRX-22. Pilot-Static System

3. Windshield anti-ice (Figure SRX-23):
• The windshield is heated with AC current to provide anti-ice and defog

capability.
• Two 3-phase 115 VAC alternators are on each engine accessory gear
box (N2 rpm) to provide current for heating of the windshields forward
and side panels. The rear side windows are not electrically heated.
• The left and right alternator bearings are monitored for wear with the
white “AC BEARING” annunciator. When illuminated, the alternator
has approximately 20 hours of operations remaining.
• One 3-gang AC circuit breaker for each alternator is in the baggage

compartment.
• Left and right DC controllers regulate the respective alternators and

control temperature via embedded windshield sensors. The left and
right controller circuit breakers are on cockpit left CB panel.
• Two left temperature sensors and two right temperature sensors are
used by the controller to regulate windshield temperature at 110°F
(43°C). If the system malfunctions and windshield temperature reaches
140°F (60°C), system overtemperature circuitry deactivates AC current
to the entire left or right system. This condition causes the amber W/S
O’HEAT and the W/S FAULT annunciators to flash.
After temperature cool down of the affected side to 115°F (46°C), the
system can automatically reset and again apply AC current for heating.
If the side reaches the overtemperature limit again, the system will shut
down. This on/off condition is called “cycling.” AFM “Abnormal
• Left or right system malfunctions other than overheat conditions

activate the amber W/S FAULT annunciator only. AFM “Abnormal
• Windshield sensors are tested with the WS TEMP rotary test knob per
AFM “Normal Procedures.” See “Rotary Test” section for more
information.
NOTE
The W/S FAULT annunciator may not test after cold
soak at extremely cold temperatures. If this occurs,
repeat the test after the cabin has warmed up. The test
must be completed prior to flight.
I f t h e w i n d s h i e l d i s h e a t s o a ke d a b ove + 5 6 ° C
(+134°F), the test results in a W/S FAULT annunci-
ator illuminating.

SECONDARY
140°F/60°C SENSOR
S S
W/S
O'HEAT PRIMARY
L R P P SENSOR

110°F/43°C
(NORM TEMP)
DC DC
CONTROLLER W/S CONTROLLER
FAULT
L R
WINDSHIELD
LH RH
ALTERNATOR L R ALTERNATOR
O'RIDE
ON
LEGEND AC
BEARING OFF
LH ALTERNATOR
L R
SRX-51
RH ALTERNATOR
Figure SRX-23. Windshield Anti-Ice System

• The left and right WINDSHIELD switches on the tilt panel have three
positions:
• O’RIDE—Allows the controllers to heat the windshield to 110°F

at a faster heating rate than normal. The controller continues to
regulate windshield temperature at 110°F.
• ON (normal operating position)—Allows the controller to heat the

windshield to 110°F at a normal “ramp-up” heating rate. The
controller continues to regulate windshield temperature at 110°F.
• OFF—Deactivates the controller and the respective alternator.
4. Engine anti-ice (Figure SRX-24):
• Engine anti-ice heats the fan nose cone, T1 and T0 probes, the nacelle
lip, and stator vanes.
• The following are heated anytime the engine is operating:
• Fan nose cone—P2.8 bleed air
• T1 probe—P3 bleed air
• With the engine or engine/wing anti-ice switch ON, the following are
heated:
• T0 probe—Electrically
• Nacelle lip—P3 bleed air
• First two sets of stator vanes—P3 bleed air
• With an engine or engine/wing anti-ice switch ON, the nacelle and

stator anti-ice bleed air valves (PRSOVs) deenergize open and the T0
probe receives electricity. The amber ENGINE ANTI-ICE annunciator
illuminates steady until the nacelle temperature rises above 60°F
(15°C). Above 60°F, the annunciator extinguishes indicating normal
operation.
NOTE
If ambient temperature is approximately 59°F (15°C)
or warmer, the ENG ANTI-ICE L–R annunciators
may not illuminate when anti-ice is selected ON. To
ensure that bleed air is flowing to the engine inlet,
the crew should observe a momentary small decrease
in N 2 when ENGINE ON is selected.
• In flight with the engine or engine/wing anti-ice switch ON, if the

nacelle temperature does not exceed 60°F or the stator valve does not
open within 4 minutes 45 seconds, the steady amber engine anti-ice
annunciator begins to flash indicating a malfunction. AFM “Abnormal
Procedures” are required.

WING WING WNG XFLOW WING/ENGINE
ON L R
ANTI-ICE XFLOW O'HEAT ON
VALVE (106°C)
220° 220°
L R 220° (N/C) 220° L R OFF
ON
OFF ENGINE
160° 160°

160° 160°
(71°C) EMER
L WING PRESS
ANTI-ICE VALVE
PRSOV (N/C)
(N/O)
L PRECOOLER
ENG
ANTI-ICE 60° (15°C)
60° R NACELLE
L R ANTI-ICE PRSOV (N/O)
P3 P3
560° 560°
LEGEND R STATOR ANTI-ICE PRSOV (N/O)
PURGE AIR BLD AIR

O'HEAT
P3 BLEED AIR
L R
RAM AIR
WING BLEED-AIR SHUTOFF CAPABILITY

DUE TO AN O'HEAT CONDITION WING AND ENGINE ANTI-ICE ON
SRX-53
Figure SRX-24. Wing/Engine Anti-Ice System

• During ground operations, the annunciator illuminates steady only

during warmup or for malfunction. The engine T 1 probe is not
monitored for electrical current or heat.
• With engine or engine/wing anti-ice switch ON, engine ignition

activates continuously.
• With a loss of main DC power (emergency bus condition), the nacelle

and stator anti-ice valves fail-safe open. The T0 probe is not heated.
• The T1-FCU sensor (flush-mounted sensor) is heated indirectly with

nacelle heat.
• Engine anti-ice must be selected when temperature is +10°C or less

with visible moisture (ground or in flight).
5. Wing anti-ice (Figure SRX-24 and SRX-25):
• Precooled engine bleed air (P3) is used to heat the wing leading edge.
The bleed-air pressure is regulated at 16 psi.
• With the engine/wing anti-ice switch ON, the wing anti-ice bleed air
valve (PROSV) deenergizes open and the amber WING ANTI-ICE
annunciator illuminates steady. When the wing leading edge
temperature exceeds 220°F (110°C), the annunciator extinguishes
indicating normal operation.
• When initially turned on, the WING ANTI-ICE annunciator begins

flashing if the wing does reach operating temperature within 4 minutes
and 45 seconds. AFM “Abnormal Procedures”are required.
• The wing undertemperature switches are enabled with engine/wing

anti-ice switch ON.
• On ground, the WING ANTI-ICE annunciator illuminates steady only.
• With a loss of main DC power (emergency bus condition), the wing

anti-ice valves fail-safe open.
• Wing anti-ice is not recommended for use during ground operations.
NOTE
The wing anti-ice valve is held closed as a result of
a bleed air overheat condition on the respective side.
This automatic action protects the wing leading edge
from excessive heat.

WING
O'HEAT
L R
160°F SWITCH FUEL BOUNDARY

HEA
T SH
IELD
PUR
GE P
ASS
AIR A
FLO GE
W
BLE
DEFLECTOR SHIELD ED
AIR
SRX-55
Figure SRX-25. Wing Leading Edge Cross Section

Wing crossflow (see Figure SRX-24):
• Wing crossflow is selected ON with WING XFLOW switch. When

selected ON, the Xflow valve energizes open and connects the left and
right wing bleed air systems.
• Wing crossflow is used per AFM “Abnormal Procedures” conditions.
• The crossflow valve fail-safes closed.
• There is no annunciation during crossflow use.
Wing overheat protection (see Figure SRX-24):
• Wing overheat protection is provided by three overtemperature

switches. Two 160°F (71°C) spar sensors are in the cooling purge
passage airflow between the heated leading edge and the fuel tank and
one 230°F (110°C) switch on a bracket in wing cuff of the inboard
leading edge.
• On ground or in flight, if an overheat condition is sensed by either of

the three switches, the amber WING O’HEAT L–R annunciator
illuminates flashing. This condition automatically closes the respective
side wing anti-ice valve (PRSOV) causing the wing to become cold, or
not heat up if wing anti-ice is selected.
As the overheat condition cools below the 160°F or 230°F value, the
wing anti-ice valve automatically reactivates if its switch is ON. This
OFF–ON activation is called “cycling.” AFM “Abnormal Procedures”
must be consulted.
• The wing overtemperature sensors are activated with or without wing

anti-ice selected ON.
6. Deice boots (Figure SRX-26):
• The tail deice system for the horizontal stabilizer is a pneumatic boot
system.
• Engine bleed air, service air (23 psi), is used to inflate and deflate the
boots.
• The system consists of a control switch in the cockpit, dual timer/logic

PC boards, two control valves, two pressure switches, two rubber deice
boots, and deice annunciators.
• The TAIL AUTO–OFF–MANUAL switch is on the ANTI ICE/DEICE

switch panel. The system is powered by main DC power through the
TAIL DEICE circuit breaker on the left CB panel.

NOTE:
XL USES A SINGLE LOGIC BOARD
XL/XLS
LOGIC
AUTO MODE–ONE
18 SECOND CYCLE BOARDS
EVERY 3 MINUTES L R PRECOOLED
SERVICE BLEED
AIR PRESSURE
TAIL (ENG OR APU)

TAIL
DEICE
AUTO
5 OFF 23 PSI
PRESSURE
REGULATOR
MANUAL
LEGEND VACUUM
BELOW
RIGHT GENERATOR 16 PSI
16 PSI
PRESSURE
VACUUM PRESSURE SWITCH
SERVICE AIR
P P
16 PSI & ABOVE
COMBINATION VACUUM
EJECTOR/SOLENOID VALVES (NC)
L BOOT R BOOT
SRX-57
Figure SRX-26. Tail Deice System

• Selecting AUTO starts the automatic 18-second inflation cycle. The left
boot inflates during the first 6 seconds (white TL DEICE PRESS–L
advisory light illuminates), and then returns to the vacuum position,
extinguishing the annunciator light. After a 6-second pause, the right
boot inflates (white TL DEICE PRESS–R advisory light illuminates)
during the last 6 seconds, then extinguishes. Approximately 3 minutes
later, the cycle repeats itself.
• Placing the control switch to MANUAL, bypasses the timer logic and
simultaneously inflates both deice boots. The boots remain inflated as
long as the switch is held in the MANUAL position. Recommended
inflation time is 6 to 8 seconds and should be repeated at 3- to 5-minute
intervals as long as icing conditions are encountered. Both white
advisory TL DEICE PRESS L–R lights illuminate simultaneously as
both boots inflate.
• The amber TL DEICE FAIL L–R annunciator illuminates flashing if the

control switch is in AUTO and boot inflation pressure does not reach 16
psi, or the boots do not cycle properly due to timer or control valve
failure. The annunciator also illuminates flashing if the control switch has
been turned OFF and the timer and/or control valve(s) are still energized.
• Loss of normal DC power prevents tail deice function.
• The tail deice boots should not be activated at indicated RAT

temperatures below –35°C/–31°F at airspeeds at or above 150 KIAS or
–40°C/F at airspeeds below 150 KIAS. Boot cracking may result.
PNEUMATICS/AIR CONDITIONING
• Hot, P3 engine bleed air is used for environmental/pressurization, wing
anti-ice, and service air. This bleed air must be reduced in temperature by
use of a cross-flow type heat exchanger or “precooler.” Engine anti-ice
does not use precooled air; it uses raw bleed air off the side of the engine.
• On ground and during flight, the XLS uses cool engine fan bypass air
to flow into the precooler and extract heat from the engine hot bleed air
as it flows through the precooler and into the bleed-air manifolds. The
targeted bleed-air temperature after exiting the precooler is 475°F
(246°C) (Figure SRX-27).
• The amber BLEED AIR O’HEAT L–R annunciator illuminates if the

bleed-air temperature exiting the precooler reaches 560°F (293°C). This
overheat condition automatically causes the affected side wing anti-ice
shutoff valve to close, if open, or not open if later selected for use.
• Excessively hot bleed air exiting the precooler can be shut off by
selecting the opposite sides source with the source selector knob.
• When operating at or above maximum cruise thrust (CLB detent) and

RAT is approximately 0°C or warmer, selection of WING/ENGINE
ANTI-ICE ON may cause the BLEED AIR O’HEAT L–R annunciator
to illuminate. The situation is not hazardous and corrects itself within a
few seconds.

XLS PRECOOLER OPERATION ON GROUND & IN FLIGHT
LEGEND

FAN AIR
VALVE
P3 ENGINE
PRECOOLER BLEED AIR FAN AIR
CROSS-FLOW
MIXER
EXHAUST
VENT
GND
DC
AIR PRECOOLER
CONTROL
560°
BLD AIR
TO SYSTEMS L R
SRX-59
Figure SRX-27. Engine Bleed-Air Precooler

The pressurization source selector has the following positions:
• OFF—All valves are closed; bleed air is still available for service air
and anti-ice/deice.
• LH—The left flow control valve is relaxed open; the right flow control
valve is energized closed. The ACM receives air from the left engine
• NORMAL—The left and right flow control valves are relaxed open
(this is the fail-safe condition of the system), providing normal airflow
from both engines to the ACM (12 ppm total airflow).
• RH—The right flow control valve is relaxed open; the left flow control
valve is energized closed. The ACM receives air from the right engine
• EMER—The emergency pressure flow control valve is energized open;

the left and right flow control valves are energized closed. Airflow to
the ACM is stopped and control of temperature is with the left throttle.
Emergency pressurization is not available on the ground. Emergency
pressurization is provided by the left engine only and airflow is
diverted to the forward portion of the dropped aisle ducts on the left
side of the cabin.
Temperature Control:
• Normal DC power is required for automatic and manual modes.
• Temperature control is provided separately for both cockpit and cabin.
• Temperature is controlled in the cockpit and cabin by mixing constant

temperature cool air from the ACM system with unconditioned warm
bleed air. This moderated warm air is directed to the warm air
distribution network for the cockpit and cabin areas.
• Cold air for the overhead WEMACS is supplied directly from the ACM
system.
• Components common to the cockpit and cabin temperature control

systems include temperature control valves (TCVs), mixing muffs,
temperature sensors, zone sensors, duct overheat switches, and the
temperature controller.
• AIR DUCT O’HEAT CKPT/CAB annunciators illuminate if the aft

end of the respective air supply duct (cabin or cockpit) temperature has
overheated (300°F/149°C).
• An additional source of cold or hot air is provided by the auxiliary

power unit (APU) (Figure SRX-28).

