Pilot Training Manual V2- Challenger 605

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CL 601-3A/R PILOT TRAINING MANUAL VOLUME 2
Record of Revision No. .01
This is a revision of the CL 601-3A/R Volume 2 Pilot Training Manual.
The portion of the text or figure affected by the current revision is indicated by a solid vertical
line in the margin. A vertical line adjacent to blank space means that material has been delet-
ed. In addition, each revised page is marked “Revision .01” in the lower left or right corner.
The changes made in this revision will be further explained at the appropriate time in the
training course.
FlightSafety
international
COURSEWARE SUPPORT—HURST 8900 Trinity Blvd. Hurst, Texas 76053 (817) 276-7500 Fax (817) 276-7501
the best safety device in any aircraft is a well-trained crew...
9555 Ryan Avenue
Dorval, Quebec, Canada H9P 1A2
(800) 573-4025
www. flightsafety.com
Canadair
CHALLENGER CL-601-3A/R 
MODEL CL-600-2B16 
PILOT TRAINING MANUAL
VOLUME 2
AIRCRAFT SYSTEMS
FlightSafety Canada LteeLtd.'
FlightSafety Canada LteeLtd.'
Courses for the Canadair Challenger Model CL-600-2B16 and other Canadair
aircraft are taught at:
FlightSafety Canada Ltd.
Montreal Learning Center
9555 Ryan Avenue
Dorval, Quebec
Canada H9P 1A2
(800) 573-4025 • Fax (514) 631-2263
FlightSafety Texas
Houston Learning Center
7525 Fauna Street
Houston, TX 77061
(800) 927-1521 • Fax (713) 644-2118
FlightSafety International
Tucson Learning Center
1071 E. Aero Park Blvd.
Tucson, AZ 85706
(800) 203-5627 • Fax (602) 889-9619
Copyright © 2003 by FlightSafety International, Inc. All rights
reserved. Printed in the United States of America.
FOR TRAINING PURPOSES ONLY
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
SYLLABUS
Chapter 1 AIRCRAFT GENERAL
Chapter 2 ELECTRICAL POWER SYSTEMS
Chapter 3 LIGHTING
Chapter 4 MASTER WARNING SYSTEM
Chapter 5 FUEL SYSTEM
Chapter 6 AUXILIARY POWER UNIT
Chapter 7 POWERPLANT
Chapter 8 FIRE PROTECTION
Chapter 9 PNEUMATICS
Chapter 10 ICE AND RAIN PROTECTION
Chapter 11 AIR CONDITIONING
Chapter 12 PRESSURIZATION
Chapter 13 HYDRAULIC POWER SYSTEMS
Chapter 14 LANDING GEAR AND BRAKES
Chapter 15 FLIGHT CONTROLS
Chapter 16 AVIONICS
Chapter 17 MISCELLANEOUS SYSTEMS
WALKAROUND
APPENDIX
ANNUNCIATOR PANEL
INSTRUMENT PANEL POSTER
MODIFYING YOUR PTM VOLUME 2
Please note that the Challenger Model CL-600-2B16 Pilot Training Manual Volume 2 includes
a compilation of both the CL 601-3A and CL 601-3R.
Where information is standard for both models, the footer shall denote CL 601-3A/R. If information
is specific to one model the footer shall read CL 601-3A or CL 601-3R as appropriate.
The following chapters are specifically affected by differences between models, please take the
indicated actions to make your manual correspond to the model of your aircraft:
Chapters Affected Action
Chapter 2—Electrical Pages 2-21 through 2-29
Discard Appropriate Pages
Chapter 5—Fuel System Discard Appropriate Pages
Chapter 7—Powerplant Discard Appropriate Pages
Chapter 9—Pneumatics Discard Appropriate Pages
Chapter 11—Air Conditioning Discard Appropriate Pages
Chapter 12—Pressurization Discard Appropriate Pages
Annunciator Panel Discard Appropriate Pages
CL 601-3A/R SYL-i
SYLLABUS
CONTENTS
Page
COURSE INFORMATION ............................................................................................... SYL-1
Learning Center Information...................................................................................... SYL-1
Description of Training Facility ................................................................................. SYL-1
Type of Aircraft .......................................................................................................... SYL-7
Category of Training .................................................................................................. SYL-7
Duty Position.............................................................................................................. SYL-7
Curriculum Title ......................................................................................................... SYL-7
Curriculum Prerequisites............................................................................................ SYL-7
Course Objectives .................................................................................................... SYL-10
Training Schedule (Typical)..................................................................................... SYL-10
Simulator and Flight Training.................................................................................. SYL-13
Completion Standards .............................................................................................. SYL-13
AIRCRAFT GROUND TRAINING CURRICULUM SEGMENT................................ SYL-14
Curriculum Segment Outline ................................................................................... SYL-14
Training Module Outlines ........................................................................................ SYL-15
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FLIGHT TRAINING CURRICULUM SEGMENT ....................................................... SYL-21
Training Hours ......................................................................................................... SYL-21
Flight Training Module Outlines ............................................................................. SYL-21
Completion Standards .............................................................................................. SYL-25
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
SYL-ii CL 601-3A/R FOR TRAINING PURPOSES ONLY
CL 601-3A/R SYL-iii
ILLUSTRATIONS
Figure Title Page
SYL-1 Montreal Facility Floor Plan............................................................................... SYL-2
SYL-2 Houston Facility Floor Plan................................................................................ SYL-3
SYL-3 Tucson Facility Floor Plan.................................................................................. SYL-5
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
SYLLABUS
COURSE INFORMATION
LEARNING CENTER INFORMATION
FlightSafety International is an aviation training company that provides type-specific training
programs for over 50 different models of aircraft, using a fleet of over 150 simulators. FlightSafety
operates 38 Learning Centers, including Centers in Europe and Canada.
Training for the Challenger series aircraft is conducted at the Montreal Learning Center in Dorval,
Quebec, Canada, the Houston Learning Center in Houston, Texas, and the Tucson Learning Center
in Tucson, Arizona. The Centers are owned and operated by FlightSafety International.
Listed below are the addresses of the three Centers:
FlightSafety Canada Ltd FlightSafety International
Montreal Learning Center Houston Learning Center
9555 Ryan Avenue 7525 Fauna at Airport Boulevard
Dorval, Quebec, Canada H9P 1A2 Houston, Texas 77601
FlightSafety International
Tucson International Airport
1071 E. Aero Park Blvd.
Tucson, AZ 85706
DESCRIPTION OF TRAINING FACILITY 
Each classroom and briefing room is adequately heated, lighted, and ventilated to conform to
local building, sanitation, and health codes. Thebuilding construction prevents any distractions
from instruction conducted in other rooms or by flight operations and maintenance operations
on the airport.
Classrooms are equipped for presentation of 35mm slides by front- or rear-screen projection,
controlled from a lectern. A standard overhead projector is available for use in the classroom.
Some overhead projectors are equipped with computer graphic animated motion for displaying
schematics and diagrams. Cockpit panel posters and/or cockpit mockups are also available at
most locations.
Briefing rooms are equipped with cockpit panel posters, a white liquid chalkboard, a table, and
chairs for individual or small-group briefings. Floor plans of the Montreal, Houston and Tuscon
Learning Centers follow.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R SYL-1FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
SYL-2 CL 601-3A/R FOR TRAINING PURPOSES ONLY
PARKING
AREA
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Canadair Challenger
Learning Center
BR
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✚ FIRST AID KIT LOCATIONS
 BASEMENT STORE ROOM
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Figure SYL-1. Montreal Facility Floor Plan
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R SYL-3FOR TRAINING PURPOSES ONLY
Figure SYL-2. Houston Facility Floor Plan (Sheet 1 of 2)
LOUNGE
CLASSROOM
BRIEFING
ROOM
COMPUTER
ROOM
ATR-42
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ROOM
PROGRAM MANAGERS
CUSTOMER SUPPORT
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ROOM 143 ROOM 142 ROOM 141 ROOM 140
EXIT
FIRE
EXIT
INSTRUCTORS
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RECEIVING
LOUNGE ROOM 137 ROOM 138 ROOM 139
MEN
ROOM 123 ROOM 121
ROOM 117 ROOM 119
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ELECTRICAL
SUPPLY SUPPLY
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FlightSafety
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FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
SYL-4 CL 601-3A/R FOR TRAINING PURPOSES ONLY
 7526 WYNLEA (Back)
HOUSTON, TEXAS
FlightSafety
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Figure SYL-2. Houston Facility Floor Plan (Sheet 2 of 2)
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R SYL-5FOR TRAINING PURPOSES ONLY
Figure SYL-3. Tucson Facility Floor Plan (Sheet 1 of 2)
BALCONY
SIMULATOR ROOM
117
SIMULATOR ROOM
132
COMPUTER ROOM
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COMPUTER ROOM
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BRIEF
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VEST
FIRST FLOOR
Learjet Learning Center
Tucson, Arizona
FlightSafety
international
LEAR 31
LEAR 35
(FC350)
LEAR 60
LEAR 25 LEAR 55 LEAR 35
(200)
LEAR 45
CHALLENGER 601-3R
PROGRAM
MANAGER
113
DOSPROGRAMMANAGER
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CL-600-2B16 PILOT TRAINING MANUAL
SYL-6 CL 601-3A/R FOR TRAINING PURPOSES ONLY
Figure SYL-3. Tucson Facility Floor Plan (Sheet 2 of 2)
CLASSROOM 
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Learjet Learning Center
Tucson, Arizona
FlightSafety
international
TYPE OF AIRCRAFT
The Canadair Challenger CL-600 series, which includes the Challenger 600, 601-1A, 601-3A,
and 601-3R.
CATEGORY OF TRAINING 
Initial Equipment and Transition training for a CL-600 type rating added to an existing pilot
certificate or the issuance of an Airline Transport Pilot Certificate with a CL-600 type rating.
DUTY POSITION 
Pilot-in-Command (PIC)
CURRICULUM TITLE 
Challenger Series Pilot Initial Equipment/Transition Training Course.
CURRICULUM PREREQUISITES 
Core Training Curriculum Prerequisites
§61.63
A pilot may enroll in this course and complete all of the items of the practical test required for
a CL-600 type rating that are authorized to be accomplished in the flight simulator, then com-
plete the items not approved for flight simulator in flight in a CL-600 Series airplane, if the pilot:
1. Holds a private pilot certificate with an airplane rating.
2. Holds an instrument rating or concurrently completes the instrument course.
3. Has a minimum of 1,000 hours flight experience in airplanes as a pilot (May be waived at
the discretion of the Center Manager).
4. Holds a MEL category rating without centerline thrust limitation.
§61.157
A pilot who meets the above requirements of §61.63 may concurrently apply for an Airline
Transport Pilot certificate with a CL-600 type rating, providing the pilot:
1. Holds a commercial pilot certificate or an ICAO recognized Airline Transport Pilot or
Commercial Pilot license without restrictions.
2. Meets the eligibility requirements of §61.151.
3. Has passed the written test required by §61.153.
4. Meets the experience requirements of §61.155.
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CL-600-2B16 PILOT TRAINING MANUAL
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Specialty Training Curriculum Prerequisites
§61.55
A pilot may enroll in the SIC course and complete all of the training that is authorized to be ac-
complished in the Level A through D flight simulator, then complete the items required in flight
in a CL-600 Series aircraft, if the pilot: 
1. Holds a private or commercial pilot certificate with an airplane rating.
2. Holds an instrument rating.
NOTE
Training completed under §61.63 or §61.157 core curriculums will also satisfy this cur-
riculum, except the requirement for one (1) takeoff and one (1) landing in the aircraft.
§61.58
A pilot may enroll in the CL-600 Series §61.58 course and complete all of the items required
fora PIC Check required by §61.58 that are authorized to be completed in a flight simulator, if
the pilot:
1. Holds a pilot certificate with a CL-600 type rating.
Completion Methods
The completion methods are as follows:
1. 100% Flight Simulator with no Limitations
2. 100% Flight Simulator with 15 hours SOE Limitation
3. 100% Flight Simulator with 25 hours SOE Limitation
4. Combination of Flight Simulator and Aircraft with no Limitation
100% Flight Simulator with No Limitations
A pilot may complete all of the practical test required for a CL-600 type rating in an approved
Level C or D flight simulator, except for the preflight inspection which must be completed in
either a static airplane or by using an approved pictorial means, if the pilot:
1. Holds a type rating in a multiengine turbojet airplane; or
2. Has been appointed by a military service as a pilot-in-command of a turbojet multiengine
land airplane; or
3. Has at least 2,000 hours of actual flight time, of which 500 hours must be in turbine-
powered multiengine airplanes; or
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4. Has at least 500 hours of actual flight time in CL-600 Series airplanes; or
5. Has at least 1,000 hours of flight time in at least two different airplanes requiring a type
rating.
100% Flight Simulator with 15 Hours SOE Limitation
If a pilot does not meet the above prerequisites, he is eligible for 100% flight simulator with 15
hours of SOE limitation, if the pilot:
1. Holds a type rating in a propeller-driven airplane; or
2. Has, since the beginning of the preceeding 12 calendar months, logged at least 100 hours of
flight time in multiengine airplanes that require a type rating and at least 25 hours of the
flight time were in CL-600 Series aircraft.
A pilot qualifying under this subparagraph may not act as PIC of a CL-600 Series airplane and
will be issued a CL-600 type rating, or an ATP certificate with a CL-600 type rating, as appro-
priate, with the limitation, “This certificate is subject to 15 hour supervised operating experi-
ence pilot-in-command limitations for the CL-600.”
100% Flight Simulator with 25 Hours SOE Limitation
If a pilot does not meet any one of the prerequisites listed above, the pilot may be eligible to
complete 100% in the flight simulator and receive a certificate with the following limitations:
A pilot who qualifies under this subparagraph may not act as PIC of a CL-600 Series airplane
and will be issued a CL-600 type rating, or an ATP certificate with a CL-600 type rating, as ap-
propriate, with the limitation, “This certificate is subject to 25 hour supervised operating ex-
perience pilot-in-command limitations for the CL-600.”
Combination of Flight Simulator and Aircraft with No Limitation
A pilot may use the Combination Method of Completion if:
1. A Level A or B flight simulator was used to complete the flight simulator modules. The pilot
must accomplish training and checking on the prescribed items of the Practical Test Standards
in the aircraft, to receive a certificate with no limitations.
2. The training was completed in a Level C or D flight simulator, and the pilot elects to com-
plete aircraft training in lieu of SOE limitations, he/she must successfully complete on a static
airplane, or in flight, the following:
A. Preflight Inspection
B. Normal Takeoff
C. Normal ILS Approach
D. Missed Approach
E. Normal Landing
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Completion Standards
Completion is based on proficiency. Syllabus times are estimates. Pilots must demonstrate sat-
isfactory performance through formal and informal examinations in the classroom and flight
simulator, and in flight to ensure they meet the knowledge and skill requirements necessary to
meet the course objectives. The Minimum Acceptable Performance Guidelines are as follows:
1. Each pilot shall fly the flight simulator and/or aircraft within the appropriate standard.
Depending on the type of operation, passenger seating, configuration within the aircraft, and/or
pilot’s level of certification, the tolerances of the appropriate standard will be specified in
one of the following publications:
A. Commercial Pilot Practical Test Standards
B. Instrument Rating Practical Test Standards
C. Airline Transport Pilot and Type Rating Practical Test Standards
2. The Instructor and/or Training Center Evaluator will determine the applicable standards prior
to the start of any training or evaluation session. The required standards will be discussed
with the pilot being trained.
COURSE OBJECTIVES
Upon the completion of this course, the pilot will have the necessary knowledge and skills to
demonstrate that he/she is the master of the aircraft, with the successful outcome of a proce-
dure or maneuver never in doubt, and to meet or exceed the requirements/standards listed in the
Airline Transport Pilot and Type Rating Practical Test Standards.
TRAINING SCHEDULE (TYPICAL) 
Listed below is a typical schedule for the pilot training curriculum. Additional days may be re-
quired if qualifying in the aircraft. On occasion, the schedule may be rearranged to meet the
needs of the client or Center. Days off will be scheduled as per each training center.
Ground Training
Hours
Day 1 Classroom ............................................................................................................................ 6.5
Aircraft General
CRM
Master Warning Systems
Avionics/FMS
Day 2 Classroom ............................................................................................................................ 6.0
Review
Avionics/FMS
Day 3 Classroom ............................................................................................................................ 6.0
Review
Electrical Systems
Lighting System
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Day 4 Classroom ............................................................................................................................ 6.0
Review
Hydraulics Systems
Landing Gear and Brakes
Day 5 Classroom ............................................................................................................................ 6.0
Review
Flight Controls
Fuel System
APU
Day 6 Classroom ........................................................................................................................... 6.0
Review
Powerplant
Thrust Reversers
Fire Protection
Day 7 Classroom ........................................................................................................................... 6.0
Review
Pneumatics System
Ice and Rain Protection Systems
Air Conditioning System
Pressurization/Oxygen Systems
Day 8 Classroom ........................................................................................................................... 7.0
Entire System Review
Examination
Day 9 Classroom ........................................................................................................................... 6.0
Airplane Flight Manual
Flight Planning
Performance
Weight and Balance
Day 10 Classroom ........................................................................................................................... 6.0
3A/R Avionics/FMS
Systems Integration
Day 11 Classroom ........................................................................................................................... 6.0
3A/R Avionics/FMS
Systems Integration
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Simulator Training (Based on Training as a Crew)
Hours
Day 12 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period One .............................................................................................4.0
Day 13 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period Two............................................................................................. 4.0
Day 14 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period Three .......................................................................................... 4.0
Day 15 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period Four ............................................................................................ 4.0
Day 16 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period Five............................................................................................. 4.0
Day 17 Briefing Room Oral and Pre/Post Briefings ........................................................................ 2.5
Flight Simulator—Period Six
Simulator Practical Test ....................................................................................................... 4.0
Aircraft—Preflight Inspection ............................................................................................. 2.0
Day 18 Simulator Pre/Post Briefings................................................................................................ 1.5
Flight Simulator—Period Seven (LOFT) ............................................................................ 2.5
or
Aircraft—Flight Training (if required) ................................................................................ 1.5
Day 19 Aircraft Practical Test (if required)...................................................................................... 1.0
NOTE
Flight time may vary due to weather and air traffic conditions.
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SIMULATOR AND FLIGHT TRAINING 
Pilot performance during simulator and flight training shall be graded as: Proficient (1), Normal
Progress (2), Additional Training Required (3), Unsatisfactory (4), or Discussed (D).
The criteria for grading shall be as follows:
• Proficient (1)—The pilot is able to easily perform the procedure or maneuver and is the
obvious master of the aircraft, with the successful outcome of the maneuver never in doubt.
• Normal Progress (2)—The pilot is making satisfactory progress toward proficiency in the
procedure or maneuver but still requires assistance from the instructor. However, the in-
structor is satisfied that, with additional practice as provided in the standard syllabus, the
pilot will become fully proficient in the maneuver or procedure.
• Additional Training Required (3)—The pilot’s progress is not satisfactory. However, the
instructor is of the opinion that additional training over and above that specified in the
syllabus will enable the pilot to become proficient.
• Unsatisfactory (4)—The pilot shows basic deficiencies, such as lack of knowledge, skill,
or ability to perform the required procedures or maneuvers. If the present level of per-
formance or progress is maintained, it is doubtful that the pilot will become proficient.
Further training shall be given only after review by the Center Manager.
• Discussed (D)—This designation indicates that the item was discussed and not performed
in the simulator or aircraft. The discussion revealed a satisfactory knowledge of the ap-
propriate procedure, aircraft system, etc.
• Trained (T)—Trained in maneuvers for procedures only, no flight training credit taken.
Simulator and Flight Training—The pilot is required to achieve a grade of 1 (proficient) by the
completion of training. Additional training will be provided in the portion of the flight in which
the pilot experienced difficulty. Decision to terminate training for a pilot who demonstrates sub-
standard performance will be made by the Center Manager.
COMPLETION STANDARDS 
The pilot must demonstrate through written examination and simulator/flight practical tests that
he/she meets the qualification standards for each segment of the course:
• Aircraft Ground Training—The pilot must demonstrate adequate knowledge of the
Challenger series airplane to pass a written final exam with a minimum passing grade of
70% corrected to 100%.
• Simulator/Flight Training—The pilot will meet the standards of the Airline Transport Pilot
and Type Rating Practical Test Standards (FAA-S-8081-5).
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AIRCRAFT GROUND TRAINING
CURRICULUM SEGMENT
CURRICULUM SEGMENT OUTLINE
Objective: To provide pilots with the necessary knowledge for understanding the basic func-
tions of aircraft systems, the use of the individual system controls, and the integration of air-
craft systems with operational procedures to sufficiently prepare them to enter the flight training
curriculum segment.
FlightSafety Administration
General Operational Subjects Modules
A. Weight and Balance
B. Performance
C. Flight Planning
D. Approved Flight Manual
E. Windshear (optional)
F. High Altitude Training (if required)
Aircraft Systems Modules
A. Aircraft General
B. Electrical Power
C. Lighting
D. Master Warning
E. Fuel
F. APU
G. Powerplant
H. Thrust Reversers
I. Fire Protection
J. Pneumatics
K. Ice and Rain Protection
L. Air Conditioning
M. Pressurization
N. Hydraulic Power
O. Landing Gear and Brakes
P. Flight Controls
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Q. Avionics
R. Oxygen
S. Systems Review and Examination
Completion Standards
A. Systems—The pilot must demonstrate adequate knowledge of the aircraft systems, per-
formance, and flight planning by successfully completing a written examination with a min-
imum score of 70%, that is corrected to 100%.
B. Systems Integration—The pilot must be able to describe, locate, and identify aircraft sys-
tems, and perform normal, abnormal, and emergency checklists. 
TRAINING MODULE OUTLINES
General Operational Subjects
The subject of ground training, referred to as “general operational subjects,” includes instruc-
tion on certain operational requirements that are specific to the FAR 135 certificate holder and
to the aircraft in which the training is being conducted. Training in general operational sub-
jects are not conducted by FlightSafety unless specifically pertinent to this course.
A. Weight and Balance Module
1. General Elements
a. Principles and Methods of Weight and Balance Determination
2. Limitation Elements
3. Operational Elements
B. Aircraft Performance Module
1. General Elements
a. Use of Charts, Tables, Tabulated Data, and other related material
b. Performance Problems
c. Performance Limiting Factors
C. Flight Planning Module
1. General Elements
a. Flight Planning Charts
2. Operational Elements
3. Limitation Elements
D. Approved Flight Manual Module
1. Applicability and Description of the AFM
2. Normal, Abnormal, and Emergency Procedures Sections
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3. Limitations Section
4. Maneuvers and Procedures Section
5. General Performance Section
6. Systems Description
7. Appendices and Bulletins
E. Windshear (optional)
1. Windshear Weather
2. High Altitude Meteorology
3. Lessons Learned from Windshear Encounters
a. Encounter During Takeoff—After Lift-off
b. Encounter During Takeoff—On Runway
c. Encounter On Approach
d. Windshear Effects on Airplanes and Systems
e. Development of Wind Models
4. Model of Flight Crew Actions
a. Evaluate the Weather
b. Avoid Known Windshearc. Consider Precautions
(1) Takeoff Precautions
(2) Approach Precautions
d. Follow Established Standard Operating Techniques
e. Windshear Recovery Techniques
(1) Encounter During Takeoff—After Lift-off
(2) Encounter on Approach
(3) Encounter During Takeoff—On Runway
F. High Altitude Training [IAW §61.31 (f)(l)(i)] (as required)
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency Procedure Elements
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Aircraft Systems Modules
A. Aircraft General Module
1. General Elements
a. System Description
b. Controls and Components
c. Annunciators
d. Miscellaneous
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
B. Electrical Module
1. General Elements
a. System Description
(1) AC System
(2) DC System
(3) Circuit-Breaker Panels
b. Controls and Components
c. Annunciators
d. Miscellaneous
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
C. Lighting Module
1. General Elements
a. System Description
b. Controls and Components
c. Miscellaneous
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
D. Warning Systems Module
1. General Elements
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
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E. Fuel Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
F. Auxiliary Power Unit (APU) Module
1. General Elements
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
G. Powerplant Module
1. General Elements
a. System Description
b. Controls and Components
c. Indicators/Indications
d. Annunciators
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
H. Thrust Reversers Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
I. Fire Protection Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
J. Pneumatics Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
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K. Ice and Rain Protection Module
1. General Elements
a. System Description
b. Controls and Components
c. Annunciators
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
L. Air Conditioning Module
1. General Elements
2. Operational Elements
3. Limitations Elements
4. Emergency/Abnormal Procedure Elements
M. Pressurization Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
N. Hydraulics Module
1. General Elements
a. System Description
b. Controls and Components
c. Indicators/Indications
d. Annunciators
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
O. Landing Gear and Brakes Module
1. General Elements
a. System Description
b. Controls and Components
c. Indicators/Indications
d. Annunciators
e. Servicing/Preflight
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2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
P. Flight Controls Module
1. General Elements
a. System Description
b. Controls and Components
c. Indicators/Indications
d. Annunciators
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
Q. Avionics Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
R. Oxygen Module
1. General Elements
2. Operational Elements
3. Limitation Elements
4. Emergency/Abnormal Procedure Elements
S. Review Module 
1. Written Examination with a Passing Grade of 70%, Corrected to 100%.
Systems Integration Modules
Systems Integration provides the pilot with instruction on aircraft systems interrelationships
with respect to normal, abnormal, and emergency procedures. Pilots will be introduced to, and
will exercise in, the elements of crew resource management as part of the integration process,
including, but not limited to such elements as the following:
• Situational Awareness and the Error Chain
• Synergy and Crew Concept
• Workload Assessment and Time Management
Pilots will become familiar with the cockpit layout, checklists, maneuvers, and procedures.
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NOTE
An individual who has experience as a SIC in the aircraft and has previously completed
an equivalent training curriculum without obtaining a type rating may complete Systems
Integration Training during the prebriefing sessions prior to each Simulator Module.
FLIGHT TRAINING CURRICULUM SEGMENT
TRAINING HOURS
Each flight simulator period is approximately 2.0 hours in length and incorporates an additional
0.5 hour for prebriefing and 0.5 hour for debriefing. Prebrief and postbrief times are reflected
in the Ground School Training Hours Summary section of Chapter 5.
Training is generally conducted as a crew; however, a pilot training alone may complete the course.
The training hours for pilots training as a crew are specified in the table below. The training
hours for a pilot training as a single crew member are specified in subsequent tables.
FLIGHT TRAINING MODULE OUTLINES
Flight Training Curriculum Segment
A. Objective: With the use of an approved flight simulator, cockpit checklist, and appropri-
ate instrument approach and airport charts, the pilot will be able to accomplish the Normal
and Emergency/Abnormal checklists, perform selected maneuvers and procedures, and im-
plement Cockpit Resource Management techniques.
B. Training Equipment and Location
1. Simulators
a. Montreal
b. Houston
c. Tucson
2. Cockpit Poster Panel—Briefing Room
3. Cockpit Checklists—Simulator and Briefing Room
4. Instrument Approach and Airport Charts—Simulator and Briefing Room
C. Maneuvers and Profiles
NOTE
Training includes, but is not limited to the following maneuvers and procedures:
1. Preparation
a Prestart Procedures
b. Performance Limitations
c. Full Cockpit Check
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2. Surface Operation
a. Starting
b. Taxi
c. Pretakeoff Checks
3. Takeoff and Departure
a. Normal
b. Area Departure
c. Rejected
d. Instrument
e. Crosswind
f. Powerplant Failure V1
g. Takeoff with Lower than Standard Minimum
4. Instrument/Area Departure
a. Climb—Normal
b. Climb—One Engine Inoperative to Enroute Altitude
c. Navigation Equipment and Assigned Radials
5. Enroute
a. Steep Turns
b. Approach to Stall—Enroute Configuration
c. Approach to Stall—Takeoff and Approach Flaps
d. Approach to Stall—Landing Configuration
e. Unusual Attitude
f. In-flight Powerplant Shutdown
g. In-flight Powerplant Restart
h. High Speed Handling Characteristics
6. Descent
a. Normal Descent
b. Rapid Decompression
c. Emergency Descent
7. Approaches/Instrument Arrivals
a. Navigation Equipment and Assigned Radials
b. Holding
c. STAR
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d. Precision Approach
(1) ILS
(2) ILS with One Engine Inoperative
(3) Missed
(4) Missed Approach—with a Powerplant Failure
e. Nonprecision Approach
(1) VOR
(2) NDB
(3) LOC
(4) One Engine Out
(5) Missed
(6) Missed Approach—with a Powerplant Failure
(7) GPS
f. Coupled Approach
g. Circling Approachh. Visual Approach
8. Landings
a. Normal Landing
b. Landing from a Precision Approach
c. Crosswind
d. Maneuver to Landing with Powerplant Failure
e. Landing with an Engine Out—Full Stop
f. Rejected Landing to a Missed Approach
g. Landing from a Circling Approach
h. Zero Flap Landing
i. After Landing Procedures
j. Parking and Securing
9. Other Flight Procedures
a. ATC Procedures/Phraseology
b. Ice Accumulation on Airframe
c. Windshear
d. Emergency Evacuation
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10. System Procedures (Normal, Abnormal, and Emergency)
a. Air Conditioning
b. Aircraft Fires
c. Anti-icing and Deicing
d. APU
e. Autopilot
f. Cabin Fire/Smoke
g. Communication Equipment
h. Electrical
(1) AC
(2) DC
i. Fire Protection
j. Flap Systems
k. Flight Controls
l. Flight Instrument System
m. FMS/Automatic or Other Approach and Landing Systems
n. Fuel
o. Hydraulics
p. Landing Gear
q. Navigation Systems
r. Pneumatics
s. Powerplant
t. Pressurization
u. Stall Warning Devices
11. Human Factors
a. CRM
b. Attitude
c. Judgement
d. Checklist/QRH
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COMPLETION STANDARDS
The pilot must perform all procedures and maneuvers to the tolerances listed in the Airline
Transport Pilot/Type Rating Practical Test Standards. It is expected, during the earlier simula-
tor flights, that the tolerance for completion of a maneuver or procedure be greater than during
the later simulator flights. In all cases, it is expected that the pilot should strive to meet the tol-
erances listed below.
1. Takeoff
a. Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
b. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5 KIAS
2. Departure, Cruise, Holding, and Arrival
a. Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
b. Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10°
c. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10 KIAS
3. Steep Turns
a. Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
b. Rollout Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10°
c. Bank Angle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
d. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10 KIAS
4. Approach to Stall
a. Recognize Perceptible Stall or Stall Warning
b. Recover at First Indication of Stall
c. Strive for Minimum Altitude Loss
5. IFR Approaches (Prior to Final Approach)
a. Nonprecision Approach
(1) Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
(2) Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
(3) Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5 KIAS
b. Precision Approach
(1) Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
(2) Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
(3) Airspeed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10 KIAS
6. IFR Approaches (During Final Approach)
a. Nonprecision Approach
(1) CDI . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±1/2 Scale Deflection
(2) RMI . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5° Deviation
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(3) Bearing Pointer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5° Deviation
(4) MDA. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +50, –0 Feet
(5) Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +5, –0 KIAS
b. Precision Approach
(1) Glide Slope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±1/4 Scale Deflection
(2) Localizer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±1/4 Scale Deflection
(3) DH . . . . . . . . . . . . . . . . . . . . . . –0 Feet Prior to Initiating Missed Approach
(4) Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +5, –0 KIAS
7. Circling to Land
a. Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +100, –0 Feet
b. Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
c. Bank Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Should Not Exceed 30°
d. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5 KIAS
8. Missed Approach
a. Altitude. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
b. Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5°
c. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±5 KIAS
9. Landing
a. Final Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +5, –0 KIAS
10. Powerplant Failure-Multiengine Aircraft
a. Altitude. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±100 Feet
b. Heading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10° (±5˚)*
c. Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±10 KIAS (±5 KIAS)*
* During Takeoff
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CL 601-3A/R 1-i
CHAPTER 1
AIRCRAFT GENERAL
CONTENTS
Page
INTRODUCTION ................................................................................................................... 1-1
GENERAL............................................................................................................................... 1-1
STRUCTURES........................................................................................................................ 1-6
Fuselage ........................................................................................................................... 1-6
Doors................................................................................................................................ 1-9
Wing............................................................................................................................... 1-13
Airplane Parking and Mooring ...................................................................................... 1-14
AIRPLANE SYSTEMS ........................................................................................................ 1-14
Electrical System ........................................................................................................... 1-14
Lighting..........................................................................................................................1-15
Warning Systems ........................................................................................................... 1-15
Fuel System.................................................................................................................... 1-15
Auxiliary Power Unit ..................................................................................................... 1-16
Powerplants.................................................................................................................... 1-16
Fire Protection................................................................................................................ 1-16
Pneumatic System.......................................................................................................... 1-16
Ice and Rain Protection.................................................................................................. 1-17
Air Conditioning ............................................................................................................ 1-17
Pressurization................................................................................................................. 1-17
Hydraulic Power Systems .............................................................................................. 1-18
Landing Gear and Brakes .............................................................................................. 1-18
FOR TRAINING PURPOSES ONLY
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Flight Controls ............................................................................................................... 1-18
Avionics ......................................................................................................................... 1-19
Oxygen System.............................................................................................................. 1-19
1-ii CL 601-3A/R FOR TRAINING PURPOSES ONLY
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CL 601-3A/R 1-iii
ILLUSTRATIONS
Figure Title Page
1-1 Canadair Challenger ................................................................................................. 1-2
1-2 Dimensions and Ground Clearances ........................................................................ 1-2
1-3 Turning Radii............................................................................................................ 1-4
1-4 Danger Areas............................................................................................................ 1-5
1-5 Flight Compartment ................................................................................................. 1-6
1-6 Structural Subassemblies.......................................................................................... 1-7
1-7 Nose Landing Gear Assembly.................................................................................. 1-8
1-8 Passenger Compartment (Typical) ........................................................................... 1-8
1-9 Engine and Pylon...................................................................................................... 1-9
1-10 Vertical and Horizontal Stabilizers........................................................................... 1-9
1-11 Passenger and Crew Door ...................................................................................... 1-10
1-12 Overwing Emergency Exit ..................................................................................... 1-11
1-13 Baggage Compartment Door.................................................................................. 1-12
1-14 Rear Equipment Bay Door ..................................................................................... 1-13
1-15 Winglet ................................................................................................................... 1-13
1-16 Landing Gear Locking Pins ................................................................................... 1-14
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
INTRODUCTION
This training manual provides a description of the major airframe and engine systems
installed in the Canadair Challenger CL-600-2B16, model CL-601-3A/R.
This chapter covers the structural makeup of the airplane and provides a general overview
of the systems. The material has been prepared from the basic design data and is not
meant to supersede any of the manufacturer’s system or operating manuals. All subse-
quent changes in airplane appearance or system operation will be covered during aca-
demic training and in subsequent revisions to this manual.
GENERAL
The airplane is manufactured by Canadair
Limited. It is a sweptwing, twin-engine mono-
plane designed to accommodate a crew of 3
and a maximum of from 8 to 19 passengers in
spacious comfort.
The airplane is powered by two General Elec-
tric CF-34 turbofan engines and is certificated
in accordance with FAR 25 and FAR 36 with
appropriate amendments.
CHAPTER 1
AIRCRAFT GENERAL
CL 601-3A/R 1-1FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
Figure 1-1 shows a completed Challenger, and
Figure 1-2 displays the dimensions and ground
clearances of the airplane.
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CL-600-2B16 PILOT TRAINING MANUAL
Figure 1-1. Canadair Challenger
Figure 1-2. Dimensions and Ground Clearances (Sheet 1 of 2)
3 FT 1 IN.
5 FT 4 IN.
61 FT 7 IN.
60 FT 0 IN. (No Tail Tank)
68 FT 5 IN.
20 FT 8 IN.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 1-3FOR TRAINING PURPOSES ONLY
19 FT 2 IN.
7 FT 6 IN.
64 FT 4 IN.
10 FT 5 IN.
12 FT 2 IN. 
9 FT 4 IN. 
STATIC GROUND LINE
8 FT 10 IN. 20 FT 4 IN.
27 FT 8 IN.
13 FT 1 IN.
4 FT 2 IN.
CABIN INTERIOR DIMENSIONS
LENGTH
WIDTH (MAX DIA)
HEADROOM (FLOOR TO
MAX HEIGHT)
VOLUME
FLOOR AREA
FLOOR WIDTH
28 FT 3 IN.
8 FT 2 IN.
6 FT 1 IN.
1,150 CU FT
202.5 SQ FT
7 FT 2 IN.
Figure 1-2. Dimensions and Ground Clearances (Sheet 2 of 2)
Figure 1-3 shows the turning radii applicable
to the maximum nosewheel steering angle of
55 degrees, the minimum taxi strip width re-
quired for a 180-degree turn at the maximum
steering angle, and the turning radii for the
maximum towing angle of 90 degrees.
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CL-600-2B16 PILOT TRAINING MANUAL
MAXIMUM
STEERING
ANGLE
8 IN
MINIMUM TURNING RADII USING NOSEWHEEL STEERING
(MAXIMUM STEERING ANGLE 52° TO 55°)
46
 F
T
26 FT
15 FT
33 FT
40 
FT
53
 FT
NOTE:
LESSER STEERING ANGLES
REQUIRE WIDER TAXI STRIP
FOR 180-DEGREE TURN.
THEORETICAL TURNING POINT
WITH NOSEWHEEL AT 52° TO 55°
61-FOOT MINIMUM
TAXI STRIP WIDTH FOR
180-DEGREE TURN
52°
TO 55°
26 FT
33
 F
T
38
 F
T
NOTE:
MINIMUM TURNING RADII WITH
NOSEWHEEL AT 90°
(AIRPLANE MAXIMUM TOWING ANGLE)
90°
THEORETICAL TURNING POINT
WITH NOSEWHEEL AT 90°
Figure 1-3. Turning Radii
While engines are running, there are dan-
ger areas to the front and rear of the en-
gines . Figure 1-4 shows the temperature
and dis tance cr i ter ia .
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 1-5FOR TRAINING PURPOSES ONLY
MAXIMUM
THRUST
IDLE
THRUST
18 FT
MAXIMUM
THRUST
IDLE
THRUST
180
170
160
150
140
130
120
110
100
90
80
70
60
50
40
30
20
10
0
FEET
180
170
160
150
140
130
120
110
100
90
80
70
60
50
40
30
20
10
0
FEET
532° C
(990° F)
393° C
(740° F)
60° C
(140° F)
27° C
(80° F)
227° C
(440° F)
60° C
(140° F)
27° C
(80° F)
400 FT/SEC
800 FT/SEC
200 FT/SEC
200 FT/SEC
150 FT/SEC
100 FT/SEC
50 FT/SEC
30 FT/SEC
100 FT/SEC
50 FT/SEC
18 FT
12 FT
25
 F
T
Figure 1-4. Danger Areas
STRUCTURES
FUSELAGE
The fuselage is an all-metal, semimonocoque
structure comprising nose, center, and tail
sections riveted together.The nose section is
primarily the flight compartment areas. The
center section includes the passenger cabin
area and the avionics bay. The wing is bolted
to the fuselage below the avionics bay. The tail
section is primarily an equipment bay to which
the engines and empennage are attached.
Figure 1-6 illustrates the subassemblies of the
basic structure.
Nose Section
The nose section is effectively divided by the
flight compartment floor into upper and lower
halves. The upper half comprises the flight
compartment and forward avionics bay.
The flight compartment (Figure 1-5) contains
the airplane instruments, caution warning
lights, controls, circuit-breaker panels, two
crew seats, control columns, and pedals. On
the center windshield post there is a pilot’s eye
locator to enable seat adjustment for optimum
field of vision.
Miscellaneous items in the flight compart-
ment include a portable fire extinguisher, two
oxygen masks, two cup holders, storage boxes,
pouches for checklists, and smoke goggles etc.
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Figure 1-5. Flight Compartment
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 1-7FOR TRAINING PURPOSES ONLY
Figure 1-6. Structural Subassemblies
GE CF-34
ENGINE COWLINGS
ENGINE
PYLON
MAIN ENTRANCE
DOOR
EMERGENCY EXIT
BAGGAGE
DOOR
PRESSURIZED PASSENGER 
COMPARTMENT
(FUSELAGE CENTER 
SECTION)
FLIGHT 
COMPARTMENT
FORWARD 
AVIONICS BAY
(UNPRESSURIZED)
HYDRAULICALLY
OPERATED NOSE
GEAR DOORS
FORWARD HYDRAULICS
EQUIPMENT BAY
RADOME
AIR-DRIVEN GENERATOR
COMPARTMENT
NOSE GEAR REAR DOOR
NOSE LANDING GEAR
FLIGHT COMPARTMENT FLOOR
UNDERFLOOR AVIONICS
BAY (PRESSURIZED)
FWD
AUX TANK
PASSENGER
CABIN FLOOR
AFT AUXILIARY TANK
REAR PRESSURE
BULKHEAD
REAR FUSELAGE UNDER-
FLOOR AREA (PRESSURIZED)
MAIN LANDING GEAR WHEEL BAY
AFT EQUIPMENT BAY
(UNPRESSURIZED)
TAIL CONE
RUDDER
HORIZONTAL
STABILIZER
ELEVATORS
VERTICAL 
STABILIZER
VERTICAL STABILIZER/REAR 
FUSELAGE FRAME 
STRUCTURE
WING-TO-FUSELAGE
REAR FAIRING
WINGLET
FLIGHT
SPOILER
GROUND
SPOILER
FLAPS
MAIN LANDING GEAR
AND DOORS
WING-TO-FUSELAGE
FRONT FAIRING
WING CENTRAL BOX
STRUCTURE
REMOVABLE LEADING EDGES
WINGLET
AILERON
The lower half of the nose section comprises
compartments for the brake accumulators and
brake valve control mechanism, the air-driven
generator, the flight control forward mechanisms,
and the nose wheel well and mounting structure.
A weather radar antenna pedestal, mounted in
front of the upper and lower nose section
halves, is enclosed by a cone-shaped Kevlar
radome. The radome is provided with a sys-
tem of conductive paths to reduce the risk of,
and damage from, lightning strikes.
The nose landing gear (NLG) assembly (Fig-
ure 1-7), mounted on the underside of the nose
section lower half at the rear of the nose wheel
well, is a conventional oleopneumatic, shock-
absorbing strut fi t ted with two steerable
wheels. The NLG retracts forward into the
well and is enclosed within the well by hy-
draulically actuated doors.
Center Section
The center section is divided by a single-level
floor into the pressurized passenger com-
partment (cabin area) (Figure 1-8) and the un-
derfloor area. The passenger compartment
incorporates the passenger-crew entrance door
and the baggage door on the left side, an over-
wing emergency exit door on the right side, and
six windows on each side including one in the
emergency exit door.
The underfloor area is divided by pressure
bulkheads into three sections:
1. The pressurized avionics bay, which
houses various electronic components
2. The unpressurized main landing gear
bay, which houses (1) bins to accept the
main landing gear wheels when retracted
and (2) the reservoir and other compo-
nents of the No.3 hydraulic system.
3. The pressurized rear fuselage under-
floor section
Tail Section
The unpressurized tail section comprises the
rear equipment bay, the vertical stabilizer/rear
fuselage frame structure, and tail cone. The
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CL-600-2B16 PILOT TRAINING MANUAL
Figure 1-7. Nose Landing Gear Assembly
Figure 1-8. Passenger Compartment 
(Typical)
rear equipment bay houses the auxiliary power
unit (APU), two air-conditioning units (ACU),
the reservoirs and other components of hy-
draulic systems No. 1 and No. 2, and the air-
plane battery. An access door is provided at the
bottom center of the rear equipment bay.
