Aeroplane Flight Simulator Evaluation Handbook Volume 1

Transcription

Third Edition June 2005
ISBN 1 85768 154 1
Digitally printed by
Colourtech, Ashford, Kent
www.colourtechgroup.com
INTERNATIONAL STANDARDS FOR
THE QUALIFICATION OF AEROPLANE
FLIGHT SIMULATORS
AEROPLANE FLIGHT SIMULATOR
EVALUATION HANDBOOK
THIRD EDITION
JUNE 2005
Evaluation Handbook 3rd Edition
PREFACE
In spite of the technological advances in both the gathering and processing of aircraft
flight test data and the development and proving of mathematical models for flight
simulators, there are still many unknown factors present within any flight test data which
cause considerable difficulties when those data are used in the design of a flight
simulator. One purpose of this document, therefore, is to assist all sections of the
industry in applying "engineering judgment" to those situations. Hopefully, it will also
provide guidance for constructing a Qualification Test Guide and conducting simulator
evaluation tests. This Handbook, known as Volume 1, is complemented by a second,
separate volume aimed at providing guidance in the conduct of the Functions and
Subjective Tests - as distinct from the Validation Tests, which are the primary subject
of this volume.
In 1995, the ICAO Manual was published by the International Civil Aviation Organisation
(ICAO) as the “Manual of Criteria for the Qualification of Flight Simulators” (hereafter
known as the ICAO Manual). During 2001 an international working group under the
joint chairmanship of the European Joint Aviation Authorities and the United States
Federal Aviation Administration reviewed and modernised the Standards contained in
the Manual. The majority of the validation tests of Appendix B of the ICAO Manual
(what was Appendix 2 in the original RAeS International Standards document) were
revised. A few tests were added, and about an equal number were deleted. This Third
Edition of the Evaluation Handbook is intended to be a companion to the Second
Edition of the ICAO Manual. There are additional improvements to expand the
usefulness of the Handbook both for flight simulator evaluation and for planning and
conducting validation tests.
The document, in common with the general atmosphere of cooperation found within the
flight simulation industry has been generated from contributions received from
interested parties worldwide. Nevertheless it is not intended that the Evaluation
Handbook is referred to or quoted as being the definitive reference source for
determining policy - that function remains with the regulatory authorities themselves.
It is hoped that this Third Edition of the Handbook will continue to develop as a useful
and 'living' document, and provide some of the background and details on testing and
evaluation needed by those engineers, pilots, managers and regulatory authorities who
are entrusted with the complex task of evaluating an aeroplane flight simulator.
Thanks are once again due to many individuals and organisations within the flight
simulation industry and it seems fair to acknowledge their valued co-operation, without
which this Third Edition could not have been produced.
M I Blackwood
June 2005
i
DISCLAIMER
The Royal Aeronautical Society does not accept responsibility for the
technical accuracy nor for the opinions expressed within this publication.
Published by
The Royal Aeronautical Society , 4 Hamilton Place, London, W1J 7BQ
© Royal Aeronautical Society 2005
ii
AMENDMENT RECORD
REV
DESCRIPTION/SHEETS DATE
AFFECTED
COMMENT
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
Original
All
01/01/92
First Draft
A
All
05/05/92
Second Draft
B
All
01/04/93
First
Edition
Circulation)
2nd Edition
(Total Re-issue)
01/02/95
First Full Publication
2nd Edition
(Reformatted for PDF)
01/02/02
First Issue to the World Wide
Web
3rd Edition
(Total Re-issue)
01/06/05
Major Update as a result of
changes to the ICAO
Standards in 2001.
(Limited
iii
LIST OF CONTENTS
CHAPTER TITLE
PAGE
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
LIST OF FIGURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . vii
1.0
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xx
1,1
BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xx
1.2
SIMULATOR STANDARDS . . . . . . . . . . . . . . . . . . . . . . xx
1.3
VALIDATION TESTS . . . . . . . . . . . . . . . . . . . . . . . . . . . xx
1.4
QTG REVIEW TECHNICAL EVALUATION . . . . . . . . xxiii
1.5
FUNCTIONS AND SUBJECTIVE TESTS . . . . . . . . . xxvi
2.0
AUTOMATIC TESTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.1
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.2
BASIC METHODOLOGY . . . . . . . . . . . . . . . . . . . . . .
2.3
PASS/FAIL CRITERIA . . . . . . . . . . . . . . . . . . . . . . . .
3.0
INTEGRATED TESTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxx
3.1
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxx
3.2
BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxx
3.3
ENGINE MODELS . . . . . . . . . . . . . . . . . . . . . . . . . . . xxxii
3.4
AVIONICS FITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxxiii
4.0
MANUAL TESTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.1
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2
OBJECTIVES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3
TECHNIQUES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.0
FUNCTIONS AND SUBJECTIVE TESTS . . . . . . . . . . . . . . xxxvii
5.1
DISCUSSION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxxvii
5.2
TEST REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . xxxvii
6.0
COMPUTER CONTROLLED AEROPLANES . . . . . . . . . . . xxxix
6.1
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxxix
6.2
DISCUSSION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxxix
6.3
APPLICABLE DEFINITIONS . . . . . . . . . . . . . . . . . . . . . xl
6.4
ADDITIONAL FLIGHT TESTS . . . . . . . . . . . . . . . . . . . . xli
6.5
ABBREVIATIONS USED . . . . . . . . . . . . . . . . . . . . . . . . xli
6.6
NOTES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xli
7.0
PRESENTATION OF SIMULATOR TEST RESULTS . . . . . . . xlii
iv
xxvii
xxvii
xxvii
xxix
xxxv
xxxv
xxxv
xxxv
7.1
7.2
7.3
7.4
7.5
7.6
7.7
ACCURACY OF TABULATED DATA . . . . . . . . . . . . .
PLOT SCALES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
FLIGHT TEST DATA PROBLEMS . . . . . . . . . . . . . . .
SIMULATOR/FLIGHT-TESTED AIRCRAFT . . . . . . . .
CONFIGURATION DIFFERENCES
NAMES FOR AEROPLANE VARIABLES . . . . . . . . . .
SELECTING FLIGHT TEST RESULTS . . . . . . . . . . . .
TEST PARAMETERS TO BE RECORDED . . . . . . . . .
xlii
xlii
xliii
xliv
xliv
xlvi
xlvi
8.0
CONFIGURATION CONTROL . . . . . . . . . . . . . . . . . . . . . . . .
8.1
GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.2
MAINTENANCE CONSIDERATIONS . . . . . . . . . . . .
8.3
ENGINEERING CHANGE CONTROL SYSTEM . . . .
8.4
SOFTWARE CONFIGURATION CONTROL . . . . . . .
8.5
AEROPLANE CONFIGURATION CONTROL . . . . . .
9.0
REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l
10.0
LIST OF CONTRIBUTORS . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10.1 AIRCRAFT MANUFACTURERS . . . . . . . . . . . . . . . . . .
10.2 SIMULATOR/VISUAL MANUFACTURERS . . . . . . . . . .
10.3 SIMULATOR OPERATORS . . . . . . . . . . . . . . . . . . . . .
10.4 REGULATORY AUTHORITIES . . . . . . . . . . . . . . . . . . .
10.5 OTHERS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10.6 CONTRIBUTIONS TO THIS HANDBOOK . . . . . . . . . . .
11.0
TYPICAL QTG TEST INDEX . . . . . . . . . . . . . . . . . . . . . . . . . . . lvi
12.0
EVALUATION NOTES: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . lxviii
12.1 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . lxviii
12.2 OTHER COMMENTS . . . . . . . . . . . . . . . . . . . . . . . . lxviii
SECTION 1
SECTION 2
SECTION 3
SECTION 4
SECTION 5
PERFORMANCE . . . . . . . . . . . . . . . . . . . . . .
HANDLING QUALITIES . . . . . . . . . . . . . . . . .
MOTION SYSTEM . . . . . . . . . . . . . . . . . . . . .
VISUAL SYSTEM . . . . . . . . . . . . . . . . . . . . . .
SOUND SYSTEM . . . . . . . . . . . . . . . . . . . . . .
xlvii
xlvii
xlvii
xlviii
xlviii
xlviii
liii
liii
liii
liii
liv
liv
liv
1-1
2-1
3-1
4-1
5-1
APPENDIX A
FLIGHT TEST DATA CONSIDERATIONS . . . . . . . . . A-1
APPENDIX B
DYNAMIC DATA ANALYSIS . . . . . . . . . . . . . . . . . . . B-1
APPENDIX C
EXAMPLE COMPLIANCY STATEMENTS . . . . . . . . . C-1
v
APPENDIX D
MOTION SYSTEM ENVELOPE . . . . . . . . . . . . . . . . . D-1
APPENDIX E
DISCUSSION OF MATH PILOTS . . . . . . . . . . . . . . . . E-1
APPENDIX F
THE ELECTRONIC QTG . . . . . . . . . . . . . . . . . . . . . . F-1
vi
LIST OF FIGURES
There follows a list of diagrams in the order in which they appear. Note that the
diagrams are included for general information only and have been deemed to
form a reasonable cross-section of the listed tests. There is no intention to
suggest that any of them fulfil the precise requirements of the test to which they
belong - indeed many of them have been specifically chosen because they fail
to fulfil those requirements. In any case the reader should be aware they will
differ in many details from those plots encountered during the evaluation of a
given simulator. The comments associated with the diagrams are intended to
either provide insight into that test, or to make a comment which can be applied
more generally across a range of tests.