CKPT TEMP SEL CABIN TEMP SEL
AIRDUCT
AUTO 7 1 AUTO O'HEAT ENG P3
BLEED AIR
CKPT CAB
COLD HOT COLD HOT CKPT CAB
SEL SEL
PRECOOLER
SUPPLY SUPPLY
MANUAL MANUAL

T
OZONE
CONVERTER
T T T APU
COCKPIT ARM REST
ZONE Z
SENSOR FLOOR
R FLOW
CONTROL
TCV (16 PSI)
(NO)
FOOT WARMERS
COCKPIT AREA
WEMACS
WATER SEPARATOR APU BAV
TCV
T
EMER
ACM
PRESS
T
WEMACS TCV ACM
O'HEAT
CABIN ZONE
SENSOR Z
AISLE
MIXING (NO)
FLOOR
MUFF
ARM REST BLD AIR
T
O'HEAT
475°F
LEGEND T
EMER L R
EMER (PRSOV) (NC) 560°F
PRECOOLED BLEED AIR PRESS
ANTI-SKID
COLD ACM AIR INOP
CABIN/COCKPIT UNDER-FLOOR DUCTING
STATIC FLOW
SRX-61
Figure SRX-28. Air Conditioning System

PRESSURIZATION
• Normal DC power and 23-psi (service air) air/vacuum are required for
both AUTO and ISOBARIC MODE operation. AUTO mode also
requires input from the No. 1 ADC (Figure SRX-29).
• ISOBARIC MODE is the result of loss of No. 1 ADC input. It is indicated

by an amber LED on the face of the cabin pressure controller. Control
pressurization using FL and CA mode in the altitude select window.
• MANUAL MODE can be selected at any time. It requires no DC

power source nor normal vacuum to operate. Cherry picker uses cabin
pressure for closing the outflow valves and nose wheel well low
pressure static air for vacuum to open the outflow valves. Will not
override the 14,500 ± 500 cabin altitude limiter valves, nor will it
override the MAX cabin differential protection of 9.5 ± 0.1 psid.
• Provides a sea level cabin to 25,230 feet, with a 9.3 ± 0.1 psid.
Provides a 6,800 feet cabin altitude at 45,000 feet.
• Normal DC power and vacuum is required to operate the cabin dump

valve. Cabin dump will not override the cabin altitude limit valves.
• Airplane is depressurized on the ground (left squat switch) with less

than 85% throttle position angle.
• Above 85% throttle position angle on the ground, the pressurization

controller goes into the takeoff/prepressurization mode. This increases
cabin pressure to preclude pressure bumps upon takeoff. Normal
prepressurization will descend the cabin to approximately 50 feet
below takeoff field elevation during the takeoff run.
• The pressurization controller (Figure SRX-30) provides both normal

(8,000 feet and below) and high altitude (8,100 to 14,000 feet)
autoschedule pressurization control, depending upon what value set
landing pressure altitude is selected in the SET ALT window of the
controller (Figure SRX-31). High altitude mode prevents the CAB ALT
annunciator from illuminating at a cabin pressure of 10,000 feet while
descending or climbing below FL250 (Figures SRX-32 and SRX-33).
However, it will illuminate if cabin altitude should reach 14,500 feet.
• The controller is programmed to limit cabin climb and descent rates to

+600/–500 fpm respectively.
• High altitude mode climb and descent rates are limited to a maximum
of +2,500/–1,500 fpm respectively.

PRESS SYSTEM SELECT
UP EMER DUMP 15 20
MANUAL M ON
10 25
A SET ALT 4 5 6
3 PSI 7
N FL EXER 8 30
2
U 1 9
A 5 0 10 35
AUTO L DIFF 40
NORM PRESS 50
DOWN 0
CABIN ALT
RATE X1000 FT
DEPRESSURIZE CABIN BEFORE LANDING

OUTSIDE
STATIC
28 VDC SOURCE
#1 AIR CABIN AIR

DATA
COMPUTER
NOSE WHEEL
WELL STATIC HIGH
SOURCE ALT
SIGNAL CABIN AIR
VACUUM
EJECTOR
> 1.5 PSID
CABIN AIR
CABIN AIR SHUTTLE

VALVE CABIN AIR
1.5 PSI
ORIFICE CABIN AIR
OUTSIDE
STATIC
SOURCE VACUUM
FLAPS
CAB ALT 23 PSI
LEGEND
UP 0°
BLEED AIR
TRIM TO
NOSE
DOWN
CLB T.O.
200 KIAS 7° STATIC PRESSURE
CRU
T T
O H T.O. &
R APPR 15°
O
T
T
200 KIAS
SERVICE AIR
L
E
NOSE
UP IDLE
SPEED
CUT
LAND
175 KIAS
35°
CABIN AIR
BRAKE
OFF
ENGINE SYNC
LH RH MUST BE OFF
RETRACT & LANDING
VACUUM
SRX-63
EXTEND
Figure SRX-29. Pressurization System

SRX-64
ANTI-ICE / DEICE PRESSURIZATION
PRESS SYSTEM SELECT

PITOT & WINDSHIELD WINDSHIELD
WING INSP UP EMER DUMP 15 20

STATIC AIR
ON L
O'RIDE
R ON ON MANUAL M ON 18 10
4 5 6
25
A
SET ALT 3 PSI 7

ON N FL EXER 8 30
2
U 1 9
A 5 0 10 35
OFF L DIFF 40
OFF OFF OFF AUTO
DOWN
NORM 0 0
PRESS 50
CABIN ALT
WNG XFLOW WING/ENGINE RATE X1000 FT
TAIL
ON L ON R AUTO DEPRESSURIZE CABIN BEFORE LANDING
PRESS SOURCE
OFF OFF NORM
ON CKPT TEMP SEL CABIN TEMP SEL
LH RH
OFF ENGINE MANUAL
LIGHTS
PASS GND REC/ AUTO AUTO
SAFETY NAV TAIL FLOOD ANTI-COL OFF EMER
ON ON ON ON
GND CKPT CAB
OFF REC COLD HOT COLD HOT
ON SEL SEL
SEAT BELT OFF OFF OFF SUPPLY SUPPLY
ON MANUAL MANUAL
Figure SRX-30. Pressurization Control Panel

Max Delta P Limit Autoschedule Boundary
45000

35000
Aircraft 30000 Descent to SLA

Altitude
25000
(FT) Climb to FL410
20000
Cabin @ SLA
15000 1500 ft above SLA
Take off
10000 from 1000 FT
5000 Negative Delta P Limit
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)

SRX-65
Figure SRX-31. Auto Schedule Boundary

SRX-66
45000
Aircraft climbs to
35000

Cabin Holds @ 78000 ft until
Cabin Climbs Acft descends below FL 245

30000
Aircraft to and maintains
7800 ft. at 600 FPM
Altitude 25000 Cabin Climbs to
Landing Field
(FT)
20000 (NLT 1500 AGL)
15000 Takeoff from

3000 ft
10000
5000
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)
Figure SRX-32. High Altitude Landing Graph

45000
35000

Cabin will reach 8000 ft with

30000 Acft at approx. FL 250
Aircraft
Altitude 25000
(FT) Descent to
20000
SLA
Climb to Takeoff from
15000
FL 450 14000 ft
10000
5000
0
0 2000 4000 6000 8000 10000 12000 14000
Cabin Altitude (FT)
SRX-67
Figure SRX-33. High Altitude Departure Graph

SERVICE AIR
• Bleed air supplied by the engines or APU.
• Regulated at 23 psi.
• Used for (Figure SRX-34):
• Horizontal stabilizer deice boots, inflation pressure.
• Pressurization outflow valve operation.
• Cabin entrance primary door seal and acoustic door seals.
• Throttle detents, EECs AUTO mode.

FLAPS
UP 0°
THROTTLE TRIM
NOSE
DOWN
TO
CLB T.O.
200 KIAS 7°
DETENTS
CRU
T T
O H T.O. &
R APPR 15°
O 200 KIAS
T
T
L
E
NOSE
UP IDLE
LAND 35°
175 KIAS
SPEED
BRAKE CUT
OFF
ENGINE SYNC
LH RH MUST BE OFF
RETRACT & LANDING
EXTEND
DOOR SEALS
VACUUM EJECTOR
FOR OUTFLOW VALVES
23 PSI
PRECOOLER REGULATOR
PRECOOLER
L FLOW ACM
CONTROL
VALVE P3 ENG
BLEED AIR
LEGEND APU
BAV
SERVICE AIR
VACUUM APU
BLEED AIR
BLEED AIR TO DEICE
SYSTEM
Figure SRX-34. Service Air System

OXYGEN
• A 76-cubic-foot bottle is standard and is in the right side of the lower
nose compartment (Figure SRX-35).
• The bottle pressurization green arc is marked from 1,600 to 1,800 psi.
This does not ensure oxygen availability to the crew or cabin. The
valve at the bottle must be checked safety wired open.
• Quick-donning EROS crew masks are stowed in a retainer below the

crewmember side windows. The masks have an integral microphone
and pressure regulator. Three positions are afforded: EMER (for
pressure breathing), 100%, and diluter demand. Masks must be stowed
properly to qualify as quick-donning masks.
• Passenger masks are stowed in overhead containers. Passenger oxygen

selector on the pilot console has three positions: OFF (crew only);
AUTO (masks will automatically drop if cabin pressure exceeds
approximately 14,500 feet, with normal DC power available); ON
(manual drop).
• With the OXYGEN selector in AUTO, if cabin altitude exceeds 14,500

feet, passenger masks drop automatically. If cabin pressure is restored
to normal values, the solenoid valve is deenergized at approximately
12,000 feet cabin altitude, shutting off oxygen flow to the passenger
masks.
• Oxygen cylinder is serviced through a service port in the lower aft sill
of the right nose compartment (aviator breathing oxygen only!).
• A green overboard discharge indicator (disc) is below the aft edge of

the nose compartment door. A missing or ruptured disc indicates the
oxygen cylinder has overpressurized and maintenance must be
performed before flight.

14,500 +/- 500
FILLER VALVE &
PROTECTIVE CAP
Cabin Altitude
OVERBOARD
COPILOTS
DISCHARGE FACE MASK
INDICATOR CYLINDER
PRESSURE GAUGE 5 AMP
OXYGEN
CB
28 - VOLT

DC
SHUTOFF
VALVE ALTITUDE
PRESSURE
SWITCH
PRESSURE
REGULATOR OVERHEAD
OXYGEN CHECK
CYLINDER VALVE DROP BOX
SOLENOID
PILOTS FACE
MASK
LEGEND
PASS OXY
OXYGEN SUPPLY (HI PRESS) ON OFF AUTO
OFF ON
OXYGEN CYLINDER
PASS OXY
OXYGEN SUPPLY (REG MED PRESS) AUTO
STATIC FLOW OXYGEN CONTROL VALVE

Oxygen System
SRX-71
Automatic Deploy
Figure SRX-35. Oxygen System
AUXILIARY POWER UNIT (APU)

The Allied Signal Model RE100-XL is a fully automatic, constant speed gas
turbine engine mounted in a titanium steel fireproof enclosure in the tail
cone. It utilizes a single-stage centrifugal impeller and a single-stage turbine.
The APU requires main DC power, fuel from the right tank, and control sig-
nals from the aircraft for operation. Service is conducted through the APU
panel on the right-rear fuselage above the engine pylon. The following lists
some general APU highlights:
• The APU is optional equipment. The unit is installed in place of the

standard vapor cycle air conditioner.
• APU installation increases aircraft weight by approximately 100

pounds plus any required nose ballast.
• The APU generator provides 28 VDC power and bleed air for ground
and in-flight use.
• Maximum altitude for in-flight start is 20,000 feet.
• Maximum altitude for in-flight operation is 30,000 feet.
• The APU produces no thrust.
• The APU is not certified for unattended use.
ELECTRONIC CONTROL UNIT (ECU)

The ECU is responsible for automatic APU operations. The ECU supplies the
following functions:
• The ECU is powered with the APU MASTER switch ON.
• Built-in test during power-up
• Automatic start control
• Speed control
• ECU autorelight function applies ignition at 94% rpm to prevent

flameout.
• Protective shutdown capability—Excessive EGT, rpm overspeed, low

oil pressure (LOP), high oil temperature
• Start inhibit capability
• Fault code storage
• Fault reporting to the field service monitor (FSM). The FSM provides
download capability.

FUEL SYSTEM
• Fuel is normally supplied from the right wing fuel tank except during
left to right crossfeed operations.
• Right boost pump operates continuously during APU start and APU
operation. If crossfeeding from left to right, the left boost pump
supplies fuel for APU operations (the right boost pump deenergizes).
• When the right boost pump is operating for APU operations only, the
amber FUEL BOOST–R annunciator does not illuminate.
• Fuel flow is 110 pph during loaded operation (generator online and
bleed valve open).
• Fuel flow indications are available in the FMS.
• APU fuel valve opens during the start sequence and closes for normal /
abnormal shutdown including APU fire.
OIL SYSTEM
• Oil reservoir is in the accessory gearbox. Oil quantity is approximately
1.5 US quarts.
• APU normally uses the same oil as the engines.
• Oil service is through the small door on the outside access panel.
• The oil reservoir is cooled with compressor intake
• APU oil level should be checked within 5 minutes after the APU has
been shutdown.
• The APU service panel in the tail cone is used to check oil level
electrically. Following a successful panel LAMP TEST, select PRE
FLT position:
• No illuminating lights indicates full oil.
• Amber illumination indicates 300 cc low of oil. APU operation is
permitted. Service at next opportunity.
• Red and amber illumination indicates 550 cc low of oil. APU
operation is prohibited. Oil service is required.
• APU service panel is battery bus powered.
• Low oil pressure (LOP) switch signals the ECU to initiate a protective
shutdown. The amber APU FAIL annunciator illuminates on the far
right cockpit panel (Figure SRX 36).
• High oil temperature signals the ECU to initiate a protective shutdown.
The amber APU FAIL annunciator illuminates.
• Magnetic chip collector is inspected by maintenance only.