An engine pylon is secured to the rear equip-
ment bay above the horizontal centerline on
each side (Figure 1-9). A pressure bulkhead
is built into the front of the rear equipment
bay to withstand the pressure in the fuselage
center section.
The vertical stabilizer/rear fuselage frame
structure is secured to the rear equipment bay.
A fully cantilevered, sweptback vertical sta-
bilizer projects upward from the rear fuselage
structure and is surmounted by a sweptback
trimmable horizontal stabilizer (Figure 1-10).
A single rudder is hinged to the rear of the ver-
tical stabilizer, and an elevator is hinged to
each side of the horizontal stabilizer, trailing
edge. A tail cone is attached to streamline the
rear frame structure and can contain an op-
tional fuel tank.
DOORS
The airplane is provide with a passenger-crew
door, a baggage compartment door, and rear
equipment bay door. An overwing emergency
exit door is provided on the right side of the
passenger compartment.
The passenger-crew door is downward opening,
with the stairs incorporated in the door structure.
The baggage compartment and emergency exit
doors are of the plug type which open inward;
all other doors open outward. All doors are
flush with the airplane outer skin when closed.
Passenger-Crew Door
The entrance door is electrically or manually
operated and downward opening with its stairs
forming an integral part of the door structure.
Its movement is controlled by a counterbalance
system of gas springs and a spring-loaded
cable drum.
The latch mechanism is operated by an inside
single-lever handle, located on the forward
side of the stairs, an inside T-handle which is
recessed in a riser of the stair, and an outside
T-handle which is recessed in the door. (See
Figure 1-11.) Once closed, the door is latched
from the inside by pushing the single-lever
handle downward. The internal T-handle is
then pulled out of its recess to stow the external
handle, which clicks audibly when stowed.
Verification that the external handle is stowed
can be made by checking that the internal,
single-lever handle cannot be pulled up. After
this check, the internal T-handle must be
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 1-9FOR TRAINING PURPOSES ONLY
Figure 1-9. Engine and Pylon
Figure 1-10. Vertical and Horizontal
StabilizersCF-34 ENGINE
stowed in the riser of the stair. Four red visual
alignment marks located on the center latch
cams (2) and the upper latch cams (2) should
be checked to ascertain door security.
The door is unlocked from the inside when the
single-handle lever is pulled upward, releasing
the outer T-handle from its recess. The door is
then unlatched by continuing the rotation of the
inner handle upward. As the door opens, the
handrails unfold upward. A door pull-up assist
handle is attached to the rear handrail to assist
in opening and closing the door.
An electrical power assist system provides an
optional means of closing the door from inside
the aircraft. A control switch, labeled “CABIN
DOOR,” is located just forward of the en-
trance door. Holding the switch to the RAISE
position activates the system, which raises
the door from any open position up to the
closed position. Once the switch is released,
the door can still be operated manually.
The door unlocks from the outside by the op-
eration of a pushbutton in the handle itself,
which releases the T-handle from its recess. To
unlatch the door, the T-handle is turned45
degrees counterclockwise. A pullout handle is
also provided to assist the operator in open-
ing and closing the door from the outside.
When the door is fully open, a support leg ex-
tends to the ground to stabilize the door.
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ROLLER SUPPORT LEG
UPPER
LATCH CAM
CENTER
LATCH CAM
TENSION FITTING
TENSION
FITTING
GUIDE
PLATE
COVER PLATE
PULL-IN LEVER
EXTERNAL HANDLE
INTERNAL
HANDLE
PULL-OUT
HANDLE
PULL-IN
HANDLE
T-HANDLE
GAS SPRING
RUBBER
SEAL
ELECT
DOOR
ASSIST
SWITCH
CABIN
DOOR
SWITCH
STOP
TENSION
BUTTON
TENSION
BUTTON
UPPER LATCHSTOP SPIGOTS
CENTER
LATCH
SPIGOT
CAM
ROLLER
HANDRAILS
HANDRAIL
RESTRAINT
BRACKET
HAND GRIP
SWIVEL
PULLEY
GAS
SPRING
PULLEY
CABLE
Figure 1-11. Passenger and Crew Door
Emergency Exit
An overwing emergency exit is provided on
the right side of the cabin over the wing (Fig-
ure 1-12).
The emergency exit is 23 inches wide, opens
inward, and can be unlatched from the inside
or the outside. The inside unlatching handle
has a Betalight sign which displays “EXIT
PULL” and is readily visible during day-
light or darkness. A handgrip is located im-
mediately below the window, set in the door.
The grip is provided to support the door when
opening it from inside the cabin. The outer
push plate is captioned “PUSH IN FLAP,”
“PUSH DOOR INWARD.”
Baggage Compartment Door
The baggage compartment door is located
on the left side of the airplane immediately
aft of the passenger compartment. The door
opens inward and upward on two sets of
tracks attached to the structure. It is assisted
during opening by balance springs with cable
attachments to the door and structure. The
door is held closed by two plungers which are
operated by either the internal or external
handle.
Two plunger-actuated proximity switches
are installed to provide an indication in the
flight compartment when the door is not
safely closed.
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CL-600-2B16 PILOT TRAINING MANUAL
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Figure 1-12. Overwing Emergency Exit
PLUNGER
FITTING (DOOR-LOCKING
PLUNGER RECEPTACLE)
INNER
HANDLE
HANDLE UNSTOW BUTTON
HANDLE KEY LOCK
GUIDE
TRACKS
BALANCE SPRING BOXES
The baggage compartment door can be opened
or closed from both inside and outside. The ex-
ternal handle incorporates two locking mech-
anisms, one to hold the handle flush with the
door and the other to secure the handle to pre-
vent unauthorized opening. The door handle
security lock is operated by a key, and the
flush hold lock is a single catch which oper-
ates automatically when the handle is aligned
with a recess in the door and pushed into that
recess. The handle flush hold lock is released
by operation of an integral pushbutton on the
handle. When released, the handle moves out
of its recess under spring force, and can then
be turned to release the locking plungers and
open the door (Figure 1-13).
Rear Equipment Bay Door
The rear equipment bay door located at the bot-
tom of the rear fuselage (Figure 1-14) provides
access to the APU, air-conditioning units, ser-
vice points for No. 1 and No. 2 hydraulic sys-
tems, and the airplane battery. The door opens
downward and is secured at the forward edge
by two hinges which are equipped with quick-
release pins to facilitate easy removal.
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Figure 1-13. Baggage Compartment Door
The rear equipment bay door is held closed
by two plungers at the rear edge. Handle op-
eration is the same as for the baggage com-
partment door, with the exception that the
rear equipment bay door must be supported
during opening and closing.
WING
The airplane wing is an all-metal, advance
technology airfoil structure manufactured
as a single unit and bolted to the underside
of the fuselage. The wing comprises left and
right sweptback airfoils connected by a cen-
ter box structure. Ailerons, flaps, spoilers,
integral fuel tanks, and the support struc-
ture of the main landing gear are incorporated
into the wing. Wing leading edges are of alu-
minum alloy and incorporate thermal anti-
icing. Wingtips are of organic fiber (Kevlar)
to save weight and to facilitate repairs. The
Challenger CL-600-2B16 employs winglets
(Figure 1-15).
Fuel for the airplane is stored in three tank
areas. The deep section of the airfoil on each
side is used as a main tank, and the center sec-
tion is used as an auxiliary tank. A fuel-tight
bulkhead each side of the center section sepa-
rates the tanks, and all tanks are compartmented
to ensure a continuous supply of fuel to the en-
gines under all normal flight conditions.
The main landing gear assemblies, mounted
under the left and right airfoils inboard of the
trailing-edge flaps, are conventional oleop-
neumatic, shock-absorbing struts, each fit-
ted with two wheels. The assemblies retract
inward into bins in a well on the underside
of the fuselage.
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CL-600-2B16 PILOT TRAINING MANUAL
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Figure 1-14. Rear Equipment Bay Door
Figure 1-15. Winglet
AIRPLANE PARKING
AND MOORING
When the airplane is stationary on the ground,
precautions must be taken to ensure safety of
personnel and equipment. The extent of safety
measures to be observed depends upon the pre-
vailing or expected weather conditions and the
expected length of time the airplane will be
stationary. The airplane must be parked or
moored into the wind with landing gear lock-
ing pins installed, as illustrated in Figure 1-16.
AIRPLANE SYSTEMS
ELECTRICAL SYSTEM
The Challenger was the first executive jet to
use AC power as its primary electrical system.
Its DC requirements are met through the use
of transformer-rectifier units which convert
115 VAC to 28 VDC. A single nickel-cad-
mium battery is used for starting the APU and
as an emergency DC backup.
AC System
The AC system is divided into four subsystems:
1. Primary system
2. Auxiliary system
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Figure 1-16. Landing Gear Locking Pins
3. External system
4. Emergency system
Primary AC system power normally supplies
all the airplane electrical needs in flight. Each
engine drives an integrated drive generator
(IDG) which supplies 3-phase, 115-/200-volt,
400-Hz AC power, rated at 30 kva.
Auxiliary AC system power is supplied by an
APU whose output has the same rating as ei-
ther IDG. It is used to supply all electrical re-
quirements on the ground and can be used as
an emergency supply in the air.
External AC system power is received through
a receptacle located on the right-hand nose
section of the fuselage. If either main gen-
erator or the APU generator is brought on
line, external power automatically reverts to
standby status.
Emergency AC system power is supplied by
an air-driven generator which can be deployed
manually or automatically when both main
generators fail or are not available in flight.
Once deployed, it cannot be restowed until
after the airplane lands.
DC System
DC power can be supplied from two sources,
static conversion and the battery. On the
ground, DC power can also be supplied through
an external power receptacle on the right rear
of the fuselage. In normal operation, primary
DC power is derived from the AC system
through the use of four transformer-rectifier
units (TRU’s). The battery system provides
power to start the APU and to supply DC power
when all other sources of DC power have failed
or are not available. It also supplements the
available DC power supply when the ADG is
the only source of airplane power.
LIGHTING
The lighting system, controlled from the cock-
pit, provides illumination both externally and in-
ternally throughout the airplane. The exterior
lighting includes four landing lights and two
taxi/recognition lights; two rotating red bea-
cons;standard red, green , and white navigation
lights; wing-ice inspection lights; and anticol-
lision/strobe lights. The interior lighting in-
cludes boarding and dome lights to illuminate
the passenger door area, service compartment
lights, and flight compartment lighting.
In addition, the airplane is equipped with an
emergency lighting system which illuminates
the right wing and passenger door area for
emergency evacuation.
WARNING SYSTEMS
Warning systems provide the flight crew with
18 visual indications of systems malfunctions
through the master caution system and draw
attention to 8 significant events through the
aural warning system.
The Annunciator Section in this manual dis-
plays all light indicators, and page ANN-1
should be folded out and referred to while
studying this manual.
FUEL SYSTEM
The Challenger uses a wet-wing box struc-
ture which forms two main tanks in the out-
board wing sections and an auxiliary tank in
the wing center section. Maximum fuel ca-
pacity is approximately 17,900 pounds with
a tail tank.
Ejector pumps are used to ensure the delivery
of fuel to the collector tanks and to supply it
to the engines. Electric standby fuel pumps op-
erate automatically during engine starting and
in the event of main ejector pump failure.
A cross flow valve may be opened to allow
gravity flow to correct a fuel imbalance be-
tween the main tanks.
Fueling may be accomplished by gravity feed
but it is normally done through a single-point
pressure refueling connection located in the
right wing root.
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AUXILIARY POWER UNIT
The auxiliary power unit (APU) installed in the
Challenger is a self-contained gas turbine en-
gine manufactured by the Garret Turbine En-
gine Company and is designated as “GTCP
36-100 (E).”
The APU is installed within a fireproof enclo-
sure in the aft equipment bay, behind the rear
pressure bulkhead. It is provided with indepen-
dent fire detection and extinguishing systems.
The primary functions of the APU are:
• To provide compressed air for engine
starting and for cabin and cockpit heat-
ing and cooling through the environ-
mental control system.
• To drive a generator for the supply of AC
electrical power. In-flight use is for
emergencies only, when both main gen-
erators have failed.
The APU is independent of all airplane sys-
tems with the exception of a DC power sup-
ply for starting and a fuel supply from the
airplane fuel system.
POWERPLANTS
The Challenger CL-600-2B16 uses two Gen-
eral Electric CF-34 powerplants. Each is a
high-bypass front fanjet engine with a 6.2 to
1 bypass ratio. It incorporates a 14-stage axial-
flow compressor driven by a 2-stage air-
cooled, high-pressure turbine, an annular
combustion section, a single-stage independent
front fan driven by a 4-stage , low-pressure tur-
bine, a fixed-area concentric exhaust section,
and an integrated control system.
The high pressure single-spool compressor
incorporates variable-inlet guide vanes and
five stages of variable-stator vanes that enable
the engine to make stall-free accelerations.
The front fan, which increases mass airflow
and decreases jet velocity, gives the CF-34 a
large increase in thrust over that available
from a comparable turbojet, while consuming
the same amount of fuel. This significantly in-
creases the range capability of the airplane.
At sea level on a standard day, the CF-34 de-
livers 8,729 pounds of thrust at takeoff power
with 9,220 pounds of thrust available through
an automatic power reserve system.
Thrust reversers are installed for ground use.
FIRE PROTECTION
The fire protection system provides a means
of detecting and extinguishing fires in the en-
gines and the auxiliary power unit. It may be
considered as two systems, a fire detection
system and a fire-extinguishing system. The
fire detection system consists of three sepa-
rate sensing loops which provide visual and
aural warnings for detected fires. The fire-ex-
tinguishing system consists of three bottles
manually activated from the cockpit. In addi-
tion, the airplane is equipped with a bleed-air
leak detection system, a main landing gear
overheat warning system, and engine jet
pipe/pylon overheat detection systems. A
portable fire extinguisher is mounted on the
flight deck.
PNEUMATIC SYSTEM
The pneumatic system distributes bleed air for
use in engine starting, anti-icing systems, air-
conditioning, pressurization, and thrust-re-
verser operation. Compressed bleed air may
be extracted from both airplane engines and/or
the APU. There is also provision for connect-
ing an external air source while on the ground.
Whatever the source of the bleed air, it is
routed into a bleed air manifold. The flow of
air through the manifold is controlled by elec-
trically selected, pneumatically operated
valves. Control and monitoring of the mani-
fold air supply is performed by the pilot, using
the bleed air control panel located overhead
in the cockpit.
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ICE AND RAIN PROTECTION
Ice and rain protection is divided into two cat-
egories, depending on whether they use bleed
air or electrical power as a source of energy.
The anti-iced areas and their sources are:
• Pneumatic (hot bleed air)
• Wing leading edge
• Engine inlet area
• Electric
• Windshields
• Air data sensors
The empennage does not require anti-icing.
The wing leading edges and engine inlet areas
are anti-iced using engine bleed air. This is
drawn from the 14th compressor stage.
The windshields are anti-iced using the AC
electrical power system which also provides
a defog capability to the windshield and side
window panels.
Seven air data sensors are electrically anti-iced
by 115-VAC power heating integral elements
in each sensor.
All anti-icing operations are electrically con-
trolled with switches located in the cockpit.
Rain protection is provided by continuous en-
gine ignition and by an acrylic windshield on
which water beads; therefore, the use of wipers
is not required.
AIR CONDITIONING
The air-conditioning system uses the air cycle
system of cooling. It supplies air to the cabin
and flight compartment for heating, cooling,
ventilation, and pressurization.
The air-conditioning system includes the pres-
sure regulating and shutoff valves, refrigerated
and conditioned air ventilations system, and
separate cabin and flight compartment tem-
perature control systems.
Two identical air-conditioning units (ACUs)
are located in the aft equipment bay. Air drawn
through a ram-air intake at the base of the ver-
tical stabilizer is ducted through the units as
a cooling agent. A ducting system associated
with each air-conditioning unit is responsible
for delivering conditioned air to the cabin area
from the right ACU and to the flight deck and
cabin area from the left ACU. Cabin and flight
compartment temperature is controlled from
the cockpit in an automatic or manual mode.
Normal pressurization of the airplane is
achieved with conditioned air from the ACU’s.
On aircraft 5001-5134 emergency pressuriza-
tion is provided if both ACUs are unavailable.
PRESSURIZATION
The pressurized area of the airplane extends
from the bulkhead immediately forward of
the windshield to the pressure bulkhead at the
rear of the cabin, including the baggage com-
partment and the underfloor area. The pressure
in this area is controlled by two outflow valves
in the rear pressure bulkhead. The outflow
valves are operated by an automatic controller
with a manual pneumatic controller provided
as a backup. The pressurized area of the air-
plane is maintained at a selected altitude of
from -1,000 feet to +10,000 feet. Normal cabin
differential pressure is maintained at 8.8 psi.
Failure of the automatic mode necessitates
the use of the manual mode. In this mode, the
outflow is controlled pneumatically without
the need for electrical power. Operationin the
manual mode is accomplished by adjusting a
manual regulator and monitoring the response
on the cabin altitude indicator.
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Override controls are incorporated into the
outflow valves, which will provide for:
• Overpressure limiting, which limits the
maximum cabin differential pressure to
9.1 ±0.1 psi, regardless of the selector
setting
• Altitude limiting, which prevents the
cabin altitude from exceeding 12,500
±500 feet
• Negative pressure relief, which prevents
ambient pressure from exceeding cabin
pressure (negative pressure) by more
than -0.5 psid
HYDRAULIC POWER SYSTEMS
The Challenger has three independent hy-
draulic systems which are designated “No.
1,” “No. 2,” and “No. 3.” These systems pro-
vide hydraulic power to operate the primary
flight controls, flight and ground spoilers,
landing gear, nosewheel steering, and the
wheel brakes. All systems operate at a work-
ing pressure of 3,000 psi.
Each system contains its own reservoir and op-
erates continuously to supply operating pres-
sure to its respective subsystems.
System No. 1 is pressurized by the left engine-
driven hydraulic pump and/or by an AC elec-
tric pump (designated pump “1B”) located on
the left side of the aft equipment bay.
System No. 2 is similar and is pressurized by
the right engine-driven hydraulic pump and/or
by an AC electric pump (designated pump
“2B”) located on the right side of the aft
equipment bay.
System No. 3 is pressurized by two AC elec-
tric pumps, designated pump “3A” and pump
“3B,” which are located in the rear wing root
on each side of the airplane.
The hydraulic system control panel, located
overhead in the cockpit, provides the pilots
with the switches, lights, and gauges for sys-
tem management.
LANDING GEAR AND BRAKES
The Challenger employs three retractable land-
ing gear assemblies, each fitted with dual
wheels and an oleopneumatic shock strut.
The main landing gear retracts inward into
recesses in the wing and center fuselage. The
nose landing gear retracts forward. Normal
extension and retraction is electrically con-
trolled and hydraulically operated. For emer-
gency landing gear operation, the gear may be
extended by pulling the landing gear manual
release T-handle in the flight compartment.
The pilot’s handwheel provides 55° nose-
wheel steering each side of center through
hydraulic actuators or, with hydraulic pressure
removed, the airplane can be turned using
differential braking.
Each main wheel contains a hydraulic multi-
disc brake unit. The inboard brakes are pow-
ered by the No. 3 system, and the outboard
brakes by the No. 2 system. Nitrogen-charged
brake accumulators provide limited braking
pressure if one of the hydraulic systems is not
available. A parking brake handle can be set
to maintain pressure to the brake units.
An antiskid system independently modulates
the hydraulic pressure to each brake to give op-
timum braking performance under all condi-
tions. The system incorporates locked wheel
protection, touchdown protection, and a built-
in test function.
FLIGHT CONTROLS
Hydraulic power is used to operate the ailerons,
elevators, and rudder (the primary flight con-
trols.) Both flight and ground spoilers are also
hydraulically operated. Trailing-edge flaps are
electrically controlled and actuated. The trim
controls are actuated from pilot input switches
and autopilot input signals through electri-
cally driven actuators. Stall protection includes
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a stick shaker, autopilot disconnect, aural and
visual warning, and a stick pusher actuation.
Roll, pitch, and yaw changes are generated
from the control wheels, control columns, and
rudder pedals, conveyed by mechanical means
to separate and independent hydraulic power
control units which move the flight surfaces.
The ailerons and elevators are each powered
by two of the three hydraulic systems, while
the rudder is powered by all three systems.
Loss of any single hydraulic system will not
affect flight control operation.
Lift modulation is accomplished by the de-
ployment of spoilers or flaps. Flight spoilers
consist of two hydraulically actuated panels, one
hinged to each upper wing surface forward of
the outboard flaps. They can be extended to, and
held in, any position between 0° and 40°.
Ground spoilers are located on the upper wing
surface forward of the inboard flaps. They
have two operating positions, fully retracted
or fully deployed (45°) for ground use.
The flap system consists of externally hinged
inboard and outboard slotted flap panels which
are mounted on the wing trailing edges. The
panels are electrically selected and electri-
cally driven. A flap control lever quadrant on
the center pedestal has detents at the 0°, 20°,
30°, and 45° positions.
Aileron and rudder trim are actuated electri-
cally. There are no trim tabs. Pitch trim inputs
are electromechanically applied to an actua-
tor which varies the angle of incidence of the
horizontal stabilizer.
AVIONICS
The following equipment is fitted to the Chal-
lenger before the completion center modifi-
cations are added:
• Pitot-static system
• Flight instruments
• Navigation systems
• Automatic flight control system
• Communications system
OXYGEN SYSTEM
The Challenger is delivered to the customer
with a completely installed cockpit oxygen sys-
tem. A single cylinder supplies oxygen to both
quick-donning, diluter-demand pilot masks.
Each mask contains its own flow regulator.
The oxygen cylinder is ground rechargeable.
It incorporates one pressure-reducing valve
and two relief valves, including an overboard
discharge indicating disc.
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CL 601-3A/R 2-i
CHAPTER 2
ELECTRICAL POWER SYSTEMS
CONTENTS
Page
INTRODUCTION ................................................................................................................... 2-1
GENERAL............................................................................................................................... 2-1
AC SYSTEM ........................................................................................................................... 2-2
Primary AC System.......................................................................................................... 2-2
Auxiliary AC System ....................................................................................................... 2-4
External AC System......................................................................................................... 2-6
Emergency AC System .................................................................................................... 2-8
AC Power Distribution..................................................................................................... 2-9
Control and Monitoring ................................................................................................. 2-12
DC SYSTEM......................................................................................................................... 2-13
DC Static Conversion..................................................................................................... 2-14
Battery System............................................................................................................... 2-16
Battery Overload............................................................................................................ 2-16
Battery Contactor........................................................................................................... 2-16
External DC ................................................................................................................... 2-17
Control and Monitoring .................................................................................................2-19
CIRCUIT-BREAKER PANEL LOCATIONS....................................................................... 2-21
QUESTIONS......................................................................................................................... 2-29
FOR TRAINING PURPOSES ONLY
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ILLUSTRATIONS
Figure Title Page
2-1 AC System Components Locations.......................................................................... 2-2
2-2 Integrated-Drive Generator....................................................................................... 2-3
2-3 Main AC Power (Power Off).................................................................................... 2-4
2-4 Main AC Power APU Feed ...................................................................................... 2-5
2-5 Main AC Power Normal GEN Feed......................................................................... 2-5
2-6 Main AC Power GEN 1 Feed (GEN 2 INOP).......................................................... 2-6
2-7 AC External Power Connector ................................................................................. 2-7
2-8 External AC Power Feed (All GENs INOP) ............................................................ 2-7
2-9 Emergency AC Power Feed (All GENs INOP)........................................................ 2-8
2-10 Air-Driven Generator Location ................................................................................ 2-8
2-11 Ram-Air Turbine ...................................................................................................... 2-9
2-12 Total AC Power ...................................................................................................... 2-10
2-13 Essential AC Power Panel ...................................................................................... 2-11
2-14 ADG Controls......................................................................................................... 2-11
2-15 Electrical Control Panel ......................................................................................... 2-12
2-16 DC System Components Locations ....................................................................... 2-13
2-17 Primary DC Power Feed (Normal Operation)........................................................ 2-14
2-18 Primary DC Power Feed TRU No. 2 INOP ........................................................... 2-15
2-19 Primary DC Power Feed ESS TRU No. 1 INOP ................................................... 2-15
2-20 Battery Master Switch OFF (All Power Off) ......................................................... 2-16
2-21 DC External Power Connector............................................................................... 2-17
2-22 Battery Master Switch ON (W on W).................................................................... 2-17
2-23 External DC Battery Master Switch ON (W on W)............................................... 2-18
FOR TRAINING PURPOSES ONLY
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2-24 Both ESS TRUs Failed (W off W) ......................................................................... 2-18
2-25 DC Electrical Controls ........................................................................................... 2-19
2-26 BATTERY MASTER Switch................................................................................. 2-19
2-27 Total Electrical System........................................................................................... 2-20
2-28 CBP-A Main AC and DC Bus 1............................................................................. 2-22
2-29 CBP-A Battery Bus ................................................................................................ 2-23
2-30 CBP-B Main AC and DC Bus 2............................................................................. 2-24
2-31 CBP-B Battery Bus ................................................................................................ 2-25
2-32 CBP-C AC Essential Bus ....................................................................................... 2-26
2-33 CBP-D DC Essential Bus....................................................................................... 2-27
2-34 CBP-E Battery Direct Bus ..................................................................................... 2-28
TABLES
Table Title Page
2-1 Generator Trip Levels............................................................................................... 2-3
2-2 Circuit-Breaker Panels ........................................................................................... 2-21
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INTRODUCTION
The Canadair Challenger CL-600-2B16, model CL601-3A/R has an electrical system that
is comparable to that of a modern airliner. It was the first business jet to have a primary
electrical system based on AC power. The 115-VAC system is lighter and more reliable for
a given power output than DC-based systems.
Twenty-eight Volts DC is derived from the AC system through the use of transformer-
rectifier units (TRU).
A battery is provided for APU starting and emergency DC backup.
GENERAL
The electrical system consists of two engine-
driven AC generators, one APU-driven AC
generator, one ram-air turbine-driven AC gen-
erator, four transformer-rectifier units, one
nickel-cadmium battery, plus devices for in-
terconnection and control. Control is exer-
cised primarily through the electrical control
panel located on the center console.
Provision is made to connect both AC and DC ex-
ternal power through separate receptacles. The AC
receptacle is located on the right nose section
and the DC receptacle on the aft right fuselage.
Emergency in-flight AC power is provided
automatically by an air-driven generator
(ADG). The generator is driven by a ram-air
turbine (RAT) and supplies power directly to
the ADG bus.
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CHAPTER 2
ELECTRICAL POWER SYSTEMS
AC SYSTEM
The AC electrical system is divided into
four subsystems: primary, auxiliary, exter-
nal, and emergency.
PRIMARY AC SYSTEM
The primary AC system normally supplies all of
the airplane electrical needs in flight. Each en-
gine drives an integrated drive generator which
consists of a constant-speed drive and a matched
generator. The generator supplies 3-phase, 115-
/200-volt, 400-Hz AC power and is rated at 30
kilovolt-amps (kva). Other components include
generator control units, generator line contac-
tors, and generator transfer contactors. The AC
system components locations and interconnec-
tion are shown in Figure 2-1.
Integrated-Drive Generators
(IDG)
Each IDG (Figure 2-2) is made up of a con-
stant speed drive (CSD) and an AC generator
that are assembled into one unit.
Constant-Speed Drive (CSD)
The CSD provides the mechanical interface be-
tween the engine accessory gearbox and the
generator. The CSD uses a self-contained oil
supply as an operating medium. Through a
series of hydromechanical devices, the vari-
able engine rpm is converted to a constant
12,000-rpm output to the generator. Thermal
and torque protection is provided by discon-
nect features built into the CSD.
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APU GEN
IDG 2
IDG 1 
MAIN AC BUS
EXTERNAL AC
CONNECTION
ADG
MAIN ELECTRICAL
DISTRIBUTION PANEL
AC ESSENTIAL BUS
Figure 2-1. AC System Components Locations
The CSD oil also acts as a cooling agent for
the CSD and the generator. The oil is cooled
in a heat exchanger located on the engine
pylon service assembly. A sight glass on the
CSD allows for oil quantitychecking. An oil
filter with a pop-out indicator button is pro-
vided to signal the need for a filter change. A
magnetic chip detection device is incorpo-
rated into the CSD drain plug.
Generator
The AC generator is a 30-kilovolt-amp (kva),
115-/200-volt, 3-phase, 400-Hz, brushless al-
ternator that is lubricated, cooled, and driven
by the CSD.
Generator Control Unit (GCU)
Each generator has its own generator control
unit to monitor, regulate, and control genera-
tor output. In conjunction with the other
GCU's, circuit protection and switching is
provided for the generator line contactors and
the generator transfer contactors.
The GCU will turn off the generator output if any
of the trip levels listed in Table 2-1 are exceeded.
Generator Line Contactor (GLC)
Each generator line contactor is responsible for
supplying its own main bus with power from:
• Its own generator through a generator
line contact; or
• A secondary source through a trans-
fer contactor
The secondary source is used only when the pri-
mary is unavailable. If power is not available
at either contact, the main bus is not powered
and the appropriate MAIN BUS OFF light on
the electrical control panel illuminates.
Generator Transfer Contactor
(GTC)
The GTC operates to supply power to the GLC
on demand. The priority of its sources are: (1)
the APU generator and (2) the opposite IDG.
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GENERATOR
CONSTANT SPEED
DRIVE (CSD)
ENGINE
ACCESSORY
GEARBOX
Figure 2-2. Integrated-Drive Generator
PARAMETER TRIP LEVEL TIME DELAY
Voltage High(125 volts) 10 seconds at 125 volts
0.1 second at 150 volts
Low (100 volts) 4.25 seconds
Frequency High (425 Hz) 0.1 second
Low (375 Hz) 3.5 seconds
Power 64.5 kva 6.5 seconds
Table 2-1. GENERATOR TRIP LEVELS
Provided no airplane generators are on line,
the GTC is capable of accepting power from
a suitable external AC source (Figure 2-3).
AUXILIARY AC SYSTEM
Auxiliary generation is accomplished by a
single APU-driven system which may be used
to feed either or both AC main buses. The
APU generator has the same ratings as an en-
gine integrated-drive generator.
APU Generator Adapter
The generator adapter provides the mechani-
cal interface between the APU gearbox and the
AC generator. Since the APU is operated at a
governed speed of 58,737 rpm, a constant-
speed drive is not necessary. Instead, a straight-
through adapter is utilized to operate the
generator at 12,000 rpm to maintain the out-
put frequency at 400 Hz. A self-contained oil
system supplies oil for adapter and generator
lubrication and cooling. The adapter oil is
cooled by APU fuel through a heat exchanger.
The auxiliary power system GCU provides
the same functions as the primary system
GCU. The auxiliary power system GCU pro-
vides the same functions as the primary sys-
tem GCU. The auxiliary power/external power
(AP/EP) contactor (Figure 2-3) supplies power
to the GTC from either the APU or an exter-
nal AC source. Its function is similar to that
of a GLC.
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LEGEND
UNPOWERED
IDG
1
IDG
1
GCUGCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
GLC
2
GLC
1
GTC
2
GTC
1
AP/EP
CONT
MAIN AC BUS 1 MAIN AC BUS 2
Figure 2-3. Main AC Power (Power Off)
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LEGEND
APU GENERATOR POWER
UNPOWERED
POWERED BUS
IDG
1
IDG
1
GCUGCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
GLC
2
GLC
1
GTC
2
GTC
1
AP/EP
CONT
MAIN AC BUS 1 MAIN AC BUS 2
LEGEND
IDG 1 POWER
UNPOWERED
POWERED BUS
IDG 2 POWER
IDG
1
IDG
1
GCUGCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
GLC
2
GLC
1
GTC
2
GTC
1
AP/EP
CONT
MAIN AC BUS 1 MAIN AC BUS 2
Figure 2-4. Main AC Power APU Feed
Figure 2-5. Main AC Power Normal GEN Feed
EXTERNAL AC SYSTEM
AC external power is supplied through a stan-
dard six-pin AC power receptacle located on
the right-hand nose section of the fuselage
(Figure 2-7). An external power monitor en-
sures that the external AC power is safe and
within specified limits before power is ap-
plied to the airplane buses (Figure 2-8). If the
parameters of phase rotation, frequency (400
± 25 Hz), and voltage (115 ± 9 VAC) are sat-
isfactory, the monitor causes a green AVAIL
light to illuminate on the electrical power sec-
tion of the overhead panel. Provided no other
airplane generator is on line, movement of the
G PWR switch to the ON position extinguishes
the AVAIL light and illuminates the amber IN
USE light. This indicates that external AC
power is now connected to the airplane buses.
If the APU generator or any main generator is
brought on line, external power automatically
reverts to an available status.
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LEGEND
IDG 1 POWER
UNPOWERED
POWERED BUS
IDG
1
IDG
1
GCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
EXT
AC
EXT
AC
MONITORMONITOR
GLC
2
GLC
1
GTC
2
GTC
1
AP/EP
CONT
MAIN AC BUS 1 MAIN AC BUS 2
Figure 2-6. Main AC Power GEN 1 Feed (GEN 2 INOP)
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Figure 2-7. AC External Power Connector
LEGEND
EXTERNAL AC POWER
UNPOWERED
POWERED BUS
IDG
1
IDG
1
GCUGCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
EXT
AC
EXT
AC
MONITORMONITOR
GLC
2
GLC
1
GTC
2
GTC
1
AP/EP
CONT
MAIN AC BUS 1 MAIN AC BUS 2
Figure 2-8. External AC Power Feed (All GENs INOP)
EMERGENCY AC SYSTEM
The emergency AC power system provides
115-/200-volt, 3-phase, 400-Hz AC power au-
tomatically in flight in the event of loss of all
primary and auxiliary AC power. The air-
driven generator (ADG) is stowed in a com-
partment on the right side of the nose (Figure
2-10) and can be deployed either automatically
or manually. Once deployed, the ADG cannot
be retracted until the airplane is on the ground.
When deployed, the ADG provides power di-
rectly to the ADG bus. ADG power is rated at
15 kva and is controlled through a GCU also
rated at 15 kva.
The ADG consists of a ram-air turbine (RAT)
and an AC electrical generator mounted on a
trunnion-pivoted support leg. The ram-air tur-
bine is a two-bladed 19-inch propeller (Figure
2-11) with a variable-pitch mechanism that is
used to achieve the constant speed necessary to
operate the generator at a constant frequency.
The AC generator is similar in construction
and operation to the other three generators but
is rated at 15 kva. It is air cooled and therefore
subject to lower temperature restrictions.
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LEGEND
ADG POWER
UNPOWERED
POWERED BUS
IDG
1
IDG
1
GCUGCU
IDG
2
IDG
2
GCUGCU
APU
GEN
APU
GEN
GCUGCU
ADGADG
GCUGCU
EXT
AC
EXT
AC
MONITORMONITOR
GLC
2
GLC
1
A23 B23
C5EMERG
AC TC
ESS
PWR
TC
GTC
2
GTC
1
AP/EP
CONT
ADG BUS
3B
ESS AC BUSMAIN AC BUS 1 MAIN AC BUS 2
Figure 2-9. Emergency AC Power Feed (All GENs INOP)
Figure 2-10. Air-Driven Generator Location
The ADG control unit is responsible for mon-
itoring and regulating generator output and
for bringing the ADG bus on line. The GCU
provides an output as the generator frequency
passes 380 Hz. Thereafter, the output is avail-
able as long as the frequency remains in the
250- to 450-Hz range. A frequency outside
these limits will cause the ADG to be discon-
nected from its bus until the frequency again
passes 380 Hz increasing, or 430 Hz decreas-
ing. There is also an overvoltage trip at 130
volts with an automatic reset when the volt-
age decreases below 120 volts.
The GCU provides automatic connection of the
ADG bus to hydraulic pump 3B and to the AC
essential bus. An additional output to the elec-
trical control panel allows monitoring of ADG
voltage and frequency.An ADG loadmeter is
not provided.
AC POWER DISTRIBUTION
Refer to Figure 2-12 while reading this sec-
tion. There are a total of eight AC buses.
1. AC bus 1
2. AC bus 2
3. AC essential bus
4. ADG bus
5. 26-volt AC bus 2
6. 26-volt AC essential bus
7. AC utility bus 1
8. AC utility bus 2
AC Bus 1 and AC Bus 2
Normal AC electrical power is distributed
from the two main AC buses-AC bus 1 and AC
bus 2. In flight, each main bus is normally
supplied by its own IDG. Generator No. 1 sup-
plies AC bus 1 and generator No. 2 supplies
AC bus 2. Either bus may be powered by the
APU-driven generator if its IDG is not on line.
Both buses may be powered by a single gen-
erator if only one generator is on line. While
on the ground, if no generators are on line, both
buses may be fed by external AC power. A
parallel system feed is not available; therefore,
no AC bus may be fed from more than one
source at any one time.
The supply of power to the two main buses is
automatically controlled by the generator con-
trol units (GCU) through generator line con-
tactors (GLC), generator transfer contactors
(GTC), and the auxiliary power/external power
contactor (AP/EP C).
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Figure 2-11. Ram-Air Turbine
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AC Essential Bus
The AC essential bus is normally supplied
from AC bus 1. Should the voltage and/or fre-
quency of the power from AC bus 1 exceed spe-
cific tolerances, the essential power transfer
contactor switches over to allow power from
AC bus 2 to feed the AC essential bus. This ac-
tion causes a green ALTN light to illuminate
on the essential power panel. Depressing the
ALTN switchlight also causes the essential
power transfer contactor to move over and
supply the AC essential bus from AC bus 2. If
the voltage on the AC essential bus falls below
90 volts, a FAIL light on the AC essential
power pane l i l lumina tes (F igure 2-13) .
Whenever the ADG is operating, the ADG AC
emergency transfer contactor supplies power
to the AC essential bus.
ADG Bus
Whenever both main AC buses are unpowered
in flight and both IDGs are off line, the ADG
is deployed automatically. Within four seconds
of deployment, the ADG starts producing 115-
/200-volt, 3-phase, 400-Hz power. The power
is automatically routed to the AC essential
bus, hydraulic pump 3B, and the voltage and
frequency meters of the electrical control
panel. If automatic deployment fails, the ADG
may be deployed manually (Figure 2-14). the
ADG auto deploy control unit contains a test
function and an override pushbutton. If one or
more of the main AC buses have been recov-
ered, the pilot may press the POWER TXFR
OVERRIDE button to transfer the AC essen-
tial bus and hydraulic pump 3B back to their
normal power supplies. The manual deploy
assembly must be in the stowed position be-
fore the override button is pressed.
IDG
1
APU
GEN
EXT
AC ADG
MAIN AC BUS 1 ESS AC BUS
ESS 26 VAC
AC UTIL BUS 1 AC UTIL BUS 2
GCU
GLC 
1
GTC 
1
GTC 
2
GLC 
2
AP/EP
CONT
A 23 B 23
ESS
PWR
TC
C 5 
EMERG
AC TC
GCU MONITOR GCU
IDG
2
GCU
26 VAC 2
ADG BUS
3B
UNPOWERED
LEGEND
MAIN AC BUS 2
Figure 2-12. Total AC Power
26-VAC Bus 2 and 26-VAC
Essential Bus
There are two transformers which convert AC
bus 2 and AC essential bus 115 VAC to 26
VAC and power the airplane flight instru-
ments. The 26-VAC bus 2 supplies the copi-
lot's instruments and the 26-VAC essential
bus supplies the pilot's instruments.
AC Utility Buses
There are two AC utility buses provided. One
is powered from AC bus 1; the other is pow-
ered from AC bus 2. The utility buses supply
power for customer-installed equipment under
two conditions:
1 With airplane weight on wheels and
flaps selected to 0, and any single
source of AC power available
2. With airplane airborne and any two
generators on line
Automatic load shedding of both utility buses
occurs in flight anytime only one generator is
on line.
ADG Control Unit
Checks the following two functional tests.
Lamp Test
When selected, will indicate that the BATT DI-
RECT BUS is powered.
Unit Check
• Cont inu i ty o f the UPLOCK Squib
Circuit
• Continuity of three transfer contactors
• Logic circuits
• Weight on wheel circuits
• Any 2 main generators on-line
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 2-11FOR TRAINING PURPOSES ONLY
FAIL ALTN
PRESS TO
TRANSFER
ESSENTIAL
AC POWER
Figure 2-13. Essential AC Power Panel
Figure 2-14. ADG Controls
Indication of a Good Test
Ground Mode: 1 sec blank then green light on
for 2 secs then out.
Airborne Mode: 5 sec blank then green light
on for 2 sec then out.
CONTROL AND MONITORING
Cockpit control and monitoring of the elec-
trical system is achieved by use of the elec-
trical control panel (Figure 2-15). This panel,
just forward of the throttle levers, contains
switches to test and bring on line the main gen-
erators and meters for monitoring frequency,
voltage, and electrical load in various parts of
the system.
Generator Loadmeter
Three generator loadmeters allow for moni-
toring of the output of generator No. 1, gen-
erator No. 2 and the APU generator The range
of readings is from 0 to 50 kva.
Generator OFF Light
Illumination of a GEN OFF light indicates
that the generator is not switched on, has no
output, is out of frequency tolerance, or is out
of voltage tolerance.
OVLD Light
The overload (OVLD) light illuminates when
the generator load exceeds 34.5 kva.
Generator Control Switch
Each generator is controlled by its own gen-
erator control switch which has three posi-
tions: ON, OFF/RESET, and TEST. When in
the center OFF/RESET position, the switch in-
hibits the generator control unit, preventing au-
tomatic startup of the generator. This position
also resets the generator fault relay, enabling
the generator to be switched on again follow-
ing a fault trip. The TEST position, in con-
junction with the voltage and frequency
meters, permits a generator to be monitored be-
fore it is brought on line. In the ON position,
the switch activates the generator on line or to
allow for automatic startup.
2-12 CL 601-3A/R FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
Figure 2-15. Electrical Control Panel
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 2-13FOR TRAINING PURPOSES ONLY
MAIN BUS OFF Light
The MAIN BUS OFF light illuminates when
both sides of a generator line contactor are
deenergized. It is an indication that there is no
power on a main AC bus.
AUTO OFF/FAIL Switchlight
During manual operation, when pressed in,
the switchlight latches and AUTO OFF illu-
minates, indicating that the bus can only be fed
from its own generator.
During automatic operation, the FAIL light il-
luminates when a bus fault is detected and
isolated by the system. Both FAIL lights il-
luminate if the APU generator is feeding the
faulty bus.
Voltage and Frequency
Metering
A rotary selector allows for the monitoring of
AC voltage and frequency on separate meters.
The rotary selector positions are labeled "EXT
PWR," "GEN 1," "APU," "GEN 2," and "ADG."
DC SYSTEM
DC supply can be subdivided into two systems,
static conversion and battery. In normal op-
eration, primary DC power is derived from
the AC system through the use of four trans-
former-rectifier units (TRU) located in the
forward unpressurized avionics bay (Figure 2-
16). The battery system provides power to
MAIN BATTERY
 CHARGER
EXTERNAL DC
 CONNECTION
AUXILIARY BATTERY
MAIN BATTERY
AFT ELECTRICAL
DISTRIBUTION BOX (JB4)
NO. 1 MAIN DC
BUS (CBP-A)
LEFT BATTERY
BUS (CBP-A)
RIGHT BATTERY
 BUS (CBP-B)
DC ESSENTIAL
 BUS (CBP-D)
FROM AC
BUSES
TRANSFORMER
RECTIFIER UNITS
LEGEND
BATTERY AND EXTERNAL POWER
AC POWER
DC STATIC CONVERSION
NO. 2 MAIN DC
 BUS (CBP-B)
Figure 2-16. DC System ComponentsLocations
2-14 CL 601-3A/R FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
start the APU and to supply backup DC power
when a primary DC power source has failed
or is not available. It also supplements the
available DC power supply when the ADG is
the only source of airplane power.