FIGURE #
TITLE
PAGE
1a1-1
Example of Simulator Test Results for Minimum
Radius Turn
1A-4
1a1-2
Example of OPS Manual Data for Minimum Turn
Radius
1A-5
1a2-1
Example of Simulator Test Results for Rate of Turn
versus Nosewheel Steering Angle
1A-8
1b1-1
Example of Simulator Test Results for Ground
Acceleration Time & Distance (Part 1)
1B-5
1b1-2
Example of Simulator Test Results for Ground
Acceleration Time & Distance (Part 1)
1B-6
1b2-1
Control Speed, Ground
1B-11
1b3-1
Unstick Speed
1B-14
1b4-1
Example of Simulator Test Results for Normal
Takeoff
1B-18
1b5-1
Example of Simulator Test Results for Engine
Inoperative Takeoff
1B-22
Section 1
vii
1b6-1
Example of Simulator Test Results for Crosswind
Takeoff
1B-26
1b7-1
Example of Simulator Test Results for Rejected
Takeoff
1B-29
1b8-1
Example of Simulator Test Results for Dynamic
Engine Failure After Takeoff
1B-33
1c1-1
Example of Simulator Test Results for Climb in Clean
Configuration
1C-4
1c2-1
Inoperative Climb, Second Segment
1C-8
1c3-1
Inoperative Enroute Climb
1C-11
1c4-1
Inoperative Climb, Approach
1C-14
1d1-1
Example of Simulator Test Results for Level Flight
Acceleration
1D-4
1d2-1
Example of Simulator Test Results for Level Flight
Deceleration
1D-7
1d3-1
Example of Simulator Test Results for Cruise
Performance
1D-10
1d4-1
Example of Simulator Test Results for Idle Descent
1D-13
1d5-1
Example of Simulator Test Results for Emergency
Descent (Manual)
1D-16
1e1-1
Example of Simulator Test Results for Stopping Time
& Distance, Dry Runway
1E-4
1e2-1
Example of Simulator Test Results for Reverse
Thrust Stopping Time & Distance (1)
1E-7
1e2-2
Example of Simulator Test Results for Reverse
Thrust Stopping Time & Distance (2)
1E-8
viii
1e3-1
& Distance, Wet Runway
1E-11
1e4-1
& Distance, Icy Runway
1E-14
1f1-1
Example of Aeroplane Manufacturer's Proof of Match
Data (Engine Acceleration)
1F-4
1f1-2
Acceleration
1F-4
1f2-1
Acceleration & Deceleration (Combined)
1F-7
2a1-1
Example of Simulator Test Results for Pitch
Controller Force versus Position Calibration
2A-5
2a1-2
Example of Simulator Test Results for Elevator
versus Pitch Controller Position Calibration
2A-6
2a2-1
Example of Simulator Test Results for Roll Controller
Force versus Position Calibration
2A-9
2a2-2
Example of Simulator Test Results for Aileron &
Spoiler versus Roll Controller Position Calibration
2A-10
2a3-1
Example of Simulator Test Results for Rudder Pedal
Force versus Position Calibration
2A-13
2a4-1
Example of Simulator Test Results for Nosewheel
Steering Controller Force versus Position
2A-16
2a5-1
Example of Simulator Test Results for Rudder Pedal
Steering Calibration
2A-18
2a7-1
Example of Simulator Test Results for Pitch Trim
Rate Test
2A-23
Section 2
ix
Example of Simulator Test Results for Cockpit
Throttle Lever versus EPR
2A-26
2a9-1
Example of Simulator Test Results for Brake Pedal
Calibration
2A-29
2b-1
Underdamped Step Response
2B-5
2b-2
Critically Damped Step Response
2B-6
2b1-1
Example of Simulator Test Results for Pitch Control
Dynamics
2B-10
2b2-1
Example of Simulator Test Results for Roll Control
Dynamics
2B-13
2b3-1
Example of Simulator Test Results for Yaw Control
Dynamics
2B-16
2b4-1
Example of Simulator Test Results for Small Control
Inputs, Pitch
2B-19
2b5-1
Inputs, Roll
2B-22
2b6-1
Inputs, Yaw
2B-25
2c1-1
Example of Simulator Test Results for Power Change
Dynamics
2C-4
2c2-1
Example of Simulator Test Results for Flap Change
Dynamics, Retraction
2C-5
2c2-2
Example of Simulator Test Results for Flap Change
Dynamics, Extension
2C-10
2c3-1
Example of Simulator Test Results for Speedbrake
Change Dynamics, Extension
2C-13
x
2a8-1
2c3-2
Example of Simulator Test Results for Speedbrake
Change Dynamics, Retraction
2C-16
2c4-1
Example of Simulator Test Results for Gear Change
Dynamics, Retraction
2C-19
2c4-2
Example of Simulator Test Results for Gear Change
Dynamics, Extension
2C-21
2c5-1
Example of Simulator Test Results for Longitudinal
Trim
2C-24
2c6-1
Manoeuvring Stability
2C-28
2c7-1
Static Stability
2C-31
2c8-1
Example of Simulator Test Results for Stall
Characteristics
2C-35
2c9-1
Example of Simulator Test Results for Phugoid
Dynamics
2C-39
2c10-1
Example of Simulator Test Results for Short Period
Dynamics
2C-42
2d1-1
Control Speed, Air
2D-5
2d1-2
Control Speed, Air - Time History
2D-6
2d2-1
Example of Simulator Test Results for Roll Response
(Cruise Condition)
2D-9
2d3-1
Example of Simulator Test Results for Step Input of
Cockpit Roll Controller
2D-12
2d4-1
Example of Simulator Test Results for Spiral Stability
2D-15
xi
2d5-1
Inoperative Trim
2D-18
2d6-1
Example of Simulator Test Results for Rudder
Response (Yaw Damper Off)
2D-21
2d6-2
Example of Simulator Test Results for Rudder
Response (Yaw Damper On)
2D-22
2d7-1
Example of Simulator Test Results for Dutch Roll
2D-26
2d8-1
Example of Simulator Test Results for Steady State
Sideslip
2D-30
2e1-1
Example of Simulator Test Results for Normal
Landing
2E-5
2e2-1
Example of Simulator Test Results for Minimum Flap
Landing
2E-8
2e3-1
Example of Simulator Test Results for Crosswind
Landing
2E-12
2e4-1
Inoperative Landing
2E-15
2e4-2
Inoperative Landing, Alternate Engine Fit
2E-16
2e5-1
Example of Simulator Test Results for Autopilot
Landing
2E-19
2e6-1
Example of Simulator Test Results for Go-Around, All
Engines Operating
2E-22
2e7-1
Example of Simulator Test Results for One Engine
Inoperative Go-Around
2E-25
2e8-1
Example of Simulator Test Results for Directional
Control with Symmetric Reverse Thrust (1)
2E-28
xii
2e8-2
Control with Symmetric Reverse Thrust (2)
2E-29
2e9-1
Control with Asymmetric Reverse Thrust
2E-32
2f1-1
Example of Simulator Test Results for Ground Effect
Demonstration (Snapshots)
2F-6
2g-1
Wind Training Aid Model #1
2G-4
2g-2
2G-5
2g-3
2G-6
2g-4
2G-7
2g-5
Wind Training Aid Wind Factor Chart
2G-8
2g-6
United Kingdom Royal Aerospace Establishment
(now Qinetiq) Microburst Vortex Ring Air Flow Model
2G-9
2g1-1
Example of Simulator Test Results for Takeoff
Windshear Demonstration
2G-12
2h1-1
Example of Simulator Test Results for Overspeed
Protection Function
2H-5
2h2-1
Speed Protection Function
2H-8
This space left intentionally blank
2h4-1
Example of Simulator Test Results for Pitch Angle
Protection Function
2H-14
xiii
2h5-1
Example of Simulator Test Results for Bank Angle
Protection Function
2H-18
2h6-1
Example of Simulator Test Results for Angle of Attack 2H-21
Protection Function
Section 3
3-1
Flight Simulator Six-Axis Synergistic Motion System
3-2
3-2
Simulator Motion System Drive Block Diagram
(Simplified)
3-3
3a-1/3a-2
Frequency Response Results Example 1
3A-4
3a-3/3a-4
3A-5
3b-1
Example of Simulator Test Results for Motion System
Cross-Drive (Leg Balance) at 0.5 Hz
3B-4
3b-2
Cross-Drive (Leg Balance) at 3.0 Hz
3B-5
3c-1
Turn Around (Platform Heave Motion versus
Reference Drive Demand at 0.5 Hz)
3C-4
3e-1a
Example of Motion System Repeatability Test Results 3E-4
3e-1b
3e-1c
3e-1d
3e-1e
xiv
3e-1f
Example of Motion System Repeatability test Results
3E-9
3f-1a
Motion Cueing Performance Signature - Normal
Takeoff
3F-5
3f-1b
Takeoff
3F-6
3f-1c
Takeoff
3F-7
3f-1d
Takeoff
3F-8
3f-1e
Takeoff
3F-9
3f-1f
Takeoff
3F-10
3f-1g
Takeoff
3F-11
3f-1h
Takeoff
3F-12
3f-1i
Takeoff
3F-13
3f-1j
Takeoff
3F-14
3g-1
Vibration Analysis Windowing Functions
3G-6
3g-2a
APSD Plot, Processed using 0.25 Hz Bandwidth
3G-7
3g-2b
3G-7
xv
3g-3a
Single Pass Analysis
3G-8
3g-3b
Multiple Pass Analysis
3G-8
3g-4
Example of Simulator Test Results for Flap Buffet
Amplitude Time History (Y- and Z-Axes Only)
3G-15
3g-5
Example of Aeroplane Manufacturer Data for Flap
Buffet - Amplitude Time History
3G-15
3g-6
Example of Simulator Test Results for Flap Buffet PSD Plots (Y- and Z-Axes Only)
3G-16
3g-7
Example of Aeroplane Manufacturer Data for Flap
Buffet - PSD Plots
3G-17
3g-8
Example of PSD Plot Obtained from Stand-alone Test 3G-18
Equipment
3g-9
Example of Time History Plot Obtained from Standalone Test Equipment
3G-18
4a-1
Example of Simulator System Response (Latency)
Results
4A-3
4a-2
Example of Simulator System Response (Transport
Delay) Results
4A-4
4a1-1a
Example of Simulator Test Results for Transport
Delay (Yaw) Part 1
4A-10
4a1-1b
Example of Simulator Test Results for Transport
Delay (Yaw) Part 2
4A-11
4b1-1
Example of Spherical Grid Test Pattern
4B-3
Section 4
xvi
4b2-1
Example of Spherical Grid Test Pattern with example
angular measurement
4B-5
4b3-1
Example of Surface Contrast Checkerboard Test
Pattern
4B-7
4b4-1
Example of Surface ResolutionTest Pattern
4B-9
4b5-1
Example of Surface ResolutionTest Pattern
4B-11
4b6-1
Example of Lightpoint Size Test Pattern
4B-13
4b7-1
Example of Lightpoint Array Test Pattern
4B-15
4c-1
Visual Ground Segment Horizontal and Vertical
Distances - Pilot/Glideslope Antenna/Main Gear
4C-4
4c-2
Visual Ground Segment Horizontal and Vertical
Distances - Aeroplane to Ground
4C-5
4c-3
Visual Ground Segment Diagram Example 1
4C-7
4c-4
4C-8
5-1
Example of Simulator Calibration Results - Quiet
Room with Low Noise Fan
5-6
5-2
Example of Simulator Calibration Results - Air
Conditioning On versus Off
5-7
5-3
Example of Simulator Calibration Results Adjustment for Air Conditioning
5-8
Section 5
xvii
5a-1
Example of Simulator Test Results for Landing
Condition Sound Test
5A-4
5d-1
Recommended Maximum Simulator Background
Noise
5D-3
5e-1
Example of Recurrent Frequency Response Test
Tolerance
5E-3
B-1
Example of Critical Damping (Statically and
Dynamically Stable)
B-4
B-2
Example of Positive Damping (Statically and
Dynamically Stable)
B-5
B-3
Example of Negative Damping (Statically Stable but
Dynamically Unstable)
B-5
B-4
Example of Simple Divergence (Statically and
Dynamically Unstable)
B-6
B-5
Method of Determination of Time to Half Amplitude of
a Second Order Oscillation
B-7
E-1
Flap Change Match - Open Loop Roll Axis
E-3
E-2
Flap Change Match - Closed Loop Roll Axis
E-4
E-3
Open Loop Landing Match
E-6
E-4
Closed Loop Landing Match
E-6
E-5
Closed Loop Elevator Difference (Simulator - Flight
Test)
E-7
E-6
Open Loop Landing Match Modified Function
E-8
Appendices
xviii
E-7
Comparison of Pitching Moment Ground Effect
Coefficient Increment
E-9
E-8
Closed Loop Landing Match 20% Reduction in
Elevator Effectiveness
E-10
E-9
Elevator Error for Closed Loop Match 20% Reduction
in Effectiveness
E-11
xix
1.0
INTRODUCTION
1.1
BACKGROUND
The evaluation of aeroplane simulators for qualification under the
“International Standards for the Qualification of Airplane Flight Simulators”
developed by the Royal Aeronautical Society and amended by the 2nd
Edition of the ICAO “Manual of Criteria for the Qualification of Flight
Simulators” (hereafter referred to as the “ICAO Manual”, Reference 19),
is a complex and demanding technical task. The task often requires the
application of sound engineering judgement to determine if the simulator
performance in a given area meets the requirements of the ICAO Manual.