MASTER MASTER
WARNING CAUTION
RESET RESET
HDG NAV APR BC VNAV ALT VS FLC
200
FMS LNAV VASEL VGP 100 300
160 AP ENG 9000 0 400

DC AMPS
160 E 20 20
6 APU RELAY
10 10 4 ENGAGED
140 10000 2
APU FAIL
6 20
1
125 98 00
120
4 10 10 APU
2 1
2 FIRE
N560FS
R 20 20 4
100 9500 6
1
–950
30 WX BARO PRE
BARO HSI TCAS NAV FMS
MIN TERR RAD VIEW
.261 M 200 STD
AOA DME
HDG FMS1 RW0IL NAV ADF NAV ADF
349 023 VOR1 OFF FMS OFF FMS
329 13 23.0 NM PUSH
PUSH
STD
.50 12 MIN OFF TO TEST
BARO
33 N 157 KTS BRG PFD DIM MINIMUMS SET BRG
KHUT FMS STATUS
BRG PTR
WPT
3 COM 1 COM 2 CABIN EMER
FMS1 +13 MSG
30
002
ADF APPR
ET MICROPHONE
DR I
0:00:00 WPT N
001 WIND P
CLOCK H
-04 NAV 1 NAV 2
19:39:07 5 39 COM 1 COM 2 ADF 1 ADF 2 DME 1 DME 2 BOTH
WEATHER V
I O
WX/R/T TAWS D I
C
T4.5° A E
STAB TGT RA 9.8NM+13 TERRAIN MLS 1 MLS 2 MUTE MKR
LX/ON TA 4.5NM-04 INHIBIT S
P
H
D
K P
R H
COCKPIT VOICE
Honeywell RECORDER
HOLD
5 SEC
TEST HEADSET ERASE
CKPT RECIRC AHRS 2

HI DC L SLEW MIC OXY MASK
HDG REV ATT REV ADC REV
S
O L
F A
F V
E
LO TEST R SLEW MIC HEAD SET
Figure SRX-36. APU Annunciators—Copilot Panel
PNEUMATIC SYSTEM
• A main duty of the APU is to provide supplemental bleed air to the
aircraft environmental/pressurization and all service air systems.
• The ACM, TCVs and underfloor ducting, and deice boots are major
APU bleed-air users.
• The APU cannot supply bleed air to the anti-ice systems.
• Bleed air from the APU is supplied through a bleed-air valve (BAV)
(see Figures SRX-28 and SRX-34).
• The BAV is controlled by the ECU and the BLEED AIR MAX
COOL–ON–OFF switch on the APU control panel.

• After start, with the BLEED AIR switch ON, the ECU opens the BAV
halfway and supplies regulated bleed air to the aircraft bleed-air
manifold. When the BAV is open (or other than closed), the white
BLEED VAL OPEN annunciator illuminates.
• MAX COOL switch selection opens the BAV full open.
• If an ACM O’HEAT condition occurs, the ECU commands the BAV to

the mid position until the condition is cleared.
• Bleed air is regulated by the ECU according to EGT and inlet ambient
temperatures. As EGT increases, bleed air is reduced to maintain a safe
EGT. If EGT reaches 690°C, the ECU provides a protective shutdown
(Figure SRX-37).
ELECTRICAL SYSTEM
• One 28 VDC, 300 constant ampere starter-generator is on the gearbox.
APU generator load has priority over bleed-air load. The ECU reduces
bleed air as required to maintain 100% shaft rpm for generator
operation.
• The generator is controlled via its GCU and generator switch on the
APU control panel (Figure SRX-37). The GCU and three-position
generator switch that operates identically to engine generator switches.
• When selected online, generator current is supplied to the system at the

crossfeed bus (Figure SRX-38). The APU generator and the engine
generators can all simultaneously supply power to the aircraft bus
system.
• Generator load is indicated with an amperage gauge on the far right

side of the cockpit control panel (Figure SRX-36).
• Maximum generator loads (red lines) are 200A on ground and 230A in
flight up to 30,000 feet.
• The APU and engine generators are not interchangeable.

Figure SRX-37. APU Control Panel

FIRST ENGINE START (R) USING APU GEN & BATTERY - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
GCU GCU
APU 1
RELAY 7
L APU RELAY R 5
STARTER STARTER
GPU
APU GENERATOR FIELD GEN BATT RELAY OVER (32.5 VDC) GEN FIELD
RELAY
BATTERY GPU INTERIOR POWER
INPUT
SRX-77
Figure SRX-38. First Engine Start (R)—APU Generator On Line

FIRE PROTECTION
• Fire detection—Uses a gas-filled fire detection loop inside the fireproof
APU enclosure. As heat increases, the gas expands and causes a
pressure sensor to activate the red APU FIRE switchlight on the far
right cockpit control panel (see Figure SRX-36). Upon fire detection
the following occur:
• ECU automatically initiates an automatic shutdown.
• APU generator goes offline (field relay opens).
• APU fuel shutoff valve closes.
• Fuel boost deenergizes.
• The APU fire bottle is armed.
• An APU fire protective shutdown is stored in the ECU memory.
• Fire extinguishing—One dedicated fire bottle is above the baggage

compartment ceiling.
• Fire bottle arming by the ECU is indicated with the illumination of the
red APU FIRE switchlight. Pressing the red switchlight fires the
contents of the bottle into the APU compartment.
• If the red switchlight is not pressed by the crew, the ECU fires the
bottle 8 seconds after the light illuminates.
• The fire detection loop and bottle is continuously monitored by the

ECU. If the loop malfunctions or bottle becomes low, the ECU
automatically shuts down the APU or inhibits its start. The amber APU
FAIL annunciation illuminates for either malfunction.
EXTERIOR PREFLIGHT
• Check APU air inlets on the upper right rear fuselage—Check CLEAR
(compressor inlet, cooling inlet for the starter-generator).
• APU exhaust—Check CLEAR.
• Tail cone ram air inlet on the right rear fuselage below the pylon—
Check CLEAR.
• APU drains on the bottom of the rear fuselage.
• Check oil quantity lights on the service panel in the tail cone

APU CONTROL PANEL AND ANNUNCIATOR FUNCTIONS

Prior to APU starting, the aircraft battery switch must be ON (see Figure
SRX-37).
NOTE
APU starts on the ground may be aircraft battery
starts only, EPU starts only (battery disconnect relay
opens during start), or aircraft generator(s) assisted
battery starts.
In-flight APU starts are battery only starts (squat
switch logic prevents generator-assisted APU starts).
In-flight starts are prohibited above 20,000 feet.
In-flight APU starts are prohibited after dual gener-
ator failure.
APU MASTER switch—The MASTER switch is placed ON to provide elec-

trical power to the ECU. The ECU performs APU power-up tests. After the
power-up tests are completed, the ECU accomplishes the prestart built-in test
equipment (BITE) test to ensure no faults exist that would inhibit a start. If
a fault is detected, the APU FAIL light illuminates.
APU FAIL light—Illuminates for an APU fault of low fire bottle pressure. APU
start attempt is prohibited when the APU FAIL light is illuminated.
APU TEST button—Performs a lamp test of annunciators (FIRE WARNING,

APU FAIL, APU RELAY ENGAGED, BLEED VALVE OPEN, READY TO
LOAD), digital indicators (RPM-50, EGT-500, DC VOLTS-00.0) and in-
tegrity of the APU fire system.
APU GENERATOR switch—ON position allows generator power to connect

to the airplane crossfeed bus after the READY TO LOAD light illuminates.
OFF position disconnects the APU generator from the crossfeed bus. RESET
position allows a possible reset of an APU generator tripped field relay.
APU BLEED AIR VALVE switch—ON position opens the BAV valve to the
mid-position while MAX COOL position opens the BAV to full (BLEED VAL
OPEN illuminates with either position). OFF position closes the BAV. Prior
to shutdown, the APU should be unloaded. The APU BLEED AIR switch is
selected OFF. The BLEED VAL OPEN light extinguishes when the BAV closes.
NOTE
Any time the APU is operating, the service air sys-
tem is pressurized whether or not the bleed-air valve
is open or closed.
APU START/STOP switch—The ECU provides automatic starting after plac-

ing the MASTER switch “ON” followed by momentarily placing the APU
START switch to “START.” The ECU controls ignition and fuel automatically
during start as required for ambient conditions.

The aircraft right boost pump activates (FUEL BOOST–R annunciator remains
extinguished; R LO FUEL PRESS extinguishes).
If the APU start is an engine generator(s) assisted start (ground only), the en-
gine start relay(s) close (engine start button(s) illuminates), and the APU start
logic commands the battery isolation relay open to protect the 225-amp cur-
rent limiters.
At 5% rpm, the ECU powers the ignition unit, fuel torque motor, and the APU
fuel solenoid valve (open). During start, the ECU controls fuel scheduling,
and continually monitors engine speed and EGT limits as determined by am-
bient conditions (T2). If scheduled limits are exceeded, the ECU executes a
precautionary shutdown (APU FAIL illuminates). The fault code is stored in
memory for ease of maintenance during troubleshooting.
The STOP position initiates a simulated overspeed signal to the ECU to ini-
tiate an immediate shutdown. After commanding shutdown using the APU
START–STOP switch, the ECU remains powered until the APU MASTER
switch is placed OFF.
Following an APU shutdown for any reason, a restart must not be attempted
until 30 seconds after the rpm indicator displays 0%
APU RELAY ENGAGED light—Illuminates then extinguishes prior to the
READY TO LOAD light illuminating. At 50% speed, the speed sensor signals
the GCU to deenergize the start relay and the APU RELAY ENGAGED extin-
guishes. If the speed sensor fails and/or the GCU fails to open the start relay
at 50%, the ECU backs up the GCU and opens the start relay at 60% rpm.
READY TO LOAD light—At 95% rpm the start counter records the start.
At 95% rpm plus 4 seconds, the ECU shifts to onspeed control. The READY
TO LOAD illuminates (start is complete). The APU may now be loaded elec-
trically and pneumatically.
At 99% rpm, the ignition unit is deenergized.
At 100% rpm, the APU is considered onspeed. At 100% rpm, the ECU main-
tains constant rotor speed rpm at 100% plus or minus 1.0% (70,200 rpm), EGT
within limits and the DC VOLTAGE indicator should display 28.5 VDC.
If APU speed drops below 94%, the ignition unit automatically reenergizes,
unless the APU is in a protective or normal shutdown mode.
The programmed ECU onspeed EGT and overspeed shutdown limits are es-
tablished at 690°C (1275°F) and 108% respectively.
APU GENERATOR—After the READY TO LOAD illuminates, the APU gen-
erator may be placed online. Placing the APU generator switch ON, energizes
the APU generator power relay to connect the APU generator output to the
airplane crossfeed bus. The APU ammeter on the copilot instrument panel
should reflect an amperage load.
APU GEN OFF light—Indicates the APU generator relay is open with the APU
running onspeed.

APU AMMETER—200 amp maximum on ground, 230 amp maximum in flight.
HOBBS METERS—At the bottom of the APU panel. Begins recording APU
operation when normal oil pressure is sensed by the ECU. This meter is used
for generator maintenance.
APU FIRE light/button—Alerts the crew of an APU fire in the APU enclo-
sure. APU immediately shuts down. Pressing the button activates the extin-
guisher. Extinguisher automatically activates 8 seconds after the light
illuminates if the button is not pressed.
START POWER LOGIC

• APU BATTERY START ON-GROUND – GPU or engine generators
are not available (Figure SRX-39).
• Battery supplies current through the APU relay for starting.
• APU GPU START ON GROUND—GPU is connected and available

(Figure SRX 40).
• GPU supplies current for start through the APU relay. The battery
disconnect relay opens and the battery does not supply current.
• FIRST ENGINE START, APU GENERATOR ON LINE—APU

generator and battery available (see Figure SRX-38).

APU relay and the isolation relay.
• Battery current is supplied through the APU relay for engine start.
• SECOND ENGINE START ON GROUND, APU AND ENGINE

GENERATORS ON LINE—Both generators and the battery are
available (Figure SRX-41):

APU relay.
• Engine generator current is supplied for cross start through the left
and right start relays.
• The battery supplies engine start current through the left and right
start relays.