DC STATIC CONVERSION
In normal operation, primary DC power is ob-
tained from the AC system through the use of
four separate and independent transformer-
rectifier units (TRU). Each TRU converts 3-
phase, 115-/200-VAC power into unregulated
28-volt DC. TRU 1, rated at 100 amps, is pow-
ered by AC bus 1 and feeds DC bus 1 (Figure
2-17). TRU 2, rated at 100 amps, is powered
by AC bus 2 and feeds DC bus 2. Essential TRU
1 is powered from the AC essential bus and es-
sential TRU 2 is powered by AC bus 2. Both
essential TRUs are limited at 30 amps and
feed the DC essential and the battery bus. Loss
of DC power on any bus is detected by bus
sensing relays which cause illumination of
the appropriate BUS OFF light on the electrical
control panel (Figure 2-15). If essential TRU
1 or 2 fails, the respective ESS TRUS 1 OFF
or 2 OFF light illuminates ( Figure 2-15).
TRUs 1 and 2 normally supply their own as-
sociated buses. If a TRU failure occurs, the
power distribution to DC buses 1 and 2, except
the utility buses (Figure 2-17), can be main-
tained by energizing DC bus ties which cross-
connect the outputs of TRU 1 and TRU 2.
These bus ties are controlled by switchlights
on the electrical control panel labeled "BUS
TIE CLOSED." Both BUS TIE CLOSED
switchlights would have to be illuminated to
replace either TRU 1 or TRU 2 (Figure 2-15).
When a DC tie relay is closed, the respective
utility bus is load shed.
Essential TRUs 1 and 2 are dedicated to the DC
essential bus and cannot be used to supply ei-
ther the No. 1 DC bus or No. 2 DC bus. If es-
sential TRU 1 or 2 falls, the DC essential bus
can be supplied from the operating essential
TRU (Figure 2-17), or in an emergency con-
dition (all AC power lost), by deploying the
ADG. During ADG operation, the battery bus
and the DC essential bus are connected.
During transfer to ADG, when the aircraft is in
a weight-off-wheels condition and the battery bus
is unpowered, the DC essential bus is supplied
from the battery direct bus via the DC emer-
gency tie control 2 contactor (Figure 2-24).
MAIN AC BUS 1
DC BUS 1
ESS DC BUS
DC UTILITY BUS 1 DC UTILITY BUS 2
DC BUS 2
MAIN AC BUS 2ESS AC BUS
TRU
1
A17
C2
A69
A68
DC TIE
CONT
D10
EMER
DC 1 TC
D5
D4
D7
B69 B68
D8D9
B29 B17
DC TIE
CONT
ESS
TRU 2
ESS
TRU 1
TRU 
2
POWERED BUS
IDG 1 POWER
LEGEND
UNPOWERED BUS
IDG 2 POWER
TRU 1 POWER
TRU 2 POWER
ESSENTIAL TRU 1 POWER
ESSENTIAL TRU 2 POWER
BATT BUS
Figure 2-17. Primary DC Power Feed (Normal Operation)
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 2-15FOR TRAINING PURPOSES ONLY
MAIN AC BUS 1
DC BUS 1
DC UTILITY BUS 1 DC UTILITY BUS 2
DC BUS 2
MAIN AC BUS 2ESS AC BUS
TRU
1
A17
C2
A69
A68
DC TIE
CONT
D10
EMER
DC 1 TC
D5
D4
D7
B69 B68
D8D9
B29 B17
DC TIE
CONT
ESS
TRU 2
ESS
TRU 1
TRU 
2
UNPOWERED
POWERED BUS
IDG 1 POWER
LEGEND
IDG 2 POWER
TRU 1 POWER
TRU 2 POWER
ESSENTIAL TRU 1 POWER
ESSENTIAL TRU 2 POWER
ESS DC BUS
BATT BUS
MAIN AC BUS 1
DC BUS 1
ESS DC BUS
DC UTILITY BUS 1
DC BUS 2
MAIN AC BUS 2
TRU
1
A17
C2
A69
A68
DC TIE
CONT
D10
EMER
DC 1 TC
D5
D4
D7
B69 B68
D8D9
B29 B17
DC TIE
CONT
ESS
TRU 2
ESS
TRU 1
TRU 
2
UNPOWERED
POWERED BUS
IDG 1 POWER
LEGEND
IDG 2 POWER
TRU 1 POWER
TRU 2 POWER
ESSENTIAL TRU 1 POWER
ESSENTIAL TRU 2 POWER
DC UTILITY BUS 2
ESS AC BUS
BATT BUS
Figure 2-18. Primary DC Power Feed TRU No. 2 INOP
Figure 2-19. Primary DC Power Feed ESS TRU No. 1 INOP
2-16 CL 601-3A/R FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
BATTERY SYSTEM
The battery supplies the battery direct bus via
the external DC contactor. The battery is a
nickel-cadmium (nicad) with a nominal out-
put of 24 volts and a rating of 43 ampere-
hours . The ba t te ry i s loca ted in the a f t
equipment bay (Figure 2-16) and is kept
charged by a separate battery charger unit.
The battery charger is powered from AC bus
2 (Figure 2-20). There are two circuits: a
charging circuit and a monitoring circuit. The
charging circuit is capable of bringing the bat-
tery from a 30% charge to a 100% charge in
one hour. The battery-charging circuit is au-
tomatically inhibited if the monitor circuit
detects a cell imbalance, high battery tem-
perature, or a faulty battery temperature sen-
sor. The battery charger is installed in the aft
equipment bay, above and outboard of the bat-
tery. Illumination of the CHARGER light on
the overhead panel indicates that:
• Power is not available to the charger unit.
• There is a battery cell imbalance.
• There is a battery temperature abnormality.
• The battery contactor is closed and the
battery is being used.
BATTERY OVERLOAD
Should an overload occur when the battery is
connected to the battery bus, the battery re-
mote-control circuit breaker (RCCB) trips and
takes the battery off line. The RCCB trips at
50 amps to protect the battery from a short cir-
cuit on the battery contactor or the battery
bus (Figure 2-20). The RCCB is physically lo-
cated in the aft equipment bay. When it trips,
it also trips the BATT RCCB CONT circuit
breaker on the DC essential bus in the cock-
pit. Resetting this breaker provides a reset
function to the RCCB.
BATTERY CONTACTOR
When there is no AC power on the airplane and
essential TRUs 1 and 2 are unpowered, the
battery supplies power to the battery bus
through the battery contactor. When AC power
is available, essential TRUs 1 and 2 supply
the DC essential bus and the battery bus. In
order for the battery to supply power to the bat-
tery bus, the battery master switch must be in
the ON position and both essential TRUs must
be unpowered. In this case, the sense relay
which monitors the output from both essential
TRUs is deenergized and causes the battery
contactor to close. If the ADG is the only
source of power, the sense relay will be deen-
ergized causing the battery contactor and the
MAIN AC BUS 2
ESS DC BUS
BATT BUS BATT DIR BUS
BATT
CHGR
B32
MAIN
BATT
ESS
TRU 2
D9 D8
D7
D4
EMER
DC 2 TC
W OFF W
E8 
EXT
DC TCD6
BATTERY
MASTER
SENSOR
RELAY
BATTERY CONT
EMER
DC 1 TC
D10
D5
ESS
TRU 1
EXT DC
UNPOWERED
POWERED BUS
BATTERY POWER
LEGEND
Figure 2-20. Battery Master Switch OFF (All Power Off)
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 2-17FOR TRAINING PURPOSES ONLY
emergency DC2 transfer contactor (W off W
only) to close. This will allow the BATTERY
DIRECT BUS to provide backup power to
both the BATTERY BUS and the DC ES-
SENTIAL BUS.
EXTERNAL DC
External DC power can be connected through
the DC external contactor on the right rear
side of the airplane (Figure 2-21). External
power is used to power the battery direct bus
during ground servicing and for starting the
APU. The external DC power replaces the bat-
tery as a source of power. Whenever the ex-
ternal contactor is energized, the battery is
disconnected and external DC is connected
to the battery direct bus and to the APU start-
ing circuit. It also supplies a ground which
causes the IN USE light on the overhead panel
to illuminate. Figure 2-21. DC External Power Connector
MAIN AC BUS 2
ESS DC BUS
BATT DIR BUS
BATT
CHGR
B32
MAIN
BATT
ESS
TRU 2
D9 D8
D7
D4
EMER
DC 2 TC
W OFF W
E8
EXT
DC TC
D6
BATTERY
MASTER
SENSOR
RELAY
BATTERY
CONT
EMER
DC 1 TC
D10
D5
ESS
TRU 1
EXT
DC
UNPOWERED
POWERED BUS
BATTERY POWER
LEGEND
BATT BUS
Figure 2-22. Battery Master Switch ON (W on W)
2-18 CL 601-3A/R FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
MAIN AC BUS 2
ESS DC BUS
BATT DIR BUS
BATT
CHGR
B32
MAIN
BATT
ESS
TRU 2
D9 D8
D7
D4
EMERDC 2 TC
W OFF W
E8
EXT
DC TC
D6
BATTERY
MASTER
SENSOR
RELAY
BATTERY
CONT
EMER
DC 1 TC
D10
D5
ESS
TRU 1
EXT
DC
UNPOWERED
POWERED BUS
EXTERNAL DC POWER
LEGEND
BATT BUS
MAIN AC BUS 2
ESS DC BUS
BATT DIR BUS
BATT
CHGR
B32
MAIN
BATT
ESS
TRU 2
D9 D8
D7
D4
EMER
DC 2 TC
W OFF W
E8
EXT
DC TC
D6
BATTERY
MASTER
SENSOR
RELAY
BATTERY
CONT
EMER
DC 1 TC
D10
D5
ESS
TRU 1
EXT
DC
UNPOWERED
POWERED BUS
IDG 2 POWER
LEGEND
BATTERY POWER
BATTERY CHARGER BATT BUS
Figure 2-23. External DC Battery Master Switch ON (W on W)
Figure 2-24. Both ESS TRUs Failed (W off W)
CONTROL AND MONITORING
A rotary selector on the DC metering portion
of the electrical control panel (Figure 2-25) al-
lows for the monitoring of voltage and load
conditions. Bus voltage is monitored on DC
bus 1, DC bus 2, and the DC essential buses.
TRU load is displayed in amps. The fifth con-
tact on the rotary selector allows for monitor-
ing of battery condition and is also the point at
which external DC power voltage is read. The
voltmeter scale is between 15 and 35 volts. The
current load is displayed on a loadmeter cali-
brated from 0 to 100 amps. A BUS OFF light
illuminates on the electrical control panel when
any of the four DC buses is unpowered.
A BATTERY light on the overhead panel
(Figure 2-26) illuminates to indicate that bat-
tery power is not available for use by the bat-
tery bus, possibly due to one of the following
conditions:
• The BATTERY MASTER switch is in
the OFF position.
• The battery has failed.
Figure 2-27 shows the total electrical system
in simplified form and illustrates AC and DC
system interconnection.
Figures 2-28 through 2-34 show the circuit-
breaker panels and locations.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A/R 2-19FOR TRAINING PURPOSES ONLY
10
20
25
30
35 DC
VOLTS
0
20
40 60
80
100DC
AMPS
TRU LOAD
BUS VOLTS
ESS
1
ESS
2
1
2
BATT
VOLTS
1 OFF
2 OFF
BATT
BUS
OFF
ESS
BUS
OFF
MAIN
BUS 1
OFF
MAIN
BUS 2
OFF
BUS
TIE
CLOSED
DC METERING
DC POWER
BUS
TIE
CLOSED
TRU 1 TRU 2 ESS TRU BATT
Figure 2-26. BATTERY MASTER Switch
ENGINE START
RL
BATTERY
CHARGER
ON
OFF
IGNITION ELECT PWR
GPWR
BATTERY
MASTER
START
CONT
IGN
STOP
IN
FLIGHT
START
START
STOP
IN
FLIGHT
START ON
OFF
AVAIL
IN USE
IGN A
ON
IGN B
ON
Figure 2-25. DC Electrical Controls
2-20 CL 601-3A/R FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
LEGEND
IDG 1 POWER
APU POWER
BATTERY POWER
ADG POWER
IDG 2 POWER
EXTERNAL AC POWER
TRU 1 POWER
ESSENTIAL TRU 1 POWER
ESSENTIAL TRU 2 POWER
TRU 2 POWER
EXTERNAL DC POWER
BATTERY CHARGER
IDG
1
IDG
1
GCU
IDG
2
IDG
2
GCU
APU
GEN
GCU
ADG
GCU
EXT
AC
MONITOR
GLC
2
GLC
1
A23 B23
C5
C2
A47
A17
A69 B69A68 B68
E8
D6
RCCB
EXT
DC TC
EMER
DC 2 TC
W OFF W
EMER
DC 1 TC
DC TIE
CONT
DC TIE
CONT
BATTERY
MASTER
BATTERY
CONT
SENSOR
RELAY
D5
D4
EMERG
AC TC
ESS
PWR
TC
GTC
2
GTC
1
AP/EP
CONT
D9D10
ADG BUS
D8 D7
B29
B17
B32
B47
DC UTIL BUS 2DC UTIL BUS 1
ESS
TRU 1
TRU
1
TRU
2
BATT
CHGR
MAIN
BATT
EXT
DC
ESS
TRU 2
3B
ESS AC BUSMAIN AC BUS 1
ESS 26 VAC
AC UTIL BUS 1
26 VAC 2
AC UTIL BUS 2
MAIN AC BUS 2
DC BUS 2
ESS DC BUS
DC BUS 1
BATT BUS BATT DIR BUS
Figure 2-27. Total Electrical System
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY
CIRCUIT-BREAKER
PANEL LOCATIONS
Table 2-2 lists circuit-breaker panels by name
and describes their locations.
CL 601-3A 2-21
BUS NAME LOCATION
Battery Direct Bus Aft fuselage, close to battery
• AC Essential Bus Just forward of the pilot's side console
• 26-VAC Essential Bus
DC Essential Bus Just forward of the copilot's side console
• Main Bus 1 (AC and DC) Inboard on the bulkhead behind the pilot's seat
• Utility Bus 1 (AC and DC)
• Main Bus 2 (AC and DC)
• Utility Bus 2 (AC and DC) Inboard on the bulkhead behind the copilot's seat
• 26-VAC Bus 2
Battery Bus Two panels (one electrical bus), one outboard on
the bulkhead behind each pilot's seat
Table 2-2. CIRCUIT-BREAKER PANELS
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
2-22 CL 601-3A
Figure 2-28. CBP-A Main AC and DC Bus 1
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CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3A 2-23
Figure 2-29. CBP-A Battery Bus
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
2-24 CL 601-3A
Figure 2-30. CBP-B Main AC and DC Bus 2
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CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3A 2-25
Figure 2-31. CBP-B Battery Bus 
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
2-26 CL 601-3A
Figure 2-32. CBP-C AC Essential Bus
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3A 2-27
1
1
1
1
NOTE:
RED IDENTIFICATION COLLAR INSTALLED
(POST SB 601–0379)
Figure 2-33. CBP-D DC Essential Bus
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
2-28 CL 601-3A
Figure 2-34. CBP-E Battery Direct Bus
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY
1. The primary Challenger electrical sys-
tem is:
A. 115-volt AC
B. 30-volt DC
C. 24-volt DC
D. 30-volt AC
2. A constant generator speed is ensured by
the:
A. GCU
B. ADG
C. CSD
D. RAT
3. The number of AC generators is:
A. 1
B. 2
C. 3
D. 4
4. Emergency in-flight AC power is pro-
vided by:
A. IDG
B. ADG
C. GCU
D. CSD
5. The CSD oil quantity is checked by:
A. Looking at a pop-out indicator
B. Reading a gage
C. Viewing a sight glass
D. Monitoring a light.
6. While on the ground, AC power can be
supplied by the engines or by a(n):
A. APU generator
B. External DC connector
C. ADG
D. CSD
7. A source of emergency in-flight AC power
is:
A. IDG
B. CSD
C. ADG
D. TRU
8. If the ADG is inadvertently deployed, it
can be retracted:
A. By depressing the reset switch
B. With a manual crank
C. By holding the deploy button down for
3 seconds
D. Only on the ground
9. If the OVLD light illuminates:
A. Generator load has exceeded 34.5 kva.
B. Automatic operation of GTC's is in-
hibited.
C. Both sides of the GLC are deener-
gized.
D. Both A and B are correct.
10. Primary DC power is supplied by a(n):
A. Inverter
B. TRU
C. Converter
D. ADG
CL 601-3A 2-29
QUESTIONS
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY
CIRCUIT-BREAKER PANEL
LOCATIONS
Table 2-2 lists circuit-breaker panels by name
and describes their locations.
CL 601-3R 2-21
BUS NAME LOCATION
Battery Direct Bus Aft fuselage, close to battery
• AC Essential Bus Just forward of the pilot's side console
• 26-VAC Essential Bus
DC Essential Bus Just forward of the copilot's side console
• Main Bus 1 (AC and DC) Inboard on the bulkhead behind the pilot's seat
• Utility Bus 1 (AC and DC)
• Main Bus 2 (AC and DC)
• Utility Bus 2 (AC and DC) Inboard on the bulkhead behind the copilot's seat
• 26-VAC Bus 2
Battery Bus Two panels (one electrical bus), one outboard on
the bulkhead behind each pilot's seat
Table 2-2. CIRCUIT-BREAKER PANELS
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
2-22 CL 601-3R
Figure 2-28. CBP-A Main AC and DC Bus 1
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FOR TRAINING PURPOSES ONLY CL 601-3R 2-23
Figure 2-29. CBP-A Battery Bus
FOR TRAINING PURPOSES ONLY
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2-24 CL 601-3R
Figure 2-30. CBP-B Main AC and DC Bus 2
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FOR TRAINING PURPOSES ONLY CL 601-3R 2-25
Figure 2-31. CBP-B Battery Bus
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
2-26 CL 601-3R
Figure 2-32. CBP-CAC Essential Bus
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3R 2-27
1
1
1
1
NOTE:
RED IDENTIFICATION COLLAR INSTALLED
(POST SB 601–0379)
Figure 2-33. CBP-D DC Essential Bus
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
2-28 CL 601-3R
S5KJ
OIL FILTER
DS3KJ DS2KJ
LH RH
TEST
LH
TEST
RH
S3KJ S4KJ
CHIP DETECT
RESET
DS5KJ DS4KJ
LH
ENG
IGN B
ENG
OIL CONT
BYPASS
IND
ENG
OIL
POWER
BATT
SHUNT
FUSES
RH
B
A
T
B
U
S
B
A
T
T
E
R
Y
 D
IR
E
C
T
 B
U
S
B
A
T
T
E
R
Y
 D
IR
E
C
T
 B
U
S
20
3
10
50
5
3
3
3
3
3
5
5
7.5 7.5
75
3
E7
E6
E9
BOARD
BAT
CONT
EXT AC
PWR
CONT
ESS
PWR
CONT
AUTO
APR
CONT
FUEL
DE-
FUEL
APU
BACKUP
PWR
CONT
IND
APU
START
MAN
ADG
DEPLOY
CONT
SERV
LIGHTS
CBP–E
Figure 2-34. CBP-E Battery Direct Bus
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CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3R 2-29
QUESTIONS
1. The primary Challenger electrical sys-
tem is:
A. 115-volt AC
B. 30-volt DC
C. 24-volt DC
D. 30-volt AC
2. A constant generator speed is ensured by
the:
A. GCU
B. ADG
C. CSD
D. RAT
3. The number of AC generators is:
A. 1
B. 2
C. 3
D. 4
4. Emergency in-flight AC power is pro-
vided by:
A. IDG
B. ADG
C. GCU
D. CSD
5. The CSD oil quantity is checked by:
A. Looking at a pop-out indicator
B. Reading a gage
C. Viewing a sight glass
D. Monitoring a light.
6. While on the ground, AC power can be
supplied by the engines or by a(n):
A. APU generator
B. External DC connector
C. ADG
D. CSD
7. A source of emergency in-flight AC power
is:
A. IDG
B. CSD
C. ADG
D. TRU
8. If the ADG is inadvertently deployed, it
can be retracted:
A. By depressing the reset switch
B. With a manual crank
C. By holding the deploy button down for
3 seconds
D. Only on the ground
9. If the OVLD light illuminates:
A. Generator load has exceeded 34.5 kva.
B. Automatic operation of GTC's is in-
hibited.
C. Both sides of the GLC are deener-
gized.
D. Both A and B are correct.
10. Primary DC power is supplied by a(n):
A. Inverter
B. TRU
C. Converter
D. ADG
CL 601-3A/R 3-i
CHAPTER 3
LIGHTING
CONTENTS
Page
INTRODUCTION ................................................................................................................... 3-1
GENERAL............................................................................................................................... 3-1
INTERIOR LIGHTS .............................................................................................................. 3-2
Cockpit Lights.................................................................................................................. 3-2
Passenger Compartment Lights ....................................................................................... 3-4
Service Lights .................................................................................................................. 3-4
EXTERIOR LIGHTS .............................................................................................................. 3-4
Landing and Taxi-Recognition Lights ............................................................................. 3-5
Wing Ice Inspection Lights .............................................................................................. 3-6
Navigation Lights............................................................................................................. 3-6
Anticollision Strobe Lights and Beacon Lights ............................................................... 3-6
EMERGENCY LIGHTS ......................................................................................................... 3-6
QUESTIONS ........................................................................................................................... 3-7
FOR TRAINING PURPOSES ONLY
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CL 601-3A/R 3-iii
ILLUSTRATIONS
Figure Title Page
3-1 Interior Lighting Controls ........................................................................................ 3-2
3-2 Boarding Lights Control........................................................................................... 3-4
3-3 Ordinance Lights Controls ....................................................................................... 3-4
3-4 Exterior Lighting Locations ..................................................................................... 3-5
3-5 Landing and Taxi-Recognition Lights Controls ....................................................... 3-5
3-6 External Lighting Control Panel............................................................................... 3-6
3-7 Emergency Lighting Control .................................................................................... 3-6
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
INTRODUCTION
The Canadair Challenger CL-600-2B16, model CL-601-3A/R lighting system provides
exterior and interior illumination. Interior lighting groups include cockpit lights, pas-
senger compartment lights, and service lights. Exterior lighting includes the standard
required package of airplane lights.
An emergency lighting system is provided to automatically illuminate routes used for
emergency evacuation.
Most lighting controls are located in the cockpit; however, some lights have control
switches either integral or located near the individual lighting devices.
GENERAL
Included in the interior lighting groups are
boarding and dome lights to illuminate the
passenger door area, service compartment
lights, and cockpit lighting.
Exterior lighting consists of one recogni-
tion/taxi light and one landing light in the
leading edge of each wing and two landing
lights in the radome for approach and ground
visibility, standard red, green and white nav-
igation lights, anticollision strobe and bea-
con lights to provide airborne identification,
and wing ice inspection lights.
In addition, the airplane is equipped with an
emergency lighting system which illuminates
the right wing and passenger door area for
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CHAPTER 3
LIGHTING
CL 601-3A/R 3-1FOR TRAINING PURPOSES ONLY
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emergency evacuation. This self-contained,
battery-powered system is automatically ac-
tivated by loss of airplane power.
Wiring, logic, and power are provided to allow
installation of ordinance signs.
INTERIOR LIGHTS
COCKPIT LIGHTS
The cockpit lights consist of incandescent
lighting and fluorescent and integral panel
lighting. Figure 3-1 illustrates interior light-
ing controls and their locations.
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
MISC LTS
BOARDING DOME
STBY
COMP SERVICE
DIM
OFF
ON
OFF
ON
LANDING LTS
EXTERNAL LTS
LEFT NOSE RIGHT
OFF BRT
RECOG/
TAXI LT
NAV WING ANTI COLL
OVERHD 
PANEL
WARN
LTS
FLOOD
LTS
TEST
OFF
BRIGHT
DIM
WARN
LTS
FLOOD
LTS
TEST
OFF
BRIGHT
DIM
DIGITS FLOODS
LIGHTING
INST
OFF BRT OFF BRT OFF BRT
BRT
OFF
DIM
OFF BRT OFF BRT
BRT
OFF
DIM
ON
OFF
FLOOD FLOOR
LIGHTING
INST
Figure 3-1. Interior Lighting Controls
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Incandescent Lighting
Incandescent lighting consists of floor lights
and map reading lights. Floor lights are sup-
plied from 28-VDC bus 1. Control switches are
located on the pilot’s and copilot’s side light-
ing panels. Illumination is provided for each
pilot in the area of the rudder pedals.
Map reading lights are located on either side
of the overhead panel. The lights are individ-
ually adjustable for direction and intensity.
The intensity of each light is controlled with
individual rheostats located on the overhead
lighting panel. Both map lights arepowered
from the battery bus with the R map light hav-
ing DC Bus 2 as a redundant power supply.
Panel Lighting
Panel lighting is of two types: Fluorescent
and integral.
Fluorescent flood lighting is provided along
the glareshield to illuminate the center instru-
ment panel, pilot and copilot instrument panels,
side panels, fascia panels, and side consoles.
Each instrument panel and side console has
two fluorescent strip lights, while side pan-
els have only a single strip each. One of the
light strips for each instrument panel and both
lights for each side console are controlled
with a BRT-OFF-DIM switch. The other light
for each instrument panel is controlled with
a potentiometer.
The controls for the floodlights are located on
the center pedestal. Side console floodlight con-
trols are located on the associated fascia panel.
Power for the pilot’s floodlights is supplied
from the DC essential bus. The copilot’s flood-
lights are powered from the battery bus.
Integral lighting includes edge-lit panel light-
ing and instrument lighting.
The integral lighting in the cockpit is con-
trolled in two individual areas:
• Overhead panel
• Center pedestal
Integral lighting is supplied with up to 5 VAC
by incandescent lamp dimmers controlled
from the various lighting panels.
Overhead Panel
Under normal conditions, the overhead panel
lighting is supplied with 5 VAC stepped down
from AC bus 2. In the event that AC bus 2 is
not powered, such as before the APU is
started, the overhead panel receives power
from the battery bus through a DC incan-
descent lamp dimmer.
Normal intensity of the overhead panel light-
ing is controlled with a potentiometer located
on the cabin and miscellaneous lighting panel.
Left and right battery bus circuit-breaker panel
lighting is also controlled with the overhead
panel supply.
Center Instrument and Center
Pedestal Panels
The center instrument panel, center pedestal,
and main buses 1 and 2 circuit-breaker panels
are supplied with 5 VAC stepped down from
the AC essential bus. Intensity of this lighting
is controlled with a potentiometer located on
the center pedestal lighting control panel.
Pilot and Copilot Panels
The pilot instrument panel, side panel, fascia,
side console, and AC essential bus circuit-
breaker panel use 5 VAC stepped down from the
AC essential bus. The copilot panels, includ-
ing the DC essential bus circuit-breaker panel,
use 5 VAC stepped down from AC bus 2.
The intensity of pilot and copilot lighting is con-
trolled with a potentiometer on the associated
side panel of the center pedestal panel.
PASSENGER COMPARTMENT
LIGHTS
Passenger compartment lights consist of dome
lights, boarding lights, and ordinance lights.
Dome and boarding lights are contained in a
single assembly above the main entrance door.
The dome light is powered by the battery bus
and is controlled with a switch on the overhead
lighting panel. The boarding lights are pow-
ered by the battery direct bus and controlled
from either the overhead lighting panel or a
switch located just aft of the main entrance
door. (See Figure 3-2.)
Ordinance light wiring is provided so that a
completion center can install appropriate signs
controllable from the cockpit. There are sep-
arate NO SMKG and SEAT BLTS switches
on the center pedestal (Figure 3-3). All ordi-
nance light power except relay control is sup-
plied from AC bus 1. Relay control is powered
from DC bus 1. Both control switches have
three positions and are labeled “ON-OFF-
AUTO.” When AUTO is selected, both signs
illuminate whenever (1) cabin altitude is above
10,400 feet, (2) the landing gear is selected
down, or (3) the test button on the LDG GEAR
control panel is pressed.
The FASTEN SEAT BELT sign also illumi-
nates if the flaps are not at 0°.
SERVICE LIGHTS
Service lights are located in the nosewheel bay,
the avionics bay, and the rear equipment bay.
All service lights are powered by the battery
direct bus. The nosewheel bay service light is
controlled from the overhead lighting panel
(Figure 3-1), while service lights in the other
areas are controlled with switches at their re-
spective locations.
EXTERIOR LIGHTS
The airplane is equipped with landing lights,
taxi-recognition lights, wing ice inspection
lights, navigation lights, anticollision strobe and
beacon lights, and emergency lights. Figure 3-
4 illustrates exterior lighting locations.
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VESTIBULE
LIGHTS
BOARDING
LIGHTS
AISLE
LIGHTS
Figure 3-2. Boarding Lights Control
CABIN SIGNS
SEAT BLTSNO SMKG
OFF
AUTO
ON
Figure 3-3. Ordinance Lights Control
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LANDING AND TAXI-
RECOGNITION LIGHTS
Landing and taxi - recogni t ion l ights are
mounted in each wing root’s lower leading
edge. Two addit ional landing l ights are
mounted in the radome and are controlled with
the NOSE LANDING LT switch on the over-
head panel. The lights are controlled with the
L and R LANDING LT, and RECOG/TAXI
LT switches on the overhead l ight panel
(Figure 3-5). Positioning a landing light switch
to ON causes both the landing and the taxi-
recognition lights on the side selected to il-
luminate. Positioning the RECOG/TAXI LT
switch to ON causes both taxi lights to illu-
minate. The left lights are powered by AC bus
1 and the right lights by AC bus 2.
EMERGENCY
LIGHT
NAVIGATION
AND STROBE LIGHTS
ROTATING
BEACON (OPTIONAL)
NAVIGATION
LIGHT
TAXI-RECOGNITION
AND LANDING
LIGHTS
ROTATING
BEACON (OPTIONAL)
TWO
LANDING
LIGHTS
WING
INSPECTION
LIGHT
Figure 3-4. Exterior Lighting Locations
Figure 3-5. Landing and Taxi-Recognition
Lights Control
OFF
ON
OFF
ON
OFF
ON
OFF
ON
LANDING LTS
LEFT NOSE RIGHT
RECOG/
TAXI LT
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WING ICE INSPECTION LIGHTS
Wing ice inspection lights are mounted in both
sides of the fuselage just above the wing. They
are controlled with the WING switch located
on the overhead external lighting panel (Figure
3-6). Power is supplied by DC bus 1.
NAVIGATION LIGHTS
Dual unit navigation lights are mounted in the
left and right wingtips, tail cone, and vertical
stabilizer bullet. The lights are powered by AC
bus 1 via a step-down transformer and are
controlled with the NAV switch on the over-
head lighting panel.
ANTICOLLISION STROBE
LIGHTS AND BEACON LIGHTS
Anticollision strobe lights are integral with each
navigation light except in the vertical stabilizer
bullet. The wingtip lights are powered by DC bus
2 and the tail cone light by DC bus 1. There are
also two red beacon lights (not on green
airplanes), one on the top of the vertical stabilizer
and another on the lower fuselage. They are
powered by DC bus 1 and DC bus 2. Control is
with the ANTI-COLLISION light switch located
on the overhead lighting panel (Figure 3-6). The
switch has three positions which are
BCN/STROBE, OFF, and BEACON. In the
BEACON position only the rotating beacons are
powered. The BCN/STROBE position routes
power to the strobe lights and the beacons.
EMERGENCY LIGHTS
Four emergency lights are provided to illu-
minate escape routes during emergency evac-
uation. There are three lights in the fuselage
above the right wing and one light on the left
side just forward of the main entrance door.
Two battery packs independently supply
power. Emergency lighting is controlled from
the emergency lighting panel on the copilot’s
side console (Figure 3-7). The switch labeled
“EMERGENCY LIGHTING” has three posi-
tions: ON, OFF, and ARM.
In the ARM position , the system charges the
battery packs and automatically illuminates the
emergency lights if DC essential power (charg-
ing supply) fails.
When positioned to ON, the battery packs do
not charge and the lights illuminate, along
with a white light labeled “EMER LTS ON”
located next to the switch.
Positioningthe switch to OFF causes the emer-
gency lighting system to deenergize and the
amber EMER LTS OFF light to illuminate.
This light warns the crew that the emergency
lights will not be activated automatically.
To prevent automatic illumination on shut-
down, the EMERGENCY LIGHTING switch
must be turned off while AC power is still
available to power the DC essential bus. (AC
power is converted by essential TRU 1 and 2
to feed the DC essential bus.)
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF
ON
OFF BRT
MISC LTS
BOARDING DOME
STBY
COMP SERVICE
OVERHD
PANEL
DIM
OFF
ON
LANDING LTS
EXTERNAL LTS
LEFT NOSE RIGHT
RECOG/
TAXI LT
NAV WING ANTI COLL
Figure 3-6. External Lighting
Control Panel
Figure 3-7. Emergency Lighting Control
EMER
LTS
ON
EMER
LTS
OFF
OFF
ARM
ON
EMERGENCY LIGHTING
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1. The control switches for the floor lights
are located on the:
A. Pilot’s fascia
B. Overhead panel
C. Center pedestal
D. Pilot and copilot side panels
2. When the floodlight switch on the center
pedestal is moved to BRT, only one cen-
ter instrument panel fluorescent strip light
illuminates. The first corrective action
should be:
A. Check the battery bus voltage.
B. Reposition the switch to OFF.
C. Rotate the center pedestal poten-
tiometer clockwise.
D. Reposition the switch to DIM.
3. The control potentiometer for overhead
panel lighting is located on the:
A. Cabin and miscellaneous lighting
panel
B. Center pedestal
C. PiIot’s side panel
D. Copilot’s fascia
4. The boarding light is controlled:
A. With a switch on the door
B. With a switch aft of the door
C. With a switch on the overhead panel
D. Both B and C
5. When the ordinance sign switches are in
AUTO, the seat belt and no smoking
signs illuminate:
A. When cabin altitude is above 10,400
feet
B. When the landing gear is selected
down
C. When LDG GEAR test is initiated
D. All the above
6. The nosewheel bay service light control
is located:
A. On the overhead panel
B. Near the entrance door
C. In the avionics compartment
D. In the nosewheel bay
7. If the left landing light switch is on and
the taxi light switch is off, the light(s) that
illuminate are:
A. Both landing lights
B. The left landing light
C. The left taxi light
D. Both B and C
8. If DC bus 2 is unpowered, the anticol-
lision strobe lights that are inoperative
are the:
A. Left wingtip lights
B. Both wingtip lights
C. Tail cone light
D. Both B and C
9. The emergency lights switch position
tha t a l lows the ba t t e ry packs to be
recharged is:
A. ON
B. OFF
C. ARM
D. AUTO
10. If AC power is turned off while the emer-
gency lights switch is in ARM:
A. The emergency lights illuminate.
B. The EMER LTS OFF annunciator
flashes.
C. The switch automatically repositions
to OFF.
D. A horn sounds.
QUESTIONS
CL 601-3A/R 4-i
CHAPTER 4
MASTER WARNING SYSTEM
CONTENTS
Page
INTRODUCTION ................................................................................................................... 4-1
GENERAL............................................................................................................................... 4-1
MASTER CAUTION SYSTEM.............................................................................................. 4-2
General ............................................................................................................................. 4-2
Components ..................................................................................................................... 4-4
Lamp Test Switches ......................................................................................................... 4-4
Illumination Causes.......................................................................................................... 4-5
AURAL WARNING SYSTEM............................................................................................. 4-13
General........................................................................................................................... 4-13
Components ................................................................................................................... 4-13
QUESTIONS......................................................................................................................... 4-16
FOR TRAINING PURPOSES ONLY
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CL 601-3A/R 4-iii
ILLUSTRATIONS
Figure Title Page
4-1 Master Caution System ............................................................................................ 4-3
4-2 Lamp Test Switches.................................................................................................. 4-4
4-3 System Warning Annunciators............................................................................... 4-10
4-4 System Status Annunciators................................................................................... 4-12
4-5 Landing Gear MUTE HORN Button ..................................................................... 4-13
4-6 Aural Warning System ........................................................................................... 4-15
TABLES
Table Title Page
4-1 Caution Annunciators............................................................................................... 4-5
4-2 Miscellaneous Annunciators .................................................................................. 4-10
4-3 Warning Annunciators............................................................................................ 4-11
4-4 Nondimmable Annunciators .................................................................................. 4-12
4-5 Aural Warnings ...................................................................................................... 4-14
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
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INTRODUCTION
The Canadair Challenger CL-600-2B16, model CL-601-3A/R master warning system pro-
vides the crew with information on malfunctions of airplane equipment, unsafe operat-
ing conditions which require immediate attention, or the operation of a particular system
that is not normally used. A system of aural tones is used to draw attention to certain
significant situations that may have safety implications.
GENERAL
The airplane is designed to operate on the
“black cockpit” concept. Each illuminated
light indicates a system or situation status.
Red lights provide warnings of faults which
require immediate attention. Amber lights de-
note cautionary items of a less urgent nature.
Green and white lights indicate conditions
which are normal when in use.
The two basic central warning systems are (1)
the master caution system and (2) the aural
warning system. The master caution system
provides the crew with up to 18 visual annun-
ciators for airplane system malfunctions. Ad-
ditional information on the specific malfunction
may then be found on the associated system con-
trol panel. The master caution annunciator pan-
els are located just beneath the overhead panel.
When any annunciator panel light illuminates,
there is a brief delay before the two master
caution lights on the glareshield start to flash.
The delay compensates for transient warning
illuminations. The master caution lights can
be extinguished and reset by depressing either
TEST
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CHAPTER 4
MASTER WARNING SYSTEM
CL 601-3A/R 4-1FOR TRAINING PURPOSES ONLY
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4-2 CL 601-3A/R FOR TRAINING PURPOSES ONLY
of the two amber switchlights labeled “MAS-
TER CAUTION PRESS TO RESET.” The lights
in the annunciator panels are also extinguished
when the MASTER CAUTION switchlights
are reset. They may be recalled (if the system
fault is not corrected) by selecting the RE-
CALL positionon the 10 channel annunciator
panel test-recall switch.
The aural warning system provides a distinct au-
dible tone for each of eight significant events.
Provision is made to test the master caution
lights, all system annunciator lights, and the
audible tones with a series of switches and a
control panel.
All annunciator lights, system malfunction
lights, their colors, and their locations are shown
in the Annunciator Section of this manual.
MASTER CAUTION 
SYSTEM
GENERAL
Most airplane system malfunctions that are not
of an urgent nature are indicated by the illu-
mination of an amber light on the individual
system’s control panel which, in turn, illumi-
nates one of 18 annunciator panel lights and
the two flashing MASTER CAUTION lights
(See Annunciator Section).
The MASTER CAUTION lights are switch-
lights designed to draw attention to the sys-
tem annunciators. They also function as a
means of canceling or resetting the master
caution system.
The system provides power to the various cau-
tion, advisory, and warning lights located
throughout the cockpit.
An annunciator light, in conjunction with the
flashing MASTER CAUTION lights, illumi-
nates to indicate a malfunction or improper con-
dition in each of the following airplane systems:
• Anti-icing
• Auxiliary Power Unit
• Environmental Control
• Hydraulics
• Electronic Circuit-Breaker Channel Fail-
ure
• Antiskid
• Main Landing Gear Bay Overheat Fail-
ure
• Navigation
• Doors
• Electrical
• Engine
• Flight Controls
• Fuel
• Fire Fault
• Weight On Wheels
• Emergency Lights
There are two spare annunciator lights.
These 18 annunciator lights are arranged in two
panels: a 10 channel annunciator panel and an
8 channel annunciator panel, both located just
below the overhead panel where they can be
easily viewed by both crewmembers.
When a malfunction or improper condition
occurs in any of the systems being monitored,
an individual amber caution light illuminates
on the control panel for that system. In addi-
tion, the appropriate annunciator light illu-
minates and, after a brief delay, the MASTER
CAUTION lights, located on the glareshield
directly in front of each pilot, start flashing.
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A flashing MASTER CAUTION light directs
the pilots’ attention to the 10 and 8 channel an-
nunciator panels to determine which system
is at fault, and then to the specific system
panel to determine what the actual fault is.
Once the fault has been identified, the MAS-
TER CAUTION lights and the annunciator
lights on the 10 or 8 channel annunciator pan-
els can be extinguished and reset by pressing
either MASTER CAUTION switchlight.
By selecting RECALL on the test switch lo-
cated on the 10 channel annunciator panel,
any annunciators associated with a system cur-
rently malfunctioning illuminate on the 10 and
8 channel annunciator panels. Both MASTER
CAUTION lights also flash after a short delay.
This provides the crew with an effective sys-
tem status check. The individual system an-
nunciators remain illuminated and cannot be
reset as long as the malfunction exists.
This is not a memory system; it will not recall
a system annunciator where no faults con-
tinue to exist.
ELECTRONIC
CIRCUIT-
BREAKER
BOX
TEST/RECALL
SWITCH
AIRPLANE SYSTEMS
(ILLUMINATE WHEN MALFUNCTION EXISTS
OR WHEN LIGHTS ARE BEING TESTED)
AMBER MALFUNCTION LIGHT
RED WARNING LIGHTS*
GREEN STATUS LIGHTS
WHITE STATUS LIGHTS
AMBER STATUS LIGHTS**
DIMMING
MODULE
(1 OF 32)
D-41
DC ESS BUS
B-161
BATT BUSA-161
BATT BUS
4.5 SEC
DELAY
A-160
BATT BUS
B-87
DC BUS 2
BRT/DIM
SWITCH
WARNING LIGHTS
TEST SWITCHES
PROVIDE GROUND
* THE ABOVE ONLY APPLIES
TO RED ANNUNCIATORS
THAT ARE POWERED FROM 
DIMMING MODULES.
AFFECTED SYSTEM
PROVIDES GROUND
** ONLY APPLIES TO AMBER
LIGHTS THAT DO NOT ILLUM-
INATE MASTER CAUTION.
AIRPLANE SYSTEMS
MALFUNCTIONS
DETECTED BY:
• PRESSURE SWITCHES
• PROXIMITY SWITCHES
• THERMAL SWITCHES
ALL MALFUNCTIONS
PROVIDE A GROUND
1 OF 40
POWER OUTPUTS
POWER
SUPPLY
MASTER
CAUTION
PRESS TO
RESET
D-42
DC ESS
BUS
4.5 SEC
DELAY
A-159
BATT BUS
MASTER
CAUTION
PRESS TO
RESET
10-CHANNEL SYSTEM
ANNUNCIATOR PANEL
8-CHANNEL SYSTEM
ANNUNCIATOR PANEL
Figure 4-1. Master Caution System
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COMPONENTS
The master caution system (see Figure 4-1)
contains the two annunciator panels, an elec-
tronic circuit-breaker unit, two MASTER
CAUTION lights, and up to 32 individual dim-
ming modules.
Electronic Circuit Breaker
(ECB) Box
This unit has 40 individual power outputs.
Two outputs power the 8 and 10 channel an-
nunciators. the other outputs provide power to
the dimming modules, which illuminate most
cockpit caution, warning, and advisory lights.
A failure of any of the 40 power outputs which
make up this ECB is indicated by the illumi-
nation of the Auto CB fail annunciator light on
the 10 channel annunciator panel. There is an
internal reset circuit which causes the annun-
ciator light to extinguish if the ECB is suc-
cessful in reestablishing power to that circuit.