It is the purpose of this Handbook to provide guidance for the application
of judgement in the evaluation as well as guidance in the conduct of the
evaluation. In addition, this guidance may be useful to the aircraft
manufacturer or other data provider for planning and conducting
validation tests, data recording and presentation. This Handbook first
addresses the task in general and then addresses each test individually.
1.2
SIMULATOR STANDARDS
Appendix A of the ICAO Manual discusses the required performance for
qualification of the highest level of flight simulator. In many instances
these requirements are open to some interpretation; therefore, the
authorities would like to promote discussions between owners, operators,
simulator manufacturers and data providers early in a project life cycle to
assist with interpretations. Consequently, it would seem inappropriate to
address these requirements from the perspective of this Handbook. The
original International Standards were developed between September
1989 and January 1992, by an international working group. As a result
of these standards being put into use over the succeeding years, the
small number of anomalies present in the original document came to light,
along with an impression that more information would be useful and
clarifying the viewpoint of the regulatory authorities over the intent of
many of the tests. Thus a similar Working Group was convened during
2001 to address these issues.
1.3
VALIDATION TESTS
1.3.1
General
These tests, found in Appendix B of the ICAO Manual, compare simulator
performance to aeroplane performance. The performance should match
within the tolerances specified for the test to pass. During both initial and
xx
recurrent evaluations, the baseline simulator performance is established
against the standard, i.e., the aeroplane flight test (or other acceptable)
data. For recurrent evaluations, simulator performance must also be
compared to the baseline performance, i.e. the data in the Master
Qualification Test Guide (MQTG), to identify any change in performance.
This comparison is used as a check that the configuration control system
used by the simulator operator is being utilised properly and also allows
any drifts or deviations to be identified by the operator's technicians or
engineers at an early stage, preferably prior to the next recurrent
evaluation. Unless there is good reason for the operator or data provider
to change the modelling or testing in a particular area - and such changes
have been discussed and agreed with the regulatory authority, any
deviations in the automatic test results must be corrected such that the
results of recurrent evaluations are indistinguishable from those contained
in the master QTG, which is itself a reflection of the aeroplane flight test
data.
1.3.2
Parameters
The parameter list for a test may vary from simulator to simulator
depending on the way the flight test was conducted, the method of data
gathering used and also of the basic aeroplane configuration. To provide
flexibility to the industry in meeting the Standard, a required parameter list
for each test was not part of the ICAO Manual. (A list of recommended
test parameters for validation tests is included in the IATA Document,
“Flight Simulator Design and Performance Data Requirements”, Appendix
D, Reference 12). In general, parameters plotted must include all those
which have a tolerance applied, all those required to confirm initial
conditions or verify the flight condition, all input parameters throughout the
time history, and any other parameters that may affect the test result. In
essence, all those parameters which are necessary and helpful to
determine the outcome of the test.
The parameters which are being used as the test drivers, for example
pilot control positions, must also be noted. For each Validation Test case
discussed in this Handbook, a suggested list of plot parameters is
provided, though it should be noted that these lists are by no means
definitive in every case. One other comment worth making is that it is not
necessary to provide vast numbers of plots for every test - though in
general the issue has in the past been that too few, rather than too many
plots have been provided.
xxi
1.3.3
Automatic Tests
The Standard requires proof of match within tolerance for the specified
variables. In every case where performance is out of tolerance and
correct implementation of the simulation mathematical model is ensured,
the first step should be to bring this condition to the attention of the data
provider. If the data provider is unavailable or is unable to resolve the
problem, and since it is ultimately the responsibility of the operator to
demonstrate the performance of the simulator, the operator may choose
to finely tune the mathematical model to bring the performance within
tolerance or else to search for another flight data sample which supports
a test that will meet the Standard. If fine-tuning proves impossible, or
alternative data do not exist, and at the discretion of the evaluator
(meaning the regulatory authority in charge of the evaluation),
performance mismatches may be discussed with a view to the use of
rational engineering judgement which may provide a reason for the
mismatch. Such reasons should typically only be provided for extremely
short duration excursions which can clearly be explained by reference to
other presented flight data or accounted for by an acceptable, logical
explanation. This type of procedure would address situations where short
duration variations in the flight data (for example a significant wind gust,
which may not be fully accommodated in the flight model or the test data)
cause an excursion in a performance parameter outside of the tolerance.
Out-of-tolerance results which are to be justified in this manner should not
cover a significant proportion of the test parameters, or of the time history.
Ideally, the out-of-tolerance excursion should not last more than 10% of
the pertinent portion of the time history. Such examples of the use of
engineering judgement are to be adequately documented with each test.
The main value of automatic testing is the ease and rapidity of conducting
tests and the inherent repeatability. Care must be exercised in the design
and implementation of automatic tests to ensure that the objective of
repeatability does not overshadow the objective of the test itself; i.e.
validation of a feature of the simulator. For more information, see Section
2.0 on page xxvii.
1.3.4
Manual Tests
Manual testing serves to cross check and verify automatic testing and is,
therefore, very important to simulator validation. These tests must be run
“end-to-end” and without any backdriven parameters to confirm correct
implementation of the automatic tests as well as compliant simulator
performance. The manual test procedure listed for the test must be
complete enough so that the pilot evaluator can conduct the test using
xxii
this procedure as the sole source of information, i.e. he need not
reference the flight data or any other part of the QTG to complete the test.
The Standard demands that all manual tests must meet the same
performance match as the automatic tests. To complete all the QTG
Validation Tests manually to an exact performance match would be a time
consuming task; therefore, at the discretion of the evaluator, manual test
results can be accepted that do not overlay the test data provided that a
logical interpretation of the results indicates a performance match. For
example, if a stick shaker speed is the performance parameter required,
the time history need only show that the simulator configuration and the
environment matched the test conditions, and that the aircraft was in a
steady, 1 knot/sec deceleration at the time the shaker activated. Where
multiple parameters must be in tolerance, each excursion beyond
tolerance of any parameter must be explainable by deviations in other
presented parameters. These excursions should be of short duration, i.e.
less than 10% of the pertinent time history. Failing these rationalisations
of results, the test must be rerun until a suitable match is obtained or until
it is determined that a problem with the test exists. For more information,
see Section 4.0 on page xxxv.
1.3.5
Integrated Tests
Integrated testing is important to validate the overall integrated simulator
systems. It demonstrates the response of the simulator to a stimulus at
the pilot's controls. Whilst each simulator subsystem may be satisfactorily
demonstrated when subjected to a test of only that system, it does not
demonstrate that all systems perform satisfactorily when integrated. For
example, satisfactory response of the aerodynamic model by stimulation
at the control surface does not prove that the response of the integrated
flight control system model and aerodynamic model is also satisfactory.
Integrated testing may be done either automatically or manually. Where
such tests are performed using a force input, whether automatically or
manually, the term “end-to-end” testing may also be applied to describe
the methodology. Whilst it is recognised that integrated testing can be
difficult to accomplish automatically because of the present inadequacies
of the mechanisms used to manipulate the cockpit controllers, all tests
should still be integrated except where significant complexities arise which should be discussed with the regulatory authority at an early stage.
Note that back-driving the controls during the test is not the same as
integrated testing. Manual testing then, should effectively be a verification
of the automatic test design or procedure. Where data deficiencies cause
difficulties in the matching process, a less integrated test may be
xxiii
acceptable to the regulatory authority. For more information, see Section
3.0 on page xxx.
1.4
QTG REVIEW TECHNICAL EVALUATION
1.4.1
General
The technical review of the QTG must be completed well in advance of
the on-site evaluation to permit rectification of any problems discovered
during the review before the on-site evaluation. Furthermore, a thorough
review of the QTG should prevent QTG related delays during the on-site
evaluation.
1.4.2
Step 1 - Overview
First check to see if the QTG is complete. Are all the required tests
presented, all the statements of compliance provided, and all the pertinent
information on the device provided?