SRX-82
APU START - USING BATTERY ONLY - ENGINE GENERATORS NOT AVAILABLE - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
GEN

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B
START START I
GCU GCU
APU 1
RELAY 7
L APU RELAY R 5
STARTER STARTER
GPU
BATTERY FIELD GEN BATT RELAY OVER (32.5 VDC) GEN FIELD
RELAY
GPU INTERIOR POWER
INPUT
Figure SRX-39. APU Start On Ground (Engines OFF)

APU START - USING GPU - ENGINE GENERATORS NOT AVAILABLE - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
GEN

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B
START START I
GCU GCU
APU 1
LEGEND RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
BATTERY STARTER STARTER
GPU
EXTERNAL DC RELAY DISCONNECT VOLTAGE RELAY
RELAY
INTERIOR POWER
GPU
SRX-83
Figure SRX-40. APU Start On-Ground—GPU (EPU)

SRX-84
SECOND ENGINE START (L) USING R ENG GEN, APU GEN, & BATTERY - ON GROUND - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
28.5
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
GCU GCU
LEGEND APU 1
RELAY 7
L APU RELAY R 5
BATTERY BATTERY ENGAGED A
STARTER STARTER
GPU
APU GENERATOR DISCONNECT
RELAY VOLTAGE RELAY
RELAY
GPU INTERIOR POWER
R GENERATOR INPUT
Figure SRX-41. Second Engine Start (L)—APU Generator On Line

• APU START ON GROUND, ENGINE GENERATORS ON LINE—

Both engine generators and the battery supply current for APU start
(Figure SRX 42).
• Engine generators supply current for APU start through the

respective start relays and the APU relay.
• Battery isolation relay opens to prevent current travel
• The battery supplies engine start current through the APU relay.
• APU START IN FLIGHT—Battery current only is available during

flight (Figure SRX-43):
• Battery supplies APU start current through the APU relay.
• Battery isolation relay opens to prevent engine generators from

participating.

SRX-86
APU START USING L & R ENGINE GENERATORS & BATTERY - GROUND ONLY - AVIONICS OFF
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
ON

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET V RESET R
B
START START I
LEGEND GCU GCU
APU 1
R GENERATOR RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
STARTER STARTER
L GENERATOR GPU
BATTERY RELAY
GPU INTERIOR POWER
INPUT
APU GENERATOR
Figure SRX-42. APU Start On Ground (Generator Assist)

APU START USING BATTERY ONLY - IN-FLIGHT - AVIONICS ON
ENGINE START
L DISENGAGE R
EMER SYS EMER AVN
START
50A 50A
GEN

OFF
GCU
RESET
L - AVN R - AVN
APU BUS BUS
STARTER 100
200
300
GEN AVN APU AVN

FIELD PWR GEN 0 400
PWR
RELAY RELAY RELAY
DC AMPS
RELAY
DC VOLTS
INTERIOR
0.0
60A 225A 225A 60A
MASTER

A A
SWITCH SWITCH E
BATT
ON 25A
U OFF R N
OFF OFF T
28.5 EMER 28.5 E
RESET 25V RESET R
B START START I
GCU GCU
APU 1
LEGEND RELAY 7
L APU RELAY R 5
BATTERY ENGAGED A
L GENERATOR STARTER STARTER
GPU
R GENERATOR RELAY DISCONNECT VOLTAGE RELAY
RELAY
GPU INTERIOR POWER
INPUT
BATTERY
SRX-87
Figure SRX-43. APU Start—In Flight (Battery Only)

APU OPERATING LIMITATIONS

1. APU operation is prohibited until a satisfactory APU test has been
accomplished as contained in the “Normal Procedures” section.
2. Starting the APU is prohibited whenever the APU FAIL LIGHT is

illuminated.
3. APU start attempt is prohibited after a dual generator failure.
4. Following shutdown for any reason, APU restart must not be attempted
until 30 seconds after the RPM indicator reads 0%.
5. Applying deice (anti-ice fluid of any type) is prohibited with APU

operating.
6. Deployment of the thrust reversers for more than 30 seconds with the APU
running is prohibited.
7. The APU is not approved for unattended operation.
8. The limits in Table SRX-1 apply to APU starting and operation:
Table SRX-1. APU OPERATING LIMITS
OPERATING MAX MAX NI% FUEL MAX GEN AMBIENT

CONDITION: ALT FT EGT °C TEMP°C LOAD AMPS TEMP °C
(NOTE 3) (NOTE 2)
STARTING 20,000 690 — Refer to basic AFM — –54 to 54
fuel limitations
RUNNING 30,000 690 108 Refer to basic AFM 200 GND
fuel limitations (NOTE 1) –54 to 54
230 FLT
NOTES:
1. Transient current greater than 200 amperes is approved for APU cross generator start of the main engines.
2. APU Ammeter Instrument Markings:
a. Red Triangle = 200 amperes
b. Red Line = 230 amperes
3. APU automatically shuts down if EGT limits are exceeded.

BATTERY AND APU STARTER CYCLE LIMITATIONS

Starter Limitation
Three APU start cycles per 30 minutes. Three cycles of operation with a 90-
second rest period between cycles is permitted.
Battery Limitation
Nine APU start cycles per hour. (An APU battery start counts as 1/3 of a nor-
mal engine battery start.)
Starting the main engines using a generator cross start from the APU counts
as 1/3 of a normal engine battery start.
NOTE
1. No battery cycle is counted when starting the
APU from a ground power unit.
2. Use of an external power source with voltage in

excess of 28 VDC or current in excess of 1000
amps may damage the starter. Minimum 800 amps
for start.
3. If battery limitation is exceeded, a deep cycle

including a capacity check must be accomplished
to detect possible cell damage. Refer to Chapter
24 of the Maintenance Manual for procedure.

AVIONICS
All primary avionics systems and components are DC-powered XLS Primus
1000 Control Display System. Sensor inputs include (Figure SRX-44):
• Dual Litef LCR-93 attitude and heading reference system (AHRS).
• Dual microair data computers (MADC).
MADCs are powered by ADC 1 and 2 circuit breakers on the right CB panel
and provide the following data to the high level data link control bus (HLDC):
• Pitot pressure, total and static air temperature for TAS/CAS to the IC-
615s for PFD airspeed tapes, MACH and VMO/MMO indications and
warning horn.
• Static pressure, pressure altitude, and baro-corrected altitude (inches or

hPa) for the PFD altitude tapes.
• Altitude change rate for the PFD vertical velocity indicators.
• TAS data for the FMS and MFD.
• Pressure altitude to the TCAS and EGPWS.
• Altitude information to the Kollsman pressurization controller (ADC

No. 1 only).
• Also output data for the transponder, flight data recorder, flight
director, and autopilot.
The true airspeed (TAS) temperature probe (Rosemount) provides tempera-

ture data to the MADCs only. It is electrically anti-ice protected any time the
airplane is weight-off-wheels and the avionics master power switch is on.
The RAT gauge source temperature is provided by normal DC from the EEC
temperature sensor (T TO. probe) in the right engine inlet. If the right T TO probe
fails, No. 2 MADC automatically provides temperature information to the RAT
indicator.
INTEGRATED AVIONICS COMPUTERS (IAC):

• Dual IC-615 computers provide data processing for the pilot and
copilot EFIS system. Normally, IAC No. 1 powers the pilot PFD and
MFD; the No. 2 IAC powers the copilot PFD.
• Both IACs contain a sensor interface and flight director computers.

Only the No. 1 IAC contains an autopilot computer.
NOTE
Each display unit houses its own symbol generator.

AHRS #1 AHRS #2
#1 #2
A A
H ATT MICRO AIR DATA H
ATT
R COMPUTERS R
U HDG HDG U

FLUX FLUX
VALVE VALVE
AUTOPILOT
SERVOS
DIGITAL DATA BUS
IAC IAC
#1 FD/AP #2
PFD 1
FD/AP
PITCH PFD 2
IC 615 IC 615
SENSOR INTERFACE SENSOR INTERFACE
FD COMPUTER FD COMPUTER
AUTOPILOT COMPUTER
ROLL HSI TCAS WX
TERR
BARO
RAD
PRE
VIEW
NAV FMS
MAP
PLAN
TCAS
WX
TERR
NORM EMER HSI TCAS WX
TERR
BARO
RAD
PRE
VIEW
NAV FMS
INC
NAV ADF NAV ADF ET1 ET2 PUSH TO ENTER RCL SKP NAV ADF NAV ADF
R
OFF FMS OFF FMS N OFF FMS OFF FMS
PUSH G PUSH
PUSH STD PUSH STD
OFF TO TEST DEC ST1 ST2 DATA
PAG ESC OFF TO TEST
OFF
BARO SET BARO
BRG PFD DIM MINIMUMS SET BRG MFD DIM Honeywell BRG PFD DIM MINIMUMS SET BRG
VAPP LNAV VNAV AL-VN 2000 FMS LNAV VASEL VGP

PFD MFD
242 YD OFF AP OFF CAT2 1400 160 AP ENG 9000
920 HDG FMS1 RWO1L
280
329
349 23.0 NM
1000 12 MIN 160 E 20 20
20 20 FMS
6
260 4 N 39 6
2 33 10 10 4
10 10 1500 1 140 10000 2
3 80 KHUT
3
6 1
242
240 13 60 20
30
YAW 1 1
WPT
002
125 98 00
120
10 10 4 2 10 10 1
2
2
6
220 4 WPT
6 001 R 20 20 4
W
20 20 OM I 100 9500 6
1
–950
RAD 25.0 30
200 MIN BARO
.750 M 2500 29.92 IN MIN
AOA DME
.261 M 200 STD
HDG FMS1 H ICT AOA DME
VOR1 RW0IL HDG FMS1
329 349 023 13 VOR1 RW0IL
.70
23.0 NM RW01L
9.9L 329 349 023 13 23.0 NM
12 MIN DME1 TCAS DME2
.50
ICT ICT 12 MIN
33 N 157 KTS
13 NM 13 NM N 157 KTS
BRG PTR
3 FMS STATUS
TCAS TEMP 33 KHUT FMS STATUS
3
30
MSG RAT +6°C BRG PTR

FMS1 ABOVE +13 SAT –1°C +13 WPT MSG
FMS1
30
ADF APPR 002
6
RELATIVE 40 SPEED
APPR
W
ET DR TCAS AUTO TAS 162 ADF

-04 ET DR
0:00:00 CLOCK GSPD 157
WIND 0:00:00 WPT
E
CLOCK 19:37:47 DEST 001

24
CLOCK WIND
19:39:07 39 WEATHER RW01L -04
WEATHER WX/R/T 12 MIN 19:39:07 5 39
12
WX/R/T WEATHER
21 TAWS T4.5° A TAWS
WX/R/T
T4.5° A S 15 STAB TGT TERRAIN TAWS
STAB TGT TERRAIN INHIBIT T4.5° A
INHIBIT LX/ON STAB TGT TERRAIN
LX/ON RA 9.8NM+13 INHIBIT
LX/ON TA 4.5NM-04
Honeywell Honeywell
Honeywell
PFD MFD PFD

* SYMBOL GENERATORS ARE LOCATED COURSE 1 HEADING COURSE 2
INTERNALLY IN THE DISPLAY UNITS (DU) PUSH DIR PUSH SYNC PUSH DIR Honeywell
SRX-91
Figure SRX-44. XLS Primus 1000 CDS System Block Diagram

• HDG, ATT, and ADC REV buttons enable the respective IAC to utilize
the other IAC AHRS or MADC data in the event of failure, thereby
providing redundancy.
• The SG1/NORM/SG2 selector on the RI-552 (avionics) controller

allows either IAC to power all three displays in the event of IAC or
symbol generator failure.
COMPARISON MONITOR ANNUNCIATORS (9)

When PFD 1 and PFD 2 display the same type of information but from dif-
ferent sources, displayed data is compared (Table SRX-2). An exceeded
threshold (trip point) activates an amber comparison monitor on the PFD.
Table SRX-2. COMPARISON MONITOR ANNUNCIATORS (9)
ANNUNCIATOR PARAMETER TRIP POINT LOCATION

Pitch ±5° Upper right of
PIT
attitude sphere
ROL Roll ±5° Upper right of

attitude sphere
Attitude Both PIT and ROL Upper right of
ATT
monitors triggered attitude sphere
HDG Heading ±6° Bank <5° Top right of EHSI

±12° Bank >5° compass
A Barometric ±200 ft Top right of vertical

L Altitude altimeter tape
T
I Airspeed ±5 kts Top left of vertical

A airspeed tape
S
Radio Altitude Avg (RA1 + RA2) +10 At negative 10°

RAD
8 pitch mark
(attitude sphere)
GS Glide Slope ±50 µAmp Right of negative

Approximately 1 dot 10° pitch mark
LOC Localizer ±40 µAmp Right of EHSI’s

Approximately 1/2 dot center

PRIMUS 880 WEATHER RADAR

• X-band alphanumeric digital radar with a display designed to provide
weather location and analysis, as well as ground mapping.
• Can be operated in conjunction with the EFIS and MFD equipment to

provide radar video displays.
• The radar transmitter is normally disabled on the ground. However, by

rapidly depressing the STAB switch four times within 3 seconds, the
transmitter will radiate on the ground.
PRIMUS II RADIO SYSTEM

• Dual remote radio management units (RMUs). RMU 1 is powered by
the emergency DC bus and RMU 2 is powered by main DC power.
• COM 1, NAV 1, ADF 1, etc., are controlled by the left RMU. COM 2,
NAV 2, ADF 2, etc., are controlled by the right RMU. Reversion is
provided so each RMU can control all Primus II radios.
• VHF COMM is provided by the RCZ-850 integrated communications

unit. Operates in the frequency range of 118.00 to 136.97 MHz, and
can be strapped to extend the upper frequency range to 152 MHz. It is
8.33 khz spacing capable.
• VHF NAV is provided by the RCZ-850 integrated navigation unit.