Dimming Modules
The dimming modules provide interface be-
tween the airplane systems and the annunci-
ator panels. These modules are located under
the consoles in the cockpit. Each module is
powered by a separate channel of the elec-
tronic circuit breaker and routes this power to
the associated light groups. The modules each
have eight different channels that can be used
for illuminating lights.
The BRT-DIM switch is used to illuminate
most lights powered by the dimming modules
at either of two selected light intensities.
LAMP TEST SWITCHES
Lamp test switches (Figure 4-2) are located on
the 10-channel annunciator panel, the pilot’s fas-
cia panel, the copilot’s fascia panel, and the rear
of the center pedestal. All lights can be tested by
any of the three WARN LT test switches.
WARN
LTS
WARNING LT TEST
TEST
OFF
ON
OFF
TEST
RECALL
BRT
DIM
Figure 4-2. Lamp Test Switches
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ILLUMINATION CAUSES
Table 4-1 lists the 10 and 8 channel annun-
ciators and the individual system caution an-
nunciators, shows their color, and gives the
cause for their illumination. The 10 and 8
channel annunciator panel locations are shown
in the Annunciator Section.
Table 4-1. CAUTION ANNUNCIATORS
Indicated wing anti-ice system has failed.
The wing anti-ice sensor has failed.
Indicated cowl anti-ice system has failed.
Indicated windshield heat has failed.
The ADS heater has failed.
Indicated pitot heat has failed.
Indicated ice detector system has failed.
HTR
FAIL
DUCT
FAIL
SENSOR
FAIL
ANNUNCIATOR ASSOCIATEDANNUNCIATOR CAUSE FOR ILLUMINATION
R FAIL
L FAIL
FAIL
FAIL
PITOT
HEAT
PITOT
HEAT
ON
FAIL
ON
FAIL
The APU fuel pump is inoperative.
The APU/APU generator adapter oil pressure is low.
The APU LCV has failed to close.
LO
PRESS
HI
TEMP
SOV
CLOSED
PUMP
INOP
OPEN
FAILED
TEST
NO HT
ANTI-ICE
APU
The APU/APU generator adapter oil temperature is excessive.
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Table 4-1. CAUTION ANNUNCIATORS (Cont)
Cabin pressurization has failed.
The pressurization controller has failed.
CABIN
PRESS LOW
ANNUNCIATOR ASSOCIATEDANNUNCIATOR CAUSE FOR ILLUMINATION
FAIL
OFF
FAIL
OFF
ENV CONT
HYD
DGRAD—Accuracy of display data cannot
be guaranteed.
IRS 1,2, 3—IRS backup battery charger
has failed. 
COMP MON—There is a difference between
pilot's and copilot's ATT, LOC, GS or IAS.
FAULT—Fault has occurred in the IRS system.
ON BATT—IRS is powered by backup battery.
BATT FAIL—Battery voltage less than required
for IRS operation.
The main landing gear bay overheat detection
 system has failed.
INBD
FAILINBD
TEST
OUTBD
FAIL
OUTBD
TEST
Indicated hydraulic system electric pump pressure is
below 1,800 psi with the flaps selected or pump 
pressure is below 1,800 psi with switch selected ON.
Hydraulic system 1, 2, or 3 overtemped.
ELEC
PUMP
Indicates hydraulic No. 1 system pressure 
is below 1,800 psi.
L ENG
PUMP
Indicated air-conditioning unit has been shut
down automatically.
Indicates hydraulic No. 2 system pressure 
is below 1,800 psi.
An electronic circuit-breaker channel has failed.
Indicated antiskid system has failed.
R ENG
PUMP
OVHT
WARN
FAIL
AUTO CB FAIL
ANTI-SKID
NAV
MLG BAY
OVHT FAIL
IRS
1
IRS
3
IRS
2
COMP
MON
DGRAD
ALIGN
NO AIR
BATT FAIL
NAV RDY
ON BATT 
FAULT
AUTO
FAULT
HI
TEMP
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Table 4-1. CAUTION ANNUNCIATORS (Cont)
BATTERY
CHARGER
The battery master is off, disconnected, or
has failed.
The battery charger has failed or battery is
supplying battery bus loads and o'temp sensing
not available/cell imbalance or overtemp.
The main entrance door's inner T-handle is not stowed.
The main entrance door is not locked.
The baggage door is not secure.
ANNUNCIATOR ASSOCIATEDANNUNCIATOR CAUSE FOR ILLUMINATION
DOORS
ELEC
PASS DR
UNLKD
BAG DOOR
UNSAFE
Indicated generator is not operating.
Indicated generator has overloaded.
GEN OFF
OVLD
Indicated automatic bus tie is switched off.
Indicated automatic bus tie has failed.
Main AC bus 1 or DC bus 1 is not powered.
AUTO OFF
FAIL
MAIN
BUS 1
OFF
Main AC bus 2 or DC bus 2 is not powered.
MAIN
BUS 2
OFF
The DC essential bus is off.
1 OFF
2 OFF
ESS
BUS
OFF
The battery bus is off.
BAT
BUS
OFF
Essential AC bus power has failed.FAIL
High engine vibration is present.HIGHVIB
Essential TRU 1 or 2 has failed.
The engine APR system has failed.
ENGINE
PASS DR
NOT RDY
PASS DR
READY
APR
OVHT
WARN
FAIL
UNSAFE
TO ARM
ARMED
The thrust reversers are not safe to arm.
The thrust reversers are armed.
The engine jet pipe/pylon warning system
has failed.
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Table 4-1. CAUTION ANNUNCIATORS (Cont)
ANNUNCIATOR ASSOCIATEDANNUNCIATOR
FLT CONT
CHAN 1
INOP
CHAN 2
INOP
NOT
ARMED
SEC
TRANS
OVHT
MOT 1
OVHT
MOT 2
FUEL
PITCH
YAW
SCAV
MAIN
MACH
TRIM
OFF
FLAPS
FAIL
SPLR
INOP
LOW
PRESS
Indicated pitch trim channel has failed.
Pitch trim has overspeed failure; and the
channel 1 brake has been applied.
CAUSE FOR ILLUMINATION
Indicated flap motor has overheated.
Indicated scavenge ejector fuel pump
is inoperative.
Indicated main ejector fuel pump is inoperative.
The aileron PCU control valve has jammed or
one hydraulic system is unpressurized.
The elevator PCU control valve has jammed.
The rudder PCU has jammed, or one hydraulic
system is unpressurized.
Mach trim is not engaged.
The flaps have failed.
Tail tank system not armed for auto transfer.
Tail tank system has switched to secondary
transfer (included on AC5135 and subsequent
incorporating SB 601-0355).
The ground spoiler is inoperative.
Indicated standby electric pump is inoperative.
Indicated engine fuel filter is clogged.
Indicated engine fuel pump inlet pressure is low.
OVSP
CHANGE
CHAN
ROLL
MON
SAFE
ON
INOP
VALVE
CLOSED
FILTER
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Table 4-1. CAUTION ANNUNCIATORS (Cont)
Indicated Firex bottle has low pressure.
ANNUNCIATOR ASSOCIATEDANNUNCIATOR CAUSE FOR ILLUMINATION
FIRE FAULT
WOW
Indicated fire warning system has failed.
FIRE
WARN
FAIL
LOW
PRESS
Indicates a disagreement between various WOW
outputs or if either channel power source fails.
The emergency light is off, or the flight
recorder is malfunctioning.
EMER
LTS OFF
HIGH
VIB
WOW
O/P FAIL
WOW
I/P FAIL
EMER LGT
FLT REC
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Table 4-2 lists annunciators that do not acti-
vate MASTER CAUTION annunciators.
Figure 4-3 illustrates the warning annuncia-
tors in schematic form.
Table 4-3 lists the red warning annunciators,
causes for their illumination, and their power
sources.
Table 4-2. MISCELLANEOUS ANNUNCIATORS
NOTE
These will not activate MASTER CAUTION.
ELECTRONIC
CIRCUIT-
BREAKER
BOX
AIRPLANE SYSTEMS
WARNING LIGHTS
DIMMING
MODULE
(1 OF 32)
D-41
DC ESS BUS
B-161
BATT BUS
B-87
DC BUS 2
NOTE:
 THE ABOVE ONLY APPLIES TO
 RED ANNUNCIATORS THAT ARE
 POWERED FROM DIMMING MODULES.
BRT/DIM
SWITCH
WARN LIGHTS
TEST SWITCHES
1 OF 40
POWER OUTPUTS
POWER
SUPPLY
Figure 4-3. System Warning Annunciators
Nosewheel steering is inoperative.
ANNUNCIATOR CAUSE FOR ILLUMINATION
N/W
STEER
FAIL
Inidcates disagreement between proximity switches.
WOW
O/P FAIL
WOW
I/P FAIL
AP DISC
YD OFF
The Yaw Damper is disconnected.
FAIL
ON
Auxiliary Battery failure or internal fault.
IRS
2 AC5135 and Subsequent and SB 601-0418, Auxiliary Battery not charging.
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Table 4-3. WARNING ANNUNCIATORS
A leak is detected in the L/R 10th- or
14th-stage bleed-air manifold.
A leak is detected in the wing or
fuselage ducting.
An overheat is detected in a wing
leading edge.
The IAS is within 3% of the
aerodynamic stall.
The indicated computer channel
is unserviceable.
The altitude compensator is
selected off or has failed.
Dimming modules
Dimming modules
Dimming modules
Dimming modules
Dimming modules
The nose gear doors are open.
Steady—The autopilot has been
manually disconnected
Flashing—The autopilot has
automatically disconnected
Dimming modules
The parking brake is engaged. Dimming modules
The landing gear is in transit or is not
locked in the position selected, or the
LDG GEAR TEST button is pressed.
Landing gear
control unit
Indicated engine or APU has a fire. Fire detection unit
Ice has been detected, but the
anti-icing systems for wing and
engines are not turned on.
Ice detection units
The main landing gear bay
has an overheat.
The jet pipe/pylon has an overheat.
Overheat
detection unit
Stall protect
computer power
supply
FLASHER
MASTER
WARNING
ANNUNCIATOR
ASSOCIATED
ANNUNCIATOR CAUSE FOR ILLUMINATION
POWER
SOURCE
DUCT
FAIL
FLASHER
OVHT
STALL
PUSH
OVHT
OVHT
STALL
PROTECT
FAIL
NOSE
DOOR
OPEN
PARKING
BRAKE
ALT
COMP
FAIL
LH ENG
FIRE
PUSH
RH ENG
FIRE
PUSH
APU
FIRE
PUSH
ICE
ICE
YD OFF
AP DISC
BLEED
CLOSED
DUCT
FAIL
DUCT
FAIL
SENSOR
FAIL
OVHT
ISOL
OPEN
Engine oil pressure is low. Dimming modules
LOP
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Figure 4-4 illustrates the status annunciators
in schematic form.
Table 4-4 lists the nondimmable annunciators
and their associated systems.
AIRPLANE SYSTEMS STATUS LIGHTS
GREEN
WHITE
PROVIDES GROUND
PROVIDES GROUND
PROVIDES GROUND AMBER(THAT DO NOT ILLUMINATE MASTER CAUTION)
DIMMING
MODULE
(1 OF 32)
BRT/DIM
SWITCH
WARN LIGHTS
TEST SWITCHES
1 OF 40
POWER OUTPUTS
POWER
SUPPLY
ELECTRONIC
CIRCUIT-
BREAKER
BOX
D-41
DC ESS BUS
B-161
BATT BUS
B-87
DC BUS 2
Figure 4-4. System Status Annunciators
Table 4-4. NONDIMMABLE ANNUNCIATORS
ANNUNCIATOR ASSOCIATEDSYSTEM
External AC/DC
L/R Engine
Start Switch
ASSOCIATED
SYSTEM
Jet Pipe/Pylon
Overheat Detection
Firex Bottles
(1, 2, APU)
Landing Gear
Control Handle
Main Landing
Gear Bay Overheat
Detection
Ice Detection
Fire Detection
ANNUNCIATOR
Stall ProtectionSTALLPUSH
OVHT
OVHT
START
LH ENG
FIRE
PUSH
RH ENG
FIRE
PUSH
APU
FIRE
PUSH
ICE
ICE
IN USE
AVAIL
SQUIB
SQUIB
Emergency
Lighting
EMER
LTS ON
Fire Protection–
Engines
BOTTLE 1
ARMED
PUSH TO
DISCH
BOTTLE 2
ARMED
PUSH TO
DISCH
BOTTLE
ARMED
PUSH TO
DISCH
Fire Protection–
EnginesFire Protection–
APU
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AURAL WARNING 
SYSTEM
GENERAL
The aural warning system provides an aural
warning when a hazardous situation occurs,
The system provides the flight crew with a
distinct aural warning for each of the seven
events listed in Table 4-5. The switch legends
listed are found on the test selector switch.
COMPONENTS
The system (Figure 4-6) consists of an aural
warning unit and an aural warning test panel.
Power is supplied to the system from the DC
essential bus and the battery bus. Either sup-
ply is sufficient for system operation.
Aural Warning Unit
The aural warning unit processes inputs from
various sensing devices and generates an in-
dividual warning tone electronically for the
eight events indicated above. The output is fed
via amplifiers to the pilot’s and copilot’s head-
phones and to the flight deck speaker system.
The volume to the speakers may be adjusted by
a control knob on the aural warning test panel.
The volume to the headphones is preset and
cannot be adjusted by the flight crew.
Aural Warning Test Panel
The aural warning test panel consists of a ro-
tary tone test selector switch, a volume ad-
justment, and a press-to-mute switch. This
allows the flight crew to test the individual tone
generators and also to mute those tones which
can be muted. Operation of the rotary selec-
tor switch applies an input to generate one of
the eight aural warnings. With the selector
switch in the FIRE or FLAP OVSPD posi-
tions, the tone can be muted by pressing the
press-to-mute switch, which then illuminates.
The tones cannot be muted in any of the other
positions. The system is tested by selecting
each event in turn on the rotary selector and
pressing the mute switch to ensure that only
the two sounds indicated above can be muted.
Tone Muting
After the aural warning is heard and the mute
button is pressed, the aural warning stops and
the button illuminates white and reads “TONE
MUTED.” When the fire or flap overspeed con-
dition is corrected, the TONE MUTED light
extinguishes. In order to hear what warning is
being muted, the illuminated button can be
pressed. On the landing gear control panel there
is a MUTE HORN button (Figure 4-5). this but-
ton is to mute the horn which sounds to indicate
that the landing gear is unsafe when either throt-
tle is in flight idle or lower. When pressed, the
MUTE HORN button illuminates amber.
NOTE
When flaps are extended beyond 30°,
the landing gear unsafe horn cannot
be muted.
UP
LDG GEAR
DN
TEST
MUTE
HORN
DN LCK
REL
LEFT
NOSE
RIGHT
+
Figure 4-5. Landing Gear MUTE HORN
Button
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Table 4-5. AURAL WARNINGS
WARNING SOUND DURATION CAUSE INPUT
Engine fire
APU fire
Bell As long as hazard
exists but may be
muted by TONE
MUTED switchlight.
Fire detected in:
• APU or
• LH engine or
• RH engine
• APU fire detection
• LH engine fire detection unit
• RH engine fire detection unit
Incorrect
takeoff
configuration
Intermittent
horn
Until causes are
corrected, throttles
are returned to idle
or airplane takes off.
• Landing gear control unit
• Throttles
• Flap control unit
• Horizontal stabilizer unit
• Spoiler proximity switches
Overspeed Clacker As long as hazard
exists
VMO exceeded Air data computer
Stall Warbler As long as hazard
exists
Aircraft exceeds stall
margin
Stall warning computer
SELCAL Chime One second In-coming calls Decodes unit from HF1
or HF2
Landing gear
configuration
Horn As long as hazard
exists
As long as hazard
exists
• Landing gear not down
 and locked with flaps
 greater than 30°
• Landing gear not down
 and locked with either
 throttle at idle (may be
 muted by switch on
 landing gear control
 lever panel)
Landing gear control unit
Altitude alert C chord One second • At 1,000 ft from the
 altitude selected on
 vertical navigation
 computer/controller on
 approach to that
 altitude, or
• At 250 ft from the
 altitude selected on
 VNCC on departure
 from that altitude
Vertical NAV computer/-
controller
Airspeed too
high for flap
setting
Wailer Warning begins after
half second delay
and continues as
long as hazard exists
but may be muted by
TONE switchlight.
• 232 kts with any flap
 extension, or
• 198 kts with flaps
 extended beyond 20°
 or,
• 190 kts with flaps
 extended beyond 30°
• Airplane on the ground
 with throttle(s) above
 27.5°, and
• Flaps not extended to
 20°, or
• Spoilers not stowed, or
• Horizontal stabilizer
 out of trim
• Thottle(s) greater than
 25.5°, flaps out of 0° and 
 flight spoiler not stowed
Flap control unit
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AURAL
WARNING
UNIT
FLIGHT
DECK
SPEAKERS
FIRE
OVERSPEED
LANDING
GEAR
INTERPHONE
SYSTEM
TO
HEAD-
PHONES
ALTITUDE
ALERT
FLAP
OVERSPEED
HF 1
HF 2
TOC
SELCAL
RH ENGINE FIRE
DETECTION UNIT
THROTTLE
LEVERS
LANDING GEAR
CONTROL UNIT
AIR DATA
COMPUTER
1 AND 2
VERTICAL NAV
COMPUTER/
CONTROLLER
HORIZONTAL
STABILIZER
UNIT
SPOILER
PROXIMITY
SWITCHES
SELCAL
DECODER
UNIT
LH ENGINE FIRE
DETECTION UNIT
APU FIRE
DETECTION UNIT
STALL
WARNING
COMPUTER
STALL
FLAP CONTROL
UNIT
AURAL WARNING TEST PANEL
AURAL WARNING
VOLUME CONTROL
PRESS TO MUTE
TONE
MUTED
TONE TEST
STALL
FLAP
OVSPD
TOC
SELCAL
LDG GR
FIRE
OVSPD
ALT ADV
OFF
Figure 4-6. Aural Warning System
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1. If an illuminated system malfunction light
suddenly extinguishes, it indicates:
A. Five minutes have passed.
B. The malfunction no longer exists.
C. Three minutes have passed.
D. The MASTER CAUTION lights have
been reset.
2. The MASTER CAUTION lights can be
reset:
A. Anytime, by depressing either MAS-
TER CAUTION switchlight
B. Anytime, except when a red annun-
ciator is flashing
C. Anytime, except when a fire extin-
guisher switch is pushed
D. Anytime, except when a red annun-
ciator is illuminated steady
3. Which of the following systems are rep-
resented on the 8 and 10 channel annun-
ciators:
A. Electronic circuit breaker
B. Weight on wheels
C. Emergency lights
D. All the above
4. The color light that indicates a situation
requiring attention, but not immediate
action is:
A. Red (warning)
B. Amber (caution)
C. White (advisory)
D. Green (test)
5. Which of the following aural warnings is
mutable:
A. Wailer (flap overspeed)
B. Warbler (stall)
C. C/chord (altitude alert)
D. Clacker (overspeed)
6. Most cockpit lights are dimmed:
A. Automatically by photoelectric cells
B. By depressing any test switch
C. By depressing each individual light
D. By selecting the BRT-DIM switch to
DIM
QUESTIONS
CL 601-3A 5-i
CHAPTER 5
FUEL SYSTEM
CONTENTS
Page
INTRODUCTION ................................................................................................................... 5-1
GENERAL............................................................................................................................... 5-1
FUEL STORAGE .................................................................................................................... 5-2
FUEL DISTRIBUTION .......................................................................................................... 5-3
Gravity Flow .................................................................................................................... 5-3
Scavenge Ejectors ............................................................................................................ 5-3
Main Ejectors ................................................................................................................... 5-3
Transfer Ejectors .............................................................................................................. 5-3
Standby Electric Pumps ...................................................................................................5-3
Engine-Driven Pumps ...................................................................................................... 5-3
Crossflow Valve ............................................................................................................... 5-3
Powered Crossfeed Valve................................................................................................. 5-3
APU FUEL SYSTEM.............................................................................................................. 5-8
FUEL CONTROLS AND INDICATORS ............................................................................... 5-9
General ............................................................................................................................. 5-9
Fuel Control Panel ........................................................................................................... 5-9
Fuel Quantity.................................................................................................................. 5-11
VENT SYSTEM.................................................................................................................... 5-12
REFUELING......................................................................................................................... 5-13
General........................................................................................................................... 5-13
Pressure Refueling ......................................................................................................... 5-13
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DEFUELING......................................................................................................................... 5-16
TAIL TANK........................................................................................................................... 5-17
General........................................................................................................................... 5-17
Normal Transfer............................................................................................................. 5-17
Fuel Jettison ................................................................................................................... 5-18
Refueling and Defueling................................................................................................ 5-18
QUESTIONS......................................................................................................................... 5-20
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CL 601-3A 5-iii
ILLUSTRATIONS
Figure Title Page
5-1 Fuel System—General Layout ................................................................................. 5-2
5-2 Fuel Distribution—Normal Operation (without Powered Crossfeed)...................... 5-5
5-3 Fuel Distribution—Normal Operation (with Powered Crossfeed 
Service Bulletin 601-0225) ...................................................................................... 5-5
5-4 Fuel Distribution—Abnormal Operation, Main Electrical
(without Powered Crossfeed) ................................................................................... 5-6
5-5 Fuel Distribution—Abnormal Operation, Main Electrical
(with Powered Crossfeed Service Bulletin 601-0225) ............................................. 5-6
5-6 Fuel Distribution—Engine Start (without Powered Crossfeed ..................................5-7
5-7 Fuel Distribution—Engine Start
(with Powered Crossfeed Service Bulletin 601-0225) ............................................. 5-7
5-8 APU Fuel System—Normal Operation.................................................................... 5-8
5-9 APU Fuel System—Negative G Condition .............................................................. 5-9
5-10 Fuel Controls and Indicators (with Powered Crossfeed
Service Bulletin 601-0225) .................................................................................... 5-10
5-11 Vent System............................................................................................................ 5-12
5-12 Pressure-Refueling System .................................................................................... 5-13
5-13 Exterior Fueling Components ................................................................................ 5-14
5-13 Refuel-Defuel Control Panel.................................................................................. 5-15
5-14 Refuel-Defuel Control Panel—Tail Tank............................................................... 5-15
5-15 Tail Tank Quantity Panel ........................................................................................ 5-17
5-17 Tail Tank Fuel Transfer Panel ................................................................................ 5-17
5-18 Tail Tank Flow Schematic ...................................................................................... 5-19
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INTRODUCTION
The Canadair Challenger CL-600-2B16, model CL-601-3A fuel system provides fuel
for the two turbofan engines, as well as the auxiliary power unit (APU). Fuel is also used
to cool the APU generator adapter oil and the main engine oil.
GENERAL
The Challenger uses a wet-wing box struc-
ture which forms three separate fuel tanks;
two main tanks in the outboard wing sections
and an auxiliary tank in the wing center sec-
tion. Maximum fuel capacity is approximately
16,500 pounds.
NOTE
Aircraft incorporating SB-601-0262,
maximum fuel capacity is approxi-
mately 17,900 pounds, due to the
addition of a tail tank.
Fuel is supplied to the engines from two collector
tanks. Fuel is delivered from each collector tank
to its respective engine by a main ejector pump
located within the tank. Additional scavenge and
transfer ejector pumps are located within the
main and auxiliary tanks to ensure proper fuel
distribution.
Electrically operated standby fuel pumps are pro-
vided. These pumps are operated during engine
starting or following a main ejector pump failure.
A fuel imbalance between the main tanks may
be corrected by opening a crossflow valve. This
allows the quantities in the main tanks to equal-
ize by gravity flow. For airplanes fitted with
SB-601-0225, opening a LEFT to RIGHT or a
RIGHT to LEFT powered crossfeed shutoff
valve will transfer fuel from the one main tank
to the other.
0
2
4 6
8
10
MAIN
FUEL
LBS X 100
CHAPTER 5
FUEL SYSTEM
The airplane may be refueled over the wings
by gravity. However, the normal method is
through an adapter located in the right wing
root using a single-point pressure system.
FUEL STORAGE
Using a wet-wing concept, the entire wing box
structure is sealed to form three tanks, which
carry most of the fuel (Figure 5-1). Two addi-
tional tanks are fitted under the cabin floor,
fore and aft of the auxiliary tank, which is in
the wing center section. These tanks are inter-
connected with the auxiliary tank. A tail tank
will be fitted aft of the stabilizer rear spar.
The main tanks encompass the internal wing
volume from near the wingtip to near the wing
roots. There are 16 inspection and maintenance
access panels in the lower surface of each wing.
The gravity filler port for each main tank is lo-
cated on the upper outboard wing surface.
The auxiliary tank encompasses the entire
center section of the wing. There are access
panels in the lower wing surface. The auxil-
iary tank gravity filler port is located in the
right wing root just aft of the leading edge.
Contained within the auxiliary tank are two
collector tanks which are extensions of each
main tank. They incorporate the main fuel
ejectors and feed fuel directly to each engine.Each collector tank is constantly kept full
with fuel from its respective main tank. The
standby electric fuel pumps are also housed
within the collector tanks.
FlightSafety Canada LtéeLtd.
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Figure 5-1. Fuel System—General Layout
COLLECTOR
TANKS
SINGLE-POINT
REFUEL-DEFUEL
ADAPTER
AUXILIARY TANK
FILLER CAP
AUXILIARY TANK
RIGHT MAIN
TANKRIGHT MAIN
 FILLER CAP
LEFT MAIN
TANK
LEFT MAIN
FILLER CAP
TAIL TANK*LEGEND
MAIN TANK FUEL
AUXILIARY FUEL TANK *AC INCORPORATING
 SB 601-0262
FUEL DISTRIBUTION
GRAVITY FLOW
Fuel flows by gravity within the main tanks
through lightening holes in the ribs. One inner
rib in each main tank is equipped with flap-
per valves to prevent outward flow of fuel
(Figure 5-2 or 5-3).
Fuel flows from the inboard sections of the
main tanks to the collector tanks (Figure 5-2
or 5-3). Flapper valves at the collector tank
inlets prevent reverse flow of fuel into the
main tanks.
SCAVENGE EJECTORS
Gravity flow to the collector tanks is sup-
plemented by scavenge ejectors located at
the rear of the inboard section of each main
tank. The scavenge ejectors ensure that the
collector tanks are supplied with fuel re-
gardless of airplane attitude. Failure of a
scavenge ejector will cause illumination of
an amber caution light in the cockpit.
Motive flow for operation of a scavenge ejec-
tor comes from the high-pressure side of the
two-stage engine-driven fuel pump (Figure
5-2 or 5-3).
Ejectors have no moving parts. Each operates
on the venturi principle to convert small-
volume, high-pressure motive flow at the
throat of the ejector into large-volume, low-
pressure output at the ejector nozzle.
MAIN EJECTORS
Fuel is supplied from the collector tanks to the
low-pressure side of each engine-driven pump
by a main ejector located within the tank.
Motive flow for operation of the main ejectors
is supplied by the high-pressure side of each
engine-driven fuel pump.
Each main ejector provides continuous fuel
flow to its own engine through firewall shutoff
valves. Flow to the opposite engine is not pos-
sible because of one-way check valves in the
feed lines. Failure of a main ejector will cause
illumination of an amber caution light in the
cockpit, and will activate both standby elec-
tric pumps. (Figure 5-4 or 5-5).
TRANSFER EJECTORS
When main tank fuel quantity drops below
the 93% full level, float valves open, allow-
ing the transfer ejectors to draw fuel from the
auxiliary tank to the inboard sections of the
main tanks. Motive flow for the transfer ejec-
tors is provided by the output of the associated
main ejector. No cockpit indication of fuel
transfer or transfer ejector failure is provided.
A one-way check valve in each transfer ejec-
tor prevents fuel migration from the main
tanks to the auxiliary tank.
STANDBY ELECTRIC PUMPS
Electric standby pumps are provided for en-
gine starting and as a backup in the event that
a main ejector becomes inoperative (Figure
5-4 or 5-5). The two DC-powered pumps will
then operate simultaneously to draw fuel
from their respective collector tanks and feed
a common line capable of providing fuel to
either engine.
Once armed by cockpit switches, the standby
pumps operate automatically when the output
pressure of either main ejector falls below 10
psi. During the engine start sequence, both
pumps operate until the engine-driven pumps
generate enough motive flow to operate the
main ejectors.
The left electric pump is powered by the bat-
tery bus, while the right pump receives power
from DC bus No. 2.
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ENGINE-DRIVEN PUMPS
Each two-stage engine-driven fuel pump is a
single unit containing two pumps mounted in
piggyback fashion. The first stage is a cen-
trifugal, low-pressure pump which receives
fuel from the main ejector and supplies it to
the engine and the second stage, or high-
pressure side, of the pump. This second stage
uses a positive-displacement pump to supply
high-pressure motive flow fuel to the main
and scavenge e j ec to r s . ( see F igure 5 -2 
or 5-3.)
Operation of the engine fuel system is dis-
cussed in Chapter 7, “Powerplant.”
CROSSFLOW VALVE
Should a main tank fuel imbalance occur in
flight for any reason, it can be corrected by
opening the crossflow valve (see Figure 5-2
or 5-3) which connects both main tanks and
both collector tanks. Balance is achieved
through gravity flow only. To avoid a serious
imbalance which might occur if the wings are
not level, the crossflow should not be left open
when the airplane is on the ground.
POWERED CROSSFEED VALVE
When depressing the LEFT TO RIGHT or
RIGHT TO LEFT switchlight , the associated
powered crossfeed shutoff valve (Figure 5-4
or 5-5) opens to allow fuel flow, by gravity, into
the auxiliary fuel tank. Fuel is then transferred
to the opposite tank while a quantity is re-
turned to its original tank by the transfer
ejectors when in flight or the fuel boost pumps
via the transfer ejectors when on the ground.
NOTE
Maximum imbalance is 800 pounds.
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Figure 5-2. Fuel Distribution—Normal Operation (without Powered Crossfeed)
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
COLLECTOR
TANK
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINE
MAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
COLLECTOR
TANK
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
TRANSFER
EJECTOR
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVEMAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-3. Fuel Distribution—Normal Operation (with 
Powered Crossfeed Service Bulletin 601-0225)
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ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
COLLECTOR
TANK
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINE
MAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
COLLECTOR
TANK
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
TRANSFER
EJECTOR
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
Figure 5-4. Fuel Distribution—Abnormal Operation,
Main Ejector Fail (without Power Crossfeed)
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVEMAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-5. Fuel Distribution—Abnormal Operation, Main Ejector Fail 
(with Power Crossfeed ServiceBulletin 601-0225)
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ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
COLLECTOR
TANK
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINE
MAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
COLLECTOR
TANK
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
TRANSFER
EJECTOR
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
Figure 5-6. Fuel Distribution—Engine Start (without Powered Crossfeed )
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVEMAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-7. Fuel Distribution—Engine Start (with Powered Crossfeed 
Service Bulletin 601-0225)
APU FUEL SYSTEM
The APU is normally supplied with fuel from
the right main tank by an electric fuel pump.
The pump is identical with the standby elec-
tric pumps described previously.
The APU fuel pump operates whenever it is se-
lected on by a switch on the APU control
panel. Fuel in excess of APU requirements is
routed back to the right main tank through a
fuel-oil heat exchanger which cools the APU
generator adapter oil (Figure 5-8). Operation
o f t he APU i s d i s cus sed i n Chap t e r 6 ,
“Auxiliary Power Unit.”
The APU fuel feed line is fitted with two APU
fuel shutoff valves that are synchronized and
controlled by the APU electronic control unit.
To ensure uninterrupted operation of the APU
during brief moments of negative G flight or
in case of APU fuel pump failure, fuel can be
supplied from the left engine feed line to the
APU (Figure 5-9). This line has a differential
pressure, one-way check valve which opens
whenever the main APU supply pressure drops
10 psi lower than the pressure in the left en-
gine fuel feed line. Fuel from the left engine
feed line cannot flow to the right tank or to any
heat exchanger because of a check valve in-
stalled in the main feed line.
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RIGHT MAIN TANK
APU
FUEL
PUMP
STANDBY
PUMPSFROM
LEFT MAIN
EJECTOR
TO LEFT
ENGINE
APU NEGATIVE-G
CHECK VALVE
APU PUMP PRESSURE
MAIN EJECTOR PRESSURE
APU NEGATIVE-G
SHUTOFF VALVE
APU FUEL FEED
SHUTOFF VALVE
FUEL
CONTROL
UNIT
TO PNEUMATIC
SYSTEM
APU GENERATOR
OIL HEAT
EXCHANGER
LOAD
CONTROL
VALVE
RESTRICTOR
CHECK
VALVE
FUEL CONTROL
UNIT SHUTOFF VALVE
PRESSURE
SWITCH
CENTER TANK
LEGEND
CHECK
VALVE
Figure 5-8. APU Fuel System—Normal Operation
FUEL CONTROLS
AND INDICATORS
GENERAL
The fuel controls and indicators are grouped
on the center instrument panel (Figure 5-10).
The fuel control panel is located just above the
fuel quantity panel.
The fuel control panel contains five switch-
lights, six additional annunciators, and a fuel
temperature gage. The fuel quantity panel
contains five digital readouts.
FUEL CONTROL PANEL
The standby electric fuel pumps are controlled
by a pair of switchlights labeled “PUMP.”
Pump operation is indicated by illumination
of the green ON legend in the top half of the
associated switchlight.
The bottom half of the PUMP switchlight has
an amber INOP legend which illuminates to in-
dicate that the associated pump is not selected
on or that the pump is not operating properly.
A third switchlight, labeled “X-FLOW,” con-
trols operation of the crossflow valve. The
green OPEN light illuminates to indicate that
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RIGHT MAIN TANK
APU
FUEL
PUMP
STANDBY
PUMPSFROM
LEFT MAIN
EJECTOR
TO LEFT
ENGINE
APU NEGATIVE-G
CHECK VALVE
STANDBY PUMP PRESSURE
MAIN EJECTOR PRESSURE
APU NEGATIVE-G
SHUTOFF VALVE
APU FUEL FEED
SHUTOFF VALVE
FUEL
CONTROL
UNIT
TO PNEUMATIC
SYSTEM
APU GENERATOR
OIL HEAT
EXCHANGER
LOAD
CONTROL
VALVE
RESTRICTOR
CHECK
VALVECHECK
VALVE
FUEL CONTROL
UNIT SHUTOFF VALVE
PRESSURE
SWITCH
CENTER TANK
LEGEND
Figure 5-9. APU Fuel System—Negative G Condition
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TOTAL
QUANTITYFUEL
AUX
R. MAINL. MAIN
PUSH ON OFF
PUMP PUMPX-FLOW
E
J
C
T
F
E
E
D
L
E
F
T
E
N
G
F
U
E
L
E
N
G
F
U
E
L
FUEL CONTROL
SCAV
MAIN 
VALVE
CLOSED
FILTER 
LOW
PRESS
VALVE
CLOSED
FILTER 
LOW
PRESS
SCAV
MAIN 
ON
INOP 
ON
INOP 
OPEN
LEFT
TO
RIGHT
RIGHT
TO
LEFT
E
J
C
T
F
E
E
D
R
I
G
H
T
70
-20
20
40
0
70
-20
20
60
40
0
FUEL
°C
L R
60
Figure 5-10. Fuel Controls and Indicators (with Powered 
Crossfeed Service Bulletin 601-0225)
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the valve is fully opened. The light extin-
guishes when the valve is fully closed. Since
the valve is motor operated, expect a delay of
approximately 2 seconds from the time the
switchlight is pressed until the proper indi-
cation appears. The valve normally remains in
the closed position and is opened only during
flight to correct a fuel imbalance.
The green LEFT TO RIGHT and RIGHT TO
LEFT switchlights are part of the POWERED
CROSS FEED system. Should an imbalance
between main tank fuel levels develop and
gravity equalizing with the crossflow valve not
be possible, the transfer is possible by de-
pressing the appropriate switchlight. The
depressed switchlight will illuminate steady
and will start flashing after eight minutes as
a reminder. An interlock prevents simultane-
ous operation of both switchlights.
The remaining five lights for each engine pro-
vide information as follows:
• The amber SCAV light illuminates to in-
dicate an inoperative scavenge ejector.
• The amber MAIN light illuminates to in-
dicate an inoperative main ejector. This
automatically triggers both standby elec-
tric pumps to operate if they are selected
to the ON position.
• The white VALVE CLOSED light illu-
minates to indicate that the firewall
shutoff valve has closed. Control of this
valve is from the respective FIRE PUSH
switchlight on the center glareshield.
• The amber FILTER light illuminates to
indicate an impending fuel filter bypass
or a clogged filter. (This condition is
covered in Chapter 7, “Powerplant.”)
• The amber LOW PRESS light illumi-
nates if fuel pressure at the inlet side of
the engine-driven pump falls below a
predetermined value.
• A fuel temperature indicator in the cen-
ter of the fuel control panel indicates
the temperature of the fuel as it leaves
the fuel heater at the fuel filter. (This is
a l so cove red i n Chap t e r 7 ,
“Powerplant.”) All amber lights on this
panel will activate the flashing master
caution lights and illuminate the FUEL
annunciator light.
FUEL QUANTITY
The fuel quantity in each main tank as well as
the auxiliary tank is measured by a system of
the capacitance-type transmitters located in
each tank. Quantity information is fed to a sig-
nal conditioner which displays the quantity in
pounds for each tank, as well as the total, on the
fuel quantity panel. Only usable fuel is shown.
The digital readouts are tested from the engine
instrument test switch on the center instru-
ment panel. A successful test is indicated by
the appearance of a series of eights in the dig-
ital readouts (except for the last digit which
is zero on all but the total readout wherein thelast two digits are always zero).
The probes for the fuel quantity are powered
as follows:
• L. MAIN ............... DC essential bus
• R. MAIN ........................ battery bus
• AUX and TAIL.................. DC bus 1
The fuel quantity indicators are powered via
the SDC as follows:
• RH, LH,
AUX, TAIL ......... DC essential bus
• TOTALIZER ................. battery bus
If power to the probes is lost, the affected
quantity indicator will read zero, and the
amount of fuel remaining in that tank will be
subtracted from the total.
VENT SYSTEM
Each tank is vented at two different points
through a series of vent lines which allow air
to enter or escape the tanks, depending upon
whether fuel is being used or added (Figure 5-
11). The vent lines extend from each wingtip
to common manifolds which form an inter-
connected inverted “U” in each fuselage wall
and then return to the wing area where they ter-
minate under the trailing edge on each side in
a flush-mounted NACA scoop. The scoop,
which has ice rejection capability, maintains
a slight positive tank pressure during flight due
to ram-air effect.
There are no valves or screens in the vent
lines, so dirt or ice accumulation does not nor-
mally occur. Any trapped fuel or moisture in
the vent lines is continuously purged from the
low points by a bleed line connected to each
scavenge ejector.
During pressure fueling, the fuel tank vent lines
are augmented with special vent valves which will
be described later under “Pressure Refueling.”
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AUXILIARY
TANK VENT
MAIN TANK
CLIMB VENT
MOTIVE
FLOW
PURGE
LINE
SCAVENGE FLOW
TO COLLECTOR TANK
MAIN TANK VENT
NACA SCOOP INLET
PURGE LINE
INVERTED U
VENT TUBE
SCAVENGE
EJECTOR
Figure 5-11. Vent System
REFUELING
GENERAL
All tanks are normally fueled by means of the
single-point pressure adapter located in the
right wing root (Figure 5-13). The refueling
system is controlled from a swing-out control
panel located in the fillet above the right wing.
The system has automatic fuel cutoff to pre-
vent overfilling.
Overwing or gravity fueling is also possible.
However, due to the location of the filler caps
(Figure 5-13), it is not possible to fill the main
tanks completely. A separate gravity filler
port is provided for each main tank and the aux-
iliary tank. The forward and aft tanks are
gravity fueled through the auxiliary tank.
Never open an overwing filler cap if
that main tank contains more than
4,000 pounds or if the level is not
known. Fuel in excess of 4,000 pounds
will spew from the filler if opened.
PRESSURE REFUELING
The pressure-refueling system (Figure 5-12)
consists of a single-point adapter, a pressure
manifold containing a two-way check valve,
three shutoff valves (SOV) associated with
three float-operated, full level-control valves,
and three vent valves that can be tested during
the refueling process and are utilized as backup
to normal pressure relief during refueling.
The adapter and manifold can accept a flow
rate of up to 250 gpm at a pressure of from 20
to 55 psi. The pressure-fueling process is con-
trolled from the fuel-defuel panel.
During a normal pressure-refueling opera-
tion, the sequence of events is as follows:
• Move the power switch to ON and check
that the green POWER ON light illu-
minates (Figure 5-14 or 5-15) (powered
from battery direct bus).
• Check that all three green vent valve (VV)
lights are extinguished and that all three
or four green SOV CLosed lights are 
illuminated.
CAUTION
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TANK SOV
LEVEL
CONTROL
VALVE
VENT
TEST
VALVE *
VENT
(RELIEF)
VALVE *
VENT
TEST VALVES*
FUEL
MANIFOLD
SINGLE-POINT
ADAPTER
TWO-WAY
CHECK VALVE
VENT LINES
REFUELING PRESSURE
LEGEND
*NOTE
 ENERGIZED CLOSED DURING FUELING ONLY
Figure 5-12. Pressure-Refueling System
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AUXILIARY TANK GRAVITY
REFUELING PORT
PRESSURE REFUEL- DEFUEL
CONTROL PANEL
MAIN TANK
GRAVITY REFUELING PORT
PRESSURE REFUELING
ADAPTER
Figure 5-13. Exterior Fueling Components
• Connect fueling nozzle from the truck
to the single-point adapter.
• Open the fuel nozzle valve and check
that with fuel pressure applied, all three
VV Open lights remain off.
• Rotate the MODE selector from OFF
to TEST.
• Move the tank FUEL-DEF switch of a
tank to be filled to the FUEL position and
check that the corresponding amber SOV
OPen light illuminates.
• Check that the appropriate VV OPEN
lights illuminate within 30 seconds.
• After 30–40 seconds, the SOV CLosed
light will illuminate and the OPen light
will extinguish, indicating that the au-
tomatic level control system is working
properly.
• Test other SOVs and VVs individually
for other tanks to be filled.
• Rotate the MODE selector to FUEL.
• Check that all three VV OPEN lights
extinguish.
• Ensu re t ha t t he t ank FUEL-DEF
switches of the tanks to be filled are in
the FUEL position.
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Figure 5-15. Refuel-Defuel Control Panel—Tail Tank
Figure 5-14. Refuel-Defuel Control Panel
• Check that the SOV OPen lights are il-
luminated. Add fuel to the desired level.
• When the selected tanks are full, their
shutoff valves will close automatically
and the SOV CLosed lights will illuminate.
• Move the tank switches to OFF.
• Rotate the MODE selector to OFF.
• Remove the fueling nozzle from the
adapter, reinstall the cover, and close
the door carefully.
• Check that all lights and switches are off
and stow the refuel-defuel control panel.
NOTE
Adding fuel will compress the land-
ing gear shock struts and lower the
airplane. Be sure that stands, lad-
ders, or other equipment that might
damage the airplane are removed be-
fore refueling.
Do not chock forward of the nose gear wheels
during refueling as increasing the weight will
move the nosewheels forward.
If tanks are to be only partially filled, tank lev-
els must be moni tored f rom the cockpi t
quantity panel. For this, AC electrical power
must be provided, usually from the APU or ex-
ternal AC unit.