The following points should also be considered:
xxiv
a)
The QTG test descriptions, including and especially the initial
conditions, the flight test data and the simulator data must all be
in the same units. This most often occurs with airspeed and
thrust/torque parameters. For example, if IAS is used in the flight
data, then IAS should be shown in the simulator data and the
initial conditions. If, for whatever reason, identical units are not
used, then the conversions or “deltas” must be provided and the
data annotated accordingly.
b)
QTG simulator data should be annotated to show compliance
where that is not integral to the data presentation. Similarly,
where data must be manually compared, overlay alignment points
must be clearly marked.
c)
Each test must be complete in itself or incorporated into another
test so that no time is wasted during evaluations paging back and
forth in the QTG. This means that ideally no part of a test should
refer to another test; therefore, remarks such as “Same as 3F” or
“Refer to Sections 5B and 5C” are not appropriate and should be
avoided where possible.
d)
Provision of wind time histories, both flight data and corroborating
simulator data input, is very important since these data are often
the premise upon which “sound engineering judgement” is used
to decide short term out-of-tolerance performance.
1.4.3
Step 2 - Test-by-Test Review of Validation Tests
The key to understanding a test is a comprehensive manual test
procedure description. The first step in the review of an individual
validation test is to read the manual test description and determine if:
a)
the objective of the test is met In Accordance With (IAW) the
Standard and proper flight test procedures. The FAA Advisory
Circular 25-7 (Reference 11) is an excellent guide to proper flight
test procedures
b)
the procedure matched the flight data, i.e. are the actions of the
test pilot being accurately duplicated?
c)
the procedure is complete, i.e. can the procedure be flown
without further reference to the flight data or the QTG?
d)
there are any back-driven parameters in the procedure. A general
rule of thumb would be that if the test is fundamentally a
performance test, then such methods (e.g. closed-loop controllers
/ maths pilots) are more likely to be acceptable than if the test is
predominantly a handling case.
Once it is determined that the test meets the objective then the following
should be reviewed:
i)
Initial Conditions - do the simulator initial conditions match the
flight data?
ii)
Data Sources - are the correct data sources nominated and
included in the QTG? Are the data presented adequate to prove
the test? Overplots of flight test data on simulator data is the
preferred method of test presentation. However, this is
insufficient to meet the data presentation requirements. Copies
of the reference data must also be part of the QTG. With the
advent of Electronic Qualification Test Guides, it may be that the
‘copies’ of the data are on a CD or other media only, but they
should nevertheless be included.
iii)
Tolerances - do the nominated tolerances agree with the
Standard?
xxv
1.5
iv)
Parameters - are the required parameters presented as per the
discussion above? Chapter 7.0 (page xlii) gives more information.
v)
Automatic Tests - the automatic test procedures will normally
describe computer procedures to start and run the test. When
reviewing these procedures, determine which parameters are
driven and which parts of the simulator are excluded. Normally,
these should be integrated tests, meaning control displacements
(or forces) are used as inputs, though some parameters may be
back-driven (fed back) similar to the feed back loop created by a
pilot at the controls (see Appendix E for a detailed discussion of
this subject). The actual automatic procedures are usually
irrelevant to the evaluator since the operator will run these tests.
FUNCTIONS AND SUBJECTIVE TESTS
Functions tests relate directly to systems operations. The systems must
operate as they do in the aircraft, and the simulator, within the limits of the
technology, must handle like the aircraft. If there are any differences in
operation or handling, for example response times to a switch selection
or forces on the controls at rotation, they must be below the “threshold of
observance” of the pilot under normal operating conditions. Plainly said,
this means if the pilot evaluator notices a difference from the aircraft, this
difference must be corrected to the limits of the technology. Typically,
these tests as defined in Appendix C of the ICAO Manual, are more
difficult to standardise than the objective tests contained in Appendix B
which are the main subject of this Handbook. Chapter 5.0 on page xxxvii
gives a cursory treatment of this subject, but for a more complete
treatment of Functions and Subjective testing, refer to Volume II of this
Handbook.
xxvi
2.0
AUTOMATIC TESTING
2.1
INTRODUCTION
In the early days of Approval Test Guide (“ATG”, the term has now been
superseded by Qualification Test Guide, “QTG”) evaluations, the
validation tests would all be run manually, including the setting up of the
simulator initial conditions. Better design techniques, together with the
regulatory authority requirements for recurrent testing of the simulator at
regular intervals, has enabled all validation tests to be performed
automatically, using computer-driven stimuli.
2.2
BASIC METHODOLOGY
The principles of automatic testing are not difficult to understand once the
basic testing requirements have been analysed. Essentially, the QTG
validation tests fall into two fundamental categories:
a)
Steady-state condition tests (such as longitudinal trims,
longitudinal static stability, steady sideslip, etc).
b)
Transient response tests (such as flap change dynamics, normal
takeoff, etc).
2.2.1
Steady-State Condition Tests
For the steady-state tests, the simulator is usually positioned at the
required airspeed and altitude with the gross weight, centre of gravity,
flaps and landing gear set to the values specified in the aircraft
manufacturer's flight test data. This in itself is not sufficient, because if
the simulator is unfrozen at this point, it is very improbable that it will be
in a trimmed state. Thus the airspeed and altitude (and geographical
position, if necessary) are held constant whilst, for example, the pitching
moment is nullified by driving the stabiliser, and the longitudinal
acceleration is nullified by the engine thrust being altered accordingly.
This is the basis of automatically trimming the simulator, though the
particular technique will vary, as will the rate at which the trimming
operation is performed. Naturally, when attempting to cancel out both the
pitching moment and the longitudinal acceleration, achieving absolute
zero values of both these parameters would take, in theory at least, an
infinite amount of time and so in practice an allowable error margin is
employed within which the simulator is effectively in trim. The smaller this
error margin, the better the resultant trim, but the longer it will take -
xxvii
hence a judgement must be made by the designer of the automatic test
system as to what is an acceptable margin.
Trim conditions for other steady-state tests are achieved in a similar
manner to that for longitudinal trims, but by driving whichever control
surface is necessary to nullify the appropriate acceleration. It is not
infrequent to encounter so-called “trim” conditions from aircraft flight test
data which are not strictly in trim. There may, for example, be a pitch rate
or the airspeed may not be constant. For these conditions, the simulator
rates and accelerations should be set equal to the measured values from
the aeroplane. In addition, it is generally necessary to place a small
“bias”, or difference between the simulator value and the measured flighttest value, in each degree of freedom in order to satisfy the equations of
motion of the simulation. Well-accepted trim biases include a small offset
in angle of attack to trim lift, pitch trim (e.g. stabiliser angle) to trim
pitching moment, and thrust or climb rate to trim drag. Trimming the
lateral- directional degrees of freedom is more problematic. Various
methods may be used. For example, a small offset in rudder deflection
or sideslip angle may be used to trim yawing moment. Alternatively, an
increment in aerodynamic yawing moment coefficient may be applied to
trim the simulation while setting all flight parameters equal to the
aeroplane measurements. While it may be argued that the latter method
is the most ‘honest’ approach to accounting for unmeasurable aeroplane
or thrust asymmetries, it has not been universally accepted by regulatory
evaluators.
Since concurrence with the aircraft manufacturer or other data provider
is a pre-requisite, techniques that may be used to engineer close matches
which cannot be traced back to deficiencies in the data source must not
be utilised. Where biases or offsets are employed, their use must be
clearly explained and justified. Rationales should be used to justify each
deviation.
2.2.2
Transient Response Tests
The initial phase of a test requiring the monitoring of a selection of
parameters over a period of time will always involve setting up the
simulator and trimming in the manner described above. What occurs
subsequently to the trim is the stimulation of the simulated aircraft using
the flight controls (and in some cases wind velocities or other external
disturbances). This stimulation may either be for a relatively short portion
of the test (e.g. phugoid, flap change dynamics) or it may be for the entire
test (e.g. takeoffs and landings). The exact nature of the input and its
duration are dependent on both the type of test and the way in which the
xxviii
data have been presented by the aircraft manufacturer or flight test
organisation.
2.3
PASS/FAIL CRITERIA
The evaluator will be looking for a set of results which demonstrates that
the test criteria have been met. Short term excursions exceeding the
tolerances may well be acceptable, but ideally, all validation tests will
remain within the specified tolerances for the entire duration of the test.
However, in reality this may not always be the case because of the many
vagaries of the flight test data, of the atmosphere and of the mathematical
model itself. It is under these circumstances that engineering judgement
must be used by the evaluator.
For instance, it would be difficult to rationalise a longitudinal trim test for
which the design standard is 4 units ±0.5 unit of stabiliser when the
achieved value on the simulator is 8 units. Indeed for most, if not all,
steady-state tests little or no engineering judgement need be utilised
because the tests are either in tolerance or they are not. The evaluation
of time history tests is not quite so clear cut though. For example, a flap
change dynamics test which has a tolerance of ±1.5 degrees of pitch
attitude for a test which is 30 seconds long, may be within tolerance for
the first 29 seconds but just drifting out of tolerance during the last second
prior to the end of the test. In general, it would be quite reasonable to call
this test a pass and ignore the last second or so. Also, for tests in which
there is high pilot activity (especially takeoffs and landings) it is often the
case that some of the parameters on which the tolerances are being
applied may go outside the tolerance band for short durations. Strictly
speaking the test has failed if this happens, but again trends should be
looked at, making allowances for items such as undocumented or
unmeasurable atmospheric disturbances (e.g. wind gusts) which may
affect the results.
The 2nd edition of the ICAO Manual introduces much more definitive
criteria for QTG tests which have been developed and run against
engineering resource data rather than flight test resource data.
Specifically, the tolerances to be used have been reduced to 20% of
those which would apply against actual flight test parameters. This
requirement is also relevant to, for example, tests based on engineering
data which has been used for backup purposes because the flight test
data has been supplied with a different engine or avionics fit.
xxix
3.0
INTEGRATED TESTING
3.1
INTRODUCTION
Given aircraft manufacturer's flight test data, it might be assumed that the
method of performing a particular test for a simulator QTG is obvious. In
practice, however, this is not necessarily the case, largely because of the
challenges experienced when any mathematical model is applied to
represent a real-world situation.
Overcoming these challenges is part of the task performed, on a daily
basis, by flight simulator engineers when attempting to match flight test
data in order to prove that the simulator 'flies' like the aircraft. There are
inevitably difficulties when using a complete simulator mathematical
model (with all the individual systems’ minor inaccuracies) in proving the
device to within very tight tolerances relative to the aircraft flight test data.
Fundamentally, the airframe manufacturers' data packs have to stand up
to the rigorous testing required when a fully integrated flight simulator, as
distinct from an inhouse computer with no aircraft hardware (and/or in
some cases no aircraft systems software) is the testbed.