Operates in the frequency range of 108.00 to 117.95 MHz. The system
encompasses the functions of VHF NAV, localizer and glide slope
receiver, and marker beacon receiver, as well as ADF and DME
functions.
• ADF NAV is provided by the DF-850 ADF receiver module, a

component of the RNZ-850 integrated navigation unit. Operates in the
frequency range of 100.00 to 1799.00 kHz in 0.5 increments.
• ATC TRANSPONDER function is provided by the XS-850

transponder module, a subunit of the RCZ-850 integrated
communication unit. It functions as a 4096 code mode A transponder,
as well as providing mode C (altitude) and mode S (collision
avoidance) information. Altitude information is provided by the
respective (1 or 2) AZ-850 microair data computer in the pilot or
copilot Primus 1000 system.
• Dual diversity transponders are modules in the RCZ-850 COMM unit

and are standard equipment.
• DME NAV function is provided by the optional Primus ll DME system

module. Each module is comprised of an RNZ-850 integrated navigation
unit, an NV-850 VHF NAV receiver, and a DME-850 distance measuring
module. The DME transmitter works in the L-frequency band and the
receiver frequency range is from 962 to 1213 MHz. Normal DME
function follows the VHF NAV receiver. However, a hold function allows

the tuning of military TACAN channels in order to receive the DME

portion of the TACAN signals. DME data is displayed on two DI-850
indicators; one on the pilot and one on the copilot instrument panels.
DME data can also be displayed on the pilot and copilot EHSIs.
• The STANDBY RADIO CONTROL (SRC) is normally on the center

instrument panel, to the right of the engine gauges. It contains normal
and emergency modes. The SRC is powered from the emergency DC
bus through the NAV1 circuit breaker. It acts as an additional tuning
source for the radio system (COM1 and NAV1).
RADIO ALTIMETER
• The Collins ALT-55B radio altimeter displays radio altitude up to an
absolute altitude of 2,500 feet. Altitude is displayed on the bottom
center of the attitude sphere of the EADIs. Between 200 and 2,500
feet, the display is in 10-foot increments. Below 200 feet, it is in 5-
foot increments.
• Decision height (DH) selection is displayed digitally in the lower right

side of the EADI display. The decision height range is from 0 to 990
feet in 10-foot increments. The DH display can be removed with full
counterclockwise rotation of the DH/TST knob on the DC-550 display
controller. A decision height warning horn sounds when the airplane
reaches the decision height set on the pilot EADI.
AUTOPILOT (AP)
• The autopilot and yaw damper are engaged by depressing the AP-
ENGAGE switchlight. With the flight director OFF, pitch and roll are
manually controlled with the turn knob and pitch wheel.
• With either of the dual MS-560 flight director (FD) modes selected, the
FD controls the autopilot.
• The autopilot may be switched to the pilot FD/PFD 1 or copilot

FD/PFD 2 by an illuminated selector switch (FD/AP–PFD1, FD/
AP–PFD2) on the center instrument panel.
• If a lateral mode is engaged, the autopilot follows the FD command bars.
• The autopilot/flight control system contains pitch, roll, and yaw servos
that control the airplane in accordance with manual or FD guidance to
the autopilot.
• The Primus 1000 IAC No. 1 contains the autopilot module for
autopilot control. Consequently, if IAC No. 1 fails, the autopilot is
inoperative.
• The autopilot may be temporarily disengaged by the touch control

steering (TCS) button on the yoke(s), but the yaw damper remains
engaged.

• The autopilot is normally disengaged one of four ways:

1. Depressing the AP/TRIM DISC red switch on either yoke.
2. Electrically trimming the elevator (yaw damper remains engaged).
3. Depressing the go-around button on either throttle.
4. Pressing the AP-ENGAGE button so the AP-ENGAGE light is not
illuminated.
• LOW bank limit may be selected manually by depressing the BANK

LIMIT–LOW switch on the controller (limits bank angle to 14°). Low
bank limit automatically engages climbing through 34,000 feet and
automatically disengages descending through 33,750 feet.
FLIGHT DIRECTOR (FD):

• The XLS Primus 1000 CDS system incorporates one flight director
(FD) in each IAC. Dual synchronized FD mode selectors above each
pilot PFD are used to control the FDs. Either crewmember may control
the airplane through the control panels by switching control with the
FD/AP–PFD 1 or 2 selector as discussed above. If the FD/AP control is
switched from one pilot to the other, the AP reverts to basic mode. The
FD will have to be reprogrammed.
NOTE
When the FD/AP is coupled to the VOR, another lat-
eral mode must be selected prior to switching VOR
NAV frequencies. HDG mode may be used after syn-
chronizing HDG bug to the current airplane heading.
Basic ROLL may also be used.
STALL WARNING AND AOA SYSTEM:

• The angle-of-attack system is powered by 28 VDC from the left main
DC bus and incorporates an angle-of-attack sensor, a signal summing
unit, a vane heater monitor, an angle-of-attack indicator, a stick shaker,
and an indexer.
• The full-range-type indicator on the PFDs indicate from 0.2 to 1.0 and
marked with red, yellow, and white arcs. Lift being produced is
displayed as a percentage and, with flap position information, is valid at
VREF (on-speed) and stick shaker initiation. All other points are for
reference only. The area at the lower part of the scale (0.57 to 0.2)
represents the normal operating range, except for approach and landing.
The narrow white arc (0.57 to 0.63) covers the approach and landing
range, and the middle of the white arc (0.6) represents the optimum
landing approach (VAPP or VREF). The yellow range (0.63 to 0.87)
represents a caution area where the airplane is approaching a critical
angle of attack. The red arc (0.87 to 1.0) is a warning zone. At an
indication of approximately 0.79 to 0.88 (depending on flap setting and
rate of deceleration) in the warning range, the stick shakers activate.

• If the angle-of-attack system loses power or becomes inoperative for

other reasons, a red “X” covers both AOA scales and AOA FAIL
annunciation is shown beside airspeed tapes on the PFDs.
NOTE
The airplane must not be flown if the stick shaker is
found to be inoperative on the preflight check or if
the angle-of-attack system is otherwise inoperative.
• Stick shakers are on the pilot and copilot control columns and provide
tactile warning of impending stall. The angle-of-attack transmitter
causes the stick shakers to be powered when the proper threshold is
reached.
WARNING
If the angle-of attack vane heater fails and the vane
becomes iced, the stick shaker may not operate or may
activate at normal approach speeds. AOA HTR FAIL
annunciates if this condition exists.
• The approach indexer on the pilot glareshield provides a heads-up

display of deviation from the approach reference. The display is in the
form of three illuminated symbols that indicate the airplane angle of
attack:
• When the airplane speed is on reference, the green center circle

illuminates.
• As the speed decreases from reference (.6 AOA), the circle

illumination dims and the top red chevron illumination increases until
the top chevron is illuminated and the circle is extinguished. The top
red chevron points down, indicating that the angle of attack must be
decreased to eliminate the deviation.
• When the airplane is accelerating from the onspeed reference, the

illumination of the green circle dims and illumination of the bottom
yellow chevron increases until the circle is extinguished and only the
bottom chevron is illuminated. The bottom yellow chevron points up to
indicate that the angle of attack must be increased to eliminate the
deviation.
• The indexer is active any time the nose gear is down and locked and the
airplane is not on the ground. There is a 20-second delay after takeoff
before the indexer activates.
• Stall strips on the leading edge of each wing create turbulent airflow
at high angles of attack, causing a buffet to warn of approaching stall
conditions. This system is considered a backup to the angle-of-attack
stick shaker system in case of malfunctions and electrical power
failure.

STANDBY INSTRUMENTS
Standby Flight Display (SFD, Goodrich GH-3000)
• The GH 3000 (3-inch display) is on the center instrument panel
between DU 1 (left PFD) and DU 2 (MFD). The standby HSI is just
below the GH 3000.
• The Goodrich GH 3000 features a DC-powered full-color, active

matrix LCD indicator; dimmable fluorescent backlighting and full
range of navigation interface capabilities, including the ADC-3000 air
data computer and the MAG-3000 magnetometer. It combines
precision altitude, heading, and airspeed/Mach indications into one
composite instrument (Figure SRX-45).
• Absent main DC power, the GH 3000 is powered by a Securaplane

10.5-amp/hr battery pack. The battery is on the lower rack of the
forward avionics compartment. This battery can power the SFD
approximately 180 minutes. A green test light (STBY PWR switch)
indicates at least 75% capacity.
SLIP/SKID
INDICATION
MACH
INDICATION BAROMETRIC
SETTING
VMO/MMO .71M 30.15 in

TAPE
60
10 10 1500
AIRSPEED 40
1
INDICATION
30 13 20
00
ALTITUDE
INDICATION
9 10
10 678M METRIC
1000 ALTITUDE
HEADING
TAPE 33 N
M
LIGHT MENU ADJUSTMENT

SENSOR BUTTON KNOB
Figure SRX-45. Standby Flight Display—GH 3000

• With the STBY PWR switch ON, the display operates using the menu
access button and adjustment knob. There are four main menus. Press
menu access button, rotate adjustment knob to:
• FAST ERECT—Press knob to initiate
• SET BRIGHTNESS OFFSET—Press knob for submenu, rotate

knob to adjust, press knob to finish
• NAV (ON or OFF)—Press knob to toggle for opposite of current

condition
• BARO TYPE—Press knob for submenu, rotate knob to select

type, press knob to finish
Standby Horizontal Situation Indicator (HSI):

• The standby HSI is a 3-inch instrument on the center instrument panel,
directly below the standby flight display tube. It provides navigational
guidance in case of PFD/flight director failure, and is powered by the
emergency bus (Figure SRX-46).
• The standby HSI displays compass heading (No. 2 AHRS), and

navigation inputs from NAV 1, (i.e., glide slope, localizer deviation,
and airplane position relative to VOR radials). The compass card is
graduated in 5° increments, and a lubber line is fixed at the fore and aft
positions. A fixed reference airplane is in the center of the HSI, aligned
with the lubber line markings. In addition, there is a course deviation
bar and course cursor, as well as a blue ADF needle that displays ADF
1 bearings and rotates around the outer portion of the dial (not available
with loss of normal DC power).
Figure SRX-46. Standby HSI

CLOCKS
• A digital Davtron (Model M877) clock on the center instrument panel
can display four functions: local time, GMT, flight time, and elapsed
time. Two versions of elapsed time may be selected: count up or count
down.
• The clock has two control buttons: SEL (select) and CTL (control). The
SEL button is used to select the desired function and the CTL button is
used to start and reset the selected mode.
• Enable the flight time mode with a landing gear squat switch, which
causes the clock to operate any time the airplane weight is off the
landing gear. The flight time may be reset by the pilots.
• Each PFD contains a digital clock showing GMT synchronized by the

FMS. Each PFD also contains an elapsed timer and countdown timer
controlled by ET1–2 and ST1–2 buttons on the multifunction controller
in the pedestal.
• ET1–2 starts, stops, and resets the elapsed timer and the countdown
timer. ST1 is used, along with the data set knob or the multifunction
controller, to set a countdown time.
TCAS II
• TCAS ll detects and tracks aircraft in the vicinity of your own airplane.
It interrogates the transponders of other aircraft and analyzes the
signals to range and bearing, and relative altitude if it is being reported.
It then issues visual and aural advisories so that the crew may perform
appropriate vertical avoidance maneuvers. TCAS control is provided
through the RMUs.
• The TCAS button on each display controller allows selection of TCAS

targets overlaid on the HSI in either ARC or MAP modes for either
pilot or copilot PFD.
• The TCAS button on the multifunction controller allows selection of

either MFD map overlay or MFD TCAS zoom window.
TERRAIN AWARENESS AND WARNING SYSTEM (TAWS):

• The standard Allied Signal enhanced ground proximity warning system
(EGPWS) provides visual and aural warnings of terrain in the
following basic EGPWS modes:
1. Excessive rate-of-descent with respect to terrain (Mode 1).
2. Excessive closure rates to terrain (Mode 2).
3. Negative climb before acquiring a predetermined terrain clearance

after takeoff or a missed approach (Mode 3).

4. Insufficient terrain clearance based on flap configuration (Mode 4).
5. Inadvertent descent below glide slope (Mode 5).
6. Minimums callout upon reaching DH (Mode 6).
7. SMART 500 callout—Altitude callout at 500 AGL (Mode 6).
8. Excessive bank angle alerting (Mode 6).
9. Windshear warning and windshear caution alerts (Mode 7).
In addition, the enhanced ground proximity warning system provides the fol-
lowing terrain map enhanced modes:
1. Terrain clearance floor exceedance.
2. Look-ahead cautionary terrain alerting and warning awareness.
3. Terrain awareness display. EGPWS provides display of

approximate terrain and obstacles. The terrain display is color and
intensity-coded (by density) to provide visual indication of the
relative vertical distance between the airplane and the terrain.
4. The EGPWS terrain overlay can be selected for display on either

PFD HSI or MFD MAP mode using the TERR button on either
display controller or multifunction controller.
AREA NAVIGATION
• Universal avionics systems UNS-1 Esp flight management system
(FMS) is a centralized control and master computer system, designed to
consolidate and optimize the acquisition, processing, interpretation, and
display of certain airplane navigation and performance data. The UNS-1
Esp FMS system may be installed as GPS only or multisensor system.
Digital air data information (including baro-corrected altitude and true
airspeed) and heading input is required of all installations.
• Each individual navigational sensor is specifically designed for primary

navigation. The FMS system takes advantage of the good properties of
a particular sensor, while minimizing its liabilities. The system
processes multiple range information from the DME, true airspeed data
from the air data computer, velocity and position information from the
long-range navigation sensors, and airplane heading, in order to derive
one best computed position (BCP).
• The FMS contains a memory capacity of up to 100,000 waypoints. The

stored Jeppesen database provides the capacity for complete coverage
for SIDs, STARs, approaches, high/low airways, navaids, IFR
intersections and airports with runways longer than 4,000 feet with IFR
approaches in the worldwide data base. It provides the capability for:

1. Pilot data storage

2. Company route data
3. Off-line flight planning
4. Fuel management monitoring
5. Frequency management
6. Lateral guidance and steering
7. Vertical guidance—VNAV (with FD/AP coupling authority)
• The Honeywell FMZ is optional.
LOCATOR BEACON
• The ELT 110-406 emergency locator transmitter (ELT) provides a
modulated omnidirectional signal, transmitted simultaneously on
emergency frequencies 121.50, 243.00, and 406 MHz. The system
activates by an impact of 5.0 +2/–0 g, or manually by a remote
ON–OFF switch forward of the left CB panel, and provides a GPS
navigation interface to transmit your position.
STATIC WICKS
• A static electrical charge, commonly referred to as P (precipitation)
static, builds up on the surfaces of the airplane in flight and causes
interference in radio and avionics equipment operation. The static
wicks are on the wing and empennage trailing edges, and dissipate
static electricity in flight.
• There are a total of 20 static wicks:
• One on each wingtip.
• Four on each wing trailing edge outboard of the aileron.
• One on the trailing edge of each aileron.
• Two on the trailing edge of each elevator.
• Two on the upper trailing edge of the rudder.
• One on the top of the rudder.
• One on the tail stinger.
• One or more missing static wicks cause radio P-static.
• Some static wicks may be missing for dispatch. Refer to AFM

“Normal Procedures” or MEL for conditions.