DEFUELING
The airplane may be completely defueled
without the use of electrical power via the
single-point adapter.
Fuel tender suction of negative 8 psi (max-
imum) at the manifold will open the tank
shutoff valves and withdraw fuel. When the
tanks are empty, the valves will close by
spring pressure.
The tanks may be selectively defueled using
the refuel-defuel control panel in the DEFUEL
mode as follows:
• Move the power switch (Figures 5-14
and 5-15) to ON and check that the green
POWER ON light illuminates.
• Connect the defueling nozzle from the
tender to the single-point adapter.
• Open the nozzle valve and check for a
maximum negative pressure of 8 psi.
• Rotate the MODE selector to the DE-
FUEL position.
• Move the tank switch for tank(s) to be de-
fueled to the DEF position. This allows
suction to open the associated shutoff
valve, illuminating the amber OPen light.
To prevent fuel tank damage caused
by negative pressure, always open
overwing filler port for the tank
being defueled as soon as tank quan-
tity is less than 4,000 pounds. Do
not open filler cap if tank contains
more than 4,000 pounds.
• When the tank quantity decreases to the
desired level, move the tank switch to
OFF. This will cause the shutoff valve
to close and the green CLOSED light
will illuminate.
• Rotate the mode selector to OFF.
• Move the power switch to OFF.
• Remove the defuel ing nozzle f rom
the adapter. Reinstall the cover and close
the door carefully.
• Check that all lights and switches are off
and stow the refuel-defuel control panel.
CAUTION
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
5-16 CL 601-3A FOR TRAINING PURPOSESONLY
TAIL TANK
GENERAL
An additional fuel tank, which will increase
the fuel load by some 187.7 U.S. gallons (1,250
pounds), can be fitted in the tail.
A powered fuel transfer and a backup secondary
transfer system will automatically transfer fuel
from the tail tank to the auxiliary tank.
Pressure refueling and defueling is possible
through the single-point refueling and defu-
eling panel.
The tail tank cannot be refueled unless the
auxiliary tank is full.
The tail tank is also equipped with a fuel dump
system, as well as a quantity indicator (Figure
5-16) in the cockpit.
The tail tank is located aft of the vertical sta-
bilizer rear spar. It includes two transfer pumps
(DC-powered), No. 1 and No. 2 tail tank
empty/level switches, fuel quantity transmit-
ter, dump valve, and tail tank fuel level switch.
It also has a vent system connected to the rear
air duct and exit flush at the bottom of the
fuselage via a flame arrester.
NORMAL TRANSFER
With the ARMED—OFF switch on the tail tank
fuel transfer panel (Figure 5-17) set to ARMED,
the fuel transfers automatically to the auxiliary
tank at the rate of approximately 25 pounds
per minute. When only unusable fuel (approx-
imately 1.3 U.S. gallons) is left, the shutoff
valve closes and the fuel transfer pump stops.
If the ARMED—OFF switch is set to OFF
while there is still usable fuel in the tail
tank, the NOT ARMED light illuminates. If
fuel in the auxiliary tank reaches the level
of the No. 2 auxiliary tank level switch, the
NOT ARMED l ight f lashes. Placing the
ARMED—OFF switch to the ARMED po-
sition starts the normal transfer, and the
NOT ARMED light goes out. If the NOT
ARMED light remains on, jettison fuel.
In the event of a normal fuel transfer failure,
the SEC TRANS light comes on, shutting
down the primary and starting the secondary
transfer system. The SEC TRANS light indi-
cates that secondary transfer is taking place at
approximately 25 pounds per minute. Monitor
fuel quantity during secondary transfer. If ad-
equate transfer cannot be confirmed with a
maximum of ten minutes, jettison fuel.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A 5-17FOR TRAINING PURPOSES ONLY
ARMED
T/TANK FUEL TRANS
OFF
NOT
ARMED
SEC
TRANS
DUMP
SW
ARMED
DUMP
OPEN
DUMP
SELECT
Figure 5-17. Tail Tank Fuel Transfer Panel
TOTAL
FUEL
QUANTITY
LB
AUX
L. MAIN R. MAIN
TAIL
Figure 5-16. Tail Tank Quantity Panel
The fuel transfer line has a nitrogen-pres-
surized shroud in the rotor boost zone. A
perforation in this shroud signals all open
shutoff valves to close and shuts down the
boost pump. Fuel transfer cannot be accom-
plished unless the shroud is replaced and
pressurized. Fuel in this case has to be jetti-
soned.
See Figure 5-18 for a schematic of the tail
tank system.
FUEL JETTISON
Pressing the DUMP SW ARMED switchlight
causes it to illuminate green, the white DUMP
SELECT light illuminates and arms the DUMP
SELECT/DUMP OPEN guarded switchlight.
Once armed, pressing the green DUMP OPEN
switchlight causes it to illuminate and opens
the jettison valve. Fuel discharges at a rate of
100 pounds per minute.
NOTE
Fuel jettison must only be carried
out with flaps set at 0°. If an imme-
diate landing is required, the fuel
jettison procedure should be initi-
ated immediately. However, if it is
not possible to complete the jetti-
son procedure, a landing may be
made with fuel in the tail cone tank.
Do not jettison fuel in known light-
ning conditions.
REFUELING AND DEFUELING
Refueling can only be accomplished if the
auxiliary tank is full. The refuel-defuel panel
is equipped with a tail tank refuel-defuel
switch. When placed to FUEL, the green CL
light remains on until the auxiliary is full, and
then the amber OP light comes on, indicating
the refueling is taking place.
Defueling can be carried out any time, regard-
less of fuel quantity. There is no provision for
gravity refueling.
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WARNING
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A 5-19FOR TRAINING PURPOSES ONLY
REFUELING PRESSURE
LEGEND
RAM AIR
NO. 2 AUX TANK
LEVEL SWITCH
NO. 1 AUX TANK
LEVEL SWITCH
FUEL LEVEL
SWITCH
FWD
AUX TANK
PRIMARY
AUX
TANK
REAR
AUX
TANK
TAIL
TANK
LEFT
MAIN
TANK
RIGHT
MAIN
TANK
FUELING ADAPTER
AND MANIFOLD
ROTORBURST
ZONE
TRANSFER
LINE
TRANSFER
SOV
WATER DRAIN
FULL
LEVEL
SWITCH
NO. 1 TAIL TANK
LEVEL SWITCH
NO. 2 TAIL TANK
LEVEL SWITCH
JETTISON
SOV
RAM AIR
SECONDARY
TRANSFER PUMP
SECONDARY
TRANSFER SOV
PRESSURIZED
SHROUD
FUELING
LINE
FLOW
SENSOR
Figure 5-18. Tail Tank Flow Schematic
1. The engine-driven fuel pumps are nor-
mally supplied with fuel by the:
A. Main ejectors
B. Standby electric pumps
C. Scavenge ejectors
D. Transfer ejectors
2. If the main ejectors fail, the engine-driven
fuel pumps will be supplied with fuel by:
A. Gravity
B. Scavenge ejectors
C. Standby electric pumps
D. Transfer ejectors
3. Fuel imbalance between the main tanks
is corrected by:
A. Transfer ejectors
B. Scavenge ejectors
C. Standby pumps
D. Gravity crossflow
4. The scavenge ejectors:
A. Transfer fuel from the auxiliary tanks
to the main tanks.
B. Provide a flow of fuel from the main
tanks to the collector tanks.
C. Supply fuel to the engine-driven
pumps.
D. Transfer fuel from the main tanks to
the auxiliary tank and correct a fuel
imbalance when the crossflow valve
is open.
5. Flapper valves in the main tanks prevent:
A. Reverse flow of fuel from the main
tanks to the auxiliary tank
B. Reverse flow of fuel from the auxil-
iary tank to the main tanks
C. Gravity flow of fuel from the out-
board ma in t ank sec t ions to the
inboard main tank section
D. Reverse flow of fuel from the inboard
main tank sections to the outboard
main tank sections
6. Fuel for the APU is normally supplied by:
A. The left scavenge ejector
B. An electric pump in the right main
tank
C. An electric pump in the left main tank
D. The right main ejector
7. In case of a negative G condition, fuel for
the APU is supplied by:
A. The left main ejector
B. An electric pump in the right collec-
tor tank
C. An electric pump in the left collector
tank
D. All the above
8. The pump switchlights on the fuel con-
trol panel control the:
A. APU fuel pump
B. Main ejectors
C. Standby electric pumps
D. Differential pressure check valve
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CL-600-2B16 PILOT TRAINING MANUAL
5-20 CL 601-3A FOR TRAINING PURPOSES ONLY
QUESTIONS
9. When an amber MAIN light illuminates
on the fuel control panel:
A. The firewall shutoff valves close au-
tomatically.
B. The standby electric pumps start op-
erating.
C. The crossflow valve opens.
D. The associated engine will flame out.
10. The preferred method of refueling is:
A. Gravity fueling when a full fuel load
is required
B. Single-point pressure refueling
C. Single-point pressure refueling of
main tanks only
D. Single-point pressure refueling; how-
ever, the wing tanks must be topped
off by overwing fueling.
11. The maximum pressure for single-point
refueling is:
A. 35 psi
B. 45 psi
C. 55 psi
D. 65 psi
12. Never open the overwing filler caps if
the fuel level is unknown or if the tanks
contain more than:
A. 3,000 pounds
B. 4,000 pounds
C. 3,000 gallons
D. 4,000 gallons
13. The maximum negative pressure allowed
for single-point defueling is:
A. 6 psi
B. 8 psi
C. 10 psi
D. 12 psi
14. During flight, the maximum imbalance al-
lowed between the main tanks is:
A. 800 pounds
B. 1,836 pounds
C. 2,500 pounds
D. 2,836 pounds
15. The Ta i l Tank sy s t em ambe r NOT
ARMED light, if on steady:
A. Indicated no fuel in the tank
B. Indicates fuel is in the tail tank and the
transfer system should be armed by
the crew.
C. Will produce a Master Caution indi-
cation.
D. The tail tank fuel will transfer auto-
matically when flashing.
16. To dump fuel:
A. Two switchlights need to be pressed.
B. Three switchlights need to be pressed.
C. TheDUMP SELECT/DUMP OPEN
switchlight must first be armed.
D. Both A and C are good answers.
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A 5-21FOR TRAINING PURPOSES ONLY
CL 601-3R 5-i
CHAPTER 5
FUEL SYSTEM
CONTENTS
Page
INTRODUCTION ................................................................................................................... 5-1
GENERAL............................................................................................................................... 5-1
FUEL STORAGE .................................................................................................................... 5-2
FUEL DISTRIBUTION .......................................................................................................... 5-3
Gravity Flow .................................................................................................................... 5-3
Scavenge Ejectors ............................................................................................................ 5-3
Main Ejectors ................................................................................................................... 5-4
Transfer Ejectors .............................................................................................................. 5-4
Standby Electric Pumps ................................................................................................... 5-5
Engine-Driven Pumps ...................................................................................................... 5-6
Crossflow Valve ............................................................................................................... 5-7
Powered Crossfeed Valve................................................................................................. 5-7
APU FUEL SYSTEM.............................................................................................................. 5-7
FUEL CONTROLS AND INDICATORS ............................................................................... 5-8
General ............................................................................................................................. 5-8
Fuel Control Panel ........................................................................................................... 5-8
Fuel Quantity ................................................................................................................. 5-10
VENT SYSTEM.................................................................................................................... 5-10
REFUELING ......................................................................................................................... 5-11
General ........................................................................................................................... 5-11
Pressure Refueling ......................................................................................................... 5-11
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
DEFUELING......................................................................................................................... 5-14
TAIL TANK........................................................................................................................... 5-15
General........................................................................................................................... 5-15
Normal Transfer............................................................................................................. 5-15
Tail Tank Amber Warning Lights Versus Master Caution ............................................. 5-16
Fuel Jettison ................................................................................................................... 5-16
Refueling and Defueling................................................................................................ 5-16
QUESTIONS......................................................................................................................... 5-19
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CL-600-2B16 PILOT TRAINING MANUAL
5-ii CL 601-3R FOR TRAINING PURPOSES ONLY
CL 601-3R 5-iii
ILLUSTRATIONS
Figure Title Page
5-1 Fuel System—General Layout ................................................................................. 5-2
5-2 Fuel Distribution—Normal Operation ..................................................................... 5-3
5-3 Fuel Distribution—Engine Start............................................................................... 5-4
5-4 Fuel Distribution—Abnormal Operation, Main Ejector Fail ................................... 5-5
5-5 APU Fuel System—Normal Operation.................................................................... 5-7
5-6 APU Fuel System—Negative G Condition .............................................................. 5-7
5-7 Fuel Controls and Indicators .................................................................................... 5-9
5-8 Vent System............................................................................................................ 5-10
5-9 Exterior Fueling Components ................................................................................ 5-12
5-10 Pressure-Refueling System .................................................................................... 5-13
5-11 Refuel-Defuel Control Panel.................................................................................. 5-13
5-12 Tail Tank Quantity Panel ........................................................................................ 5-15
5-13 Tail Tank Fuel Transfer Panel ................................................................................ 5-16
5-14 Tail Tank Flow Schematic ...................................................................................... 5-17
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
INTRODUCTION
The Canadair Challenger CL-600-2B16, model CL-601-3R fuel system provides fuel for
the two turbofan engines, as well as the auxiliary power unit (APU). Fuel is also used
to cool the APU generator adapter oil and the main engine oil.
GENERAL
The Challenger uses a wet-wing box struc-
ture which forms three separate fuel tanks;
two main tanks in the outboard wing sections
and an auxiliary tank in the wing center sec-
tion. Maximum fuel capacity is approximately
17,900 pounds.
Fuel is supplied to the engines from two col-
lector tanks. Fuel is delivered from each col-
lector tank to its respective engine by a main
ejector pump located within the tank. Addi-
tional scavenge and transfer ejector pumps
are located within the main and auxiliary tanks
to ensure proper fuel distribution.
Electrically operated standby fuel pumps are
provided. These pumps are operated during
engine starting or following a main ejector
pump failure.
A fuel imbalance between the main tanks
may be corrected by opening a crossflow
valve. This allows the quantities in the main
tanks to equalize by gravity flow or by open-
ing a LEFT to RIGHT or RIGHT to LEFT
powered crossfeed shutoff valve to transfer
from the main to auxiliary tank.
0
2
4 6
8
10
MAIN
FUEL
LBS X 100
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CHAPTER 5
FUEL SYSTEM
CL 601-3R 5-1FOR TRAINING PURPOSES ONLY
The airplane may be refueled over the wings
by gravity. However, the normal method is
through an adapter located in the right wing
root using a single-point pressure system.
FUEL STORAGE
Using a wet-wing concept, the entire wing box
structure is sealed to form three tanks, which
carry most of the fuel (Figure 5-1). Two addi-
tional tanks are fitted under the cabin floor,
fore and aft of the auxiliary tank, which is in
the wing center section. These tanks are inter-
connected with the auxiliary tank. A tail tank
is fitted aft of the stabilizer rearspar.
The main tanks encompass the internal wing
volume from near the wingtip to near the wing
roots. There are 16 inspection and mainte-
nance access panels in the lower surface of
each wing. The gravity filler port for each
main tank is located on the upper outboard
wing surface.
The auxiliary tank encompasses the entire
center section of the wing. There are access
panels in the lower wing surface. The auxil-
iary tank gravity filler port is located in the
right wing root just aft of the leading edge.
Contained within the auxiliary tank are two
collector tanks which are extensions of each
main tank. They incorporate the main fuel
ejectors and feed fuel directly to each engine.
Each collector tank is constantly kept full
with fuel from its respective main tank. The
standby electric fuel pumps are supplied by
the collector tanks, but actually housed in the
auxiliary tank.
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CL-600-2B16 PILOT TRAINING MANUAL
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Figure 5-1. Fuel System—General Layout
COLLECTOR
TANKS
SINGLE-POINT
REFUEL-DEFUEL
ADAPTER
AUXILIARY TANK
FILLER CAP
AUXILIARY TANK
RIGHT MAIN
TANK
RIGHT MAIN
 FILLER CAP
LEFT MAIN
 TANK
LEFT MAIN
FILLER CAP
TAIL TANKLEGEND
MAIN TANK FUEL
AUXILIARY FUEL TANK
FUEL DISTRIBUTION
GRAVITY FLOW
Fuel flows by gravity within the main tanks
through lightening holes in the ribs. One inner
rib in each main tank is equipped with flap-
per valves to prevent outward flow of fuel
(Figure 5-2).
Fuel flows from the inboard sections of the
main tanks to the collector tanks (Figure 5-2).
Flapper valves at the collector tank inlets pre-
vent reverse flow of fuel into the main tanks.
SCAVENGE EJECTORS
Gravity flow to the collector tanks is sup-
plemented by scavenge ejectors located at
the rear of the inboard section of each main
tank. The scavenge ejectors ensure that the
collector tanks are supplied with fuel re-
gardless of airplane attitude. Failure of a
scavenge ejector will cause illumination of
an amber caution light in the cockpit.
Motive flow for operation of a scavenge ejec-
tor comes from the high-pressure side of the
two-stage engine-driven fuel pump (Figure
5-2). Ejectors have no moving parts. Each op-
erates on the venturi principle to convert
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 5-3FOR TRAINING PURPOSES ONLY
GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVEMAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-2. Fuel Distribution—Normal Operation
small-volume, high-pressure motive flow at
the throat of the ejector into large-volume,
low-pressure output at the ejector nozzle.
MAIN EJECTORS
Fuel is supplied from the collector tanks to the
low-pressure side of each engine-driven pump
by a main ejector located within the tank. Mo-
tive flow for operation of the main ejectors is
supplied by the high-pressure side of each en-
gine-driven fuel pump (Figure 5-3 and 5-4).
Each main ejector provides continuous fuel
flow to its own engine through firewall shut-
off valves. Flow to the opposite engine is not
possible because of one-way check valves in
the feed lines. Failure of a main ejector will
cause illumination of an amber caution light
in the cockpit, and will activate both standby
electric pumps as long as the standby pump on
the same side as the main ejector failure is se-
lected on.
TRANSFER EJECTORS
When main tank fuel quantity drops below
the 93% full level, float valves open, allow-
ing the transfer ejectors to draw fuel from the
auxiliary tank to the inboard sections of the
main tanks. Motive flow for the transfer ejec-
tors is provided by the output of the associated
main ejector. No cockpit indication of fuel
transfer or transfer ejector failure is provided.
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GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVE
MAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-3. Fuel Distribution—Engine Start
A one-way check valve in each transfer ejec-
tor prevents fuel migration from the main
tanks to the auxiliary tank.
STANDBY ELECTRIC PUMPS
Electric standby pumps are provided for en-
gine starting and as a backup in the event that
a main ejector becomes inoperative (see Fig-
ure 5-2). The two DC-powered pumps will
then operate simultaneously to draw fuel
from their respective collector tanks and feed
a common line capable of providing fuel to
either engine.
Once armed by cockpit switches, the standby
pumps operate automatically when the output
pressure of either main ejector falls below 10
psi. During the engine start sequence, both
pumps operate until the engine-driven pumps
generate enough motive flow to operate the
main ejectors.
The left electric pump is powered by the bat-
tery bus, while the right pump receives power
from DC bus No. 2.
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GRAVITY FLOW LINES
FLOAT
VALVE
FLAPPER
VALVES
SCAVENGE
EJECTOR
TO ENGINE TO ENGINE
FIREWALL
LOW
PRESSURE
SWITCH
LOW
PRESSURE
SWITCH
TO APU
FIREWALL SOV
APU SOV
OVERFLOW LINEPOWERED CROSSFEED
SHUTOFF VALVE
POWERED CROSSFEED
SHUTOFF VALVEMAIN
EJECTOR
CROSSFLOW
VALVE
MAIN
EJECTOR
OVERFLOW LINE
GRAVITY FLOW LINES
FLAPPER
VALVES
FLOAT 
VALVE
STANDBY
PUMPS
FLAPPER
VALVE
SCAVENGE EJECTOR/OVERFLOW PRESSURE
MOTIVE FLOW PRESSURE
LEGEND
MAIN EJECTOR PRESSURE
STANDBY PUMP PRESSURE
COLLECTOR
TANK
COLLECTOR
TANK
ENGINE
DRIVEN
PUMP
ENGINE
DRIVEN
PUMP
TRANSFER
EJECTOR
Figure 5-4. Fuel Distribution—Abnormal Operation, Main Ejector Fail
ENGINE-DRIVEN PUMPS
Each two-stage engine-driven fuel pump is a
single unit containing two pumps mounted in
piggyback fashion. The first stage is a cen-
trifugal, low-pressure pump which receives
fuel from the main ejector and supplies it to
the engine and the second stage, or high-pres-
sure side, of the pump. This second stage uses
a positive-displacement pump to supply high-
pressure motive flow fuel to the main and
scavenge ejectors (see Figure 5-2). 
Operation of the engine fuel system is dis-
cussed in Chapter 7, “Powerplant.”
CROSSFLOW VALVE
Should a main tank fuel imbalance occur in
flight for any reason, it can be corrected by
opening the crossflow valve (see Figure 5-2)
which connects both main tanks and both col-
lector tanks. Balance is achieved through grav-
ity flow only. To avoid a serious imbalance
which might occur if the wings are not level,
the crossflow should not be left open when the
airplane is on the ground.
POWERED CROSSFEED VALVE
When depressing the LEFT TO RIGHT or
RIGHT TO LEFT switchlight, the associated
powered crossfeed shutoff valve opens to allow
fuel flow, by gravity, into the auxiliary fuel
tank. Fuel is then transferred to the opposite
tank while a quantity is returned to its origi-
nal tank by the transfer ejectors when in flight
or the fuel boost pumps via the transfer ejec-
tors when on the ground.
NOTE
Maximum imbalance is 800 pounds.
APU FUEL SYSTEMThe APU is normally supplied with fuel from
the right main tank by an electric fuel pump.
The pump is identical to the standby electric
pumps described previously.
The APU fuel pump operates whenever it is se-
lected on by a switch on the APU control
panel. Fuel in excess of APU requirements is
routed back to the right main tank through a
fuel-oil heat exchanger which cools the APU
generator adapter oil (Figure 5-5). Operation
of the APU is discussed in Chapter 6, “Aux-
iliary Power Unit.”
The APU fuel feed line is fitted with two APU
fuel shutoff valves that are synchronized and
controlled by the APU electronic control unit.
To ensure uninterrupted operation of the APU
during brief moments of negative G flight or
in case of APU fuel pump failure, fuel can be
supplied from the left engine feed line to the
APU (Figure 5-6). This line has a differential
pressure, one-way check valve which opens
whenever the main APU supply pressure drops
10 psi lower than the pressure in the left en-
gine fuel feed line. Fuel from the left engine
feed line cannot flow to the right tank or to any
heat exchanger because of a check valve in-
stalled in the main feed line.
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CL 601-3R 5-7FOR TRAINING PURPOSES ONLY
RIGHT MAIN TANK
APU
FUEL
PUMP
STANDBY
PUMPSFROM
LEFT MAIN
EJECTOR
TO LEFT
ENGINE
APU NEGATIVE-G
CHECK VALVE
APU PUMP PRESSURE
MAIN EJECTOR PRESSURE
APU NEGATIVE-G
SHUTOFF VALVE
APU FUEL FEED
SHUTOFF VALVE
FUEL
CONTROL
UNIT
TO PNEUMATIC
SYSTEM
APU GENERATOR
OIL HEAT
EXCHANGER
LOAD
CONTROL
VALVE
RESTRICTOR
CHECK
VALVECHECK
VALVE
FUEL CONTROL
UNIT SHUTOFF VALVE
PRESSURE
SWITCH
CENTER TANK
LEGEND
Figure 5-5. APU Fuel System—Normal Operation
RIGHT MAIN TANK
APU
FUEL
PUMP
STANDBY
PUMPSFROM
LEFT MAIN
EJECTOR
TO LEFT
ENGINE
APU NEGATIVE-G
CHECK VALVE
STANDBY PUMP PRESSURE
MAIN EJECTOR PRESSURE
APU NEGATIVE-G
SHUTOFF VALVE
APU FUEL FEED
SHUTOFF VALVE
FUEL
CONTROL
UNIT
TO PNEUMATIC
SYSTEM
APU GENERATOR
OIL HEAT
EXCHANGER
LOAD
CONTROL
VALVE
RESTRICTOR
CHECK
VALVECHECK
VALVE
FUEL CONTROL
UNIT SHUTOFF VALVE
PRESSURE
SWITCH
CENTER TANK
LEGEND
Figure 5-6. APU Fuel System—Negative G Condition
FUEL CONTROLS
AND INDICATORS
GENERAL
The fuel controls and indicators are grouped
on the center instrument panel (Figure 5-7).
The fuel control panel is located just above the
fuel quantity panel.
The fuel control panel contains five switch-
lights, six additional annunciators, and a fuel
temperature gage. The fuel quantity panel
contains five digital readouts.
FUEL CONTROL PANEL
The standby electric fuel pumps are controlled
by a pair of switchlights labeled “PUMP.”
Pump operation is indicated by illumination
of the green ON legend in the top half of the
associated switchlight.
The bottom half of the PUMP switchlight has
an amber INOP legend which illuminates to in-
dicate that the associated pump is not selected
on or that the pump is not operating properly.
A third switchlight, labeled “X-FLOW,” con-
trols operation of the crossflow valve. The
green OPEN light illuminates to indicate that
the valve is fully opened. The light extin-
guishes when the valve is fully closed. Since
the valve is motor operated, expect a delay of
approximately 2 seconds from the time the
switchlight is pressed until the proper indi-
cation appears. The valve normally remains in
the closed position and is opened only during
flight to correct a fuel imbalance.
The green LEFT TO RIGHT and RIGHT TO
LEFT switchlights are part of the POWERED
CROSS FEED system. Should an imbalance
between main tank fuel levels develop and
gravity equalizing with the crossflow valve not
be possible, the transfer is possible by de-
pressing the appropriate switchlight. The de-
pressed switchlight will illuminate steady and
will start flashing after eight minutes as a
reminder. An interlock prevents simultane-
ous operation of both switchlights.
The remaining five lights for each engine pro-
vide information as follows:
• The amber SCAV light illuminates to in-
dicate an inoperative scavenge ejector.
• The amber MAIN light illuminates to
indicate an inoperative main ejector.
This automatically triggers both standby
electric pumps to operate if they are
selected to the ON position.
• The white VALVE CLOSED light illu-
minates to indicate that the firewall shut-
off valve has closed. Control of this
valve is from the respective FIRE PUSH
switchlight on the center glareshield.
• The amber FILTER light illuminates to
indicated an impending fuel filter bypass
or a clogged filter. (This condition is
covered in Chapter 7, “Powerplant.”)
• The amber LOW PRESS light illumi-
nates if fuel pressure at the inlet side of
the engine-driven pump falls below a
predetermined value.
• A fuel temperature indicator in the cen-
ter of the fuel control panel indicates
the temperature of the fuel as it leaves
the fuel heater at the fuel filter. (This is
also covered in Chapter 7, “Power-
plant.”) All amber lights on this panel
will activate the flashing master cau-
tion lights and illuminate the FUEL
annunciator light.
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Figure 5-7. Fuel Controls and Indicators
120
FUEL
80
40
0
-40
L
°C
DUMP
OPEN
DUMP
SELECT
DUMP AT 0° FLAPS ONLY
DUMP X-FER
ARMED
OFF
TAIL TANK
PUSH ON OFF
PUMP PUMPX-FLOW
E
J
C
T
F
E
E
D
L
E
F
T
E
N
G
F
U
E
L
E
N
G
F
U
E
L
FUEL CONTROL
SCAV
MAIN 
VALVE
CLOSED
FILTER 
LOW
PRESS
VALVE
CLOSED
FILTER 
NOT
ARMED
SEC
TRANS
LOW
PRESS
SCAV
MAIN 
ON
INOP 
ON
INOP 
OPEN
LEFT
TO
RIGHT
RIGHT
TO
LEFT
E
J
C
T
F
E
E
D
R
I
G
H
T
DUMP
SW
ARMED
120
80
40
0
-40
R
TOTAL
FUEL
QUANTITY
LB
AUX
L. MAIN R. MAIN
TAIL
FUEL QUANTITY
The fuel quantity in each main tank as well as
the auxiliary tank is measured by a system of
the capacitance-type transmitters located in
each tank. Quantity information is fed to a sig-
nal conditioner which displays the quantity in
pounds for each tank, as well as the total, on the
fuel quantity panel. Only usable fuel is shown.
The digital readouts are tested from the engine
instrument test switch on the center instru-
ment panel. A successful test is indicated by
the appearance of a series of eights in the dig-
ital readouts (except for the last digit which
is zero on all but the total readout wherein the
last two digits are always zero).
The probes for the fuel quantity are powered
as follows:
• L. MAIN ............... DC essential bus
• R. MAIN ........................ battery bus
• AUX and TAIL.................. DC bus 1
The fuel quantity indicators are powered via
the SDC as follows:
• RH, LH,
AUX, TAIL ......... DC essential bus
• TOTALIZER ................. battery bus
If power to the probes is lost, the affected
quantity indicator will read zero, and the
amount of fuel remaining in that tank will be
subtracted from the total.
VENT SYSTEM
Each tank is vented at two different points
through a series of vent lines which allow air
to enter or escape the tanks, depending upon
whether fuel is being used or added. (Figure
5-8.) The vent lines extend from each wingtip
to common manifolds which form an inter-
connected inverted “U” in each fuselage wall
and then return to the wing area where they
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AUXILIARY
TANK VENT
MAIN TANK
CLIMB VENT
MOTIVE
FLOW
PURGE
LINE
SCAVENGE FLOW
TO COLLECTOR TANK
MAIN TANK VENT
NACA SCOOP INLET
PURGE LINE
INVERTED U
VENT TUBE
SCAVENGE
EJECTOR
Figure 5-8. Vent System
terminate under the trailing edge on each side
in a flush-mounted NACA scoop. The scoop,
which has ice rejection capability, maintains
a slight positive tankpressure during flight due
to ram-air effect.
There are no valves or screens in the vent
lines, so dirt or ice accumulation does not nor-
mally occur. Any trapped fuel or moisture in
the vent lines is continuously purged from the
low points by a bleed line connected to each
scavenge ejector.
During pressure fueling, the fuel tank vent lines
are augmented with special vent valves which will
be described later under “Pressure Refueling.”
REFUELING
GENERAL
All tanks are normally fueled by means of the
single-point pressure adapter located in the
right wing root (Figure 5-9). The refueling
system is controlled from a swing-out control
panel located in the fillet above the right wing.
The system has automatic fuel cutoff to pre-
vent overfilling.
Overwing or gravity fueling is also possible.
However, due to the location of the filler caps
(Figure 5-9) , it is not possible to fill the main
tanks completely. A separate gravity filler
port is provided for each main tank and the aux-
iliary tank. The forward and aft tanks are grav-
ity fueled through the auxiliary tank.
Never open an overwing filler cap if
that main tank contains more than
4,000 pounds or if the level is not
known. Fuel in excess of 4,000 pounds
will spew from the filler if opened.
PRESSURE REFUELING
The pressure-refueling system (Figure 5-10)
consists of a single-point adapter, a pressure
manifold containing a two-way check valve,
three shutoff valves (SOV) associated with
three float-operated, full level-control valves,
and three vent valves that can be tested during
the refueling process and are utilized as backup
to normal pressure relief during refueling.
The adapter and manifold can accept a flow
rate of up to 250 gpm at a pressure of from 20
to 55 psi. The pressure-fueling process is con-
trolled from the fuel-defuel panel.
During a normal pressure-refueling opera-
tion, the sequence of events is as follows:
• Move the power switch to ON and check
that the green POWER ON light illu-
minates (Figure 5-11) (powered from
battery direct bus).
• Check that all three green vent valve
(VV) lights are extinguished and that
all four green SOV CLosed lights are
illuminated.
• Connect fueling nozzle from the truck
to the single-point adapter.
• Open the fuel nozzle valve and check
that with fuel pressure applied, all three
VV OPEN lights remain off.
• Rotate the MODE selector from OFF
to TEST.
• Move the tank FUEL-DEF switch of a
tank to be filled to the FUEL position and
check that the corresponding amber SOV
OPEN light illuminates
.• Check that the appropriate VV OPEN
light illuminates within 30 seconds.
CAUTION
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Figure 5-9. Exterior Fueling Components
AUXILIARY TANK GRAVITY
REFUELING PORT
PRESSURE REFUEL- DEFUEL
CONTROL PANEL
MAIN TANK
GRAVITY REFUELING PORT
PRESSURE REFUELING
ADAPTER
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TANK SOV
LEVEL
CONTROL
VALVE
VENT
TEST
VALVE *
VENT
(RELIEF)
VALVE *
VENT
TEST VALVES*
FUEL
MANIFOLD
SINGLE-POINT
ADAPTER
TWO-WAY
CHECK VALVE
VENT LINES
REFUELING PRESSURE
LEGEND
*NOTE
 ENERGIZED CLOSED DURING FUELING ONLY
Figure 5-10. Pressure-Refueling System
Figure 5-11. Refuel-Defuel Control Panel
• After 30–40 seconds, the SOV CLosed
light will illuminate and the OPen light
will extinguish , indicating that the au-
tomatic level control system is work-
ing properly.
• Test other SOVs and VVs individually
for other tanks to be filled.
• Rotate the MODE selector to FUEL.
• Check that all three VV OPEN lights
extinguish.
• Ensu re t ha t t he t ank FUEL-DEF
switches of the tanks to be filled are in
the FUEL position.
• Check that the SOV OPen lights are il-
luminated. Add fuel to the desired level.
• When the selected tanks are full, their
shutoff valves will close automatically
and the SOV CLosed lights will illuminate.
• Move the tank switches to OFF.
• Rotate the MODE selector to OFF.
• Remove the fueling nozzle from the
adapter, reinstall the cover, and close
the door carefully.
• Check that all lights and switches are off
and stow the refuel-defuel control panel.
NOTE
Adding fuel will compress the land-
ing gear shock struts and lower the
airplane. Be sure that stands, ladders,
or other equipment that might dam-
age the airplane are removed before
refueling.
Do not chock forward of the nose gear wheels
during refueling as increasing the weight will
move the nosewheels forward.
If tanks are to be only partially filled, tank
levels must be monitored from the cockpit
quantity panel. For this, AC electrical power
must be provided, usually from the APU or
external AC unit.
DEFUELING
The airplane may be completely defueled
without the use of electrical power via the
single-point adapter.
Fuel tender suction of negative 8 psi (max-
imum) at the manifold will open the tank
shutoff valves and withdraw fuel. When the
tanks are empty, the valves will close by
spring pressure.
The tanks may be selectively defueled using
the refuel-defuel control panel in the DEFUEL
mode as follows:
• Move the power switch ( see Figure 5-
10) to ON and check that the green
POWER ON light illuminates.
• Connect the defueling nozzle from the
tender to the single-point adapter.
• Open the nozzle valve and check for a
maximum negative pressure of 8 psi.
• Rotate the MODE selector to the DE-
FUEL position.
• Move the tank switch for tank(s) to be de-
fueled to the DEF position. This allows
suction to open the associated shutoff
valve, illuminating the amber OPen light.
To prevent fuel tank damage caused
by negative pressure, always open
overwing filler port for the tank
being defueled as soon as tank quan-
tity is less than 4,000 pounds. Do
not open filler cap if tank contains
more than 4,000 pounds.
CAUTION
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• When the tank quantity decreases to the
desired level, move the tank switch to
OFF. This will cause the shutoff valve
to close and the green CLOSED light
will illuminate.
• Rotate the mode selector to OFF.
• Move the power switch to OFF.
• Remove the defuel ing nozzle f rom
the adapter. Reinstall the cover and close
the door carefully.
• Check that all lights and switches are off
and stow the refuel-defuel control panel.
TAIL TANK
GENERAL
An additional fuel tank, which will increase
the fuel load by some 187.7 U.S. gallons (1,250
pounds), is located in the tail.
A primary fuel transfer and a backup, sec-
ondary transfer system will automatically trans-
fer fuel from the tail tank to the auxiliary tank.
Pressure refueling and defueling is possible
through the single-point refueling and defu-
eling panel.
The tail tank cannot be refueled unless the
auxiliary tank is full.
The tail tank is also equipped with a fuel dump
system, as well as a quantity indicator (Fig-
ure 5-12) in the cockpit.
The tail tank is located aft of the vertical sta-
bilizer rear spar. It includes two transfer pumps
(DC-powered), No. 1 and No. 2 tail tank
empty/level switches, fuel quantity transmit-
ter, dump valve, and tail tank fuel level switch.
It also has a vent system connected to the rear
air duct and exit flush at the bottom of the
fuselage via a flame arrester.
NORMAL TRANSFER
With the ARMED—OFF switch on the tail tank
fuel transfer panel (see Figure 5-7) set to
ARMED, the fuel transfers automatically to
the auxiliary tank at the rate of approximately
25 pounds per minute. When only unusable
fuel (approximately 1.3 U.S. gallons) is left, the
shutoff valve closes and the fuel transfer pump
stops.
If the ARMED—OFF switch is set to OFF
while there is still usable fuel in the tail tank,
the NOT ARMED light illuminates. If fuel in
the auxiliary tank reaches the level of the No.
2 auxiliarytank level switch, the NOT ARMED
light flashes. Placing the ARMED—OFF
switch to the ARMED position starts the nor-
mal transfer, and the NOT ARMED light goes
out. If the NOT ARMED light remains on,
jettison fuel.
In the event of a normal fuel transfer failure,
the SEC TRANS light comes on, shutting
down the primary and starting the secondary
transfer system. The SEC TRANS light indi-
cates that secondary transfer is taking place at
approximately 25 pounds per minute. Moni-
tor fuel quantity during secondary transfer. If
adequate transfer cannot be confirmed with a
maximum of ten minutes, jettison fuel.
The fuel transfer line has a nitrogen-pressur-
ized shroud in the rotor boost zone. A perfo-
ration in this shroud signals all open shutoff
valves to close and shuts down the boost pump.
Fuel transfer cannot be accomplished unless
the shroud is replaced and pressurized. Fuel
in this case has to be jettisoned.
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TOTAL
FUEL
QUANTITY
LB
AUX
L. MAIN R. MAIN
TAIL
Figure 5-12. Tail Tank Quantity Panel
Figure 5-13 shows the location of the con-
trols and lights. Figure 5-14 has the schematic
of the tail tank system.
TAIL TANK AMBER
WARNING LIGHTS VERSUS
MASTER CAUTION
A steady NOT ARMED light indicates that
there is fuel in the tail tank but the arming
switch has not been selected to ARMED. This
will not turn on the MASTER CAUTION.
A flashing NOT ARMED light indicates that
the arming switch is still in the OFF position
and the tail tank is ready to transfer fuel (no
MASTER CAUTION light yet).
After a short while, if the arming switch is still
not selected to the ARMED position, the Mas-
ter Caution system FUEL annunciator will il-
luminate and the MASTER CAUTION lights
will flash. This will occur whenever fuel is sup-
pose to be transferring from the tail tank and
it’s not.
NOTE
NOT ARMED light wil l i l lumi-
nate if the Rotor Burst Detection
Switch in the pressurized shroud
senses less than 8.5 psi.
FUEL JETTISON
Pressing the DUMP SW ARMED switchlight
causes it to illuminate green, the white DUMP
SELECT light illuminates and arms the DUMP
SELECT/DUMP OPEN guarded switchlight.
Once armed, pressing the green DUMP OPEN
switchlight causes it to illuminate and opens
the jettison valve. Fuel discharges at a rate of
100 pounds per minute.
NOTE
Fuel jettison must only be carried
out with flaps set at 0°. If an imme-
diate landing is required, the fuel
jettison procedure should be initi-
ated immediately. However is is not
possible to complete the jettison pro-
cedure, a landing may be made with
fuel in the tail cone tank.
Do not jettison fuel in known light-
ning conditions.
REFUELING AND DEFUELING
Refueling can only be accomplished if the
auxiliary tank is full. The refuel-defuel panel
is equipped with a tail tank refuel-defuel
switch. When placed to FUEL, the green CL
light remains on until the auxiliary is full, and
then the amber OP light comes on, indicating
the refueling is taking place.
Defueling can be carried out any time, regard-
less of fuel quantity. There is no provision for
gravity refueling.
WARNING
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DUMP
OPEN
DUMP
SELECT
DUMP AT 0° FLAPS ONLY
DUMP X-FER
ARMED
OFF
TAIL TANK
NOT
ARMED
SEC
TRANS
DUMP
SW
ARMED
Figure 5-13. Tail Tank Fuel Transfer Panel
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REFUELING PRESSURE
LEGEND
RAM AIR
NO. 2 AUX TANK
LEVEL SWITCH
NO. 1 AUX TANK
LEVEL SWITCH
FUEL LEVEL
SWITCH
FWD
AUX TANK
PRIMARY
AUX
TANK
REAR
AUX
TANK
TAIL
TANK
LEFT
MAIN
TANK
RIGHT
MAIN
TANK
FUELING ADAPTER
AND MANIFOLD
ROTORBURST
ZONE
TRANSFER
LINE
TRANSFER
SOV
WATER DRAIN
FULL
LEVEL
SWITCH
NO. 1 TAIL TANK
LEVEL SWITCH
NO. 2 TAIL TANK
LEVEL SWITCH
JETTISON
SOV
RAM AIR
SECONDARY
TRANSFER PUMP
SECONDARY
TRANSFER SOV
PRESSURIZED
SHROUD
FUELING
LINE
FLOW
SENSOR
Figure 5-14. Tail Tank Flow Schematic
1. The engine-driven fuel pumps are nor-
mally supplied with fuel by the:
A. Main ejectors
B. Standby electric pumps
C. Scavenge ejectors
D. Transfer ejectors
2. If the main ejectors fail, the engine-driven
fuel pumps will be supplied with fuel by:
A. Gravity
B. Scavenge ejectors
C. Standby electric pumps
D. Transfer ejectors
3. Fuel imbalance between the main tanks
is corrected by:
A. Transfer ejectors
B. Scavenge ejectors
C. Standby pumps
D. Gravity crossflow
4. The scavenge ejectors:
A. Transfer fuel from the auxiliary tanks
to the main tanks.
B. Provide a flow of fuel from the main
tanks to the collector tanks.
C. Supply fuel to the engine-driven
pumps.
D. Transfer fuel from the main tanks to
the auxiliary tank and correct a fuel
imbalance when the crossflow valve
is open.
5. Flapper valves in the main tanks prevent:
A. Reverse flow of fuel from the main
tanks to the auxiliary tank
B. Reverse flow of fuel from the auxil-
iary tank to the main tanks
C. Gravity flow of fuel from the out-
board main tank sections to the in-
board main tank section
D. Reverse flow of fuel from the inboard
main tank sections to the outboard
main tank sections
6. Fuel for the APU is normally supplied by:
A. The left scavenge ejector
B. An electric pump in the right main
tank
C. An electric pump in the left main tank
D. The right main ejector
7. In case of a negative G condition, fuel for
the APU is supplied by:
A. The left main ejector
B. An electric pump in the right collec-
tor tank
C. An electric pump in the left collector
tank
D. All the above
8. The pump switchlights on the fuel con-
trol panel control the:
A. APU fuel pump
B. Main ejectors
C. Standby electric pumps
D. Differential pressure check valve
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QUESTIONS
9. When an amber MAIN light illuminates
on the fuel control panel:
A. The firewall shutoff valves close au-
tomatically.
B. The standby electric pumps start op-
erating.