Consequently, integrated testing is important to validate the overall
integrated simulator systems. It demonstrates the response of the
simulator to a stimulus at the pilot's controls. While each simulator
subsystem may be satisfactorily demonstrated when independently
tested, subsystem testing does not demonstrate that all subsystems
perform satisfactorily when integrated.
3.2
BACKGROUND
As time has progressed, it is clear that the validation tests which are to be
run in the simulator need to be run automatically, primarily for the
following reasons:
xxx
a)
As a regulatory body requirement, the test has to be re-run at
regular intervals for recurrent testing of the simulator.
Repeatability is, therefore, of great importance.
b)
When flying the simulator, pilots have difficulty matching the exact
inputs of the flight test pilot for a given test. This becomes very
time intensive.
c)
Shorter simulator delivery timescales mean that simulator
manufacturers needed to transfer much of their testing (including
that performed for a QTG) to their inhouse computing facilities.
The third point here is perhaps of greatest significance since, by
definition, tests that have been developed on an inhouse computer have
not been developed on an actual simulator and therefore some means
must have been employed to ‘simulate’ certain aspects of the simulator
itself, especially with regard to the pilot controls (both primary and
secondary) and the cockpit indications.
This in itself does not mean that a test run inhouse cannot be carried
across to the simulator at all, or even with a low degree of confidence.
Indeed for many QTG tests it makes very little difference whether it is run
on a fully integrated and functioning simulator if one ignores any
requirement to make the test ‘look good’ from an aesthetic point of view
in the cockpit. A prime example is a minimum radius turn, however it may
be driven, for which the results should look identical with or without the
simulator hardware.
Nevertheless, there are other considerations which impact the testing
methods used. Firstly, there was the requirement, laid out in the IATA
Simulator Data Requirements document (Reference 12), for the
mathematical model (and data) supplied by the airframe manufacturers
to the simulator manufacturers to include checks, both against itself and
also (and more significantly from the point of view of a QTG) against
aircraft manufacturer’s validation data. Thus was born the ‘Proof of
Match’ document, consisting of tests run by the airframe manufacturers’
simulator groups in order to prove that their model matched the aircraft
behaviour closely. However, the airframe manufacturers usually run
these tests not on a full flight simulator, but using an ‘Engineering
Simulation’, essentially an inhouse computer, typically with software
models of the aerodynamic characteristics and all aircraft systems that
affect flight characteristics all run in an integrated manner together with
the aircraft equations of motion. Consequently, it is the simulator
manufacturer who is the first to run the fully integrated simulation on a full
flight simulator and it is the simulator manufacturer who carries the major
load in ensuring that the tests are performed in an adequate manner from
the point of view of the regulatory authorities.
xxxi
3.3
ENGINE MODELS
The difference in aircraft performance, between two otherwise identical
airframes fitted with different engine types, is so significant that it must be
taken into consideration when producing the QTG.
In an ideal world, there would either be a single engine variant for a given
aircraft type, or else the simulator data package would contain sufficient
flight test data for a complete QTG to be generated for each
aircraft-plus-engine combination. For example, for one large jet transport
aeroplane, the flight test was performed with a Pratt & Whitney engine,
whereas some operators use General Electric engines and other use
Rolls Royce engines. The simulator data are presented in terms of thrust,
not EPR or PLA/CSA (TRA for a FADEC engine) and in fact Gross
Thrust/Ram Drag model is required for the tests to work properly. Thrust
overwrites do solve this problem, but more rigorous testing of the engine
model is really needed and can only be achieved by driving a parameter
which is much closer to the pilot's controls, i.e. Power Lever Angle,
Cross-Shaft Angle, Throttle Resolver Angle etc. Sometimes the only
solution in the simulator is to set and/or drive the throttles to arbitrary
values which produce the requisite levels of net thrust.
The disadvantages of driving the simulated engines in this way are firstly
that slight differences in an engine type (e.g. derated engine) mean that
a different TRA/PLA etc may be required for each simulator and secondly
that the QTG tests can become highly susceptible to any modifications
made to the engine system software in the simulator. Also, engine
characteristics differ slightly from one engine to another even with
engines of the same type - and they change over time. The flight test
engines may not be representative of an ‘average’ engine in the fleet.
In any case, overwriting thrust is an engineering solution that in most
cases would probably not be acceptable to the regulatory authorities.
Attachment E of the Second Edition of the ICAO Manual provides
approval guidelines for alternate engines, including those of a different
manufacturer than that of the baseline engine, and for alternate thrust
ratings. A basic summary of the tests required (one per test number) is
provided below but the reader is referred to the ICAO Manual itself for
more definitive information:
xxxii
TEST
NUMBER
1
3.4
TEST DESCRIPTION
ALTERNATE
ENGINE
TYPE
ALTERNATE
THRUST
RATING
1b(1),
1b(4)
Normal take-off/ground
acceleration time & distance
X
X
1b(2)
Vmcg. if performed for aeroplane
certification
X
X
1b(5)
Engine-out take-off
Either test
may be
performed
X
1b(8)
Dynamic engine
failure after take-off
1b(7)
Rejected take-off if performed for
aeroplane certification
X
1d(1)
Cruise performance
X
1f(1), 1f(2)
Engine acceleration and
deceleration
X
X
2a(7)
Throttle calibration 1
X
X
2c(1)
Power change dynamics
(acceleration)
X
X
2d(1)
Vmca it performed for aeroplane
certification
X
X
2d(5)
Engine inoperative trim
X
X
2e(1)
Normal landing
X
should be provided for all changes in engine type or thrust rating
AVIONICS FITS
Modern jet transport fleets rarely retain the same avionics systems
throughout their useful life, and the flight simulator will reflect this. For
example, the installation of a TCAS system in the aeroplane is almost
always followed by a similar installation in the simulator. Whilst such
systems may not affect performance or handling as such, the QTG must
be flexible enough to take account of such changes, with at least
referencing their use in the Functions and Subjective section. Other types
of avionics changes may have a significant affect on the aeroplane
xxxiii
performance, however, and in these cases the general principles applied
above for different engine configurations should be followed.
xxxiv
4.0
MANUAL TESTING
4.1
INTRODUCTION
In the past, attention has sometimes been lacking concerning the clarity
and practicality of QTG manual test procedures. Therefore, it is
considered both useful and necessary to redefine the objectives and
emphasise the techniques when performing QTG tests manually. The
regulatory authorities consider the ability to run each test manually an
important feature of simulator testing, and one which should certainly not
be discarded in favour of the tendency towards automatic testing. Both
methods should be treated with equal importance.
The following guidelines should be noted and adhered to as much as
possible. Basic guideline procedures are shown in the detailed test
descriptions (following Chapter 12) for each check and indicate an
acceptable standard. It is hoped that these can be used as a good
framework to work around.
4.2
OBJECTIVES
The general consensus is that it is not expected or practical to try and
reproduce exact flight test inputs, particularly with checks which require
high pilot activity over lengthy time periods. The exact duplication of flight
test results should not be given top priority whilst performing manual
tests, although it is reasonable to assume that many of the tests which
require only light pilot activity should still pass.
QTG tolerances should be applied as guidance to steady state values and
not stringently to each step of the time history.
4.3
TECHNIQUES
Manual test descriptions need to be written in a manner that describes the
way in which the original check may have been flown by the pilot.
Front-end inputs such as column, wheel, pedals, etc. should be utilised
whenever possible. In directing the pilot to fly a profile, for example in a
takeoff manoeuvre, it would be prudent to ask the pilot to rotate to a
particular pitch attitude rather than issuing a column position instruction.
If surface inputs are absolutely necessary then it should be made clear
where they can be monitored. The obvious practicalities - or otherwise such as trying to ‘fly’ the simulator accurately whilst trying to monitor a
control surface parameter on an instructor screen situated a metre or
xxxv
more behind the pilot's seat should be borne in mind and avoided. The
presence of a second crew member may help in these circumstances, as
it is possible that access to values that are not instrumented in the flight
deck may be necessary (e.g. rudder position, engine thrust, etc.) . The
second person may also be able to assist so that the tasks can be divided
or by giving call-outs.
For some tests, it can be helpful if the exact technique differs slightly from
the way the test is run automatically. One possible example is climb
tests, for which it is considerably easier to run the test manually if the
initial altitude is set to, say, 1000 feet below the altitude at which the test
officially begins, so as to allow the pilot to stabilise the condition prior to
recording. Also, the use of the autopilot might be considered in this case.
All instructions should ideally reference one parameter only. This will
usually be time, but can also be height or airspeed depending upon the
test being flown. Complicated or long-duration manoeuvres are usually
best described by simplifying or removing the less important details in
favour of allowing the pilot to understand what he is trying to achieve.
Whilst this may result in the pilot’s actions not precisely replicating the test
data, it is normally much easier to examine the simulator results for the
correct trends, and will often save testing time by alleviating the need for
several attempts at the same manoeuvre.
When carrying out a manual QTG check the pilot should first observe an
automatic check and note throttle and control positions, rates and any
relevant indications before starting the manual test. It may also be worth
checking that the pedals are aligned with zero rudder, as some earlier
simulators were prone to loose datum with vigorous wear.
xxxvi
5.0
FUNCTIONS AND SUBJECTIVE TESTS
5.1
DISCUSSION
Chapter 1, sub-paragraph 5 referred the reader to the second volume of
the Evaluation Handbook which is designed to give a much more
complete treatment of Functions and Subjective Testing. The comments
offered below are intended to assist the evaluator who has an engineering
rather than a pilot training bias and can only be considered as a very brief
overview of the subject matter.
Accurate replication of aeroplane systems functions must be checked at
each flight crew member position. This includes procedures using the
operator’s approved manuals, aircraft manufacturers' approved manuals
and checklists. Handling qualities, performance and simulator systems
operation must be subjectively assessed. In order to ensure the functions
tests are conducted in an efficient and timely manner, operators are
encouraged to coordinate with the appropriate regulatory authority
responsible for the evaluation so that any skills, experience or expertise
needed by the regulatory authority evaluation team are available.
At the request of a regulatory authority, the simulator may be assessed
for a special aspect of an operator's training program during the functions
and subjective portion of an evaluation. Such an assessment may
include a portion of a LOFT scenario or special emphasis items in the
operator's training program. Unless directly related to a requirement for
the current qualification level, the results of such an evaluation would not
normally affect the simulator's current status. As always, regulatory
authorities should be consulted for the particular details.