SRX-102
Antenna and drain tube locations are shown in Figure SRX-47.

ANTENNA AND DRAIN TUBE
Figure SRX-47. Excel Antenna and Drain Tube Locations
MAGNETOMETER NAV 1 & 2
FLUX VALVE
ADF 2 ACM AIR INLET

LOCATOR BCN
(OPTIONAL) COMM 1 APU EXHAUST

DIVERSITY TRANSPONDER 1 (LH SIDE) RH SIDE
DIVERSITY TRANSPONDER 2 (RH SIDE) LIGHTNING
ADF 1 DETECT APU AIR INLET

TCAS II UPPER RH SIDE
(OPTIONAL) SATCOM HF
GPS 1
GPS 2
(OPTIONAL)
RADAR
12 INCH
STORMSCOPE
(OPTIONAL)
ACM EXHAUST RH SIDE
APU FUEL DRAIN
TAILCONE FRESH AIR INLET RH SIDE
GLIDESLOPE ENGINE DRAIN
DME2 AFIS
DME1 BATTERY VENT
RADAR ALTIMETER HYDRAULIC RESERVOIR DRAIN
MARKER BEACON MAGNASTAR
TRANSPONDER 1 FWD LAVATORY
DRAIN REAR LAV / CONDENSER DRAIN
TCAS II LOWER
RADAR ALTIMETER
COM2
TRANSPONDER 2
GEAR BLOWDOWN
VENT
MASTER WARNING
CONTENTS
Page
ANNUNCIATORS........................................................................... MW-1
Master Warning Switchlights.................................................. MW-1
Master Caution Switchlights .................................................. MW-1
XLS ROTARY TEST ..................................................................... MW-17
EXCEL ROTARY TEST................................................................ MW-20
ACRONYMS ................................................................................. MW-22
FOR TRAINING PURPOSES ONLY MW-i

ILLUSTRATIONS
Figures Title Page
MW-1 XLS Annunciators ...................................................... MW-3
MW-2 Excel Annunciators...................................................... MW-5
MW-3 Rotary Test Knob ...................................................... MW-17
TABLES
Tables Title Page
MW-1 Master Warning/Caution Switchlights ........................ MW-2
MW-2 Master Warning Annunciators .................................... MW-7
MW-3 Auxiliary Annunciators ............................................ MW-14
MW-4 Thrust Reversers........................................................ MW-15
MW-5 Fire Switchlights ...................................................... MW-16
MW-6 APU Annunciators .................................................... MW-16
MW-7 Acronyms .................................................................. MW-22
FOR TRAINING PURPOSES ONLY MW-iii

MASTER WARNING
ANNUNCIATORS
Annunciators are classified as WARNING, CAUTION, and ADVISORY.
MASTER WARNING SWITCHLIGHTS

The crew is alerted to a warning condition with the two red MASTER WARN-
ING switchlights. Both MASTER WARNING switchlights flash when trig-
gered by a warning condition and must be immediately reset so they can
reilluminate if another warning occurs. Red and a limited number of amber
annunciators can trigger the red MASTER WARNING switchlights.
The MASTER WARNING switchlights do not automatically reset if trig-

gered, the switchlights must always be pressed to extinguish.
The illumination of a red LH or RH ENGINE FIRE light does not trigger the
MASTER WARNING switchlights.
The illumination of the red-flashing MASTER WARNING switchlights re-

quire the immediate attention and immediate action of the crew.
• Warning annunciators are red and are on the master warning panel.
When triggered, the red annunciators flash and cause the MASTER
WARNING switchlights to illuminate flashing. Amber annunciations
that cause the red MASTER WARNING switchlights to illuminate
include both amber GEN OFF L and R annunciators together and the
amber thrust reverser ARM and/or UNLOCK lights (in-flight only). The
previous conditions are considered serious and therefore activate the
MASTER WARNING switchlights. Pressing a MASTER WARNING
switchlight extinguishes both and causes the triggering red or amber
annunciator to cease flashing and illuminate steady.
MASTER CAUTION SWITCHLIGHTS

The crew is alerted to an amber caution condition with the amber MASTER
CAUTION switchlights. Both switchlights illuminate steady when triggered
by a “flashing” amber or white annunciator. Both MASTER CAUTION switch-
lights illuminate steady when triggered by a caution condition and must be
immediately reset so they can reilluminate if another caution occurs. Flashing
amber and white annunciators trigger the MASTER CAUTION switchlights.
The MASTER CAUTION switchlights will automatically extinguish and
reset should the triggering annunciator or condition be corrected or go away
before one of two switchlights is pressed.
FOR TRAINING PURPOSES ONLY MW-1

The illumination of the amber MASTER CAUTION switchlights requires the

immediate attention and subsequent action of the crew.
• Caution annunciators are generally amber and are on the master

warning panel. Exceptions include the white FUEL XFEED and the
white GND IDLE advisory annunciators. If an amber or white
annunciator flashes, it will trigger the steady-amber MASTER
CAUTION switchlights.
If an amber annunciator illuminates steady, it will not illuminate the

MASTER CAUTION switchlights. Amber annunciators that illuminate
steady indicate a system status and not a malfunction. Some amber
annunciators illuminate steady, but later flash indicating the respective
system’s status has changed to malfunction. See the individual
annunciator for specifics.
• Advisory Annunciators are white and located on the Master Warning

Panel. White annunciators indicate system status and do not flash or
trigger the MASTER CAUTION switchlights. Exceptions include the
white FUEL XFEED and the white GND IDLE annunciators. These
two exceptions initially illuminate steady, but can later flash and trigger
the MASTER CAUTION switchlights. See the individual annunciator
for specifics.
• General logic—If the annunciator flashes, it indicates a malfunction

regardless of color, and requires attention. If an annunciator illuminates
steady, it is advisory in nature. Color indicates the severity of the alert
and speed at which attention is required.
Table MW-1. MASTER WARNING/CAUTION SWITCHLIGHTS
ANNUNCIATOR DESCRIPTION
MASTER WARNING switchlights—Both switchlights flash to alert the
crew to a warning situation. Three red annunciators, the illumination
of the amber L and R GEN OFF annunciators (dual generator failure),
and the amber thrust reverser ARM and UNLOCK lights inflight trigger
these red-flashing switchlights. The MASTER WARNING switchlights
are not triggered by the ENGINE FIRE light. The illumination a single
MASTER WARNING switchlight indicates a malfunction of the master
warning system.
MASTER CAUTION switchlights—Both switchlights illuminate steady
to alert the crew to a caution condition. All flashing amber and white
annunciators trigger the MASTER CAUTION switchlights. Steady
amber and white annunciators do not trigger the switchlights. The
illumination of a single MASTER CAUTION switchlight indicates a
malfunction of the master warning system.
MW-2 FOR TRAINING PURPOSES ONLY

Figure MW-1. XLS Annunciators (Sheet 1 of 2)
AUXILIARY ANNUNCIATORS
MW-4
BATT LO OIL LO HYD LO HYD STAB ENG OIL GND P/S EMER AHRS ENG
O'TEMP CAB ALT PRESS FLOW LEVEL MIS COMP VIB FLTR BP IDLE HTR PRESS AUX PWR ANTI-ICE
>160 L R L R HYD SPD BRK L R L R NO ACM
PRESS EXTEND TAKEOFF L R O'HEAT 1 2 L R
FUEL LO FUEL EEC GEN AFT AC RUDDER FUEL LO BRK STBY AIR DUCT RADOME TL DEICE TL DEICE
GAUGE LEVEL MANUAL OFF J-BOX BEARING BIAS FLTR BP PRESS P/S HTR O'HEAT FAN FAIL PRESS
FIRE EXT ANTISKD AOA HTR
L R L R L R L R LMT CB L R BOTL LOW L R CKPT CAB L R L R
INOP FAIL
FUEL LO FUEL W/S W/S F/W FIRE ACC DOOR DOOR EMER BLD AIR CHECK WING WING

BOOST PRESS FAULT O'HEAT SHUTOFF DET SYS UNLOCKED SEAL EXIT O'HEAT PFD 1 O'HEAT ANTI-ICE
FUEL L R L R L R L R L R L R NOSE TAIL CABIN LAV CHECK

XFEED L R L R L R
DOOR DOOR PFD 2
APU RELAY
APU ENGAGED
READY TO LOAD
FIRE APU FAIL BLEED VAL OPEN
MASTER MASTER TERR TAWS FLAP TAWS FD/AP CABIN AUDIO

TAWS BIAS
WARNING CAUTION NORM NORM G/S PFD 1
HEATER
TEMP CTL APU GEN SPK/HPH
RESET RESET TERR TAWS FLAP CANCELED TEST FD/AP

FAIL RMT NRM OFF AUDIO
INHIB OVRD PFD 2 HPH ONLY
Figure MW-1. XLS Annunciators (Sheet 2 of 2)

AUXILIARY ANNUNCIATORS
MW-5
Figure MW-2. Excel Annunciators (Sheet 1 of 2)

MW-6
BATT LO OIL LO HYD LO HYD STAB ENG OIL GND P/S EMER AP PITCH AHRS ENG
O'TEMP CAB ALT PRESS FLOW LEVEL MIS COMP VIB FLTR BP IDLE HTR PRESS MISTRIM AUX PWR ANTI-ICE
HYD SPD BRK NO ACM AP ROLL
>160 L R L R PRESS EXTEND L R L R TAKEOFF L R O'HEAT MISTRIM 1 2 L R
FUEL LO FUEL EEC GEN AFT AC RUDDER FUEL LO BRK STBY AIR DUCT RADOME TL DEICE TL DEICE
GAUGE LEVEL MANUAL OFF J-BOX BEARING BIAS FLTR BP PRESS P/S HTR O'HEAT FAN FAIL PRESS
FIRE EXT ANTISKD AOA HTR
L R L R L R L R LMT CB L R BOTL LOW L R INOP FAIL CKPT CAB L R L R

FUEL LO FUEL W/S W/S F/W FIRE ACC DOOR DOOR EMER BLD AIR CHECK WING WING
O'HEAT O'HEAT
BOOST PRESS FAULT SHUTOFF DET SYS UNLOCKED SEAL EXIT PFD 1 O'HEAT ANTI-ICE
FUEL CABIN LAV CHECK
XFEED L R L R L R L R L R L R NOSE TAIL DOOR DOOR L R PFD 2 L R L R
Figure MW-2. Excel Annunciators (Sheet 2 of 2)

Table MW-2. MASTER WARNING ANNUNCIATORS
BATTERY O’TEMP—Flashes if the battery is too hot. Temperature
has reached 145°F. If battery temperature reaches 160°F or greater,
both annunciator segments (>160° and BATTERY O’TEMP) flash.
This annunciation is triggered by a dedicated sensor independent of
the battery temperature gauge. Because the battery temperature
gauge uses a separate sensor, the gauge can be used to check the
validity of the red annunciator.
CAB ALT—Flashes if cabin altitude reaches 10,000 feet. Illumination
occurs at 14,500 feet if the pressurization controller detects operation
out of or into a high altitude airport (8,100–14,000 feet) and the
aircraft is below 24,500 feet.
LO OIL PRESS—Flashes if oil pressure is below 20 psi. Illumination

is triggered by a dedicated pressure switch. Because the oil pressure
gauge uses a separate sensor, the pressure gauge can be used to
verify the validity of the red annunciator.
LO HYD FLOW—Flashes if the hydraulic fluid flow rate is below

normal. Illumination occurs after a 5-second delay with the engine
operating. Annunciator normally indicates a pump failure.
LO HYD LEVEL—Flashes for a low fluid quantity in the hydraulic

reservoir (fluid quantity is 74 cu. in. or below).
HYD PRESS—Steady illumination indicates the hydraulic system is

pressurized. This illumination is normal with all hydraulic subsystem
selections. Illumination extinguishes when the respective subsystem
operation is complete.
Flashing illumination inflight indicates the hydraulic system has
remained pressurized for 40 seconds.
STAB MISCOMPARE—Steady illumination occurs on ground if the
horizontal stabilizer does not agree with the flap handle position within
30 seconds. This condition contributes to the NO TAKEOFF
annunciation.
Flashing annunciator in flight indicates: (1) the horizontal stabilizer
does not agree with the flap handle within 30 seconds or, (2) the
aircraft has exceeded 200 KIAS after takeoff with the flap handle
greater than 0°.
SPD BRK EXTEND advisory—Steady illumination indicates the left

and right speedbrake panels are fully extended. This condition
contributes to the NO TAKEOFF annunciation.