C. The crossflow valve opens.
D. The associated engine will flame out.
10. The preferred method of refueling is:
A. Gravity fueling when a full fuel load
is required
B. Single-point pressure refueling
C. Single-point pressure refueling of
main tanks only
D. Single-point pressure refueling; how-
ever, the wing tanks must be topped
off by overwing fueling.
11. The maximum pressure for single-point
refueling is:
A. 35 psi
B. 45 psi
C. 55 psi
D. 65 psi
12. Never open the overwing filler caps if
the fuel level is unknown or if the tanks
contain more than:
A. 3,000 pounds
B. 4,000 pounds
C. 3,000 gallons
D. 4,000 gallons
13. The maximum negative pressure allowed
for single-point defueling is:
A. 6 psi
B. 8 psi
C. 10 psi
D. 12 psi
14. During flight, the maximum imbalance al-
lowed between the main tanks is:
A. 800 pounds
B. 1,836 pounds
C. 2,500 pounds
D. 2,836 pounds
15. The Ta i l Tank sy s t em ambe r NOT
ARMED light, if on steady:
A. Indicated no fuel in the tank
B. Indicates fuel is in the tail tank and the
transfer system should be armed by
the crew.
C. Will produce a Master Caution indi-
cation.
D. The tail tank fuel will transfer auto-
matically when flashing.
16. To dump fuel:
A. Two switchlights need to be pressed.
B. Three switchlights need to be pressed.
C. The DUMP SELECT/DUMP OPEN
switchlight must first be armed.
D. Both A and C are good answers.
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CL 601-3A/R 6-i
CHAPTER 6
AUXILIARY POWER UNIT
CONTENTS
Page
INTRODUCTION ................................................................................................................... 6-1
GENERAL............................................................................................................................... 6-1
AUXILIARY POWER UNIT (APU).......................................................................................6-2
General ............................................................................................................................. 6-2
Major Sections ................................................................................................................. 6-2
APU SYSTEMS ...................................................................................................................... 6-4
Lubricating System .......................................................................................................... 6-4
Fuel System...................................................................................................................... 6-4
Ignition System ................................................................................................................ 6-5
Instrumentation ................................................................................................................ 6-5
Bleed-Air Control System................................................................................................ 6-6
APU Protection System ................................................................................................... 6-6
APU Starting System ....................................................................................................... 6-8
APU Cold Weather Starting ............................................................................................. 6-8
APU Normal Shutdown ................................................................................................... 6-9
QUESTIONS......................................................................................................................... 6-10
FOR TRAINING PURPOSES ONLY
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CL 601-3A/R 6-iii
ILLUSTRATIONS
Figure Title Page
6-2 APU Installation ....................................................................................................... 6-2
6-2 APU Engine Cutaway............................................................................................... 6-3
6-3 APU Air Intake......................................................................................................... 6-3
6-4 APU Exhaust Outlet ................................................................................................. 6-4
6-5 APU Control Panel ................................................................................................... 6-5
6-6 APU Remote Indicating Panel ................................................................................. 6-7
6-7 Auxiliary Battery Panel ............................................................................................ 6-8
6-8 Auxiliary Battery Panel (SB 601-0418) ................................................................... 6-9
6-9 Electrical Control Panel............................................................................................ 6-9
FOR TRAINING PURPOSES ONLY
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CL-600-2B16 PILOT TRAINING MANUAL
INTRODUCTION
This chapter deals with the auxiliary power unit (APU) installed as standard equipment
on the Canadair Challenger CL-600-2B16, model CL-601-3A/R. The APU renders the
airplane completely independent of such ground servicing requirements as electrical power,
pneumatic power for engine starting, and an air supply for the environmental systems.
GENERAL
The APU is a small, lightweight, gas-turbine
engine, certificated for ground and in-flight op-
eration. The unit is manufactured by Garrett
Turbine Engine Company and designated
“GTCP-36-100(E).”
The APU is equipped with self-contained oil,
fuel, and ignition systems. During starting
and operation, the APU is monitored by an
electronic control unit (ECU). If any primary
parameters are exceeded, the ECU will auto-
matically shut down the APU.
Separate fire protection, consisting of a mon-
itored fire detection system and a fire extin-
guishing system, is provided for the APU.
The APU’s only dependencies are (1) an elec-
trical power source for starting from either
the airplane battery or from an external DC
power unit and (2) a fuel supply from the air-
plane’s main fuel system.
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CHAPTER 6
AUXILIARY POWER UNIT
CL 601-3A/R 6-1FOR TRAINING PURPOSES ONLY
AUXILIARY POWER
UNIT (APU)
GENERAL
The APU (Figure 6-1) is mounted on a skid
support assembly enclosed in a fireproof metal
container in the aft equipment bay.
MAJOR SECTIONS
The APU (Figure 6-2) can be divided into six
major sections as follows:
1. Air intake
2. Compressor
3. Combustor
4. Turbine
5. Exhaust
6. Accessory Gear
Air Intake
The main air inlet to the compressor is located
around the waist of the APU. The air intake is
shrouded and connected by a rectangular duct
to a flush screened inlet (figure 6-3) at the top
of the rear fuselage compartment.
Compressor
The compressor is a single-stage centrifugal
type, which induces air through the intake,
compresses the air, and directs the airflow for
cooling, combustion, and bleed air extraction.
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Figure 6-1. APU Installation
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DC STARTER
MOTOR
DRIVE
PADS
FUEL
CONTROL
TURBINE
IGNITER
COMBUSTOR
EXHAUST
LOAD CONTROL VALVE
COMPRESSOR
AIR INTAKE
GEARBOX
Figure 6-2. APU Engine Cutaway
Figure 6-3. APU Air Intake
Combustor
The combustor consists of a single combustion
chamber. The chamber is designed to provide
for the proper mixing of air and fuel and con-
tainment of the combustion gases.
Turbine
A single-stage radial turbine is rigidly mounted
on the compressor rotor shaft. The turbine is
designed to extract almost all the energy from
the expanding combustion gases. The major
portion of the energy is used to drive the com-
pressor and the accessory gearbox.
Exhaust
The exhaust consists of a stub exhaust pipe ex-
tending into an exhaust ejector connected to the
atmosphere through an outlet on the right side
of the rear fuselage (Figure 6-4). A flapper
door on the left side of the APU enclosure pro-
vides air circulation for cooling. This air is ex-
hausted with the APU fumes through the ejector.
Accessory Gear
The accessory gear forms an integral part of
the APU. It converts high turbine rpm to the
speeds required for the APU accessories
which include:
• APU lubricating pump
• APU fuel control unit
• AC generator adapter unit
APU SYSTEMS
The APU systems consist of the following:
• Lubricating system
• Fuel system
• Ignition system
• Instrumentation
• Bleed-air control system
• APU protection system (ECU)
• APU starting system
LUBRICATING SYSTEM
The APU lubricating system is a self-con-
tained wet-sump system. The oil system is
fully automatic, providing for lubrication of
the APU rotor bearings and the planetary gear.
Oil quantity can be checked using the dip-
stick installed in the filler cap accessible
through the APU service door.
FUEL SYSTEM
The APU fuel system is a self-contained, high-
pressure, fully automatic system. The fuel
system consists of a high-pressure pump, a
fuel controller, a solenoid shutoff valve, a
flow divider, and a duplex spray nozzle.
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Figure 6-4. APU Exhaust Outlet
Fuel metering is controlled by a torque motor,
operating in response to input signals from
the electronic control unit. Fuel is metered to
the combustor so that the power developed by
the APU is equal to the power required, thereby
maintaining a near constant rpm under vary-
ing load conditions.
(For APU fuel supply see Chapter 5, “Fuel
System.”)
IGNITION SYSTEM
The APU incorporates a high-energy ignition
system consisting of an APU-mounted ignition
exciter and a single igniter plug in the com-
bustorchamber. The ignition system is fully
automatic and operates in response to inputs
from the electronic control unit. The ignition
system is activated by the ECU at 10% rpm dur-
ing APU starting and remains operative up to
95% rpm.
INSTRUMENTATION
APU instrumentation consists of rpm and EGT
indicators (Figure 6-5) located on the APU
control panel on the overhead panel, and an
hourmeter located within the APU enclosure
in the rear fuselage compartment. The rpm
gage is calibrated in percentage of design
100% rpm, from 0% to 110%.
The rpm signal is generated by a monopole sen-
sor installed on the planetary gearbox. The
monopole transmits the rpm signal to the elec-
tronic control unit, which in turn supplies the
RPM indicator.
The EGT indicator is calibrated from 0 to 10
in degrees Celsius times 100. EGT, or exhaust
gas temperature, is sensed by a thermocouple
in the APU exhaust duct. EGT signals are sent
to the electronic control unit, which in turn
supplies the EGT indicator.
The hourmeter is located within the APU en-
closure and controlled by the ECU to record
all APU operating time above 10% rpm.
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C
O
N
T
R
O
L
A
P
U
PUSHPUSH PUSH
PWR FUEL
ON/OFF
APU
OIL
ADPTR
OIL
BLEED
AIR
START/
STOP
STARTER
APU
READY
LO
PRESS
HI
TEMP
SOV
CLOSED
PUMP
INOP
LO
PRESS
HI
TEMP
FAILED
OPEN
%RPM
100
80 0
60
40
20
EGT
°C X 100
8
6 2
4
0
10
Figure 6-5. APU Control Panel
BLEED-AIR CONTROL SYSTEM
General
The bleed-air control system consists of an
electropneumatic surge valve and an elec-
tropneumatic load control valve.
Surge Valve
The surge valve prevents compressor stalls
and surges during operation when APU bleed
air is not being used. The surge valve is opened
by a signal from the ECU at 10% rpm during
starting and remains open unless the APU
bleed-air switchlight (Figure 6-5) is latched
in when APU rpm is above 95%.
Load Control Valve (LCV)
The LCV is the APU bleed-air valve which
connects the APU air plenum to the airplane’s
pneumatic system.
The ECU arms the LCV for operation when
rpm is above 95%. Pushing to latch the APU
bleed-air switchlight (Figure 6-5) will simul-
taneously open the LCV and close the surge
valve. The LCV is monitored by the ECU when
open, and, if operating EGT exceeds design
limits, the ECU will signal the LCV to the
closed position and the surge valve to the open
position. The valves will remain in these posi-
tions until the EGT drops and the ECU allows
the valves to resume their selected positions.
To prevent hot engine bleed-air feedback into
the APU, an electrical interlock circuit will
prevent LCV opening if either or both of the
following conditions exist:
• The left engine 10th stage bleed-air
switchlight is pushed in.
• The right engine 10th stage bleed-air
switchlight and the bleed-air ISOL
switchlight are both pushed in.
An amber FAILED light works through the
electrical interlock circuit. The FAILED light
will illuminate after a five-second time delay
if the load control valve has not closed and one
of the following conditions exists:
• The left engine 10th stage bleed-air
switchlight is pushed in.
• The right engine 10th stage bleed-air
switchlight and the bleed-air ISOL
switchlight are both pushed in.
For more information on APU bleed-air con-
trol, see Chapter 9, “Pneumatics.”
APU PROTECTION SYSTEM
The APU protection system provides for
automatic shutdown for any of the follow-
ing conditions:
• APU overspeed—in excess of design
limit (110%)
• High EGT—in excess of design limit
(732° C) (or an open or shorted ther-
mocouple)
• Overcurrent—excess current demand
by the ECU or any circuit controlled by
the ECU
• APU low oil pressure—APU oil pressure
below the minimum design limit when
APU rpm is above 95%
• APU high oil temperature—APU oil
temperature in excess of design limit
for 1 second
• Generator adapter low oil pressure—
generator adapter oil pressure below
minimum design limit for 10 seconds
• Generator adapter high oil tempera-
ture—generator adapter oil temperature
above design limit
• APU OVERHEAT or fire—Overheat or
fire detected by the APU fire sensing
loop
NOTE
Pushing the APU FIRE PUSH switch
will also cause an APU shutdown.
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Remote Fault Panel
APU malfunction indication is provided on a re-
mote fault panel (figure 6-6) externally located
in a covered access on the aft left side of the rear
fuselage. The panel contains APU fault-indi-
cating flags for OVERSPEED, OVERCUR-
RENT, EGT HIGH, OIL PRESSure LOW, OIL
TEMPerature HIGH, and GENerator adapter
fault indication for OIL PRESSure LOW and
OIL TEMPerature HIGH. The normal flag in-
dication is black; the malfunction indication is
black and white triangles.
An automatic APU shutdown will cause the ap-
propriate flag to trip.
In addition to the malfunction flags, the remote
indicating panel has two push button switches,
one labeled “INDicator RESET” and the other
labeled “APU STOP.” Pushing the INDicator
RESET button with power on the battery bus
will reset any tripped flags, even if the fault
has not been corrected.
The APU STOP pushbutton, when pushed, in-
troduces a false overspeed signal to the ECU,
and the APU automatically shuts down.
NOTE
When a remote APU shutdown is
used, the APU START–STOP and
PWR–FUEL switchlights must be
pushed out to reset the ECU.
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APU FAULT
OVER-
SPEED
OVER
CURRENT
EGT
HIGH
OIL
PRESS LOW TEMP HIGH
PRESS LOW TEMP HIGH
GEN ADAPTER OIL
IND RESET APU STOP
Figure 6-6. APU Remote Indicating Panel
APU STARTING SYSTEM
The APU must not be started until the
exterior check as outlined in the
Walkaround Checklist and the APU
Start Checklist are completed and
all discrepancies corrected.
NOTE
The following is a description of
APU starting, operation, and shut-
down. It is not intended to be used
as a procedure.
After turning on the BATTERY MASTER switch,
check the APU control panel (Figure 6-5). The
white SOV CLOSED light should be illuminated.
All other lights should be extinguished.
Push in the PoWeR–FUEL ON–OFF switch-
l i gh t and check t ha t t he ambe r PUMP
INOPerative light illuminates momentarily
and the white SOV CLOSED light extin-
guishes. The APU fuel pump in the right main
fuel tank is now running at design pressure,
and the START–STOP switchlight is armed.
Push in the START–STOP switchlight and
check to see that the green STARTER light il-
luminates. The ECU is now powered, the
starter is increasing APU rpm, and the APU
fuel supply valve and the negative G fuel sup-
ply valve are both open. Monitor the APU rpm
indicator; at 10% rpm the ECU will open the
APU fuel solenoid valve and the surge valve
and activate the APU ignition. The EGT will
start to increase and the RPM indicator will
show smooth acceleration. At 60% RPM, the
ECU will deactivate the starter and the green
STARTER light will extinguish.
Acceleration will continue; at 95% rpm the
ECU will deactivate the ignition; arm the load
control valve, the ADaPTeR, the APU LO
PRESSure, and HI TEMPerature lights; and il-
luminate the APU GENerator OFF light (Figure
6-7). Four seconds later, the green APU ReaDY
light will illuminate. The rpm and EGT will
stabilize at the steady-state, no-load values.
The APU is now fully operational and ready
to supply electrical and/or pneumatic loads.
The ECU is monitoring all automatic shut-
down parameters.
NOTE
The APU will normally start and
accelerate to 100% rpm in less than
60 seconds. If the start cycle ex-
ceeds 60 seconds, the ECU will ini-
tiate a shutdown.
APU COLD WEATHER 
STARTING
On aircraft 5001 to 5134, a dedicated battery
is provided to power the APU ECU during
cold weather starts. The auxiliary battery has
a selfcontained charger and heater and is lo-
cated inthe aft equipment bay. An indicator
is located on the copilot’s side control (Figure
6-7). Two circuit breakers through which
power is supplied to the power/heater are lo-
cated on the DC No. 1 and the battery bus.
There are various relays to activate the system.
A part of this system also opens the APU fuel
shutoff valve when the PWR FUEL switchlight
is activated.
WARNING
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IRS
1
FAIL
ON
IRS
2
AUXILIARY BATTERY
APU BATT/CHARGER FAIL
IRS
3
Figure 6-7. Auxiliary Battery Panel
On aircraft with SB 601-0418 and serial num-
bers 5135 and subsequent, a backup battery is
provided to power the APU ECU during cold
weather starts. While the APU ECU is pow-
ered from the battery bus, it also receives
power from the avionics IRS No. 2 system’s
backup battery (Figure 6-8). During cold
weather APU starts, the output level from the
aircraft main battery might not be sufficiently
high to maintain ECU operation. The backup
battery will ensure a successful start.
APU NORMAL SHUTDOWN
The APU is normally shut down by:
1. First turning off the APU generator
switch (Figure 6-9) and, without fur-
ther delay,
2. Push the START-STOP switchl ight
(Figure 6-5) . The APU spooldown
should be monitored on the rpm indica-
tor until it reaches zero. The spooldown
should be smooth.
3. Press out APU BLEED AIR switchlight. 
4. Press out APU POWER FUEL ON/OFF
switchlight.
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Figure 6-9. Electrical Control Panel
IRS
1
IRS
2
AUXILIARY BATTERY
APU BATT/CHARGER FAIL
IRS
3
Figure 6-8. Auxiliary Battery Panel
(SB 601-0418)
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1. With the BATTERY MASTER switch
on, pushing in the PWR– FUEL switch-
light will:
A. Turn on all switchlights and annun-
ciators on the APU control panel.
B. Start the APU fuel boost pump and
arm the START–STOP switchlight.
C. Open the APU fuel supply valve and
the APU negative G fuel supply valve.
D. Both B and C are correct.
2. APU ignition is activated when the:
A. APU rpm reaches 10% during starting.
B. PWR–FUEL switchlight is pushed in.
C. START–STOP switchlight is pushed in.
D. A P U f u e l b o o s t p u m p p r e s s u r e
switch opens.
QUESTIONS
CL 601-3A 7-i
CHAPTER 7
POWERPLANT
CONTENTS
Page
INTRODUCTION ................................................................................................................... 7-1
GENERAL............................................................................................................................... 7-1
ENGINES ................................................................................................................................ 7-2
General ............................................................................................................................. 7-2
Major Sections ................................................................................................................. 7-2
Operation.......................................................................................................................... 7-4
ENGINE SYSTEMS ............................................................................................................... 7-5
Engine Oil System ........................................................................................................... 7-5
Engine Fuel System ......................................................................................................... 7-7
Ignition System.............................................................................................................. 7-10
Engine Power Control.................................................................................................... 7-13
Engine Instrumentation.................................................................................................. 7-15
Engine Starting .............................................................................................................. 7-17
Engine Speed Control and APR Systems ...................................................................... 7-21
Engine Vibration-Monitoring System............................................................................ 7-25
THRUST REVERSERS ........................................................................................................ 7-26
General........................................................................................................................... 7-26
Control ........................................................................................................................... 7-26
Indication ....................................................................................................................... 7-28
Protection ....................................................................................................................... 7-28
Operation ....................................................................................................................... 7-29
QUESTIONS......................................................................................................................... 7-30
FOR TRAINING PURPOSES ONLY
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ILLUSTRATIONS
Figure Title Page
7-1 CF34 Major Sections and Gas Flow......................................................................... 7-2
7-2 CF34 Engine Modules.............................................................................................. 7-3
7-3 Oil-Replenishing Control Panel ............................................................................... 7-5
7-4 Junction Box 4 (JB4)................................................................................................ 7-6
7-5 Oil Pressure and Temperature Indicators ................................................................. 7-7
7-6 Oil System Schematic .............................................................................................. 7-8
7-7 Fuel System Schematic .......................................................................................... 7-11
7-8 Start and Ignition Control Panel............................................................................. 7-12
7-9 Throttle Quadrant ................................................................................................... 7-13
7-10 Ignition System Schematic..................................................................................... 7-14
7-11 Engine Instruments................................................................................................. 7-15
7-12 Engine Instrument Control Panel ........................................................................... 7-16
7-13 APU Control Panel................................................................................................. 7-16
7-14 Bleed-Air Control Panel......................................................................................... 7-17
7-15 External Air Supply Adapter.................................................................................. 7-19
7-16 Bleed-Air Sources (First Engine Start Schematic)................................................. 7-19
7-17 Cross Bleed Start (Left Engine from Right Engine Schematic) ............................ 7-19
7-18 Maximum Allowable Start Time and Time to Stabilize Idle—Seconds ................ 7-20
7-19 Airstart Envelope.................................................................................................... 7-21
7-20 APR Control Panel ................................................................................................. 7-21
7-21 APR/Engine Speed Schematic ...............................................................................7-24
7-22 Engine Vibration-Monitoring Panel (AC 5001-5104)............................................ 7-25
7-23 Engine Vibration-Monitoring Panel (AC 5105 and Subsequent)........................... 7-26
7-24 Thrust Reversers..................................................................................................... 7-27
7-25 Reverse Thrust Control Panel................................................................................. 7-28
7-26 REVERSER UNLOCKED Lights ......................................................................... 7-28
FOR TRAINING PURPOSES ONLY
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INTRODUCTION
This chapter describes the powerplant of the Canadair Challenger CL-600-2B16, model
CL-601-3A.
In addition to the basic powerplant information, the chapter also includes information
on all powerplant-related systems, such as engine oil, engine fuel, ignition, engine
power control, instrumentation, engine starting, engine speed control and APR sys-
tems, engine vibration monitoring, and thrust reversing.
GENERAL
The Canadair Challenger CL-601-3A is pow-
ered by two aft-fuselage-mounted turbofan
engines manufactured by the General Electric
Company.
The engines are modular in design to facilitate
maintenance and reduce airplane downtime.
Each engine incorporates self-contained oil,
fuel, and ignition systems in addition to a fire
and/or overheat detection system. A fire-ex-
tinguishing system is common to both en-
gines. Pneumatically operated cascade thrust
reversers are standard equipment.
Each engine is monitored during takeoff by an
electronically controlled automatic perfor-
mance (power) reserve system (APR). It will
automatically increase the permissible tem-
perature limits and thrust on the operating en-
gine if a power loss or failure occurs on the
opposite engine.
#1 DC
GEN 
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CHAPTER 7
POWERPLANT
CL 601-3A 7-1FOR TRAINING PURPOSES ONLY
ENGINES
GENERAL
The engines (Figure 7-1) on the Canadair
Challenger are GE Series CF34. This engine
has a high bypass ratio (6.2 to 1). The CF34
3A or 3A2 version of this engine is capable of
producing 8,729 pounds of takeoff thrust up
to 21°C (70°F) under standard sea level static
conditions. If a power loss occurs on one en-
gine, the APR system will automatically in-
crease the thrust on the operating engine to
9,220 pounds.
Modular Concept
For ease of maintenance, assembly, and dis-
assembly, the engine is designed in seven sep-
arate modules (Figure 7-2). Some of these
modules can be removed and replaced with-
out engine removal from the airplane.
MAJOR SECTIONS
For the purpose of this chapter the engine will
be discussed under seven major sections:
1. Air inlet section
2. Fan section
3. Compressor
4. Combustor
5. Turbine
6. Exhaust
7. Accessory gear
Air Inlet Section
The nacelle fairing forms the main air inlet at
the front of the engine fan section.
Fan Section
The single-stage fan and integral two-piece nose
cone are installed in the front frame. The fan is
basically the low-pressure (LP) compressor of
the engine in conjunction with a row of stators
mounted in the front frame aft of the fan.
Air entering the engine air inlet is divided into
two flow paths aft of the fan; one path directs
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FAN
SUPPORT
COMPRESSOR
COMBUSTOR
TURBINE
HIGH PRESSURE
SHAFT
ACCESSORY GEARINLET GUIDE VANE
AIR INLET
N1
N1
LP SHAFT
N2N2
EXHAUST
HP LP
Figure 7-1. CF34 Major Sections and Gas Flow
air to the compressor of the core engine and the
second path directs air into the fan bypass duct.
The fan functions to accelerate a large air mass
to a moderate velocity through the bypass duct
and contributes approximately 80% of the total
thrust developed by the CF34 engine.
Compressor
The high-pressure (HP) compressor is located
aft of the front frame. This single-spool axial
compressor has 14 stages with a pressure ratio
of 14:1.
The HP compressor supplies air for cooling,
bearing and seal pressurization, hot-point
cooling, and for combustion in the core engine.
In addition, it supplies bleed air for the air-
plane’s pneumatic services.
A variable-geometry system automatically
controls the inlet guide vanes and five variable
stator vanes to maintain a safe surge margin
across the HP compressor. This helps prevent
compressor stall or surges when the engine is
operating in the low-power range or during
rapid acceleration or deceleration.
The variable-geometry system is controlled by
the fuel control unit (FCU) as a function of HP
compressor rpm (N2) and core-inlet temper-
ature (T2). The FCU includes a fuel tempera-
ture compensating system to maintain the
required variable-geometry accuracy through-
out the normal fuel temperature range. The
variable-geometry module in the FCU will di-
rect HP fuel to two variable-geometry actua-
tors to operate the inlet and stator vanes.
A feedback system relays the position of the
vanes to the FCU at all times. When the en-
gine is static and during steady-state operation
at lower power, the inlet guide vanes and the
variable stator vane are at a close position.
This res t r ic ts the airf low to the HP com-
pressor to an amount that will ensure smooth
and continuous stall-free flow through the
compressor. As compressor rpm increases
with the addition of power, the variable-geo-
metry system moves the inlet guide vanes and
the variable stator vanes to the open position,
allowing unrestricted airflow through the com-
pressor. The response of this system will en-
sure a safe surge margin for the compressor
throughout its operating envelope.
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ACCESSORY
GEARBOX
RADIAL
DRIVE
SHAFT
FAN DRIVE
SHAFT
FAN
DRIVE
SHAFT
POWER
TAKEOFF
ASSEMBLY
LOW-PRESSURE
TURBINE SECTION
HIGH-PRESSURE
TURBINE SECTION
COMPRESSOR
SECTION
COMBUSTION
SECTION
FRONT FRAME
FAN SECTION
Figure 7-2. CF34 Engine Modules
Combustor
The combustor includes a straight-flow an-
nular combustion chamber, a liner, and the
first-stage turbine inlet nozzle.
Eighteen swirl injectors are installed in the
combustion chamber to atomize the fuel.
Initial ignitions supplied by two high-energy
ignitor plugs. The combustor system ensures
proper mixing of the air-fuel mixture, air di-
lution, and flame containment.
Turbine
The turbine section consists of a HP and
LP turbine.
The two-stage HP turbine is rigidly connected
to the HP compressor by the main rotor shaft.
The turbine extracts sufficient energy from
the expanding gases to drive the HP com-
pressor and the accessory gearbox.
The HP compressor and HP turbine assem-
blies form the HP spool of the engine. The rpm
of the HP spool is designated N2.
The four-stage LP turbine located behind the HP
turbine is rigidly connected to the single-stage
forward fan by a shaft that passes through the
main rotor shaft. The energy extracted by the
LP turbine is used to drive the fan. The remain-
ing energy in the combustion gases is accelerated
rearward to the atmosphere as the core engine’s
contribution to the total engine thrust.
The fan and LP turbine combination constitute
the LP spool. The rpm of the LP spool is des-
ignated N1.
Exhaust
The exhaust frame is located aft of the LP turbine
and consists of an exhaust duct and cone assem-
bly. The exhaust system directs the combustion
gases from the core engine to the atmosphere.
Accessory Gearbox
The accessory gearbox is attached to the lower
side of the front frame. The gearbox is driven
by a tower shaft and bevel gear assembly from
the main (HP spool rotor shaft. The following
accessories are driven by the accessory gearbox:
• N1 speed control alternator
• Integrated-drive generator
• Oil pump
• Fuel pumps and FCU
• Hydraulic pump
In addition to these accessories, an air turbine
starter is mounted onthe accessory gearbox to
provide engine cranking through a clutch.
OPERATION
Air entering the nacelle inlet (Figure 7-1) is
accelerated rearward by the fan. A large por-
tion of this air is accelerated to a moderate ve-
locity through the fan bypass duct to contribute
the major portion of the thrust. Some of the air
passing through the fan enters the core en-
gine inlet duct and is progressively increased
in pressure as it passes through the 14 stages
of the HP compressor. The compressor outlet
air is directed rearward to the straight-flow an-
nular combustor. A precise amount of the air
enters the combustion chamber where fuel is
added in the proper proportion by the 18 fuel
injectors. Ignition is provided by two high-en-
ergy ignitor plugs until the engine rpm be-
comes self-sufficient. A large portion of the
air provides dilution and insulation for the
combustion liner. The expanding combustion
gases are directed rearward to the turbine sec-
tion. The two-stage HP turbine extracts enough
energy to drive the HP compressor and the
accessory gear system. The expanding gases
continue rearward to the four-stage LP turbine
which extracts sufficient energy to drive the
fan. The remaining core energy is directed to
the atmosphere by the exhaust duct to con-
tribute to the total engine thrust.
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ENGINE SYSTEMS
The engine systems and engine-related systems
of the Canadair Challenger CL-601-3A are:
• Engine oil system
• Engine fuel system
• Ignition system
• Engine control
• Instrumentation
• Engine starting
• Engine speed control and APR systems
• Engine vibration-monitoring system
ENGINE OIL SYSTEM
General
The engine oil system is completely self-con-
tained and fully automated. The engine oil
system provides for cooling and lubrication of
the engine bearings and the accessory gearbox
in addition to adding heat to the unmetered fuel
system through the oil/fuel heat exchanger. An
oil replenishment system is installed in the
rear equipment bay.
Major Components
Oil Tank
An oil tank is mounted at the 11 o’clock po-
sition on each engine. The tank contains a
gravity filler with a dipstick mounted on the
filler cap. A master chip detector forms part
of the oil tank drain plug.
Oil-Replenishing System
The oil tank can be serviced through the in-
tegral gravity filler or through the replenish-
ment system in the rear equipment bay.
Control
An oil-replenishing control panel, powered
from the battery direct bus (Figure 7-3) con-
tains a power switch, a green power ON light,
and two green oil full switchlights labeled “LH
FULL” and “RH FULL.” These two lights have
a press-to-test feature. A three-position man-
ual selector valve labeled “L,” “OFF,” and “R”
is located adjacent to the control panel. In ad-
dition to selecting the tank for servicing, the
valve controls the power supply to the replen-
ishment pump that supplies oil from the tank
to the selected engine’s oil tank.
Indication
The appropriate oil full switchlight (Figure 7-
3) illuminates when the associated engine oil
tank is full.
Oil Pump
An oil pump containing one pressure element
and six scavenge elements is driven by the ac-
cessory gearbox.
The pressure element provides lubrication of the
main engine bearings and the accessory gearbox.
The scavenge elements provide for direct scav-
enging of the compressor and turbine bearings
and the accessory gearbox.
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Figure 7-3. Oil-Replenishing Control Panel
A separate two-element scavenge pump provides
for positive scavenging of the fan bearing sump
during all flight attitudes (climb or descent).
Chip detectors are located in strategic areas of
the scavenge system and in the oil tank.
Oil Filter
A disposable filter removes solid particles
from the oil. The filter case includes a by-
pass valve and an impending bypass indica-
tor switch. When the differential pressure
across the filter element exceeds a preset
limit, it causes the impending bypass indica-
tor on JB4 (battery direct bus) to trip. The in-
dicator must be reset by a reset button on 
JB4 (Figure 7-4).
Oil Cooler
A conventional oil-to-fuel heat exchanger
mounted on the engine maintains the oil
temperature within design limits.
Indication
A transducer on the pump pressure line senses
oil pressure and transmits it to a signal data
converter (SDC).
A resistance bulb in the oil tank provides tem-
perature signals to the SDC.
The SDC divides the signals into two outputs
and transmits them to alternate fiber optic
segments that form the vertical analog scales
of the oil pressure and oil temperature indi-
cators (Figure 7-5). The fiber optic segments
are color-coded red, yellow, and green. these
colors are also painted on the instrument face
outboard of the analog scales.
A blue light at the bottom of each vertical
scale indicates a power-on condition. The oil
pressure indicators are calibrated in psi. The
oil temperature indicators are calibrated in
degrees Celsius.
Low oil pressure is sensed by a switch on the
pressure pump output line. When the switch
closes below 28 psi, the appropriate L or R
LOP light (Figure 7-5) on the lower face of the
indicator will illuminate t o i n d i c a t e t h a t
p r e s s u r e i s b e l o w d e s i g n m i n i m u m s .
NOTE
The SDC operates from two power
sources: battery bus and essential DC
bus. Lose of either power source will
result in loss of alternate segments of
the scales. The indicators will still
provide a reasonably accurate indi-
cation of pressure and temperature.
FlightSafety Canada LtéeLtd.
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Figure 7-4. Junction Box 4 (JB4)
Operation
Figure 7-6 illustrates operation of the engine
oil system. The pressure element draws oil
from the tank, develops a pressure, and di-
rects the outflow through the bypass filter. A
relief valve limits pressure to a design value.
The pressure oil is directed through the oil
cooler and is then divided into two delivery
lines. One line is directed through a restric-
tor to the accessory gearbox, the front and
rear fan bearings, and the front compressor
bearing. The second delivery line supplies
high-pressure oil to the second and third com-
pressor bearings and to the front and rear tur-
bine bearings.
The six scavenge elements of the oil pump
provide direct scavenging of all bearings ex-
cept the front three. These forward bearings
are scavenged by a dual-element pump to re-
turn oil to the tank. The common scavenge
line enters the tank through a cyclone deaer-
ator. Oil tank pressure and bearing sump pres-
sure is controlled by an oil tank relief valve
and sump vent regulator acting as a vent and
pressure regulator.
ENGINE FUEL SYSTEM
General
The engine fuel system is an integrated hy-
dromechanical-electronic system. The fuel sys-
tem meters fuel to the combustor to provide for
starting, acceleration, deceleration, and full power
requirements under all operating conditions.
In addition, the fuel system operates the vari-
able-geometry system of the compressor to po-
sition inlet guide vanes and compressor stator
vanes to provide engine stall/surge protection.
Major Components
The major components of the fuel system include
an engine-driven LP pump, heat exchanger, a by-
pass filter, a dual-element HP pump, an inte-
grated hydromechanical-electronic fuel control
unit (FCU), a fuel flow distributor, and 18 fuel
nozzles in the combustor system.
LP Engine-Driven Pump
The LP engine-driven pump receives inlet fuel
at the standby pump or main ejector pressure,
increases this pressure, and divides the output
into two flows. One output goes to the heat ex-
changer and fuel filter before reaching the
primary HP element of this three-element
pump. The second output from the LP pump
goes to the secondary HP element. The primary
HP elementdevelops the pressure necessary
for FCU operation. The secondary HP ele-
ment supplies the motive flow fuel to the pri-
mary and scavenge ejectors in the fuel tanks.
It also supplies motive flow fuel to the ecol-
ogy tank jet pump for scavenging of the fuel
drain system.
Fuel Heater
The fuel heater is an air-to-liquid heat ex-
changer. Fourteenth-stage bleed air is modu-
lated to the heater to raise the fuel temperature
to prevent water freezing problems. An auto-
matic bypass on the heater permits all fuel to
bypass if the pressure drop across the heater
exceeds a preset value or if the fuel tempera-
ture is already sufficiently high.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
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Figure 7-5. Oil Pressure and Temperature
Indicators
P
S
I
0
25
95
100
OIL
PRESS
L R
60
80
40
L
O
P
-40
0
150
155
OIL
TEMP
163
L R
90
60
30
120
°C
FlightSafety Canada LtéeLtd.
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16
3
15
5
12
0
15
0
6090 30 0 -4
0
L
O
IL
T
E
M
P
R
10
0
95 4080 60 2
5 0O
IL
P
R
E
S
S
L
R
P S I
L O P
S
U
P
P
LY
O
IL
 P
R
E
S
S
U
R
E
L
E
G
E
N
D
S
C
A
V
E
N
G
E
 O
IL
F
U
E
L
V
E
N
T
 L
IN
E
S
E
LE
C
T
R
IC
A
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M
P
 V
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G
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LA
T
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B
 S
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M
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 S
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A
R
IN
G
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C
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M
P
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E
S
S
O
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B
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A
R
IN
G
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T
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IC
T
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B
Y
P
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S
S
 V
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E
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 T
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D
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 ∆
 L
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IN
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 S
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S
A
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E
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S
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Y
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E
A
R
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IL
 P
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LY
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IL
 T
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 V
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 B
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 V
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IN
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T
U
R
B
IN
E
B
E
A
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IN
G
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C
 S
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 T
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F
ig
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re
 7
-6
. 
 O
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S
ys
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 S
ch
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at
ic
Fuel Filter
A bypass fuel filter removes solids from the
fuel. A red pop-out bypass indicator is pro-
vided, as well as a differential pressure switch.
The switch will illuminate an amber FILTER
light (Annunciator Panel Section) on the fuel
control panel and the appropriate master cau-
tion system whenever the pressure differential
across the filter exceeds a preset value. It is
from the fuel filter that fuel temperature is
sensed and displayed on the temperature in-
dicator on the fuel control panel.
Fuel Control Unit (FCU)
The fuel control unit is an engine-driven hy-
dromechanical-electronic unit that has a me-
tering section and a computing section. The
metering section includes a mechanical gov-
ernor, a fuel metering valve, a bypass valve,
a pressurizing valve, a thrust lever-operated
shutoff valve, and an electronic control unit
(ECU) for fan rpm control.
The computing section of the FCU contains re-
lief valves and servos to sense engine param-
eters such as rpm (N2), compressor discharge
pressure (P3), compressor inlet temperature
(T2C), and the position of the variable-ge-
ometry system. An amplifier (ECU) operates
a torque motor to control fan rpm (N1).
The primary function of the FCU is to control
core engine rpm (N2) as a function of thrust
lever position. In addition, the FCU modu-
lates fuel flow to control fan rpm (N1) through
the amplifier (ECU) and the torque motor on
the FCU. (See also “Engine Speed Control
and APR Systems” in this chapter.) Engine
acceleration and deceleration are controlled by
the FCU, based on internal core pressure (P3),
and inlet temperature (T2C).
The FCU also controls the variable-geometry
system as a function of core engine N2 rpm and
compressor inlet temperature (T2C).
The FCU has a fail-safe schedule in the event
of loss of T2C input. In this case, the variable-
geometry and the acceleration schedules will
revert to a fixed temperature reference. If high
thrust is set at the time of failure, a minor de-
crease in thrust may result. If failure occurs
at idle thrust, possible compressor damage
can result if an attempt is made to accelerate
the engine.
Core engine overspeed in limited three ways:
(1) the N2 governor in the FCU, (2) the com-
puter section (therefore, if the FCU governor
fails, N2 will be limited to less than maxi-
mum allowable transient rpm if the comput-
ing section is operational), and (3) if the
computer or the metering valve servo fails, a
bypass valve will open and reduce fuel flow
to the combustor.
The fan rpm control section of the FCU lim-
its fan rpm as a function of thrust lever posi-
tion (PLA) at power settings representing
takeoff, climb, and cruise. In order to minimize
the thrust lever adjustment during climb, the
fan rpm schedule is biased as a function of fan
inlet temperature (T2). The fan is the primary
thrust producer and fan rpm is used to set
thrust. Fan rpm of both engines should be
matched when the thrust levers are aligned.
Fuel Flow Transmitter
A fuel flow transmitter is located in the me-
tered fuel line from the FCU to provide a cock-
pit indication of fuel flow.
Oil Cooler Bypass
The fuel flow path incorporates a fuel bypass
which routes fuel around the oil cooler during
all engine starts. Until fuel flow reaches 400
pounds per hour (pph), fuel is forced through
the bypass valve only. As fuel flow increases
above this level, it is divided between the by-
pass valve and the oil cooler until the flow
reaches 500 pph, when all the fuel is routed
through the oil cooler.
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Fuel Flow Distributor
The fuel flow distributor is a pressure-oper-
ated unit that senses compressor discharge
pressure and fuel pressure at opposite ends of
a linear actuator. The actuator spool has six
rows of three holes in each row. The pressure
differential representing P3 and fuel pressure
moves the spool to align the holes leading to
the 18 fuel injectors in the combustor. The
area of the holes change with spool move-
ment, precisely metering the same amount of
fuel to each of the 18 fuel injectors.
Fuel Drain and Ecology Tank
During engine shutdown, as fuel pressure
drops, a check valve will shut off inlet fuel to
the fuel flow distributor. At the same time,
another check valve will open a drain, and
combustor pressure will force fuel from the in-
jectors and hoses back through the distributor
and out to the ecology drain system. An ejec-
tor in the ecology tank, with motive flow from
the HP fuel pump element, returns fuel to the
inlet side of the LP engine-driven pump.
Operation
Figure 7-7 illustrates operation of the fuel
system in its simplest form. Initial fuel pres-
sure is supplied by the collector tank standby
pump and later by the main ejector through the
open firewall shutoff valve to the LP engine-
driven pump.
The LP pump increases fuel pressure and directs
fuel through the heat exchanger and filter and
to the dual-element HP fuel pump. The #1 HP
element produces the high fuel pressure re-
quired by the FCU. The #2 HP element supplies
the motive flow for main and scavenge ejector
operation. The metering section of the FCU, in
response to the computing section signals, me-
ters fuel through the flowmeter to the fuel man-
ifold. From the fuel manifold, fuel is supplied
in precisely equal amounts through the 18 noz-
zles in the combustor.
During this operation, the variable-geometry
section of the FCU, operating in response to
inputs representing N2 rpm and compressor
discharge pressure, directs fuel pressure to
the variable-geometry actuators to position
the inlet guide vanes and stator to produce a
safe surge margin across the compressor. At
the same time, guide vane and stator position
is fed back to the FCU.
NOTE
When the engine is static, the guidevanes and stators are at their design
maximum closed position. As the en-
gine starts, the guide vanes’ and sta-
tors’ position will change until, at high
power setting, both the guide vanes
and stators will be at the design full
open position permitting maximum
airflow through the core engine.
IGNITION SYSTEM
General
The CF34-3A series engine has a dual low
tension, capacitor ignition system.
The ignition system for each engine consists
of an ignitor plug A and an ignitor plug B in
the combustor with each ignitor powered
through its own exciter.
Operation of either ignitor is sufficient to pro-
vide for a normal engine start. The ignitor cir-
cuits for each engine are identified as “ignition
A” and “ignition B.”
Ignition Modes
The ignition system has four modes, as follows:
1. Ground start ignition
2. In-flight ignition
3. Continuous ignition
4. Auto (stall) protection ignition
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FlightS
afety C
anada L
té
e
L
td
.
C
L-6
0
0
-2
B
1
6
 P
ILO
T
 T
R
A
IN
IN
G
 M
A
N
U
A
L
C
L 601-3A
 7-11
F
O
R
 T
R
A
IN
IN
G
 P
U
R
P
O
S
E
S
 O
N
L
Y
ON
INOP
VALVE
CLOSED
FILTER
LOW
PRESS
FCU
JET PUMP
BOOST PUMP
18
INJECTORS
0
200
3500
4000
3000
2000
1000
800
FUEL
FLOW
L
x10
P
P
H
R
400
600
TO RIGHT COLLECTOR TANK
TO APU FUEL LINE
FEEDBACK
LINK
FIREWALL
SOV’S
LEFT COLLECTOR TANK
INLET
GUIDE
VANES
STATOR
VANES
JET PUMP MOTIVE FLOW FUEL
ECOLOGY TANK
LP PUMP
FUEL
HEATER
AIR
OUT
14TH-STAGE
AIR IN
FILTER
BYPASSBYPASS
FUEL TEMP
INDICATOR
PUMP BYPASS TORQUE
MOTOR
HP PUMP
N1
T2
N2
AMPLIFIER
TRANSDUCER
THROTTLE
LEVER
VARIABLE
GEOMETRY
ACTUATORS
DRAIN VALVE
18 FUEL HOSES
FUEL FLOW
DISTRIBUTOR
BYPASS
BYPASS
VALVE
OIL OUT
OIL
COOLER
OIL IN
P3 T2C N2
FUEL FLOW
COMBUSTOR
SUPPLY
LP PUMP PRESSURE
HP PUMP PRESSURE
LEGEND
SIGNAL/CONTROL
ECOLOGY DRAIN
ECOLOGY RETURN
OIL
AIR
ELECTRICAL
MECHANICAL
Figure 7-7. Fuel System Schematic
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Ground Start Ignition
The ground start ignition is integrated with the
engine start system from initiation of start to
the termination of start at 55% N2.