Functions tests should be run in a logical flight sequence at the same time
as performance and handling assessments. This also permits real time
simulator running for 2 to 3 hours, without repositioning or flight or
position freeze, thereby ascertaining (at least to a degree) proof of
reliability.
5.2
TEST REQUIREMENTS
The ground and flight tests and other checks required for qualification are
listed in the ICAO Manual, Appendix C, table of functions and subjective
tests. The table includes manoeuvres and procedures to ensure that the
simulator functions and performs appropriately for use in pilot training and
checking in the manoeuvres and procedures normally required of a
training and checking program.
xxxvii
Manoeuvres and procedures are included to address some features of
advanced technology aeroplanes and innovative training programs. For
example, “high angle of attack manoeuvring” is included to provide an
alternative to “approach to stalls”. Such an alternative is necessary for
aeroplanes employing flight envelope limiting technology.
All systems functions must be assessed for normal and, where
appropriate, alternate operations. Normal, abnormal, and emergency
procedures associated with a flight phase should be assessed during the
evaluation of manoeuvres or events within that flight phase. Systems are
listed separately under “any flight phase” to ensure appropriate attention
to systems checks.
xxxviii
6.0
TESTING OF SIMULATORS FOR COMPUTER
CONTROLLED AEROPLANES
6.1
GENERAL
Computer controlled aeroplanes, also called “highly augmented” or “fly by
wire” aeroplanes, are characterised by the fact that pilot inputs to the
control surfaces are transferred and augmented via computers.
The way that the handling qualities of these aeroplanes are experienced
by the pilot may therefore depend on the operating mode of these
computers or even the sensor inputs they use or the state of the hydraulic
systems required to actually move the surfaces.
Also, manoeuvre and flight envelope protection designed into such
aeroplanes might be impaired or completely lost by failures of the
sensors, computers etc which contribute to the control path between the
cockpit controller and the aeroplane control surfaces.
Manufacturers of computer controlled aeroplanes have defined different
terms for degraded states of the control, augmentation and protection
functions unique to their aeroplanes. For the purpose of a simulator
validation standard which needs to be applicable to various aeroplane
types, all degraded control states have been covered in the ICAO Manual
under the term “non-normal” control. The control state which is not
impaired by any failures or abnormalities is called “normal” control.
As a typical training syllabus for these aeroplanes may include both
demonstration and proficiency training when ‘flying’ with non-normal
control, there is clearly a need for flight simulator testing to address the
most important of these non-normal control states.
This of course applies almost exclusively to handling qualities tests and
almost not at all to performance tests.
6.2
DISCUSSION
For the testing of computer controlled aeroplane simulators, flight test
data are required for both the normal and non-normal control states as
indicated in the ICAO Manual.
Tests in the non-normal state should always include the least augmented
state. Tests for other levels of control state degradation may be required
if significantly different handling qualities result from these states. The
xxxix
detailed requirements must be mutually agreed between the aeroplane
manufacturer and the regulatory authorities at the time of definition of
tests for specific aeroplane data.
Where applicable, test data must record:
a)
Pilot controller deflections or electronically generated inputs including the location of the input
b)
Hardware and software part numbers of flight control computers
c)
Flight control surface positions.
These recording requirements apply to both normal and non-normal
states.
6.3
APPLICABLE DEFINITIONS
6.3.1 . . . . . . . . . . . . . . . . . . . . . . . . Computer Controlled Aeroplane
An aeroplane where the pilot inputs to the primary control surfaces are
transferred and augmented via computers.
6.3.2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Natural Aeroplane
The natural aeroplane is an aeroplane on which all stability and control
augmentation systems are inactive or the augmentation is at the least
active state required to sustain flight. The least active augmentation state
is that which is required to be as reliable as the airframe itself.
6.3.3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Normal Control
Normal control is the state where the intended control, augmentation and
protection functions for the activation of the primary control surfaces are
fully available.
6.3.4 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Non Normal Control
Non-normal is the state where one or more of the intended control,
augmentation and protection functions for the activation of the primary
control surfaces are not fully available.
Note:
xl
Specific terms such as ALTERNATE, DIRECT, SECONDARY and BACK
UP, etc. may be used to define an actual level of degradation.
6.3.5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Protection Functions
Protection functions are primary flight control functions designed to
protect the aeroplane from exceeding its flight and manoeuvre envelopes.
6.4
ADDITIONAL FLIGHT TESTS
Flight and manoeuvre envelope protection functions (see Section 2h)
(1) Overspeed
N & NN ± 5 kts airspeed
(2) Minimum Speed N & NN ± 3 kts airspeed
(3)
(4)
(5)
(6)
Load Factor
Pitch Angle
Bank Angle
Angle of Attack
N & NN
N & NN
N & NN
N & NN
± 0.1g
± 1.5o
± 2o or ± 10%
± 1.5o
Cruise
Takeoff, Cruise,
Approach or Landing
Takeoff, Cruise
Cruise, Approach
Approach
2nd segment,
Approach or Landing
All tests to be time histories. "Time history response of simulator to control
inputs during entry into protection envelope limits".
6.5
ABBREVIATIONS USED
N
NN
6.6
- Normal Control
- Non-normal Control
NOTES
a)
Where only "N" appears in the performance test section, it indicates
the preferred control state. However if the test results are
independent of control state, Non-normal control data may be
substituted.
b)
Where "N" appears elsewhere it indicates the required control state.
c)
Where "NN" appears it indicates that test data must be provided for
one or more non-normal control states, including the least
augmented state, if (and only if) the envelope protection is still active,
but different in this degraded mode.
xli
7.0
PRESENTATION OF SIMULATOR TEST RESULTS
7.1
ACCURACY OF TABULATED DATA
There have been many examples of engineers losing sight of basic test
accuracies through using line printer formats that give aeroplane variables
to untold decimal places. Outside air temperature to 3 decimal places,
centre of gravity to 3 or 4 decimal places, pitch angle and angle of attack
to 3 decimal places, etc. Acceptable tabulated simulator result accuracies
should be defined in order to keep the engineers 'on track' and also give
a better hard copy result.
The following are suggested as examples:
Centre of Gravity
Weight
Temperature
Altitude
Pitch/Angle of Attack
Bank/Sideslip/Heading
Aileron/Elevator/Rudder
Rate of Climb
Airspeed
0.1 % mac
10.0 lb or kg
0.1o
1.0 ft
0.1o
0.1o
0.1o
1.0 ft/min
0.1 knot
It is recognised that some of the above examples (e.g. weight, centre of
gravity, altitude, rate of climb) exceed the accuracy with which flight test
data can be collected, but they may be mathematically useful during
simulator testing and evaluation.
7.2
PLOT SCALES
Instances also abound of data being plotted to scales that do not allow
proper or easy interpretation.
The following are therefore recommended for plotting data:
xlii
a)
Standard engineering scales - 1, 2, 4 or multiples of 10 thereof per
inch or 2 centimetres.
b)
Standard engineering graph paper - 20 divisions per inch or 2
centimetres.
c)
All specified tolerances must be easily readable. (A guideline that
regulators have found acceptable is a minimum of 3 millimetres for
the tolerance band.)
Usually the most convenient method is to plot to the same scales as the
data supplier's hard copy, but there may be occasions where expanding or
contracting the scales facilitates better understanding of the results.
7.3
FLIGHT TEST DATA PROBLEMS
It is possible that the only available flight test results have problems, or that
no flight-test data exists for some parameters or tests.
Some examples:
a)
Flight test variable obviously offset - a special note should be made
in the QTG giving the rationale. (This frequently occurs with
business jet data when pitch/angle of attack/flight path are not
compatible or pitch/angle of attack are obviously offset when the
aeroplane is on the ground).
b)
Crosswind Takeoff/Landing - data for inertial reference system, angle
of attack, sideslip, etc. with a large degree of "noise" (scatter in the
results due to gusting). The only solution is to use the mean of the
scatter and to get the general trends correct in the simulator.
c)
No flight-test data available - use alternate data (see list below for
recommended order of priority) and request exemption from the
regulatory authority.
Order of priority of alternate data:
1
2
3
4
5
d)
Aeroplane Manufacturer's Engineering Simulation Result
Aeroplane Manufacturer's Gross Data Result (before being
reduced for Aeroplane Flight Manual use)
Aeroplane Flight Manual
Aeroplane Manufacturer's Predicted Result
Simulator Manufacturer's Predicted Result (with rationale)
Certain variables not available from the flight test - request use of an
alternate variable or plot the simulator variable and use a rationale
to justify the value.
xliii
7.4
SIMULATOR/FLIGHT-TESTED
DIFFERENCES
AIRCRAFT
CONFIGURATION
These will have to be reviewed on a case by case basis but some
guidelines are given below:-
7.5
a)
A typical example is the engines. For tests involving performance,
the simulator can be compared to the flight test result using net thrust
as the key engine parameter. For tests involving dynamic engine
characteristics (VMCG, VMCA, engine inoperative takeoff) it may be
necessary to run simulator tests with the net thrust time history of the
flight test aeroplane and to repeat the test using the simulated engine
dynamic characteristics, following fuel cut (or throttle chop).
Attachment E to the ICAO Manual, “Data Requirements for Alternate
Engines B Approval Guidelines” provides additional guidance for
validation of simulators that use engines of a different manufacturer
or a thrust rating different from the baseline simulation.
b)
Engine computer - if there are differences in the engine computer
between the simulator and the aeroplane which results in different
key engine parameters for a given throttle position, it may be
acceptable to use the flight test aeroplane key engine parameter. An
example is takeoff power from full throttle for the Garrett TFE731
engine and computer; the flight tested aeroplane may have a
different N1% takeoff power rigging than the idealised values used
for the simulator.
NAMES FOR AEROPLANE VARIABLES
It may be appropriate to work towards an acceptable (common) set of
names to be used for aeroplane variables which are used on simulator
QTG plots and printer results. Engineers have a habit of using the Greek
names of the algebraic variables instead of the real world names, e.g. PHI
for bank angle, THETA for pitch angle, DELTA ELEV for elevator angle etc.
In general this is to be discouraged, as the QTG should be aiming at
maximum clarity for the reader.
A comprehensive list of recommended variable names is included below.