Table MW-2. MASTER WARNING ANNUNCIATORS (Cont.)
ENG VIB advisory—Steady illumination indicates a vibration has been
detected in the respective engine.
OIL FLTR BP—Flashes to indicate oil filter contamination. The

respective oil filter is partially or completely blocked. Bypass is
impending. The filter may or may not be bypassing.
GND IDLE Advisory—Steady illumination indicates the aircraft is on

the ground and one or both engines are in ground idle mode (N2 rpm
47% minimum). This steady illumination appears 8 seconds after
landing.
Flashing illumination occurs after takeoff if an engine has remained in
ground idle mode. The MASTER CAUTION lights illuminate with this
flashing annunciator. The EEC switches must be AUTO for any
illumination.
NO TAKEOFF—Steady illumination indicates the aircraft is onground

and not properly configured for takeoff.
The following contribute to this condition:
• Flaps are not in takeoff range (7 or 15°)
• Speedbrakes are extended
• Pitch trim is not in takeoff range
• Stab miscompare
• Parking brake set (UK registered aircraft only)
P/S HTR—Flashing illumination occurs and the gear horn sounds if

the throttles are advanced for takeoff with the NO TAKEOFF condition
active.
Steady illumination onground with the pitot static switch OFF.
Flashing illumination occurs if: (1) the throttles are advanced for
takeoff with the switch OFF, (2) inflight if the switch is OFF, or (3) a
respective primary pitot tube or static port has lost electrical current
(malfunction).
EMER PRESS—Flashes to indicate emergency pressurization is

active. The system can be manually or automatically activated.
Automatic activation includes: (1) ACM O’HEAT, (2) cabin altitude
14,500 feet, or (3) the NORM PRESS circuit breaker is out. If selected
onground, the light illuminates but the valve does not open.
ACM O’HEAT—Flashes to indicate the ACM has overheated and
shutdown. EMER PRESS annunciation will also illuminate.

Table MW-2. MASTER WARNING ANNUNCIATORS (Cont)
AP PITCH MISTRIM (Excel only)—Flashes to indicate the elevators are
out of trim with the autopilot. The green “UP” or “DN” light illuminates
on the face of the autopilot control panel indicating the out-of-trim
direction.
AP ROLL MISTRIM (Excel only)—Flashes to indicate the ailerons are
not trimmed with the autopilot.
XLS—AP PITCH and ROLL mistrim annunciations appear in the PFDs.

AHRS AUX POWER Advisory—Steady illumination indicates the
respective AHRS is powered by the auxiliary battery in the right nose
compartment. Illumination is normal during engine start with the
avionics switch OFF. With the avionics switch ON, illumination indicates
the AHRS has transferred to the auxiliary battery due to a malfunction
(i.e., circuit breaker is out).
ENG ANTI-ICE—Steady illumination indicates the system is warming up.

Flashing illumination indicates the system has not warmed up properly.
A 4-minute and 45-second warm-up period is required before the light
begins flashing. If the system warms up but later becomes inoperative,
the annunciator flashes immediately. Causes for a flashing light include
the loss of stator vane heat or the engine nacelle is too cold. This
annunciator also flashes if engine anti-ice is selected OFF and the
stator vane heating valve does not close. The engine anti-ice
monitoring sensors are enabled when wing anti-ice is selected ON.
FUEL GAUGE—Flashes to indicate a fault is detected in the fuel
gauging system. BIT lights illuminate on the fuel quantity signal
conditioner (FQCS) indicating the faulted area. The battery switch must
be left ON after landing for light to remain illuminated.
LO FUEL LEVEL—Flashes to indicate the usable fuel level in the wing

fuel tank is 360 ± 20 pounds.
Limitation: The respective fuel boost pump must be turned ON.
EEC MANUAL advisory—Illuminates steady to indicate the respective

electronic engine control (EEC) is in manual mode. The EEC switch
has been selected to MAN or the EEC has automatically reverted to
manual.
GEN OFF—Illuminates steady on the ground with the engine shut

down.
Flashing illumination indicates the generator relay is open and the
generator is off line. Flashing illumination of the L and R lights indicate
a dual generator failure. The MASTER CAUTION and MASTER
WARNING lights illuminate for a dual failure.

AFT J BOX—Flashes to indicate an open 225 amp current limiter in
the tail cone J Box.
AFT J BOX CB—Flashes to indicate an open 5 amp start control
circuit breaker in the tail cone J box.
AC BEARING advisory—Illuminates steady to indicate respective

alternator bearing failure impending within approximately the next 20
hours of operation.
RUDDER BIAS—Flashes to indicate a rudder bias system fault. The

rudder bias valve is not in its commanded position.
FIRE EXT BOTL LOW—Flashes when either engine fire extinguisher

bottle pressure is low or discharged.
FUEL FLTR BP—Flashes when the fuel filter is contaminated. The

filter may or may not be bypassing. A pressure switch has detected
differential across the filter.
LO BRK PRESS—Flashes to indicate power brake pressure is low.

The ANTISKD INOP light illuminates with all low brake pressure
conditions indicating that antiskid is also inoperative. These flashing
lights and the MASTER CAUTION lights cannot be canceled during
ground operations for SNs 5222 and on, or earlier SNs with SB 32-17
performed.
ANTISKD INOP—Steady illumination indicates a self-test in operation
or the switch is OFF.
Flashing illumination indicates the system is inoperative. For
ANTISKD illumination only, the power brakes remain operational.
Limitation: Antiskid must be operational for takeoff.
STBY P/S HTR—Steady illumination on ground indicates the pitot-
static switch is OFF.
Flashing illumination occurs if: (1) the throttles are advanced for
takeoff with the switch OFF, (2) inflight if the switch is OFF, or (3) the
standby pitot tube or static port loses electrical current (malfunction).

AOA HTR FAIL—Steady illumination on ground indicates the pitot-
static switch is OFF.
Flashing illumination indicates: (1) on ground the switch is OFF and
the throttle has been advanced for takeoff, (2) the switch is placed to
OFF inflight, or (3) the switch is ON, ground or inflight, and the AOA
vane has lost electrical current (malfunction).
AIR DUCT O’HEAT—Flashes to indicate the bleed-air temperature in

the respective cockpit or cabin under-floor supply duct is too high
(300° or more).
RADOME FAN—Flashes to indicate a failure of the radome cooling

fan.
Limitations: On-ground operations are limited to 30 minutes with
dispatch into VMC only, unless an IC HOT annunciation appears.
FDR FAIL advisory (optional)—Steady illumination indicates the

optional flight data recorder is inoperative.
TAIL DEICE FAIL—Flashes to indicate the respective horizontal

stabilizer boot has not properly inflated. Possible controller failure.
TAIL DEICE PRESS advisory—Illuminates steady to indicate the

respective horizontal stabilizer boot has inflated properly. With the
deice switch in AUTO, normal operation is indicated by an 18-second
cycle period: left light illuminates for 6 seconds, light extinguishes for
6 seconds, right light illuminates for 6 seconds. The cycle will repeat 3
minutes later. Deice switch in manual illuminates the L and R lights
simultaneously.
FUEL XFEED advisory—Steady illumination indicates the crossfeed
valve has opened after selecting CROSSFEED.
Flashing illumination indicates the crossfeed valve has failed to close
after CROSSFEED is selected OFF. The MASTER CAUTION lights
illuminate.
FUEL BOOST—Steady illumination indicates the respective boost
pump is receiving power. Steady illumination occurs during normal
operations. These operations include: (1) manual selection ON, (2)
automatic activation during engine start, or (3) crossfeed operations.
Flashing illumination occurs when the boost pump is activated
because of low fuel pressure. All automatic activations require the
FUEL BOOST switch in the NORM position.

LO FUEL PRESS—Steady illumination appears before engine start.
Flashing illumination indicates low fuel pressure in the engine fuel
supply line anytime after start.
W/S FAULT—Steady illumination appears on ground before engine

start due to the AC alternator off-line status.
Exception—On the ground and before engine start, the light begins
flashing 8 seconds after a controller has failed.
Flashing illumination, after engine start, indicates a windshield heat
system fault (i.e., controller or alternator failure, or W/S O’HEAT).
W/S O’HEAT—Flashes to indicate the respective windshield has
over-heated. The W/S FAULT also illuminates and windshield heat
shuts down. The system may automatically reactivate after cooling
followed by another system shutdown at the overheat point (cycle on
and off).
F/W SHUTOFF—Flashes to indicate the respective fuel and hydraulic

firewall shutoff valves have closed and the generator field relay has
tripped. This annunciation occurs after the engine fire switchlight has
been pressed. All three conditions are required for light illumination.
FIRE DET SYS—Flashes to indicate a failure in the respective engine

fire detection system. Fire detection failure can be verified with the
rotary test switch. Engine fire extinguishing remains operational.
ACC DOOR UNLOCK–NOSE—Flashes to indicate one of the nose

avionics doors is not properly latched. The two bottom latches on
each door are monitored (four total).
TAIL—The baggage or tail cone door is not properly latched. SNs
188 and on or otherwise modified, the battery door is also monitored
by the TAIL.
DOOR SEAL—Illuminates steady on ground with the main door open
or the main door is closed and service bleed air is not available
(engine or APU not operating).
Flashing light indicates the main door is closed, service air is
available, but the primary pressure seal has not properly inflated. The
acoustic seal is not monitored.
CABIN DOOR—Illuminates steady on ground with the main door
open.
Flashing light indicates the main door is closed and the door is not
properly locked or the internal vent door is not closed. Main door
closure without power on the aircraft may cause illumination. Visual
indicators next to the door can be checked for improper door pin, door

EMER EXIT—Flashes to indicate the emergency exit door is not
locked or the position switch has failed.
LAV DOOR—Flashes if either lavatory door is not stowed open during

ground or in-flight operations when the flaps are selected more than 0°.
BLD AIR O’HEAT—Flashes to indicate bleed air exiting the pylon pre-
cooler has exceeded temperature limits (560°). Wing anti-ice on the
affected side is inoperative.
CHECK PFD 1/2—Flashes to indicate there is a malfunction in the

respective PFD. The IAC to PFD to IAC wrap-around function
indicates a malfunction.
Limitation: The autopilot may not be used.
WING O’HEAT—Flashes to indicate a bleed-air leak into the wing

purge air passage. The affected side wing anti-ice automatically shuts
off. If wing anti-ice is in use, it reactivates when the leading edge
cools (cycle ON and OFF). Wing overheat sensors are active with or
without the anti-ice switches ON.
WING ANTI-ICE—Steady illumination, ground or inflight, indicates that

wing anti-ice has been selected ON and the surface is warming up.
Flashing illumination indicates the surface is too cold. A 4-minute and
45-second warm-up period is required before the light begins flashing.
If the surface reaches operating temperature, but later becomes too
cold, the light flashes immediately. The undertemperature sensors are
enabled when wing anti-ice is selected ON.
AP OFF or YD OFF (Excel only)—Illumination occurs when the

autopilot or yaw damper is manually disconnected by the crew or
automatically disconnected due to malfunction. This annunciator is
next to the L and R MASTER WARNING/MASTER CAUTION
switchlights.
XLS—AP and YD OFF annunciations appear in the L and R PFDs.

Table MW-3. AUXILIARY ANNUNCIATORS
APU GEN OFF advisory (XLS only)—Steady illumination indicates
the APU is operating and its generator is off line.
BIAS HEATER FAIL (Excel and XLS)—Steady illumination indicates

the rudder bias heating blanket is heating.
Flashing light indicates blanket sensor failure. Pressing the light
causes steady illumination. This annunciator does not activate the
MASTER CAUTION lights
FD/AP PFD 1 OR 2 (Excel and XLS)—Switchlight indicates the No.1
or 2 flight director is controlling the autopilot. Press the switchlight to
change flight directors. Switching flight directors with the autopilot
engaged causes the autopilot to revert to basic pitch and heading
hold modes. The flight director modes must be reselected.
TERR NORM (Excel and XLS)—Switchlight indicates the enhanced
GPWS or TAWS warnings occur normally and the terrain map is
displayed on the MFD.
TERR INHIB—When selected, inhibits the enhanced TAWS
(EGPWS) warnings and the terrain map. Modes 1-7 remain active.
TAWS FLAP NORM (XLS)—Switchlight indicates that the TOO LOW
FLAPS audio warning activates when the aircraft is below
approximately 245 feet AGL, less than 160 KIAS, and landing flaps
are not selected.
TERR INHIBIT—When pressed, the switch disarms or cancels the
audio warning for landing with flaps less than 35°. The Excel
switchlight is labeled “GPWS FLAP NORM” and “GPWS O’RIDE”.
The functions are the same.
TAWS G/S (XLS)—Switchlight indicates normal GLIDESLOPE audio
warnings are active for deviations below the glideslope. The
GLIDESLOPE warning sounds if the aircraft is below 1000 feet AGL,
descending greater than 500 fpm, and below 1.3 dots.
TAWS FLAP O’RIDE—When pressed, disables the GLIDESLOPE
audio warnings. The Excel switchlight is labeled “GPWS G/S” and
“O’RIDE.” The functions are the same.
TAWS TEST (XLS)—Pressing the switchlight initiates the TAWS
system test. This test function is inhibited inflight. The Excel
switchlight is labeled “GPWS TEST.” The functions are the same.

Table MW-3. AUXILIARY ANNUNCIATORS (Cont)

AUDIO SPK/HPH (Excel and XLS)—Indicates normal operating mode
(default position). Audio communications are active through the
cockpit speakers and crew headsets.
AUDIO HPH ONLY—Pressing the switchlight mutes all avionics audio
through the cockpit speakers including TCAS and TAWS (EGPWS).
The gear horn and NO TAKEOFF warnings are not inhibited.
PHONE CALL (Excel and XLS) (Optional)—Steady illumination for an

incoming HF radio call.
CABIN TEMP CTL–NRM (Excel and XLS)—Indicates that cabin

temperature is controlled from the cockpit temperature controller.
RMT—When pressed, transfers the cabin temperature control to the
cabin.
Table MW-4. THRUST REVERSERS
ARM—Illumination indicates pressure is available to the thrust
reverser (pressure is sensed passed the isolation valve). Illumination
is normal on ground during TR operation, but abnormal inflight.
Illumination inflight causes the red MASTER WARNING lights to flash.
UNLOCK—Illumination indicates the thrust reverser is unlocked.
Illumination is normal on ground during TR operation, but abnormal
inflight. Illumination inflight causes the red MASTER WARNING lights
to flash.
DEPLOY—Illumination of the white light indicates the thrust reverser
is deployed. Illumination is normal on ground during TR operation, but
abnormal inflight.