Either ignition A or ignition B, or both, may
be armed for operation during a ground start
cycle. It is recommended, however, that only
one ignition circuit be armed to prolong igni-
tion plug life.
Control and Indication
Ignition control and indication is provided on
the start and ignition control panel (Figure 7-
8) located on the overhead panel. Two split-
legend switchlights are used to arm A and/or
B ignition circuits for ground starting. Pushing
either switchlight will illuminate the legend
IGN A or IGN B. This indicates that the se-
lected system is armed to its associated engine
START switchlight (Figure 7-8).
Pushing a START switchlight will illuminate
the green START legend and, simultaneously,
the white ON legend of the selected ignition
switchlight, indicating that power is being ap-
plied to the selected ignition exciter. Ignition
will continue until the start cycle is termi-
nated. Ignition and start termination will be in-
dicated when the ignition ON light and the
START light extinguish.
Ignition and start may be terminated at any
time prior to 55% N2 by pushing the appro-
priate STOP switchlight.
In-flight Ignition
In-flight ignition is separate for each engine. It
is designed to provide dual ignition for wind-
milling relights or during single-engine operation.
Control and Indication
The in-flight ignition system is controlled by
a switchlight (Figure 7-8) for each engine la-
beled “IN FLIGHT START.”
Pushing in one of these switchlights will il-
luminate the green IN FLIGHT START legend
as well as the white ON legend of both igni-
tion arming switchlights, completing the cir-
cuit for operation of ignition A and ignition
B for the associated engine. It is not necessary
to arm the ignition A or ignition B systems
prior to selecting in-flight start ignition.
Continuous Ignition
and Indication
Continuous ignition is primarily used as an
anti-flameout ignition. When selected, it pow-
ers one ignition exciter continuously on both
engines. The system is activated by a single
switchlight (Figure 7-8) labeled “CONT IGN”
only if either ignition A and/or ignition B has
been armed. The green CONT IGN legend,
the green IGN legend, and the white ON leg-
ends of the selected ignition system will all il-
luminate during operation of the continuous
ignition system.
Automatic (Stall Protection)
Ignition
The automatic or stall protection ignition sys-
tem provides anti-flameout protection during
periods of engine inlet turbulence caused by
high angles of attack.
IGNITION
ENGINE START
R
START
CONT
IGN
STOP
IN
FLIGHT
START
START
STOP
IN
FLIGHT
START
L
IGN A
ON
IGN B
ON
Figure 7-8. Start and Ignition
Control Panel
Control
The automatic ignition system is controlled by
the stall protection computer using inputs
from the angle-of-attack vanes. The stall pro-
tection computer will initiate ignition A and
ignition B for both engines 3% before the
onset of the stick shaker and maintain ignition
operation until the angle of attack is reduced.
Indication
When the stall protection ignition is operat-
ing or during a stall protection system test, the
white IGN A ON and IGN B ON (Figure 7-8)
lights will be illuminated.
Power Sources
AC power at 115 volts and 400 Hertz is used for
the ignition system. Left and right engine igni-
tion A is supplied from the essential AC bus.
Left and right engine ignition B is supplied
from the battery bus through a static inverter.
Operation
Figure 7-10 is simplified schematic of the ig-
nition system used on the Canadair Challenger.
The switchlights on the ignition and start con-
trol panel provide for ignition arming and se-
lection of the ground start ignition mode, the
continuous ignition mode, and the in-flight
start ignition mode.
The stall protection system provides ignition
of the duration of stall warning regardless of
the position of all other ignition switches.
ENGINE POWER CONTROL
General
Engine power control is provided on a quad-
rant located on the center pedestal.
Thrust Levers
The individual engine thrust levers (Figure 7-
9) operate in quadrant slots from a full aft po-
sition labeled “SHUT OFF” to a full forward
position labeled “MAX POWER.” An inter-
mediate position forward of SHUT OFF is la-
beled “IDLE.”
A mechanical latch at the rear and below each
thrust lever knob must be raised before the
thrust lever can be moved to or from the SHUT
OFF position.
A go-around button is mounted in each thrust
lever knob. When either is pushed, it will dis-
engage the autopilot and place the AFCS sys-
tem in the go-around mode. Switches are
mounted in the throttle quadrant slots to provide
(1) takeoff configuration warning for flaps,
spoilers, and horizontal stabilizer, (2) pressur-
ization ground control mode, and (3) landing
configuration warning (landing gear not down
and locked at landing power settings).
Quadrant Friction Control
A single friction adjustment twist knob (Figure
7-9) is located on the quadrant aft of and between
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Figure 7-9. Throttle Quadrant
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Figure 7-10. Ignition System Schematic
AC ESSENTIAL BUS
BATTERY BUS
A IGNITER POWER
LEGEND
B IGNITER POWER
FROM STALL PROTECTION
STATIC INVERTER
FROM STALL
PROTECTION
SYSTEM
STATIC
INVERTER
AC ESSENTIAL BUS
C-28
BATTERY BUS
B-169
BA
A B
LEFT ENGINE
IGNITION
CONTROL
RELAY
the thrust levers. Clockwise rotation will in-
crease friction, and counterclockwise rota-
tion will decrease friction.
Thrust reverse control levers are mounted pig-
gyback fashion on the thrust levers.Thrust re-
versers will be discussed later in this chapter.
ENGINE INSTRUMENTATION 
General
The primary engine instruments (Figure 7-
11) are horizontally mounted at the top left side
of the center instrument panel. From left to
right these instruments are as follows:
• N1 (fan rpm)
• ITT (interturbine temperature)
• N2 (core or gas generator rpm)
• Fuel flow
Each instrument has two vertical scales: one
for the left engine and one for the right engine
which provides a nonlinear analog readout.
Below each vertical scale (except OIL TEMP
and OIL PRESS) is a three-digit digital read-
out. To increase safety factors, each indicator
is cross-powered using two power sources;
for example, the left analog scale and the right
digital scale have the same power source. A
separate power source is used for the right
analog and left digital scale. This ensures that
a single power failure will not result in total
readout loss on any engine instrument.
The analog scales are made up of separate
colored segments. These segments are pro-
gressively illuminated from groups of bulbs
with the instrument. The light is transmitted
to the scale segments by fiber optics. The col-
ored segments provide for safe (green), cau-
tion (yellow), and warning (red) indications.
The bottom segment in each vertical scale is
power indicator that will show blue if power
is available to the scale.
Power Sources
The engine instruments are powered from a
signal data converter (SDC). The SDC is sup-
plied DC power from the battery bus and the
essential DC bus. The SDC processes the in-
puts from the various engine parameters and
produces two outputs. These outputs are sup-
plied to the instrument lamp banks. Fiber op-
tics transmit the light from the lamp banks to
the colored segments of the vertical scales.
The digital displays are converted from the as-
sociated analog displays. When compared
with the nonlinear analog readout, the digital
indicators provide a more accurate indication.
Automatic Dimming
A photoelectric cell (Figure 7-12) is provided
on the engine instrument control panel to pro-
vide for automatic engine instrument dim-
ming as ambient light conditions change. A
rheostat on the same panel allows the crew to
set brilliancy to personal preferences.
Instrument Testing
The power supplies of the SDCs are tested
with a three-position TEST switch (Figure 7-
12). Selecting the switch to position 1 or 2 tests
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100
N1
%RPM
L R
98.6
96.2
90
60
30
0
104
N2
%RPM
L R
60
40
20
0
99.2
99.4
98.2
80
0
200
3500
4000
3000
2000
1000
800
FUEL
FLOW
L
x10
P
P
H
R
400
600
0
200
871
878
800
1000
ITT
°C
L
DGT
OFF
R
500
400
300
600
700
850
860
900
Figure 7-11. Engine Instruments
the corresponding SDC power supply by il-
luminating all analog and digital displays in-
cluding the fuel panel.
Indications
The amber light above the instrument test
switch (Figure 7-12) will illuminate when a
power input source to the SDC fails. In this
case, the blue power-on segments of the af-
fected scales will extinguish, and the associ-
ated analog and opposite digital display will
be lost.
N1 (Fan) RPM
N1 (fan) rpm (Figure 7-11) is sensed by a
monopole transmitter located on the engine
front frame. Electrical signals are sent via the
SDC to the appropriate analog and digital
scale. Each scale is calibrated to indicate a per-
centage of N1 rpm from 0% to 100%.
ITT (Interturbine Temperature)
Thermocouples of different lengths are equally
spaced between the LP turbine and the HP
turbine. The thermocouples are parallel-con-
nected. The ITT output is sent to the appro-
priate vertical and digital scales of the ITT in-
dicator via the SDC. The ITT indicator scales
are calibrated in degrees Celsius from 0 to
1000°C.
A red light (Figure 7-11) above each vertical
scale will illuminate if the ITT reaches 871°C.
These lights also illuminate during the in-
strument test.
A two-position switch labeled “DGT OFF”
(Figure 7-11), located at the bottom of the ITT
panel, allows the crew to extinguish all engine
parameter digital displays which might be an-
noying on extended nighttime operations.
N2 RPM
N2 rpm (Figure 7-11) is supplied by an alter-
nator driven by the accessory gear. The rpm
signals are isolated from the alternator’s power
to eliminate interference and interruption. The
rpm signals are sent to the appropriate N2
scales via the SDC.
Fuel Flow
Fuel flow (Figure 7-11) is sensed by a mass
flow transmitter located downstream of the
FCU. The transmitter output is sent to the
SDC for processing into analog and digital
readout for display on the appropriate fuel
flow indicators.
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AUX PWR
TEST
1 2
D
I
M
Figure 7-12. Engine Instrument
Control Panel
C
O
N
T
R
O
L
A
P
U
PUSHPUSH PUSH
PWR FUEL
ON/OFF
APU
OIL
ADPTR
OIL
BLEED
AIR
START/
STOP
STARTER
APU
READY
LO
PRESS
HI
TEMP
SOV
CLOSED
PUMP
INOP
LO
PRESS
HI
TEMP
FAILED
OPEN
%RPM
100
80 0
60
40
20
EGT
°C X 100
8
6 2
4
0
10
Figure 7-13. APU Control Panel
The analog scales are calibrated in pounds of
fuel per hour from 0 to 4,000. The digital dis-
plays are in pounds per hours times 10.
ENGINE STARTING
General
Engine starting is divided into ground starts,
starter-assisted airstarts, and windmilling airstarts.
Starter
The engine starter is an electrically controlled
air turbine starter (ATS). Starter output is ap-
plied through a clutch to the accessory gear
which, in turn, rotates the HP spool.
A speed sensor operated by the ATS governor
automatically terminates the starter cycle at ap-
proximately 55% N2 rpm. The average start
cycle is less than 40 seconds. A time-delay relay
is armed when a start cycle is initiated, and, if
the ATS operation continues for more than 60
seconds, the time-delay relay will open an illu-
minate the amber STOP switchlight on the start
and ignition control panel (Figure 7-8). The
STOP switchlight may be pushed to terminate
the start sequence at any time below 55% N2 rpm.
ATS Air Sources
The air source for ATS operation can be (1)
APU bleed air, (2) an external air source, or
(3) cross bleed from an operating engine. The
minimum air pressure for starting is 45 psi.
(See Chapter 9, “Pneumatics.”)
Ground Start (APU Air)
Engine starting should not be attempted
until the Walkaround checklist and the
Cockpit checklist are completed.
To initiate a ground start using APU bleed air,
push the APU bleed-air switchlight (Figure 7-
13). The OPEN light will illuminate, and the left
scale of the bleed-air pressure indicator (Figure
7-14) should show approximately 50 psi. Push
IGN A or IGN B switchlight (Figure 7-8) to
arm an ignition system, and check that the ap-
plicable green light illuminates.
Push and hold the appropriate START switch-
light for 2 seconds. The green START light will
illuminate as will the ON light in the selected
ignition switchlight. The ISOLation valve OPEN
light will also illuminate. The left and right en-
gine BLEED CLOSED lights will extinguish.
Verify engine rotation on the N2 rpm indicator
and monitor N2 until it reaches 20% minimum
and ITT below 120° C. Then move the affected
thrust lever to IDLE, check the ITT indicator for
light-off, and continue to monitor ITT, oil pres-
sure, and N2 rpm. Also check that N1 rpm is in-
creasing in relation to N2. At approximately
55% N2 rpm, the START light and the IGNition
ON light (Figure 7-8) will both extinguish as
should the ISOLation valve OPEN light. The left
and right engine BLEED CLOSED lights
(Figure 7-14) should both illuminate. Continue
to monitor all engine-related instruments until
the engine stabilizes at idle rpm (approximately
60–64% N2). The N2 variation between engines
at idle should be within 2%.
NOTE
The idle N2 rpm of CF34 engines au-
tomatically varies as a function of
compressor inlet temperature (T2C).In case of faulty T2C input, an IDLE
FLOOR STOP is provided to prevent
N2 from decreasing below 56.9%.
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WARNING
A
I
R
B
L
E
E
D
PUSH
ON/
OFF
PUSH
ON/
OFF
14TH STAGE 10TH STGR
DUCT MON
LOOP A STBY
LOOP B NORM
10TH STAGE
ISOL ACUL LR R
BOTH OFF
CKPT HEAT
OFF
FAIL
OFF
FAIL
BLEED
AIR
R
PSI
100
50
0
L
100
50
0
L
BLEED
CLOSED
DUCT
FAIL
BLEED
CLOSED
DUCT
FAIL
BLEED
CLOSED
DUCT
FAIL
OPEN
BLEED
CLOSED
DUCT
FAIL
Figure 7-14. Bleed-Air Control Panel
If idle speed stabilizes at approxi-
mately 57% N2, the engine must be
shut down immediately and the con-
dition reported to maintenance.
Do not attempt to increase idle N2 by
advancing the thrust lever because it
can result in serious damage to the
first-stage compressor blades.
Ground Start (External Air)
The procedures for engine starting using an ex-
ternal air supply are identical with those for
APU bleed-air starts. An approved external air
unit capable of 45 psi can be connected to the
adapter (Figure 7-15) located in an access on
the left side of the rear fuselage.
Figure 7-16 illustrates the use of bleed for
ground starting the first engine.
Ground Start (Cross Bleed Air)
The procedures for a ground start using a cross
bleed-air supply are similar to those for APU
bleed or external air source, except that the
APU bleed air (Figure 7-13) must be off .
Push the BLEED AIR switch (Figure 7-14)
of the operating engine and check that the
bleed-air pressure is 45 psi minimum, then
continue as for APU bleed-air start.
Figure 7-17 illustrates the availability of bleed
air during a cross bleed start of the left engine.
NOTE
Two conditions must be met before
moving the thrust lever to IDLE for
all engine starting:
1. Indicated ITT must be less than
120° C.
2. N2 rpm must be 20% minimum.
If ITT is greater than 120° C prior to start,
the engine must be dry motored until ITT
drops below 120° C.
NOTE
When using battery or external DC
power only during engine starting,
bleed-air pressure indication will
not be available.
CAUTION
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LEFT ENGINE
ATSATS
APU
10TH STAGE
BLEED-AIR
LEFT START
VALVE
LEFT 
BLEED-AIR
SOV
EXTERNAL AIR
LCV
ISOLATION
VALVE
RIGHT START
VALVE
RIGHT
BLEED-AIR
SOV
RIGHT ENGINE
10TH STAGE
BLEED-AIR
APU BLEED AIR
LEGEND
Figure 7-16. Bleed-Air Sources (First Engine Start Schematic)
LEFT ENGINE
ATSATS
APU
10TH STAGE
BLEED-AIR
LEFT START
VALVE
LEFT 
BLEED-AIR
SOV
EXTERNAL AIR
LCV
ISOLATION
VALVE
RIGHT START
VALVE
RIGHT
BLEED-AIR
SOV
RIGHT ENGINE
10TH STAGE
BLEED-AIR
10TH-STAGE BLEED AIR
LEGEND
Figure 7-17. Cross Bleed Start (Left Engine from Right Engine Schematic)
Figure 7-15. External Air Supply Adapter
Failure to Start
Light-off as indicated by rising ITT will nor-
mally occur within 10 seconds after moving
the thrust lever to IDLE.
A start should be aborted if light-off does not
occur 25 seconds after moving the thrust lever
to IDLE.
If starter operation continues for more than 60
seconds, the time delay relay will cause the
STOP light to illuminate. At temperatures
above 15° C (59° F) the start should be aborted
by pushing the STOP switchlight anytime up
to 55% N2 rpm. The thrust lever should then
be moved to SHUT OFF; then wait one minute
before attempting another start. At tempera-
tures below 15° C (59° F) the start sequence
may exceed 60 seconds (Figure 7-18).
Before attempting another start, dry motor
the engine with both ignition systems off and
the affected thrust lever at SHUT OFF.
NOTE
The a i r t u rb ine du ty cyc l e f o r
normal engine start is 3 consecutive
cycles with 5 minutes cooling be-
tween additional cycles.
For dry motoring, the ATS duty cycle
is 90 seconds with a 5-minute cool-
ing period between additional cy-
cles of 30-second duration.
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20
20
0
–20
–40
40
60
40 60 80 100 1200
O
U
T
S
ID
E
 A
IR
 T
E
M
P
E
R
A
T
U
R
E
 —
 °
F
TOTAL TIME
TO STABILIZED
IDLE
TIME FROM
THROTTLE OPENING
TO LIGHT-OFF
Figure 7-18. Maximum Allowable Start Time and Time to Stabilized Idle—Seconds
Airstarts
Airstarts are divided into starter-assisted and
windmilling airstarts.
Starter-assisted Airstarts
The procedure for starter-assisted airstarts
(cross bleed starts) are identical with those ex-
plained previously for cross bleed starts.
All in-flight starts must be performed within
the airstart envelope (Figure 7-19).
The thrust lever should not be moved to IDLE
during airstarts unless ITT is less than 90°C.
Windmilling Airstart
Windmilling airstarts are obtained at the fol-
lowing airspeeds:
Below 10,000 feet......................... 300 KIAS
10,000 to 21,000 feet ..... 300 KIAS to VMO
N2 must be stable or increasing.
Airstarts, windmilling or starter-
assisted, should not be attempted if
the flameout or shutdown is ac-
companied by unusual noise or
other indications that mechanical
damage may exist.
Prior to initiating a windmilling airstart, all
checklist items affecting the start must be
completed. Then push the appropriate IN
FLighT START switchlight. The green light
and the IGN A and B ON lights will illuminate.
Advance the thrust lever to IDLE and moni-
tor all engine-related instruments until the en-
gine is stabilized. Then push the IN FLighT
START switchlight again. The green light and
the IGN A and B ON lights will extinguish,
then complete the After Start checklist.
ENGINE SPEED CONTROL
AND APR SYSTEMS
General
The automatic performance reserve (APR) is
a solid-state system which constantly moni-
tors the thrust of both engines during takeoff.
If significant power loss occurs in either en-
gine, it will instantaneously command an N1
(thrust) increase.
Components
The APR system components include an APR
control panel (Figure 7-20), an APR controller,
and an N1 speed selector switch for each en-
gine. In addition, the APR system utilizes the
torque motors (discussed earlier in Engine
Fuel System) and the amplifiers (ECUs) as-
sociated with the FCUs.
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AIRSTART ALTITUDE LIMIT
21
10
0
0 12 13 N2% RPM
55
WINDMILL START
10,000–21,000 FT
STARTER ASSIST
BELOW 21,000 FT
WINDMILL START
BELOW 10,000 FT
Figure 7-19. Airstart Envelope
WARNING
Figure 7-20. APR Control Panel
Control and Indication
A three-position switch (Figure 7-20) with
pos i t i ons l abe l ed “ARM,” “OFF,” and
“TEST/RESET” is located on the APR control
panel. When this switch is at the ARM posi-
tion, the system is armed provided that three
conditions exist.
1. Both engine speed control switches on
2. Both engines above 79% N1
3. No faults sensed by the integral mon-
itoring system
Selecting the APR switch off deactivates the sys-
tem. The TEST/RESET position is a spring-
loaded to the off position and is used for testing.
A split legend green light labeled “L.ON” and
“R. ON” will illuminate following an APR
trigger (activating) to confirm proper response
(N1 increase) on the serviceable engine.
A green light labeled “READY” forms the
upper part of a dual legend light which illu-
minates to confirm APR readiness above 79%
N1 if the system is armed. If a subsequent APR
trigger occurs, the READY light will extin-
guish as the L. ON or R. ON light illuminates.
An amber APR light and MASTER CAUTION
lights will illuminate as a crew warning that
either (1) the APR system is not armed for
takeoff or (2) that the APR system has failed
for one or more of the following reasons:
• Either the static or dynamic test is not valid.
• The serviceable engine’s response to an
APR trigger produces less than2% N1
rpm increase within 2 seconds.
• The ECU input voltages are outside ac-
ceptable limits.
• The monitoring system detects failure
of the microcomputer or the two inter-
nal power supplies.
• Either or both N1 input signals are out
of limits.
• Battery input voltage fails.
• The two WOW inputs disagree.
• A unwarranted APR command is triggered.
NOTE
The APR system is used only for
takeoff and is then disarmed. The
APR fail light is inhibited in flight
through WOW logic and for landing
by flap 45° selection.
A green TEST light forms the lower portion
of the READY light. During testing, this light
will illuminate as the last indication in a se-
ries until the APR switch is released from the
TEST/RESET position.
Testing
Two tests are associated with the APR sys-
tem: (1) static test and (2) dynamic test.
Static Test
Holding the APR switch in the TEST/RESET
posi t ion causes the fol lowing funct ions
and indications:
1. The system is reset, the APR program
is restarted, and all previous perfor-
mance data in the memory is cleared.
2. Validates the battery direct bus voltage
input. If not present, the amber APR
light will illuminate.
3. Tests all lamps for 1 second each in
the following sequence:
a. Ready and L. ON
b. Ready and R. ON
c. Test
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d. APR
e. Test (will remain on as long as TEST
is held)
4. If any faults are detected , the amber
APR light will illuminate.
Dynamic Test
The dynamic test is done just prior to flight. It is
automatically performed for both engines by the
APR controller and verifies that the APR system
is operational. To perform the test, the APR
switch is armed. Advance both thrust levers to
obtain an indicated N1 above 83% (note the
READY light came on at 79% N1). The APR
controller samples the fan rpm continuously to
determine if they are stable; that is, fan rpm does
not vary more than a predetermined amount. The
dynamic test initiates an APR trigger to both
APR amplifiers which causes both engines to
accelerate slightly. If the test is valid, the TEST
light will momentarily flash. If outside the
permissible limit, the amber APR light will
illuminate, accompanied by the MASTER
CAUTION light.
Operation
Before takeoff, the static and dynamic tests are
performed and determined as valid, both en-
gine speed control switches are on, and the
APR switch is at ARM. The green READY
light will illuminate after 79% N1. The amber
APR light and the green L. ON and R. ON
lights and the TEST light are extinguished.
Figure 7-21 illustrates an APR trigger. Both
fuel control amplifiers (FCUs) are receiving N1,
N2, T2, and power lever angle (PLA). The right
engine N1 has decreased below N1 speed con-
trol (79% N1) and has reached the APR “trig-
ger” speed (approximately 68% N1). The APR
controller sends a signal to both amplifiers to
increase N1. The left engine responds since it
is still on N1 speed control (above 79% N1) and
illuminates the green L. ON legend when it has
increased the required amount (approximately
2% N1). The right engine does not respond
since it is not on N1 speed control.
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Figure 7-21. APR/Engine Speed Schematic
FUEL CONTROL
AMPLIFIER
(ECU)
TM
FUEL FLOW
BATTERY POWER
N1 INPUT
LEGEND
N2 INPUT
APR TRIGGER
T2 INPUT
POWER
LEVER ANGLE
FEEDBACK
APR
CONTROLLER
FLAPS
45° WOW
BATTERY DIRECT BUS
N1
DESIRED
N1
ACTUAL
N1
T2
N2
PLA
POWER
FEEDBACK
APR
MV
FCU
FUEL CONTROL
AMPLIFIER
(ECU)
TM
N1
DESIRED
N1
ACTUAL
N1
T2
N2
PLA
POWER
FEEDBACK
METERED FUEL TO COMBUSTOR
MV—METERING VALVE
TM—TORQUE MOTOR
APR
MV
FCU
APR OPERATING,
LEFT ENGINE
ENGINE VIBRATION-
MONITORING SYSTEM
General
An engine vibration-monitoring system con-
tinuously monitors the vibration level of each
engine and provides cockpit indication of
these levels. A warning system alerts the crew
if predetermined levels are exceeded.
Components—AC 5001-5104
A transducer mounted on the compressor cas-
ing of each engine generates an electrical sig-
nal proportional to the intensity of the engine
vibration. The generated signal is sent to a
signal conditioner and then transmitted to a
dual needle, dual-scale indicator (Figure 7-22)
on the pilot’s side console. The indicator is cal-
ibrated in mils from 1 through 4.
An amber HIGH VIBration switchlight (Figure
7-22) near the indicator will illuminate in con-
junction with the pilot’s master caution system
if the vibration level of either engine exceeds the
permissible limits for more than 3 seconds.
The 3-second time delay eliminates nuisance
warnings due to high transient engine vibration.
Testing
The vibration indicating system can be tested
by pressing and holding the HIGH VIBration
switchlight. The needles will show a vibration
level of 4.0 mils DA and after 3 seconds, the
amber HIGH VIB light will illuminate accom-
panied by the ENGINE annunciator and flash-
ing MASTER CAUTION lights.
NOTE
The test system does not test the en-
gine-mounted transducers or cable
circuitry. The indicator needle re-
sponse following engine starting in-
dicates proper operation of the system.
Components—AC 5105
and Subsequent
A transducer mounted on the compressor cas-
ing of each engine generates an electrical sig-
nal proportional to the intensity of the engine
vibration. The generated signal is sent to the
cockpit indicator, for fan vibration only. If
fan vibration exceeds 2.7 mils DA for more
than 3 seconds, an amber FAN annunciator
will also come on, accompanied by an EN-
GINE master caution light. If core vibration
exceeds 1.7 mils DA for more than 3 seconds,
only the amber CORE annunciator will come
on plus an ENGINE master caution light but
no needle deflection occurs on the vibration
indicator (Figure 7-23).
The 3-second time delay eliminates nuisance
warnings due to high transient engine vibration.
Each engine has their own CORE and FAN an-
nunciator lights.
Test
The vibration indicating system can be tested
by pressing and holding in each of the CORE
and FAN indicator switchlights on the EVM
indicator panel; in turn, the pointer on the
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ENGINE
VIBRATION
PRESS
TO
TEST
HIGH
VIB
4
3
2
0
1
4
3
2
0
1
L R
VIB
MILLS
D.A.
Figure 7-22. Engine Vibration-Monitoring
Panel (AC 5001-5104)
side being tested will move to 4.0 MILLS DA.
When the switchlight is held pressed for more
than 3 seconds, the respective amber CORE
and FAN lights will illuminate accompanied
by an engine M/C light.
NOTE
The test system does not test the en-
gine-mounted transducers or cable
circuitry. The indicator needle re-
sponse following engine starting in-
dicates proper operation of the system.
THRUST REVERSERS
GENERAL
The Canadair Challenger is equipped with
electrically energized, pneumatically oper-
ated thrust reversers to aid in deceleration on
the ground. The reversers are installed im-
mediately aft of the fan section.
The reversers are certificated for ground use
only, and the control circuits are wired through
the WOW system. The reverser consists of a
tracked translating sleeve that is moved aft by
an air motor driving a flexible shaft connected
to 4 ballscrew actuators. As the translating sleeve
moves aft, cascade vanes are exposed to redirect
the fan bypass airflow forward over the nose
cowl assembly. The aft motion of the translat-
ing sleeve also causes 10 blocker doors to block
the fan air exit nozzle and directs the airflow ra-
dially outward through the cascade vanes.
Forward movement of the translating sleeve
positions the blocker doors to lie flat and form
part of the fan airflow exit nozzle.
NOTE
Only fan airflow is reversed. The re-
verser system does not affect the coreengine airflow which continues on its
normal path to the atmosphere.
CONTROL
The thrust reversers are controlled by levers
(Figure 7-9) piggyback-mounted on the thrust
levers. The thrust reverser levers are latched
in the fully forward position. The latch must
be manually released by the pilot, allowing
each thrust reverser lever to be moved be-
tween three basic positions: stow (fully for-
ward), deploy (20° aft of full forward), and
reverse thrust which is a variable position.
Stow and deploy are fixed positions.
When the thrust reverser lever is unlatched,
then moved upward and aft, a solenoid is en-
gaged to prevent more than 20° of lever travel
while the air motor operates the ballscrew
jacks to drive the translating sleeve aft, ex-
posing the cascade vanes and closing the 10
blocker doors in the bypass duct. At this point,
a microswitch is operated to release the thrust
reverser lever lock solenoid. Further aft mo-
tion of the thrust reverser lever acts through
the FCU power lever to increase fuel flow and
provide proportionally more reverse thrust.
Figure 7-24 shows the CF34 engine in for-
ward thrust and reverse thrust configurations.
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Figure 7-23. Engine Vibration-Monitoring
Panel (AC 5105
and Subsequent)
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Figure 7-24. Thrust Reversers
NORMAL OPERATION
BYPASS
AIRFLOW
BYPASS
AIRFLOW
CORE
AIRFLOW
REVERSER DEPLOYED
CORE
AIRFLOW
INDICATION
Electrical control is provided by a dual-leg-
end switchlight (Figure 7-25) for each thrust
reverser. The switchlights are located on the
reverse thrust control panel on the center
pedestal. Pushing a switchlight will arm the
system and illuminate the white ARMED leg-
end. With thrust levers at IDLE and weight on
wheels or wheel spin-up, unlatching and mov-
ing the thrust reverser lever from the stowed
position will cause the translating sleeve to
move rearward which will illuminate an amber
REVERSER UNLOCKED l i gh t on t he
glareshield (Figure 7-26). When the translat-
ing sleeves reaches full aft travel, the associ-
ated green REVERSE THRUST switchlight
(Figure 7-26) will illuminate to indicate that
reverse thrust can now be applied.
An amber UNSAFE TO ARM light forming
part of the ARMED switchlight (Figure 7-25)
will illuminate in flight to indicate that it is
unsafe to arm that thrust reverser system be-
cause (1) either the thrust reverse lever is not
in the fully stowed position or (2) the de-
ploy/stow switch is faulty (possibly giving a
permanent deploy command). Arming the sys-
tem with the UNSAFE TO ARM light illumi-
nated could result in immediate deployment
of that thrust reverser at touchdown. On the
ground this light will illuminate if the asso-
ciated reverser was not armed and the thrust
reverser lever is actuated.
The amber REVERSER UNLOCKED light
(Figure 7-26) will illuminate if (1) the trans-
lating sleeve is not stowed, (2) the pneumatic
brake is released, (3) the mechanical lock is
disengaged, or (4) the translating sleeve has
moved rearward more than 1 inch.
PROTECTION
General
The thrust reverser protection system includes
autostow, emergency stow, and a mechanical
thrust lever retard system.
Autostow
If a REVERSER UNLOCKED light illumi-
nates in flight, an electrical signal is auto-
matically sent to pneumatically stow the
reverser. If it is successful, the REVERSER
UNLOCK light extinguishes.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
7-28 CL 601-3A FOR TRAINING PURPOSES ONLY
PITCH
YAW
ROLL
UNSAFE
TO ARM
ARMED
UNSAFE
TO ARM
ARMED
SERVO
MONITOR
REVERSE THRUST
LEFT RIGHT
PUSH TO ARM
MON
SAFE
Figure 7-25. Reverse Thrust Control Panel
REVERSER
UNLOCKED
REVERSE
THRUST
LH FLT
SPLR
LH GND
SPLR
RH FLT
SPLR
RH GND
SPLR
PUSH LEFT PUSH RIGHT
THRUST REVERSER
EMERG. STOW
REVERSER
UNLOCKED
REVERSE
THRUST
Figure 7-26. REVERSER UNLOCKED
Lights
Emergency Stow
If the REVERSER UNLOCKED light remains
illuminated or cycles on and off during flight,
pushing the switchlight will back up the au-
tostow system by applying a separate and con-
t inuous e lec t r ica l s tow command to the
pneumatic drive unit.
The amber REVERSER UNLOCKED light will
remain illuminated as an indication that the
emergency stow system has been actuated.
Thrust Lever Retard System
An automatic thrust lever retard system is in-
stalled and mechanically interfaced with the
translating sleeve. Uncommanded movement
beyond 11/2 inches of the translating sleeve
will result in retarding of the thrust lever to
IDLE. A lock is then applied to prevent for-
ward movement of the thrust lever. If the re-
verser can be stowed, this lock is removed.
OPERATION
As part of the Before Landing Checklist, the
thrust reverser switchlights are selected to the
ARMED position. Check that the ARMED
legends are illuminated.
Do not arm a thrust reverser in
f l ight i f the UNSAFE TO ARM
light is illuminated
After touchdown, when weight on wheels or
wheel spin-up is being sensed and the thrust
lever at idle, unlatch and raise the thrust re-
versers levers and hold gently against the
solenoid stops. Check that both amber RE-
VERSER UNLOCKED lights illuminate, fol-
lowed in approximately 2 seconds by both
green REVERSER THRUST lights. Then
move the thrust reverser levers aft to obtain re-
verse thrust proportional to the amount of aft
lever movement. When ground speed decreases
to 80 knots, move the reverser levers forward
to a minimum reverse thrust position.
When thrust reversers are no longer required,
move the thrust reverser levers to the full
stow position. The REVERSE THRUST light
and the REVERSER UNLOCKED light on the
glareshield will extinguish.
Moving the reverser lever rapidly
from the full reverse thrust position
to the stow position will increase
ground speed because to the residual
engine thrust during spooldown.
CAUTION
CAUTION
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3A 7-29FOR TRAINING PURPOSES ONLY
1. The CF 34 engine may be defined as a:
A. Single-spool, medium-bypass turbofan
B. Nonmodular, single-spool turbofan
C. Twin-spool, high-bypass turbofan
D. High-bypass geared turbofan
2. The compressor surge margin is main-
tained by:
A. Variable compressor inlet and stator
guide vanes
B. A fuel pressure-operated bleed strap
C. Variable fan inlet and compressor inlet
guide vanes
D. Pneumatically operated compressor
bleed valves
3. When checking the engine instruments
with battery power only, the indication on
the oil pressure indicator is:
A. The blue power-on segments will re-
main extinguished.
B. All segments on the left scale will be
illuminated.
C. Both scale segments will flash on and
off alternately.
D. Alternate segments will illuminate
for both left and right scales.
4. The primary thrust indicator for CF34
engine is the:
A. N1 rpm
B. ITT
C. N2 rpm
D. Fuel flow
5. Fan rpm limiting is a sole function of the:
A. Position of the variable-geometry
system
B. Jump and rate system in the FCU
C. FCU governor
D. Speed control ECU on the FCU
6. Electrical power for engine ignition is
supplied by the:
A. Battery bus and essential AC bus
B. Essential DC bus
C. AC bus 1 and AC bus 2
D. Self-contained exciters on each engine
7. Prior to engine starting, ITT is indicating
150°C; the starting procedure is:
A. Select Ignition A or B, but not both.
B. Advance the throttle as soon as start
is initiated.
C. Motor the engine with the throttle at
IDLE for 15 seconds.
D. Motor the engine without fuel or ig-
nition until ITT drops below 120°C.
8. For an in-flight engine start, the ATS may
be used if:
A. N2 rpm is greater than 55%.
B. N1 rpm is less than 55%.
C. N1 and N2 are within 5%.
D. N2 rpm is 55% or less.
9. One recommended airspeed range for a
windmilling airstart is:
A. 200 KIAS below 10,000 feet if N1 is
not indicating
B. 300 KIAS to VMO between10,000
feet and 21,000 feet
C. 250 KIAS at all altitudes
D. 300 KIAS below 10,000 feet if N2 is
stable or decreasing
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
7-30 CL 601-3A FOR TRAINING PURPOSES ONLY
QUESTIONS
CL 601-3R 7-i
CHAPTER 7
POWERPLANT
CONTENTS
Page
INTRODUCTION ................................................................................................................... 7-1
GENERAL............................................................................................................................... 7-1
ENGINES ................................................................................................................................ 7-2
General ............................................................................................................................. 7-2
Major Sections ................................................................................................................. 7-2
Operation.......................................................................................................................... 7-4
ENGINE SYSTEMS ............................................................................................................... 7-5
Engine Oil System ........................................................................................................... 7-5
Engine Fuel System ......................................................................................................... 7-7
Ignition System .............................................................................................................. 7-11
Engine Power Control.................................................................................................... 7-14
Engine Instrumentation.................................................................................................. 7-14
Engine Starting .............................................................................................................. 7-16
Engine Speed Control and APR Systems ...................................................................... 7-21
Engine Vibration-Monitoring System............................................................................ 7-24
THRUST REVERSERS ........................................................................................................ 7-24
General........................................................................................................................... 7-24
Control ........................................................................................................................... 7-25
Indication ....................................................................................................................... 7-26
Protection ....................................................................................................................... 7-26
Operation ....................................................................................................................... 7-27
QUESTIONS......................................................................................................................... 7-28
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-iii
ILLUSTRATIONS
Figure Title Page
7-1 CF34 Major Sections and Gas Flow......................................................................... 7-2
7-2 CF34 Engine Modules.............................................................................................. 7-3
7-3 Oil-replenishing Control Panel................................................................................. 7-5
7-4 Junction Box 4 (JB4)................................................................................................ 7-6
7-5 Oil Pressure and Temperature Indicators ................................................................. 7-7
7-6 Oil System Schematic .............................................................................................. 7-8
7-7 Fuel System Schematic .......................................................................................... 7-10
7-8 Start and Ignition Control Panel............................................................................. 7-12
7-9 Ignition System Schematic..................................................................................... 7-13
7-10 Throttle Quadrant ................................................................................................... 7-14
7-11 Engine Instruments................................................................................................. 7-15
7-12 Engine Instrument Control Panel ........................................................................... 7-15
7-13 APU Control Panel................................................................................................. 7-17
7-14 Bleed-Air Control Panel......................................................................................... 7-17
7-15 External Air Supply Adapter.................................................................................. 7-18
7-16 Bleed-Air Sources (First Engine Start Schematic)................................................. 7-18
7-17 Cross Bleed Start (Left Engine from Right Engine Schematic) ............................ 7-18
7-18 Maximum Allowable Start Time and Time to Stabilize Idle—Seconds ................ 7-20
7-19 Airstart Envelope.................................................................................................... 7-20
7-20 APR Control Panel ................................................................................................. 7-21
7-21 APR/Engine Speed Schematic ............................................................................... 7-23
7-22 Engine Vibration-Monitoring Panel ....................................................................... 7-24
7-23 Thrust Reversers..................................................................................................... 7-25
7-24 Reverse Thrust Control Panel................................................................................. 7-26
7-25 REVERSER UNLOCKED Lights ......................................................................... 7-26
FOR TRAINING PURPOSES ONLY
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
INTRODUCTION
This chapter describes the powerplant of the Canadair Challenger CL-600-2B16, model
CL-601-3R.
In addition to the basic powerplant information, the chapter also includes information
on all powerplant-related systems, such as engine oil, engine fuel, ignition, engine
power control, instrumentation, engine starting, engine speed control and APR sys-
tems, engine vibration monitoring, and thrust reversing.
GENERAL
The Canadair Challenger CL-601-3R is pow-
ered by two aft-fuselage-mounted turbofan
engines manufactured by the General Electric
Company.
The engines are modular in design to facilitate
maintenance and reduce airplane downtime.
Each engine incorporates self-contained oil,
fuel, and ignition systems in addition to a fire
and/or overheat detection system. A fire-ex-
tinguishing system is common to both en-
gines. Pneumatically operated cascade thrust
reversers are standard equipment.
Each engine is monitored during takeoff by an
electronically controlled automatic perfor-
mance (power) reserve system (APR). It will
automatically increase the permissible tem-
perature limits and thrust on the operating en-
gine if a power loss or failure occurs on the
opposite engine.
#1 DC
GEN 
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CL-600-2B16 PILOT TRAINING MANUAL
CHAPTER 7
POWERPLANT
CL 601-3R 7-1FOR TRAINING PURPOSES ONLY
ENGINES
GENERAL
The engines (Figure 7-1) on the Canadair Chal-
lenger are GE Series CF 34. This engine has a
high bypass ratio (6.2 to 1). The CF34 3A1 ver-
sion of this engine is capable of producing
8,729 pounds of takeoff thrust up to 21°C
(70°F) understandard sea level static condi-
tions. If a power loss occurs on one engine, the
APR system will automatically increase the
thrust on the operating engine to 9,220 pounds.
Modular Concept
For ease of maintenance, assembly, and dis-
assembly, the engine is designed in seven sep-
arate modules (Figure 7-2). Some of these
modules can be removed and replaced with-
out engine removal from the airplane.
MAJOR SECTIONS
For the purpose of this chapter the engine will
be discussed under seven major sections:
1. Air inlet section
2. Fan section
3. Compressor
4. Combustor
5. Turbine
6. Exhaust
7. Accessory gear
Air Inlet Section
The nacelle fairing forms the main air inlet at
the front of the engine fan section.
Fan Section
The single-stage fan and integral two-piece nose
cone are installed in the front frame. The fan is
basically the low-pressure (LP) compressor of
the engine in conjunction with a row of stators
mounted in the front frame aft of the fan.
Air entering the engine air inlet is divided into
two flow paths aft of the fan; one path directs
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
7-2 CL 601-3R FOR TRAINING PURPOSES ONLY
FAN
SUPPORT
COMPRESSOR
COMBUSTOR
TURBINE
HIGH PRESSURE
SHAFT
ACCESSORY GEARINLET GUIDE VANE
AIR INLET
N1
N1
LP SHAFT
N2N2
EXHAUST
HP LP
Figure 7-1. CF34 Major Sections and Gas Flow
air to the compressor of the core engine and the
second path directs air into the fan bypass duct.
The fan functions to accelerate a large air mass
to a moderate velocity through the bypass duct
and contributes approximately 80% of the total
thrust developed by the CF34 engine.
Compressor
The high-pressure (HP) compressor is located
aft of the front frame. This single-spool axial
compressor has 14 stages with a pressure ratio
of 14:1.
The HP compressor supplies air for cooling,
bearing and seal pressurization, hot-point
cooling, and for combustion in the core engine.
In addition, it supplies bleed air for the air-
plane’s pneumatic services.
A variable-geometry system automatically
controls the inlet guide vanes and five variable
stator vanes to maintain a safe surge margin
across the HP compressor. This helps prevent
compressor stall or surges when the engine is
operating in the low-power range or during
rapid acceleration or deceleration.
The variable-geometry system is controlled by
the fuel control unit (FCU) as a function of HP
compressor rpm (N2) and core-inlet temper-
ature (T2). The FCU includes a fuel tempera-
ture compensating system to maintain the
required variable-geometry accuracy through-
out the normal fuel temperature range. The
variable-geometry module in the FCU will di-
rect HP fuel to two variable-geometry actua-
tors to operate the inlet and stator vanes.