The following are suggested, but should be reviewed in relation to the flight
test/validation data presented by the aircraft manufacturer. In any case,
with modern QTG production techniques the full, unabbreviated name is
to be preferred.
xliv
CAS
IAS
PALT
RALT
TIME
ROC
BANK
PITCH
HDGT,M
ELEV
AIL
RUD
SPOILER
P TRIM
R TRIM
Y TRIM
EPR
N1
N2
N3
FF
EGT
AOA
SLIP
AOAV
WEIGHT
C.G.
OAT
TAT
RAT
SAT
MACH
IMN
HTER
FLAP
SLAT
GEAR
- Calibrated Airspeed
- Indicated Airspeed
- Pressure Altitude
- Radio Altitude
- Test Time
- Rate of Climb
- Bank Angle
- Pitch Angle
- True (or Magnetic) Heading; Yaw Angle
- Elevator Angle
- Aileron Angle
- Rudder Angle
- Spoiler Angle
- Pitch Trim (Stabilizer or Elevator Trim)
- Roll Trim (Aileron Trim)
- Yaw Trim (Rudder Trim)
- Engine Pressure Ratio (1, 2, etc. or L, R)
- Engine N1 (Fan) Speed (1, 2, etc. or L, R)
- Engine N2 (High Pressure Rotor) Speed (1, 2, etc. or L, R)
- Engine N3 Speed (1, 2, etc. or L, R)
- Fuel Flow (1, 2, etc. or L, R)
- Engine Exhaust Gas Temp (1, 2, etc. or L, R)
- Angle of Attack (of the Fuselage Reference Line)
- Sideslip Angle
- Angle of Attack (of the Vane L, R)
- Aeroplane Gross Weight
- Aeroplane Centre of Gravity
- Outside Air Temperature
- Total Air Temperature (100% Recovery)
- Ram Air Temperature (Actual Recovery)
- Static Air Temperature
- Actual Mach Number
- Indicated Mach Number
- Terrain Elevation (above sea level)
- Trailing Edge Flap Deflection (LO ,LI, etc.)
- Leading Edge Slat Deflection (L1, L2, L3, etc.)
- Landing Gear Position
H - Handle
L - Left (LI, LO)
R - Right (LI, LO)
N - Nose
B - Body (e.g. as found on the B747)
C - Centre (e.g. as found on the DC-10-30/40, MD11)
xlv
7.6
SELECTING FLIGHT TEST RESULTS
See Appendix A, “Flight Test Data Considerations”.
7.7
PARAMETERS TO BE RECORDED FOR EACH TEST
7.7.1 Basic Test Parameters
For each test a set of basic parameters should be printed out which defines
the initial conditions for the test. These conditions should include:
Gross Weight & Centre of Gravity
Pressure Altitude
Field Elevation *
Radar Altitude or Main Gear Height above Ground *
Airspeed (Calibrated, Indicated, etc. but must be that specified in the
validation data)
Trailing Edge Flap Position & Leading Edge Flap/Slat Position
Gear Handle Position
Mach Number (for cruise/high altitude condition)
Outside Air Temperature
Wind Speed & Direction *
Runway Condition *
Engine Bleed Condition
Stability Augmentation Status (each axis)
Fuel Quantity, each Tank
Key Engine Parameters (N1,EPR, Torque, etc.)
Trim Setting (Roll, Pitch, Yaw)
Linear Accelerations (each axis)
Linear Velocities (each axis)
Rotational Accelerations (each axis)
Rotational Velocities (each axis)
* Denotes extra parameters required for tests performed on or near the
ground
7.7.2 Additional Parameter Time Histories
Additional parameter time histories are required as defined for the test.
Appendix D of the IATA Document, “Flight Simulator Design and
Performance Data Requirements” (Reference 12) provides a list of
minimum parameters that are recommended for each validation test. The
main section of this Handbook also gives guidance on a test-by-test basis.
xlvi
8.0
CONFIGURATION CONTROL SYSTEMS
8.1
GENERAL
A Configuration Control System (CCS) is a specific set of procedures
designed to:
a)
Maintain an essential and continuous cognisance of the state of the
software and hardware.
b)
Ensure that changes to the originally qualified hardware and software
configuration of the simulator are installed in accordance with the
appropriate regulatory authority requirements.
c)
Monitor changes to the actual aeroplane to determine if the
modification alters the performance, handling qualities or functional
characteristics.
d)
Following review of the aeroplane change with the training
department, ensure the applicable changes are installed in the
simulator in a timely manner.
e)
Ensure significant updates to the simulator data package that alter
the flight performance, handling qualities or functional performance
are incorporated into the simulator.
f)
Ensure that changes made during maintenance, to correct
discrepancies, are correctly implemented.
The Operator's Simulator Engineering department should establish
procedures for the Configuration Control System.
The simulator operator's management should provide personnel who are
assigned the responsibility to ensure adherence to the CCS procedures.
The CCS is somewhat analogous to the airworthiness assurance functions
required of certificate holders for commercial aeroplane operation.
8.2
MAINTENANCE CONSIDERATIONS
The CCS should establish procedures for maintenance personnel to follow
that will prevent unauthorised changes to the simulator.
xlvii
The CCS should clearly establish what changes may be made to the
simulator during the course of simulator maintenance. This should cover
both hardware and software changes that deviate from the simulator
configuration as qualified. The system should provide procedures to
advise management of any maintenance items requiring modification to
correct a discrepancy and the proposed change should be reviewed to
ensure it does not alter the configuration of the simulator as qualified.
Maintenance management will generally be responsible for monitoring
compliance with the established procedures.
8.3
ENGINEERING CHANGE CONTROL SYSTEM
An ECCS is a control system that provides for review and authorisation of
engineering changes to the simulator before implementation to ensure that:
8.4
a)
The changes are technically correct.
b)
The changes are in compliance with all relevant regulations and
guidance material.
c)
Any approvals required by regulatory authorities resulting from the
changes are obtained.
d)
The operator's ECCS procedures are followed.
e)
There is an audit trail for all changes made to the simulator.
SOFTWARE CONFIGURATION CONTROL
In order to ensure an orderly system of changes to simulator software, a
system of Software Configuration Management (SCM) should be provided.
This system may be based on a system used by the computer
manufacturer or operating system developer. This system should track
changes to software and allow the user to recreate any previous version
of the software used for training.
The SCM should provide backup and recovery procedures which enable
the operator to recover from unintended software losses.
8.5
AEROPLANE CONFIGURATION CONTROL
The CCS should provide a system to monitor changes to the aeroplane
which has been simulated to determine their effects on the performance,
xlviii
handling qualities, systems operation or functional operation. This usually
requires liaison between the operator's simulator engineering function, the
aeroplane engineering function and the training function.
xlix
9.0
REFERENCES
The following list is not exhaustive but is indicative of the major documents
which were used in compiling the International Standards for the
Qualification of Aeroplane Flight Simulation published by the RAeS in 1992
and subsequently released by ICAO as the “Manual of Criteria for the
Qualification of Flight Simulators” in 1995 and revised in 2003. For
completeness, the ICAO Manual itself is included as Reference 19.
1
Advisory Circular 121-14C
"AIRPLANE SIMULATOR AND
VISUAL SYSTEM VALIDATION"
US Dept of Transportation
Federal Aviation Administration
2
Advisory Circular 120-40
31 January 1983
3
Advisory Circular 120-40A
31 July 1983
4
Advisory Circular 120-40B
"AIRPLANE SIMULATOR
QUALIFICATION"
29 July 1991
5
5
CAP453
"AEROPLANE FLIGHT
SIMULATORS: APPROVAL
REQUIREMENTS"
UK Civil Aviation Authority
1989
6
6
DGAC
"AEROPLANE FLIGHT
REQUIREMENTS"
French Direction Generale de
l'Aviation Civile
1986
7
FSD-1
"OPERATIONAL STANDARDS
AND REQUIREMENTS APPROVED FLIGHT
SIMULATORS"
Australian Civil Aviation Authority
February 1989
8
JCAB
"AEROPLANE FLIGHT
REQUIREMENTS"
Japan Civil Aeronautics Board
1986
l
Transport Canada
1989
9
TRANSPORT CANADA
"AEROPLANE FLIGHT
REQUIREMENTS"
10
"INTERNATIONAL REGULATIONS Rediffusion Simulation Ltd
January 1990
FOR FLIGHT SIMULATORS
REQUIREMENT COMPARISON
DATABASE"
11
Advisory Circular 25-7 "FLIGHT
TEST GUIDE FOR
CERTIFICATION OF TRANSPORT 9 April 1986
CATEGORY AIRPLANES"
12
"FLIGHT SIMULATOR DESIGN
AND PERFORMANCE DATA
REQUIREMENTS"
International Air Transport
Association
4th Edition, 1993 (6th Edition
published 2000)
13
"PROGRAMMABLE WIND
SPECIFICATION AND
FORMATTING FOR SIMULATOR
TESTS ON WINDSHEAR"
Stanford Research Institute,
California, USA
DOT Contract FA-75WA-3650
29 Sep 1977
14
GENERAL NOTE ON
MICROBURST WINDSHEAR
MODELLING
A A Woodfield, Royal Aerospace
Establishment, Bedfordshire,
England, May 84
15
"DEVELOPMENT AND
APPLICATION OF A
NON-GAUSSIAN TURBULENCE
MODEL FOR USE IN FLIGHT
SIMULATORS"
P M Reeves, G S Campbell, V M
Ganzer
NASA CR02451, September
1974
16
"NON-GAUSSIAN STRUCTURE
OF THE SIMULATED
TURBULENT ENVIRONMENT IN
PILOTED FLIGHT SIMULATION"
G A J van de Moesdijk,
Delft University of Technology
Memorandum M-304, April 1978
17
"WINDSHEAR TRAINING AID"
Prepared for the FAA by the Boeing
Company under Contract
DFTA0-1-86
US Department of Commerce,
National Technical Information
Service,
Springfield, Virginia, USA
February 1987.
li
Formed under the auspices of
the United Kingdom Royal
Aeronautical Society, London,
January 1992
18
"INTERNATIONAL STANDARDS
FOR THE QUALIFICATION OF
AIRPLANE FLIGHT
SIMULATORS"
19
“MANUAL OF CRITERIA FOR THE Published by the International
QUALIFICATION OF FLIGHT
Civil Aviation Organisation, 1995.
SIMULATORS”
Second Edition published 2003.
20
JAR-STD 1A “AEROPLANE
FLIGHT SIMULATORS”
21
CFR 14 PART 60 “Flight Simulation US Dept of Transportation
Device Initial and Continuing
Federal Aviation Administration,
Qualification and Use”
to be published.