Table MW-5. FIRE SWITCHLIGHTS
ENGINE FIRE—Illumination indicates high temperature is detected in
the engine nacelle. Pressing the switchlight:
1. Closes the fuel F/W shutoff valve.
2. Closes the hydraulic F/W shutoff valve.
3. Deactives the engine generator (opens the field relay).
4. Disarms the thrust reverser.
5. Arms the engine fire bottles.
BOTTLE ARMED 1/2 switchlight—Illumination of the white light

indicates the respective engine fire bottle is armed. When pressed,
the bottle discharges. The red ENGINE FIRE switchlight must be
pressed to illuminate the BOTTLE ARMED lights.
APU FIRE—Illumination indicates high temperature in the APU
compartment. The APU automatically shuts down and the APU FAIL
light illuminates. Pressing the red switchlight discharges the APU fire
bottle. If the switchlight is not pressed, the fire bottle automatically
discharges in 8 seconds.
Table MW-6. APU ANNUNCIATORS
APU RELAY ENGAGED—Illumination indicates the APU relay is
engaged during APU start. Illumination also occurs when the APU
generator participates in an engine start.
APU FAIL—Illumination indicates the APU will not start due to a

system malfunction (i.e., the APU fire bottle is low or the fire detection
system is inoperative). If the APU is operating, the light indicates the
APU is shutting down. Reasons for automatic shutdown include fire
detected in the APU compartment or the fire bottle is low.
Limitation: Starting the APU is prohibited whenever the APU FAIL light
is illuminated.
BLEED VAL OPEN advisory—Illumination indicates APU bleed air
valve (BAV) is other than closed.
READY TO LOAD advisory—Illumination indicates the APU start is

complete and at operating speed (95% rpm + 4 seconds). The APU
generator and bleed air can be selected after illumination. The light
remains illuminated during APU operation.

TEST
SPARE OFF FIRE
WRN
AVN LDG
GEAR
ANNU BATT
TEMP
ANTI
SKID STICK
SHAKER
OVER
SPEED T / REV
W/S TEMP
Figure MW-3. Rotary Test Knob
XLS ROTARY TEST

OFF—The red light is extinguished, and the test system is shut off.
FIRE WARN—Both red ENG FIRE lights illuminate, indicating continuity.
LDG GEAR—The green NOSE, LH, and RH lights and the red GEAR UN-
LOCKED lights illuminate, and the gear warning horn sounds.
BATT TEMP—BATT O'HEAT/>160° annunciator illuminates, the MASTER

WARNING lights flash (cancelable), and the battery temperature gauge in-
dicates 160°F.
STICK SHAKER—Stick shakers on both control columns will immediately

operate. The AOA gauge needle swings to the top of the red band. The AOA
indexer initially illuminates then extinguishes the amber chevron followed
by the green “donut” leaving only the red chevron flashing.
T/REV—Both thrust reverser indicators, ARM, UNLOCK, and DEPLOY

lights illuminate. MASTER WARNING lights flash (cancelable).
W/S TEMP—Windshield heat selected ON, the W/S O’HEAT L–R annunci-
ators illuminate steady for 3 to 4 seconds, then extinguish.
Conducting test prior to engine start, the W/S FAULT L–R annunciators il-
luminate steady (alternators are not operating). Conducting test with engines
operating, the W/S FAULT and W/S O’HEAT lights illuminate for 3 to 4 sec-
onds, then extinguish.

OVER SPEED—The avionics power switch must be ON for valid test indi-
cations. The following indications occur:
• Audible overspeed warning signal sounds
• PFD 1 and PFD 2:
• ADC TEST (upper left)
• 265 KIAS (with red and white barber pole)
• 5,000 feet altitude
• 2,000 feet fpm climb
ANTISKID—The ANTISKD INOP annunciator flashes for 6 seconds, then

extinguishes. The MASTER CAUTION lights illuminate steady (cancelable).
ANNU—The avionics switch must be on for valid test indications:
• All lights on the annunciator panel illuminate
• MASTER WARNING and MASTER CAUTION illuminate (not

cancelable)
• All lights on the auxiliary annunciator switch panel illuminate
• Flight director selector mode buttons illuminate left to right and remain
steady
• AP OFF
• YD OFF
• ROL TRIM
• PIT TRIM
• All five autopilot control panel lights illuminate
• A steady altitude alert tone or a pulsating aural horn (optional phone)

sounds for combination of altitude alert and phone call tone pulsating
(becomes steady when PHONE CALL button is depressed).

AVN—The avionics switch must be ON for the avionics check. The follow-
ing indications are present:
• Annunciator panel (these lights flash):
• RADOME FAN
• CHECK PFD 1
• CHECK PFD 2
• MASTER CAUTION illuminates (cancelable)
• Flight director selector mode buttons illuminate left to right and remain
on steady
• AP OFF
• YD OFF
• ROL TRIM
• PIT TRIM
• Auxiliary annunciator switch panel:
• TERR NORM/TERR INHIBIT
• TAWS FLAP NORM/TAWS FLAP O’RIDE
• TAWS G/S O’RIDE
• TAWS TEST
• FD/AP PFD1 and FD/AP PFD2
• CABIN TEMP CTL—NORM (RMT extinguished)
• AUDIO SPK/HPH and AUDIO HPH ONLY
• All five autopilot control panel lights illuminate
• A steady altitude alert tone or a pulsating aural horn (optional phone)

sounds for combination of altitude alert and phone call tone pulsating
(becomes steady when PHONE CALL button is depressed)
SPARE—Not used.

EXCEL ROTARY TEST

OFF—The red light is extinguished, and the test system is shut off.
FIRE WARN—Both red ENG FIRE lights illuminate, indicating continuity.
LDG GEAR—The green NOSE, LH, and RH lights and the red GEAR UN-
LOCKED lights illuminate, and the gear warning horn sounds.
BATT TEMP—BATT O'HEAT/>160° annunciator illuminates, the MASTER

WARNING lights flash (cancelable), and the battery temperature gauge in-
dicates 160°F.
STICK SHAKER—Stick shakers on both control columns immediately op-

erate. The AOA gauge needle swings to the top of the red band. The red
chevron in the AOA indexer flashes (on glareshield above pilot instrument
panel).
T/REV—Both thrust reverser indicators, ARM, UNLOCK, and DEPLOY

lights illuminate. MASTER WARNING lights flash (cancelable).
W/S TEMP—Windshield heat selected ON, the W/S O’HEAT L–R annunci-
ators illuminate steady for 3 to 4 seconds then extinguish.
Conducting test prior to engine start, the W/S FAULT L–R annunciators il-
luminate steady (alternators are not operating). Conducting test with engines
operating, the W/S FAULT and W/S O’HEAT lights illuminate for 3 to 4 sec-
onds then extinguish.
OVER SPEED—The avionics power switch must be ON for valid test indi-
cations. The following indications will occur:
• The audible overspeed warning signal sounds.
• PFD1/PFD2 should indicate approximately 265 KIAS.
• PFD1/PFD2 indicates Mach 0.4 (red).
• PFD1/PFD2 altitudes indicate 5,000 feet.
• PFD1/PFD2 VSIs indicate 2,000 fpm climb.
ANTISKID—The ANTISKID INOP annunciator flashes for 6 seconds then

extinguishes. The MASTER CAUTION lights illuminate steady (cancelable).
ANNU—The avionics switch must be on for valid test indications.
• All lights on the annunciator panel illuminate.
• MASTER WARNING lights flash and MASTER CAUTION lights

illuminate steady (noncancelable).
• Both red turbine overspeed lights flash.

• Engine instrument LCDs should indicate steady 8s.
• AP OFF annunciators illuminate steady.
• Flight director mode buttons illuminate left to right and remain steady.
• Annunciators to the right of the F/D mode panel should illuminate:
• FD/AP PFD 1—FD/AP PFD 2
• TERR NORM—TERR INHIBIT (optional)
• GPWS FLAP NORM—GPWS FLAP O’RIDE (optional)
• GPWS G/S—O’RIDE (optional)
• GPWS TEST (optional)
• PHONE CALL
• All autopilot control panel lights illuminate.
• Green light on the vapor cycle A/C panel illuminates.
• A pulsating aural horn sounds, combination of:
• Altitude alert horn (steady) and phone call tone pulsating,

(becomes steady when PHONE CALL button is depressed).
AVN—The avionics power switch must be ON for the avionics system test to
be valid. The following annunciators flash in the annunciator panel:
• AP PITCH MISTRIM
• AP ROLL MISTRIM
• RADOME FAN
• CHECK PFD 1, CHECK PFD 2
• Autopilot/flight director mode selector panel lights
• All annunciators to the right of the F/D mode panel illuminate.
• MASTER CAUTION lights illuminate steady (cancelable).
• Altitude alert horn sounds.
SPARE—Not used.

ACRONYMS
Table MW-7. ACRONYMS
ACONYM DEFINITION
AFM Airplane Flight Manual
AHRS Attitude and heading reference system
ALT Altimeter
AOA Angle-of-attack
AP Autopilot
APU Auxiliary power unit
B Both
BAV Bleed-air valve
BCP Best computed position
BITE Built-in test equipment
BOW Basic operating weight
DA Decision altitude
DH Decision height
DIEGME Diethylene Glycol Monomethyl Ether
ECU Electronic control unit
EDS Engine diagnostic system
EEC Electronic engine control
EGPWS Enhanced ground proximity warning system
EGME Ethylene Glycol Monomethyl Ether
ELEV Airport elevation or runway elevation
ELT Emergency locator transmitter
FAF Final approach fix
FCU Fuel control unit
FD Flight director
FMS Flight management system
GOG Ground-on-ground
FOHE Fuel-oil heat exchanger
FSM Field service monitor
HLDC High level data link control
HSI Horizontal situation indicator
IAP Instrument approach procedures
IAC Integrated avionics computers

Table MW-7. ACRONYMS (Cont)
ACONYM DEFINITION
KCAS Calibrated airspeed
KIAS Indicated airspeed
KTAS True airspeed
LOP Low oil pressure
MADC Micro air data computers
MDA Minimum descent altitude
MSA Minimum safe altitude
PCB Printed circuit board
PF Pilot flying
PIC Pilot in command
PNF Pilot not flying
PRSOV Pressure regulating shutoff valve
PTS Practical test standards
PTM Pilot Training Manual
RMU Radio management units
RTD Resistive thermal device
RWY Runway
RAT Ram air temperature
SFD Secondary flight display
SID Standard instrument departure
SLA Set landing altitude
SRC Standby radio control
STAR Standard terminal arrival route
TAS True air speed
TAWS Terrain awareness and warning system
TCI Takeoff climb increment
TCS Touch control steering
TCV Temperature control valves
TEMP Temperature
V1 Decision speed
V2 Safety climb speed
VAPP Minimum landing approach climb speed

Table MW-7. ACRONYMS (Cont)
ACONYM DEFINITION
IFR Instrument flight rules
VENR Single-engine enroute climb speed
VFR Flap retraction speed
VR Rotation speed
VREF Minimum final approach speed
WIND Wind direction
ZFW Zero fuel weight

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INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES

LIST OF EFFECTIVE PAGES

Dates of issue for original and changed pages are:

Original............0 ................. January 2006

THIS PUBLICATION CONSISTS OF THE FOLLOWING:

Page *Revision Page *Revision

*Zero in this column indicates an original page.

FOR TRAINING PURPOSES ONLY LEP-1

At the time of printing it contained then-current information. In the

We at FlightSafety want you to have the best training possible. We

FOR TRAINING PURPOSES ONLY

MANEUVERS AND PROCEDURES

WEIGHT AND BALANCE

CREW RESOURCE MANAGEMENT

FOR TRAINING PURPOSES ONLY EC-i

FOR TRAINING PURPOSES ONLY NP-i

FOR TRAINING PURPOSES ONLY AP-i

FOR TRAINING PURPOSES ONLY EP-i

FOR TRAINING PURPOSES ONLY LIM-i

MANEUVERS AND PROCEDURES

FOR TRAINING PURPOSES ONLY MAP-i

APPROACHES AND LANDING PROCEDURES..................... MAP-26

MAP-ii FOR TRAINING PURPOSES ONLY

FOR TRAINING PURPOSES ONLY MAP-iii

MANEUVERS AND PROCEDURES

PREFLIGHT AND TAXI PROCEDURES

FOR TRAINING PURPOSES ONLY MAP-1

V 1 , V R , V 2 , V FR & V ENR —Calculated V 1 , V R , V 2 and V ENR based on exist-

CLEARANCE—Space provided for copying ATC clearances and other per-

ARPT—Name of airport or ICAO identifier.

ELEV—Airport elevation or runway elevation if significantly different than

RWY—Runway in use for departures.

FlightSafety CITATION FlightSafety CITATION

TAKEOFF DATA LANDING DATA

VFR VENR FLAPS CLEARANCE

RWY LENGTH__________RWY REQ'D______ CIG________________TEMP/DP______/_____

ZFW___________T.O. WT._________________ ALT________RMKS______________________

Figure MAP-1. Takeoff and Landing Card

MAP-2 FOR TRAINING PURPOSES ONLY

ATIS—Current ATIS information code.

WIND—Wind direction and speed as reported by ATIS.

VIS—Visibility as reported by ATIS.

CIG—Clouds and significant weather as reported by ATIS.

TEMP/DP—Temperature and dew point as reported by ATIS.

ALT—Altimeter setting as reported by ATIS.

RMKS—Any additional information provided by ATIS.

RWY LENGTH—Actual length of runway to be used for takeoff.

RWY REQ’D—Charted takeoff field length. If actual runway is less, reduce

EMERGENCY RETURN INFORMATION

MSA—Minimum Safe Altitude required for obstacle clearance. May be taken

V REF - V APP —Calculated approach speeds corresponding to the appropriate

GA N 1 —Obtained from Flight Manual for go-around (TO N 1 ). It is based on

Page Revision Page Revision

RWY LENGTHRWY REQ'D CIG__TEMP/DP/_

ZFW___T.O. WT._ ALTRMKS______________