A feedback system relays the position of the
vanes to the FCU at all times. When the en-
gine is static and during steady-state operation
at lower power, the inlet guide vanes and the
variable stator vane are at a close position.
This res t r ic ts the airf low to the HP com-
pressor to an amount that will ensure smooth
and continuous stall-free flow through the
compressor. As compressor rpm increases
with the addition of power, the variable-geo-
metry system moves the inlet guide vanes and
the variable stator vanes to the open position,
allowing unrestricted airflow through the com-
pressor. The response of this system will en-
sure a safe surge margin for the compressor
throughout its operating envelope.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-3FOR TRAINING PURPOSES ONLY
ACCESSORY
GEARBOX
RADIAL
DRIVE
SHAFT
FAN DRIVE
SHAFT
FAN
DRIVE
SHAFT
POWER
TAKEOFF
ASSEMBLY
LOW-PRESSURE
TURBINE SECTION
HIGH-PRESSURE
TURBINE SECTION
COMPRESSOR
SECTION
COMBUSTION
SECTION
FRONT FRAME
FAN SECTION
Figure 7-2. CF34 Engine Modules
Combustor
The combustor includes a straight-flow an-
nular combustion chamber, a liner, and the
first-stage turbine inlet nozzle.
Eighteen swirl injectors are installed in the
combustion chamber to atomize the fuel. Ini-
tial ignitions supplied by two high-energy ig-
nitor plugs. The combustor system ensures
proper mixing of the air-fuel mixture, air di-
lution, and flame containment.
Turbine
The turbine section consists of a HP and
LP turbine.
The two-stage HP turbine is rigidly connected
to the HP compressor by the main rotor shaft.
The turbine extracts sufficient energy from
the expanding gases to drive the HP com-
pressor and the accessory gearbox.
The HP compressor and HP turbine assem-
blies form the HP spool of the engine. The rpm
of the HP spool is designated N2.
The four-stage LP turbine located behind the HP
turbine is rigidly connected to the single-stage
forward fan by a shaft that passes through the
main rotor shaft. The energy extracted by the
LP turbine is used to drive the fan. The remain-
ing energy in the combustion gases is accelerated
rearward to the atmosphere as the core engine’s
contribution to the total engine thrust.
The fan and LP turbine combination constitute
the LP spool. The rpm of the LP spool is des-
ignated N1.
Exhaust
The exhaust frame is located aft of the LP turbine
and consists of an exhaust duct and cone assem-
bly. The exhaust system directs the combustion
gases from the core engine to the atmosphere.
Accessory Gearbox
The accessory gearbox is attached to the lower
side of the front frame. The gearbox is driven
by a tower shaft and bevel gear assembly from
the main (HP spool rotor shaft. The following
accessories are driven by the accessory gearbox:
• N1 speed control alternator
• Integrated-drive generator
• Oil pump
• Fuel pumps and FCU
• Hydraulic pump
In addition to these accessories, an air turbine
starter is mounted on the accessory gearbox to
provide engine cranking through a clutch.
OPERATION
Air entering the nacelle inlet (Figure 7-1) is
accelerated rearward by the fan. A large por-
tion of this air is accelerated to a moderate ve-
locity through the fan bypass duct to contribute
the major portion of the thrust. Some of the air
passing through the fan enters the core en-
gine inlet duct and is progressively increased
in pressure as it passes through the 14 stages
of the HP compressor. The compressor outlet
air is directed rearward to the straight-flow an-
nular combustor. A precise amount of the air
enters the combustion chamber where fuel is
added in the proper proportion by the 18 fuel
injectors. Ignition is provided by two high-en-
ergy ignitor plugs until the engine rpm be-
comes self-sufficient. A large portion of the
air provides dilution and insulation for the
combustion liner. The expanding combustion
gases are directed rearward to the turbine sec-
tion. The two-stage HP turbine extracts enough
energy to drive the HP compressor and the
accessory gear system. The expanding gases
continue rearward to the four-stage LP turbine
which extracts sufficient energy to drive the
fan. The remaining core energy is directed to
the atmosphere by the exhaust duct to con-
tribute to the total engine thrust.
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ENGINE SYSTEMS
The engine systems and engine-related systems
of the Canadair Challenger CL-601-3R are:
• Engine oil system
• Engine fuel system
• Ignition system
• Engine control
• Instrumentation
• Engine starting
• Engine speed control and APR systems
• Engine vibration-monitoring system
ENGINE OIL SYSTEM
General
The engine oil system is completely self-con-
tained and fully automated. The engine oil
system provides for cooling and lubrication of
the engine bearings and the accessory gearbox
in addition to adding heat to the unmetered fuel
system through the oil/fuel heat exchanger. An
oil replenishment system is installed in the
rear equipment bay.
Major Components
Oil Tank
An oil tank is mounted at the 11 o’clock po-
sition on each engine. The tank contains a
gravity filler with a dipstick mounted on the
filler cap. A masterchip detector forms part
of the oil tank drain plug.
Oil Replenishing System
The oil tank can be serviced through the in-
tegral gravity filler or through the replenish-
ment system in the rear equipment bay.
Control
An oil-replenishing control panel, powered
from the battery direct bus (Figure 7-3) con-
tains a power switch, a green power ON light,
and two green oil full switchlights labeled “LH
FULL” and “RH FULL.” These two lights have
a press-to-test feature. A three-position man-
ual selector valve labeled “L,” “OFF,” and “R”
is located adjacent to the control panel. In ad-
dition to selecting the tank for servicing, the
valve controls the power supply to the replen-
ishment pump that supplies oil from the tank
to the selected engine’s oil tank.
Indication
The appropriate oil full switchlight (Figure 7-
3) illuminates when the associated engine oil
tank is full.
Oil Pump
An oil pump containing one pressure element
and six scavenge elements is driven by the ac-
cessory gearbox.
The pressure element provides lubrication of the
main engine bearings and the accessory gearbox.
The scavenge elements provide for direct scav-
enging of the compressor and turbine bearings
and the accessory gearbox.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-5FOR TRAINING PURPOSES ONLY
Figure 7-3. Oil-replenishing Control Panel
A separate two-element scavenge pump provides
for positive scavenging of the fan bearing sump
during all flight attitudes (climb or descent).
Chip detectors are located in strategic areas of
the scavenge system and in the oil tank.
Oil Filter
A disposable filter removes solid particles
from the oil. The filter case includes a bypass
valve and an impending bypass indicator
switch. When the differential pressure across
the filter element exceeds a preset limit, it
causes the impending bypass indicator on
JB4 (battery direct bus) to trip. The indica-
tor must be reset by a reset button on JB4
(Figure 7-4).
Oil Cooler
A conventional oil-to-fuel heat exchanger
mounted on the engine maintains the oil tem-
perature within design limits.
Indication
A transducer on the pump pressure line senses
oil pressure and transmits it to a signal data
converter (SDC).
A resistance bulb in the oil tank provides tem-
perature signals to the SDC.
The SDC divides the signals into two outputs
and transmits them to alternate fiber optic
segments that form the vertical analog scales
of the oil pressure and oil temperature indi-
cators (Figure 7-5). The fiber optic segments
are color-coded red, yellow, and green. These
colors are also painted on the instrument face
outboard of the analog scales.
A blue light at the bottom of each vertical
scale indicates a power-on condition. The oil
pressure indicators are calibrated in psi. The
oil temperature indicators are calibrated in
degrees Celsius.
Low oil pressure is sensed by a switch on the
pressure pump output line. When the switch
closes below 28 psi, the appropriate L or R
LOP light (Figure 7-5) on the lower face of the
indicator will illuminate to indicate that pres-
sure is below design minimums.
NOTE
The SDC operates from two power
sources: battery bus and essential DC
bus. Lose of either power source will
result in loss of alternate segments of
the scales. The indicators will still
provide a reasonably accurate indi-
cation of pressure and temperature.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
7-6 CL 601-3R FOR TRAINING PURPOSES ONLY
S5KJ
OIL FILTER
DS3KJ DS2KJ
LH RH
TEST
LH
TEST
RH
S3KJ S4KJ
CHIP DETECT
RESET
DS5KJ DS4KJ
LH
ENG
IGN B
ENG
OIL CONT
BYPASS
IND
ENG
OIL
POWER
BATT
SHUNT
FUSES
RH
B
A
T
B
U
S
B
A
T
T
E
R
Y
 D
IR
E
C
T
 B
U
S
B
A
T
T
E
R
Y
 D
IR
E
C
T
 B
U
S
20
3
10
50
5
3
3
3
3
3
5
5
7.5 7.5
7.5
3
E7
E6
E9
BOARD
BAT
CONT
EXT AC
PWR
CONT
ESS
PWR
CONT
AUTO
APR
CONT
FUEL
DE-
FUEL
APU
BACKUP
PWR
CONT
IND
APU
START
MAN
ADG
DEPLOY
CONT
SERV
LIGHTS
CBP–E
Figure 7-4. Junction Box 4 (JB4)
Operation
Figure 7-6 illustrates operation of the engine
oil system. The pressure element draws oil
from the tank, develops a pressure, and di-
rects the outflow through the bypass filter. A
relief valve limits pressure to a design value.
The pressure oil is directed through the oil
cooler and is then divided into two delivery
lines. One line is directed through a restric-
tor to the accessory gearbox, the front and
rear fan bearings, and the front compressor
bearing. The second delivery line supplies
high-pressure oil to the second and third com-
pressor bearings and to the front and rear tur-
bine bearings.
The six scavenge elements of the oil pump
provide direct scavenging of all bearings ex-
cept the front three. These forward bearings
are scavenged by a dual-element pump to re-
turn oil to the tank. The common scavenge
line enters the tank through a cyclone deaer-
ator. Oil tank pressure and bearing sump pres-
sure is controlled by an oil tank relief valve
and sump vent regulator acting as a vent and
pressure regulator.
ENGINE FUEL SYSTEM
General
The engine fuel system is an integrated hy-
dromechanical-electronic system. The fuel sys-
tem meters fuel to the combustor to provide for
starting, acceleration, deceleration, and full power
requirements under all operating conditions.
In addition, the fuel system operates the vari-
able-geometry system of the compressor to po-
sition inlet guide vanes and compressor stator
vanes to provide engine stall/surge protection.
Major Components
The major components of the fuel system include
an engine-driven LP pump, heat exchanger, a by-
pass filter, a dual-element HP pump, an inte-
grated hydromechanical-electronic fuel control
unit (FCU), a fuel flow distributor, and 18 fuel
nozzles in the combustor system.
LP Engine-driven Pump
The LP engine-driven pump receives inlet fuel
at the standby pump or main ejector pressure,
increases this pressure, and divides the output
into two flows. One output goes to the heat ex-
changer and fuel filter before reaching the
primary HP element of this three-element
pump. The second output from the LP pump
goes to the secondary HP element. The primary
HP element develops the pressure necessary
for FCU operation. The secondary HP ele-
ment supplies the motive flow fuel to the pri-
mary and scavenge ejectors in the fuel tanks.
Heat Exchanger
The fuel is heated by a liquid-to-liquid (oil-
to-fuel) heat exchanger. It is located down-
stream of the fuel pump LP boost stage element
and upstream of the fuel filter.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-7FOR TRAINING PURPOSES ONLY
-40
0
150
155
OIL
TEMP
163
L R
90
60
30
120
°C
0
25
115
130
OIL
PRESS
L R
60
80
40
30
P
S
I
100
L
O
P
Figure 7-5. Oil Pressure and Temperature
Indicators
FlightS
afety C
anada L
té
e
L
td
.
C
L-6
0
0
-2
B
1
6
 P
ILO
T
 T
R
A
IN
IN
G
 M
A
N
U
A
L
7-8
C
L 601-3R
F
O
R
 T
R
A
IN
IN
G
 P
U
R
P
O
S
E
S
 O
N
L
Y
SUPPLY
OIL PRESSURE
LEGEND
SCAVENGE OIL
FUEL
VENT LINES
ELECTRICAL
SUMP VENT
REGULATOR
B SUMPA SUMP
FAN
BEARINGS
COMPRESSOR
BEARINGS
S
SS
RESTRICTOR
BYPASS VALVE
PRESSURE TRANSDUCER
AND ∆ LOW-PRESSURE
WARNING SWITCH
CHIP
DETECTOR
CHIP
DETECTOR
CHIP
DETECTORS
ACCESSORY
GEARBOX
OIL PUMP
ASSEMBLY
DEAERATOR
OIL TANK
RELIEF VALVE
CHIP
DETECTOR
FILTER
FUEL IN FUEL OUT
OIL
COOLER
IMPENDING BYPASS
INDICATOR
RELIEF VALVE
S S S S S P
COMPRESSOR
BEARINGS
TURBINE
BEARINGS
C SUMP
-40
0
150
155
OIL
TEMP
163
L R
90
60
30
120
0
25
115
130
OIL
PRESS
L R
60
80
40
30
P
S
I
100
L
O
P
OIL TANK
Figure 7-6. Oil System Schematic
Fuel Filter
A bypass fuel filter removes solids from the
fuel. A red pop-out bypass indicator is pro-
vided, as well as a differential pressure switch.
The switch will illuminate an amber FILTER
light (Annunciator Panel Section) on the fuel
control panel and the appropriate master cau-
tion system whenever the pressuredifferential
across the filter exceeds a preset value. It is
from the fuel filter that fuel temperature is
sensed and displayed on the temperature in-
dicator on the fuel control panel.
Fuel Control Unit (FCU)
The fuel control unit is an engine-driven hy-
dromechanical-electronic unit that has a me-
tering section and a computing section. The
metering section includes a mechanical gov-
ernor, a fuel metering valve, a bypass valve,
a pressurizing valve, a thrust lever-operated
shutoff valve, and an electronic control unit
(ECU) for fan rpm control.
The computing section of the FCU contains re-
lief valves and servos to sense engine param-
eters such as rpm (N2), compressor discharge
pressure (P3), compressor inlet temperature
(T2C), and the position of the variable-ge-
ometry system. An amplifier (ECU) operates
a torque motor to control fan rpm (N1).
The primary function of the FCU is to control
core engine rpm (N2) as a function of thrust
lever position. In addition, the FCU modulates
fuel flow to control fan rpm (N1) through the
amplifier (ECU) and the torque motor on the
FCU. (See also “Engine Speed Control and
APR Systems” in this chapter.) Engine accel-
eration and deceleration are controlled by the
FCU, based on internal core pressure (P3),
and inlet temperature (T2C).
The FCU also controls the variable-geometry
system as a function of core engine N2 rpm and
compressor inlet temperature (T2C).
The FCU has a fail-safe schedule in the event of
loss of T2C input. In this case, the variable-ge-
ometry and the acceleration schedules will revert
to a fixed temperature reference. If high thrust
is set at the time of failure, a minor decrease in
thrust may result. If failure occurs at idle thrust,
possible compressor damage can result if an at-
tempt is made to accelerate the engine.
Core engine overspeed in limited three ways:
(1) the N2 governor in the FCU, (2) the com-
puter section (therefore, if the FCU governor
fails, N2 will be limited to less than maxi-
mum allowable transient rpm if the comput-
ing section is operational), and (3) if the
computer or the metering valve servo fails, a
bypass valve will open and reduce fuel flow
to the combustor.
The fan rpm control section of the FCU lim-
its fan rpm as a function of thrust lever posi-
tion (PLA) at power settings representing
takeoff, climb, and cruise. In order to minimize
the thrust lever adjustment during climb, the
fan rpm schedule is biased as a function of fan
inlet temperature (T2). The fan is the primary
thrust producer and fan rpm is used to set
thrust. Fan rpm of both engines should be
matched when the thrust levers are aligned.
Fuel Flow Transmitter
A fuel flow transmitter is located in the me-
tered fuel line from the FCU to provide a cock-
pit indication of fuel flow.
#1 Fuel Manifold
The fuel manifold consists of two separate
180° halves which encircle the combustion
chamber frame. Integral with the continuous
ring are eighteen fuel injector hoses which
connect to eighteen fuel injectors. The fuel in-
jectors are dual-orifice injectors. Each injec-
tor has two independent fuel flow passages, a
primary path for startup and idle conditions and
a secondary path for above idle and higher
power settings. This system allows better fuel
distribution and atomization during entire
range of power settings.
#2 Fuel Drains
Fuel drains from the variable geometry actu-
ators, the FCU, and the combustor and turbine
sections are routed overboard.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-9FOR TRAINING PURPOSES ONLY
FlightS
afety C
anada L
té
e
L
td
.
C
L-6
0
0
-2
B
1
6
 P
ILO
T
 T
R
A
IN
IN
G
 M
A
N
U
A
L
7-10
C
L 601-3R
F
O
R
 T
R
A
IN
IN
G
 P
U
R
P
O
S
E
S
 O
N
L
Y
ON
INOP
VALVE
CLOSED
FILTER
LOW
PRESS
FCU
JET PUMP
BOOST PUMP
18
INJECTORS
120
FUEL
80
40
0
-40
L
120
80
40
0
-40
R
TO RIGHT COLLECTOR TANK
TO APU FUEL LINE FEEDBACK
LINK
FIREWALL
SOV’S
LEFT COLLECTOR TANK
INLET
GUIDE
VANES
STATOR
VANES
JET PUMP MOTIVE FLOW FUEL
LP PUMP
FUEL
HEATER
FILTER
BYPASS
FUEL TEMP
INDICATOR
PUMP BYPASS TORQUE
MOTOR
HP PUMP
N1
T2
N2
AMPLIFIER
TRANSDUCER
THROTTLE
LEVER
VARIABLE
GEOMETRY
ACTUATORS
P3 T2C N2
FUEL FLOW
COMBUSTOR
SUPPLY
LP PUMP PRESSURE
HP PUMP PRESSURE
LEGEND
SIGNAL/CONTROL
OIL
AIR
ELECTRICAL
MECHANICAL
0
200
3500
4000
3000
2000
1000
800
FUEL
FLOW
L
x10
P
P
H
R
400
600
Figure 7-7. Fuel System Schematic
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
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Operation
Figure 7-7 illustrates operation of the fuel
system in its simplest form. Initial fuel pres-
sure is supplied by the collector tank standby
pump and later by the main ejector through the
open firewall shutoff valve to the LP engine-
driven pump.
The LP pump increases fuel pressure and di-
rects fuel through the heat exchanger and fil-
ter and to the dual-element HP fuel pump. The
#1 HP element produces the high fuel pressure
required by the FCU. The # 2 HP element sup-
plies the motive flow for main and scavenge
ejector operation. The metering section of the
FCU, in response to the computing section
signals, meters fuel through the flowmeter to
the fuel manifold. From the fuel manifold,
fuel is supplied in precisely equal amounts
through the 18 injectors in the combustor.
During this operation, the variable-geometry
section of the FCU, operating in response to
inputs representing N2 rpm and compressor
discharge pressure, directs fuel pressure to
the variable-geometry actuators to position
the inlet guide vanes and stator to produce a
safe surge margin across the compressor. At
the same time, guide vane and stator position
is fed back to the FCU.
NOTE
When the engine is static, the guide
vanes and stators are at their design
maximum closed position. As the en-
gine starts, the guide vanes’ and sta-
tors’ position will change until, at
high power setting, both the guide
vanes and stators will be at the design
full open position permitting maxi-
mum airflow through the core engine.
IGNITION SYSTEM
General
The CF34-3A1 series engine has a dual high-
energy, capacitor discharge type ignition.
The ignition system for each engine consists
of an ignitor plug A and an ignitor plug B in
the combustor with each ignitor powered
through its own exciter.
Operation of either ignitor is sufficient to pro-
vide for a normal engine start. The ignitor cir-
cuits for each engine are identified as “ignition
A” and “ignition B.”
Ignition Modes
The ignition system has four modes, as follows:
1. Ground start ignition
2. In-flight ignition
3. Continuous ignition
4. Auto (stall) protection ignition
Ground Start Ignition
The ground start ignition is integrated with the
engine start system from initiation of start to
the termination of start at 55% N2.
Either ignition A or ignition B, or both, may
be armed for operation during a ground start
cycle. It is recommended, however, that only
one ignition circuit be armed to prolong igni-
tion plug life.
Control and Indication
Ignition control and indication is provided on
the start and ignition control panel (Figure 7-
8) located on the overhead panel. Two split-
legend switchlights are used to arm A and/or
B ignition circuits for ground starting. Push-
ing either switchlight will illuminate the leg-
end IGN A or IGN B. This indicates that the
selected system is armed to its associated en-
gine START switchlight (Figure 7-8).
Pushing a START switchlight will illuminate
the green START legend and, simultaneously,
the white ON legend of the selected ignition
switchlight, indicating that power is being ap-
plied to the selected ignition exciter. Ignition
will continue until the start cycle is terminated.
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CL-600-2B16 PILOT TRAINING MANUAL
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Ignition and start termination will be indi-
cated when the ignition ON light and the
START light extinguish.
Ignition and start may be terminatedat any
time prior to 55% N2 by pushing the appro-
priate STOP switchlight.
In-flight Ignition
In-flight ignition is separate for each engine. It
is designed to provide dual ignition for wind-
milling relights or during single-engine operation.
Control and Indication
The in-flight ignition system is controlled by
a switchlight (Figure 7-8) for each engine la-
beled “IN FLIGHT START.”
Pushing in one of these switchlights will il-
luminate the green IN FLIGHT START legend
as well as the white ON legend of both igni-
tion arming switchlights, completing the cir-
cuit for operation of ignition A and ignition
B for the associated engine. It is not necessary
to arm the ignition A or ignition B systems
prior to selecting in-flight start ignition.
Continuous Ignition
and Indication
Continuous ignition is primarily used as an
anti-flameout ignition. When selected, it pow-
ers one ignition exciter continuously on both
engines. The system is activated by a single
switchlight (Figure 7-8) labeled “CONT IGN”
only if either ignition A and/or ignition B has
been armed. The green CONT IGN legend,
the green IGN legend, and the white ON leg-
ends of the selected ignition system will all il-
luminate during operation of the continuous
ignition system.
Automatic (Stall Protection)
Ignition
The automatic or stall protection ignition sys-
tem provides anti-flameout protection during
periods of engine inlet turbulence caused by
high angles of attack.
Control
The automatic ignition system is controlled by
the stall protection computer using inputs
from the angle-of-attack vanes. The stall pro-
tection computer will initiate ignition A and
ignition B for both engines 3% before the
onset of the stick shaker and maintain ignition
operation until the angle of attack is reduced.
Indication
When the stall protection ignition is operat-
ing or during a stall protection system test, the
white IGN A ON and IGN B ON (Figure 7-8)
lights will be illuminated.
Power Sources
AC power at 115 volts and 400 Hertz is used for
the ignition system. Left and right engine igni-
tion A is supplied from the essential AC bus.
Left and right engine ignition B is supplied
from the battery bus through a static inverter.
Operation
Figure 7-9 is simplified schematic of the ig-
nition system used on the Canadair Challenger
CL-601-3R.
IGNITION
ENGINE START
R
START
CONT
IGN
STOP
IN
FLIGHT
START
START
STOP
IN
FLIGHT
START
L
IGN A
ON
IGN B
ON
Figure 7-8. Start and Ignition
Control Panel
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
FOR TRAINING PURPOSES ONLY CL 601-3R 7-13
Figure 7-9. Ignition System Schematic
AC ESSENTIAL BUS
BATTERY BUS
A IGNITER POWER
LEGEND
B IGNITER POWER
FROM STALL PROTECTION
STATIC INVERTER
FROM STALL
PROTECTION
SYSTEM
STATIC
INVERTER
AC ESSENTIAL BUS
C-28
BATTERY BUS
B-169
BA
A B
LEFT ENGINE
IGNITION
CONTROL
RELAY
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The switchlights on the ignition and start con-
trol panel provide for ignition arming and se-
lection of the ground start ignition mode, the
continuous ignition mode, and the in-flight
start ignition mode.
The stall protection system provides ignition
of the duration of stall warning regardless of
the position of all other ignition switches.
ENGINE POWER CONTROL
General
Engine power control is provided on a quad-
rant located on the center pedestal.
Thrust Levers
The individual engine thrust levers (Figure 7-
10) operate in quadrant slots from a full aft po-
sition labeled “SHUT OFF” to a full forward
position labeled “MAX POWER.” An inter-
mediate position forward of SHUT OFF is la-
beled “IDLE.”
A mechanical latch at the rear and below each
thrust lever knob must be raised before the
thrust lever can be moved to or from the SHUT
OFF position.
A go-around button is mounted in each thrust
lever knob. When either is pushed, it will dis-
engage the autopilot and place the AFCS sys-
tem in the go-around mode. Switches are
mounted in the throttle quadrant slots to provide
(1) takeoff configuration warning for flaps,
spoilers, and horizontal stabilizer, (2) pressur-
ization ground control mode, and (3) landing
configuration warning (landing gear not down
and locked at landing power settings).
Quadrant Friction Control
A single friction adjustment twist knob (Figure
7-10) is located on the quadrant aft of and be-
tween the thrust levers. Clockwise rotation will
increase friction, and counterclockwise rota-
tion will decrease friction.
Thrust reverse control levers are mounted pig-
gyback fashion on the thrust levers. Thrust re-
versers will be discussed later in this chapter.
ENGINE INSTRUMENTATION 
General
The primary engine instruments (Figure 7-
11) are horizontally mounted at the top left side
of the center instrument panel. From left to
right these instruments are as follows:
• N1 (fan rpm)
• ITT (interturbine temperature)
• N2 (core or gas generator rpm)
• Fuel flow
Each instrument has two vertical scales: one
for the left engine and one for the right engine
which provides a nonlinear analog readout.
Below each vertical scale (except OIL TEMP
and OIL PRESS) is a three-digit digital read-
out. To increase safety factors, each indicator
is cross-powered using two power sources;
Figure 7-10. Throttle Quadrant
for example, the left analog scale and the right
digital scale have the same power source. A
separate power source is used for the right
analog and left digital scale. This ensures that
a single power failure will not result in total
readout loss on any engine instrument.
The analog scales are made up of separate
colored segments. These segments are pro-
gressively illuminated from groups of bulbs
with the instrument. The light is transmitted
to the scale segments by fiber optics. The col-
ored segments provide for safe (green), cau-
tion (yellow), and warning (red) indications.
The bottom segment in each vertical scale is
power indicator that will show blue if power
is available to the scale.
Power Sources
The engine instruments are powered from a
signal data converter (SDC). The SDC is sup-
plied DC power from the battery bus and the
essential DC bus. The SDC processes the in-
puts from the various engine parameters and
produces two outputs. These outputs are sup-
plied to the instrument lamp banks. Fiber op-
tics transmit the light from the lamp banks to
the colored segments of the vertical scales.
The digital displays are converted from the as-
sociated analog displays. When compared
with the nonlinear analog readout, the digital
indicators provide a more accurate indication.
Automatic Dimming
A photoelectric cell (Figure 7-12) is provided
on the engine instrument control panel to pro-
vide for automatic engine instrument dim-
ming as ambient light conditions change. A
rheostat on the same panel allows the crew to
set brilliancy to personal preferences.
Instrument Testing
The power supplies of the SDCs are tested
with a three-position TEST switch (Figure 7-
12). Selecting the switch to position 1 or 2 tests
the corresponding SDC power supply by il-
luminating all analog and digital displays in-
cluding the fuel panel.
Indications
The amber light above the instrument test
switch (Figure 7-12) will illuminate when a
power input source to the SDC fails. In this
case, the blue power-on segments of the af-
fected scales will extinguish, and the associ-
ated analog and opposite digital display will
be lost.
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CL-600-2B16 PILOT TRAINING MANUAL
CL 601-3R 7-15FOR TRAINING PURPOSES ONLY
100
N1
%RPM
L R
98.6
96.2
90
60
30
0
104
N2
%RPM
L R
60
40
20
0
99.2
99.4
98.2
80
0
200
3500
4000
3000
2000
1000
800
FUEL
FLOW
L
x10
P
P
H
R
400
600
0
200
900
928
800
1000
ITT
°C
L
DGT
OFF
R
500
400
300
600
700
860
970
Figure 7-11. Engine Instruments
AUX PWR
TEST
1 2
D
I
M
Figure 7-12. Engine Instrument
Control Panel
N1 (Fan) RPM
N1 (fan) rpm(Figure 7-11) is sensed by a
monopole transmitter located on the engine
front frame. Electrical signals are sent via the
SDC to the appropriate analog and digital
scale. Each scale is calibrated to indicate a per-
centage of N1 rpm from 0% to 100%.
ITT (Interturbine Temperature)
Thermocouples of different lengths are equally
spaced between the LP turbine and the HP
turbine. The thermocouples are parallel-con-
nected. The ITT output is sent to the appro-
priate vertical and digital scales of the ITT
indicator via the SDC. The ITT indicator scales
are calibrated in degrees Celsius from 0 to
1000°C.
A red light (Figure 7-11) above each verti-
cal scale will illuminate if the ITT reaches
899° C. These lights also illuminate during
the instrument test.
A two-position switch labeled “DGT OFF”
(Figure 7-11), located at the bottom of the ITT
panel, allows the crew to extinguish all engine
parameter digital displays which might be an-
noying on extended nighttime operations.
N2 RPM
N2 rpm (Figure 7-11) is supplied by an alter-
nator driven by the accessory gear. The rpm
signals are isolated from the alternator’s power
to eliminate interference and interruption. The
rpm signals are sent to the appropriate N2
scales via the SDC.
Fuel Flow
Fuel flow (Figure 7-11) is sensed by a mass
flow transmitter located downstream of the
FCU. The transmitter output is sent to the
SDC for processing into analog and digital
readout for display on the appropriate fuel
flow indicators.
The analog scales are calibrated in pounds of
fuel per hour from 0 to 4,000. The digital dis-
plays are in pounds per hours times 10.
ENGINE STARTING
General
Engine starting is divided into ground starts,
starter-assisted airstarts, and windmilling airstarts.
Starter
The engine starter is an electrically controlled
air turbine starter (ATS). Starter output is ap-
plied through a clutch to the accessory gear
which, in turn, rotates the HP spool.
A speed sensor operated by the ATS governor
automatically terminates the starter cycle at
approximately 55% N2 rpm. The average start
cycle is less than 40 seconds. A time-delay
relay is armed when a start cycle is initiated,
and, if the ATS operation continues for more
than 60 seconds, the time-delay relay will
open an illuminate the amber STOP switchlight
on the start and ignition control panel (Figure
7-8). The STOP switchlight may be pushed to
terminate the start sequence at any time below
55% N2 rpm.
ATS Air Sources
The air source for ATS operation can be (1)
APU bleed air, (2) an external air source, or
(3) cross bleed from an operating engine. The
minimum air pressure for starting is 45 psi.
(See Chapter 9, “Pneumatics.”)
Ground Start (APU Air)
Engine starting should not be at-
t e m p t e d u n t i l t h e Wa l k a r o u n d
checklist and the Cockpit checklist
are completed.
To initiate a ground start using APU bleed
air, push the APU bleed-air switchl ight
(Figure 7-13). The OPEN light will illu-
minate, and the left scale of the bleed-air
pressure indicator (Figure 7-14) should
show approximately 50 psi. Push IGN A or
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WARNING
IGN B switchlight (Figure 7-8) to arm an igni-
tion system, and check that the applicable green
light illuminates.
Push and hold the appropriate START switch-
light for 2 seconds. The green START light will
illuminate as will the ON light in the selected
ignition switchlight. The ISOLation valve
OPEN light will also illuminate. The left and
right engine BLEED CLOSED lights will ex-
tinguish. Verify engine rotation on the N2 rpm
indicator and monitor N2 until it reaches 20%
minimum and ITT below 120° C. Then move
the affected thrust lever to IDLE, check the ITT
indicator for light-off, and continue to moni-
tor ITT, oil pressure, and N2 rpm. Also check
that N1 rpm is increasing in relation to N2. At
approximately 55% N2 rpm, the START light
and the IGNition ON light (Figure 7-8) will
both extinguish as should the ISOLation valve
OPEN light. The left and right engine BLEED
CLOSED lights (Figure 7-14) should both il-
luminate. Continue to monitor all engine-re-
lated instruments until the engine stabilizes at
idle rpm (approximately 60–64% N2). The N2
variation between engines at idle should be
within 2%.
NOTE
The idle N2 rpm of CF34 engines
automatically varies as a function of
compressor inlet temperature (T2C).
In case of faulty T2C input, an IDLE
FLOOR STOP is provided to prevent
N2 from decreasing below 56.9%.
If idle speed stabilizes at approxi-
mately 57% N2, the engine must be
shut down immediately and the con-
dition reported to maintenance.
Do not attempt to increase idle N2 by
advancing the thrust lever because it
can result in serious damage to the
first-stage compressor blades.
Ground Start (External Air)
The procedures for engine starting using an ex-
ternal air supply are identical with those for
APU bleed-air starts. An approved external air
unit capable of 45 psi can be connected to the
adapter (Figure 7-15) located in an access on
the left side of the rear fuselage.
Figure 7-16 illustrates the use of bleed for
ground starting the first engine.
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
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C
O
N
T
R
O
L
A
P
U
PUSHPUSH PUSH
PWR FUEL
ON/OFF
APU
OIL
ADPTR
OIL
BLEED
AIR
START/
STOP
STARTER
APU
READY
LO
PRESS
HI
TEMP
SOV
CLOSED
PUMP
INOP
LO
PRESS
HI
TEMP
FAILED
OPEN
%RPM
100
80 0
60
40
20
EGT
°C X 100
8
6 2
4
0
10
Figure 7-13. APU Control Panel
A
I
R
B
L
E
E
D
PUSH
ON/
OFF
PUSH
ON/
OFF
14TH STAGE 10TH STGR
DUCT MON
LOOP A
LOOP B
10TH STAGE
ISOL ACUL LR R
BOTH
OFF
FAIL
OFF
FAIL
BLEED
AIR
R
PSI
100
50
0
L
100
50
0
L
BLEED
CLOSED
DUCT
FAIL
BLEED
CLOSED
DUCT
FAIL
BLEED
CLOSED
DUCT
FAIL
OPEN
BLEED
CLOSED
DUCT
FAIL
Figure 7-14. Bleed-Air Control Panel
CAUTION
FlightSafety Canada LtéeLtd.
CL-600-2B16 PILOT TRAINING MANUAL
7-18 CL 601-3R FOR TRAINING PURPOSES ONLY
LEFT ENGINE
ATSATS
APU
10TH STAGE
BLEED-AIR
LEFT START
VALVE
LEFT 
BLEED-AIR
SOV
EXTERNAL AIR
LCV
ISOLATION
VALVE
RIGHT START
VALVE
RIGHT
BLEED-AIR
SOV
RIGHT ENGINE
10TH STAGE
BLEED-AIR
APU BLEED AIR
LEGEND
Figure 7-16. Bleed-Air Sources (First Engine Start Schematic)
LEFT ENGINE
ATSATS
APU
10TH STAGE
BLEED-AIR
LEFT START
VALVE
LEFT 
BLEED-AIR
SOV
EXTERNAL AIR
LCV
ISOLATION
VALVE
RIGHT START
VALVE
RIGHT
BLEED-AIR
SOV
RIGHT ENGINE
10TH STAGE
BLEED-AIR
10TH-STAGE BLEED AIR
LEGEND
Figure 7-17. Cross Bleed Start (Left Engine from Right Engine Schematic)
Figure 7-15. External Air Supply Adapter
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Ground Start (Cross Bleed Air)
The procedures for a ground start using a cross
bleed-air supply are similar to those for APU
bleed or external air source, except that the
APU bleed air (Figure 7-13) must be off .
Push the BLEED AIR switch (Figure 7-14)
of the operating engine and check that the
bleed-air pressure is 45 psi minimum, then
continue as for APU bleed-air start.
Figure 7-17 illustrates the availability of bleed
air during a cross bleed start of the left engine.
NOTE
Two conditions must be met before
moving the thrust lever to IDLE for
all engine starting:
1. Indicated ITT must be less than
120°C.
2. N2 rpm must be 20% minimum.
If ITT is greater than 120°C prior to start,
the engine must be dry motored until ITT
drops below 120°C.
NOTE
When using battery or external DC
power only during engine starting,
bleed-air pressure indication will
not be available.
Failure to Start
Light-off as indicated by rising ITT will nor-
mally occur within 10 seconds after moving
the thrust lever to IDLE.
A start should be aborted if light-off does not
occur 25 seconds after moving the thrust lever
to IDLE.
If starter operation continues for more than 60
seconds, the time delay relay will cause the
STOP lightto illuminate. At temperatures
above 15°C (59°F) the start should be aborted
by pushing the STOP switchlight anytime up
to 55% N2 rpm. The thrust lever should then
be moved to SHUT OFF; then wait one minute
before attempting another start. At tempera-
tures below 15°C (59°F) the start sequence may
exceed 60 seconds (Figure 7-18).
Before attempting another start, dry motor
the engine with both ignition systems off and
the affected thrust lever at SHUT OFF.
NOTE
The a i r t u rb ine du ty cyc l e f o r
normal engine start is 3 consecutive
cycles with 5 minutes cooling be-
tween addition cycles.
For dry motoring, the ATS duty cycle is 90 sec-
onds with a 5-minute cooling period between
additional cycles of 30-second duration.
Airstarts
Airstarts are divided into starter-assisted and
windmilling airstarts.
Starter-assisted Airstarts
The procedure for starter-assisted airstarts
(cross bleed starts) are identical with those ex-
plained previously for cross bleed starts.
All in-flight starts must be performed within
the airstart envelope (Figure 7-19).
The thrust lever should not be moved to IDLE
during airstarts unless ITT is less than 90°C.
Windmilling Airstart
Windmilling airstarts are obtained at the fol-
lowing airspeeds:
Below 10,000 feet......................... 300 KIAS
10,000 to 21,000 feet ..... 300 KIAS to VMO
N2 must be stable or increasing.
Airstarts, windmilling or starter-
assisted, should not be attempted if
the flameout or shutdown is ac-
companied by unusual noise or
other indications that mechanical
damage may exist.
Prior to initiating a windmilling airstart, all
checklist items affecting the start must be
completed. Then push the appropriate IN
FLighT START switchlight. The green light
and the IGN A and B ON lights will illuminate.
Advance the thrust lever to IDLE and moni-
tor all engine-related instruments until the en-
gine is stabilized. Then push the IN FLighT
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AIRSTART ALTITUDE LIMIT
21
10
0
0 10 13 N2% RPM
55
WINDMILL START
10,000–21,000 FT
STARTER ASSIST
BELOW 21,000 FT
WINDMILL START
BELOW 10,000 FT
Figure 7-19. Airstart Envelope
WARNING
20
20
0
–20
–40
40
60
40 60 80 100 1200
O
U
T
S
ID
E
 A
IR
 T
E
M
P
E
R
A
T
U
R
E
 —
 °
F
TOTAL TIME
TO STABILIZED
IDLE
TIME FROM
THROTTLE OPENING
TO LIGHT-OFF
Figure 7-18. Maximum Allowable Start Time and Time to Stabilized Idle—Seconds
START switchlight again. The green light and
the IGN A and B ON lights will extinguish,
then complete the After Start checklist.
ENGINE SPEED CONTROL
AND APR SYSTEMS
General
The automatic performance reserve (APR) is
a solid-state system which constantly moni-
tors the thrust of both engines during takeoff.
If significant power loss occurs in either en-
gine, it will instantaneously command an N1
(thrust) increase.
Components
The APR system components include an APR
control panel (Figure 7-20), an APR controller,
and an N1 speed selector switch for each en-
gine. In addition, the APR system utilizes the
torque motors (discussed earlier in Engine
Fuel System) and the amplifiers (ECUs) as-
sociated with the FCUs.
Control and Indication
A three-position switch (Figure 7-20) with
pos i t i ons l abe l ed “ARM,” “OFF,” and
“TEST/RESET” is located on the APR control
panel. When this switch is at the ARM posi-
tion, the system is armed provided that three
conditions exist:
1. Both engine speed control switches on
2. Both engines above 79% N1
3. No faults sensed by the integral mon-
itoring system
Selecting the APR switch off deactivates
the system. The TEST/RESET position is
spring-loaded to the off position and is used
for testing.
A split legend green light labeled “L.ON” and
“R. ON” will illuminate following an APR
trigger (activating) to confirm proper response
(N1 increase) on the serviceable engine.
A green light labeled “READY” forms the
upper part of a dual legend light which illu-
minates to confirm APR readiness above 79%
N1 if the system is armed. If a subsequent APR
trigger occurs, the READY light will extin-
guish as the L. ON or R. ON light illuminates.
An amber APR light and MASTER CAUTION
lights will illuminate as a crew warning that
either (1) the APR system is not armed for
takeoff or (2) that the APR system has failed
for one or more of the following reasons:
• Either the static or dynamic test is not valid.
• The serviceable engine’s response to an
APR trigger produces less than 2% N1
rpm increase within 2 seconds.
• The ECU input voltages are outside ac-
ceptable limits.
• The monitoring system detects failure
of the microcomputer or the two inter-
nal power supplies.
• Either or both N1 input signals are out
of limits.
• Battery input voltage fails.
• The two WOW inputs disagree.
• A unwarranted APR command is triggered.
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Figure 7-20. APR Control Panel
NOTE
The APR system is used only for
takeoff and is then disarmed. The
APR fail light is inhibited in flight
through WOW logic and for landing
by flap 45° selection.
A green TEST light forms the lower portion
of the READY light. During testing, this light
will illuminate as the last indication in a se-
ries until the APR switch is released from the
TEST/RESET position.
Testing
Two tests are associated with the APR sys-
tem: (1) static test and (2) dynamic test.
Static Test
Holding the APR switch in the TEST/RESET
posi t ion causes the fol lowing funct ions
and indications:
1. The system is reset, the APR program
is restarted, and all previous perfor-
mance data in the memory is cleared.
2. Validates the battery direct bus voltage
input. If not present, the amber APR
light will illuminate.
3. Tests all lamps for 1 second each in
the following sequence:
a. Ready and L. ON
b. Ready and R. ON
c. Test
d. APR
e. Test (will remain on as long as TEST
is held)
4. If any faults are detected, the amber
APR light will illuminate.
Dynamic Test
The dynamic test is done just prior to flight.
It is automatically performed for both engines
by the APR controller and verifies that the
APR system is operational. To perform the
test, the APR switch is armed. Advance both
thrust levers to obtain an indicated N1 above
83% (note the READY light came on at 79%
N1). The APR controller samples the fan rpm
continuously to determine if they are stable;
that is, fan rpm does not vary more than a pre-
determined amount. The dynamic test initiates
an APR trigger to both APR amplifiers which
causes both engines to accelerate slightly. If
the test is valid, the TEST light will momen-
tarily flash. If outside the permissible limit,
the amber APR light will illuminate, accom-
panied by the MASTER CAUTION light.
Operation
Before takeoff, the static and dynamic tests are
performed and determined as valid, both en-
gine speed control switches are on, and the
APR switch is at ARM. The green READY
light will illuminate after 79% N1. The amber
APR light and the green L. ON and R. ON
lights and the TEST light are extinguished.
Figure 7-21 illustrates an APR trigger. Both
fuel control amplifiers (FCUs) are receiving N1,
N2, T2, and power lever angle (PLA). The right
engine N1 has decreased below N1 speed con-
trol (79% N1) and has reached the APR “trig-
ger” speed (approximately 68% N1). The APR
controller sends a signal to both amplifiers to
increase N1. The left engine responds since it
is still on N1 speed control (above 79% N1) and
illuminates the green L. ON legend when it has
increased the required amount (approximately
2% N1). The right engine does not respond
since it is not on N1 speed control.
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Figure 7-21. APR/Engine Speed Schematic
FUEL CONTROL