22
JAR 25-16 Certification of Large
Aeroplanes
European Joint Aviation
Authorities, Hoofdorp, The
Netherlands. Amendment 3
published July 2003.
23
ARINC Report 436 “Guidelines for
Electronic Qualification Test Guide”
Aeronautical Radio Inc.,
Annapolis, Maryland, USA
European Joint Aviation
Authorities, Hoofdorp, The
Netherlands. Amendment 3
published July 2003.
This Handbook recognises that several of the above documents are now
essentially obsolete, but they have been left in for the benefit of those readers
who may wish to know the history of the ICAO Manual and its requirements.
lii
10.0 LIST OF CONTRIBUTORS
This Handbook is the product of much thought and effort by a large number
of people within the flight simulator industry. The international working
group convened under the auspices of the Royal Aeronautical Society in
1990 met on a total of 5 occasions plus additional meetings of various
subcommittees. In 2001, an international working group was convened
under the co-sponsorship of the European Joint Aviation Authorities and
the U.S. Federal Aviation Administration to develop the Second Edition of
the ICAO Manual. The total membership of both these working groups
consisted of the following organisations:
10.1 AIRCRAFT MANUFACTURERS (and providers of flight test data for
simulators)
Aerospatiale/Airbus Industrie (EADS Airbus)
Boeing Commercial Airplane Group
Bombardier Aerospace
British Aerospace Commercial Aircraft
Fokker Aircraft
McDonnell Douglas Corporation
Kohlman Systems Research
10.2 SIMULATOR/VISUAL MANUFACTURERS
BAe Simulation Ltd
CAE Electronics Ltd
Evans & Sutherland
Ferranti International Simulation and Training
FlightSafety International
Link-Miles Ltd
McDonnell Douglas Electronic Systems Co
Microflight
Reflectone Inc
Thales Training & Simulation Ltd (formerly Rediffusion Simulation Ltd)
Thomson CSF
USSR Simulators Design and Manufacturing Plant
10.3 SIMULATOR OPERATORS
Aeroformation
Air France
American Airlines
Ansett Airlines of Australia
liii
British Airways
Delta Airlines
Deutsche Lufthansa
FlightSafety Boeing
GECAT
KLM
Monarch Airlines
Northwest Airlines
QANTAS
SAS
Simuflite Training International
Swissair
United Airlines
United Parcel Service
US Air
10.4 REGULATORY AUTHORITIES
Civil Aviation Authority (Australia)
Civil Aviation Authority (United Kingdom)
Department of Avionika and Simulators (USSR)
Direction Generale de l'Aviation Civile (France)
Federal Aviation Administration (USA)
FOCA (Switzerland)
Joint Aviation Authorities (Europe)
Luftfahrt Bundesamt (Germany)
RLD (Netherlands)
Transport Canada (Canada)
10.5 OTHERS
Flight Research Institute of the USSR
International Air Transport Association
International Civil Aviation Organisation
Royal Aeronautical Society
10.6 ORGANISATIONS CONTRIBUTING TO THIS HANDBOOK
National Aeronautics and Space Administration
The author would also especially like to thank the following individuals for
the valuable input they have given to this 3rd edition:
Ian Bateman
Bob Curnutt
liv
Gerry Elloy
Murph Morrison
Ken Neville
Andy Ramsden
Ron Sarich
Dave Shikany
lv
11.0 TYPICAL QTG TEST INDEX
Each test prescribed in Appendix B of the ICAO Manual is listed in the INDEX on
the following pages, and described in the Sections that follow the index. In both
the index and the Sections following, each test is assigned a number or
designator which corresponds to the designation in the ICAO Manual "Table of
Validation Tests". For example, in the ICAO Manual, the first main test heading
is "1. PERFORMANCE". The first sub-heading is "a) taxi", and the first test
listed is "(1) Minimum Radius Turn".
In this Handbook, the first section is Section 1, Performance and the contents of
Section 1 are 1a Taxi, 1b Takeoff and so on. Section 1a Taxi lists tests 1a(1),
Minimum Radius Turn and 1a(2), Rate of Turn versus Nosewheel Steering Angle.
This Handbook is arranged, therefore, so that each test designator corresponds
to the same test in the ICAO Manual of Criteria for the Qualification of Flight
Simulators (Reference 19).
The "PAGE" referred to below is with reference to the sections of this handbook
which follow, the purpose of which is to give general guidance on individual tests.
SEC.
#
TEST TITLE
1
PERFORMANCE
1-1
1a
TAXI
1A-1
(1)
Minimum Radius Turn
N
Ground
1A-2
(2)
Rate of Turn vs Nosewheel Steering Angle N
Ground
1A-6
1b
TAKEOFF
(1)
Ground Acceleration Time and Distance
N
Takeoff
1B-3
(2)
Minimum Control Speed, Ground (VMCG)
N
Takeoff
1B-7
(3)
Min. Unstick Speed (VMU) or Equivalent
N
Takeoff
1B-12
(4)
Normal Takeoff
N
Takeoff
1B-15
(5)
Critical Engine Failure on Takeoff
N
Takeoff
1B-19
(6)
Crosswind Takeoff
N
Takeoff
1B-23
lvi
C.C.A.
STATE
FLIGHT
COND.
PAGE
1B-1
(7)
Rejected Takeoff
N
Takeoff
1B-27
(8)
Dynamic Engine Failure after T/O
N & NN Takeoff
1B-30
1c
CLIMB
(1)
Normal Climb All Engines Operating
N
Clean
1C-2
(2)
One Engine Inoperative 2nd Segment
Climb
N
2nd seg
climb
1C-5
(3)
One Engine Inoperative Enroute Climb
N
Clean
1C-9
(4)
One Engine Inoperative Approach Climb
N
Approach 1C-12
1d
CRUISE / DESCENT
(1)
Level Flight Acceleration
N
Cruise
1D-2
(2)
Level Flight Deceleration
N
Cruise
1D-5
(3)
Cruise Performance
N
Cruise
1D-8
(4)
Idle Descent
N
Clean
1D-11
(5)
Emergency Descent
N
Per AFM
1D-14
1e
STOPPING
1E-1
(1)
Deceleration Time and Distance, Manual Wheel
Brakes, Dry Runway, No Reverse Thrust
(a) Medium weight
N
(b) Near maximum weight
N
1E-2
(2)
1C-1
1D-1
Deceleration Time and Distance, Reverse
Thrust, No Wheel Brakes, Dry Runway
(a) Medium weight
N
(b) Near maximum weight
N
Landing
Landing
1E-5
Landing
Landing
(3)
Stopping Distance, Wheel Brakes,
Wet Runway
N
Landing
1E-9
(4)
Stopping Distance, Wheel Brakes,
Icy Runway
N
Landing
1E-12
lvii
SEC. TEST TITLE
#
C.C.A. FLIGHT
STATE COND.
PAGE
1F-1
1f
ENGINES
(1)
Acceleration
N
Approach 1F-2
or Landing
(2)
Deceleration
N
Ground
(Takeoff)
lviii
1F-5
SEC. TEST TITLE
#
C.C.A. FLIGHT
STATE COND.
PAGE
2
HANDLING QUALITIES
2-1
2a
STATIC CONTROL CHECKS
2A-1
(1)
Pitch Controller Position vs Force and Surface
Position Calibration
(a) Pitch control force
N
(b) Pitch Controller vs Elevator
N
2A-3
(2)
(3)
(4)
Roll Controller Position vs Force and Surface
(a) Roll controls force
N
(b) Roll Controller vs Aileron
N
(c) Roll Controller vs Spoiler
N
(d) Speedbrake vs Spoiler
N
Rudder Pedal Position vs Force and
Surface Position Calibration
(a) Yaw control force
(b) Pedal vs Rudder
(c) Pedal vs nosewheel steering
Ground
Ground
2A-7
Ground
Ground
Ground
Ground
2A-11
N
N
N
Ground
Ground
Ground
Nosewheel Steering Force and Position
Calibration
(a) Nosewheel steering control force
(b) Nosewheel steering
2A-14
N
N
Ground
Ground
(5)
Rudder Pedal Steering Calibration
N
Ground
2A-17
(6)
Pitch Trim Indicator vs Surface Position
Calibration
N
Ground
2A-19
(7)
Pitch Trim Rate
N
Ground
2A-21
and Approach
(8)
Alignment of Cockpit Throttle Lever vs
selected Engine Parameter
N
Ground
2A-24
lix
SEC. TEST TITLE
#
C.C.A. FLIGHT
STATE COND.
(9)
Brake Pedal Position vs Force and Brake N
System Pressure Calibration
2b
DYNAMIC CONTROL CHECKS
2B-1
(1)
Pitch Control
(a) Pitch Control Dynamics
(b) Pitch Control Dynamics
(c) Pitch Control Dynamics
2B-8
(2)
(3)
Roll Control
(a) Roll Control Dynamics
(b) Roll Control Dynamics
(c) Roll Control Dynamics
Yaw Control
(a) Yaw Control Dynamics
(b) Yaw Control Dynamics
(c) Yaw Control Dynamics
N
N
N
Ground
PAGE
2A-27
Takeoff
Cruise
Landing
2B-11
N
N
N
Takeoff
Cruise
Landing
2B-14
N
N
N
Takeoff
Cruise
Landing
(4)
Small Control Inputs, Pitch
N & NN Approach 2B-17
or Landing
(5)
Small Control Inputs, Roll
or Landing
(6)
Small Control Inputs, Yaw
or Landing
2c
LONGITUDINAL
(1)
Power Change Dynamics
(2)
Flap Change Dynamics
(a) Retraction
(b) Extension
lx
2C-1
N & NN Approach 2C-2
N & NN Takeoff
2C-5
through
Initial Flap
Retraction
to Landing
SEC. TEST TITLE
#
(3)
(4)
(5)
Spoiler/Speedbrake Change Dynamics
(a) Extension
(b) Retraction
Gear Change Dynamics
(a) Retraction
(b) Extension
Longitudinal Trim
(a) Longitudinal Trim
(b) Longitudinal Trim
(c) Longitudinal Trim
C.C.A. FLIGHT
STATE COND.
PAGE
N & NN Cruise
N & NN Cruise
2C-11
2C-14
N & NN Takeoff
2C-17
to Landing
2C-22
N or NN Cruise
N or NN Approach
N or NN Landing
(6)
Longitudinal Manoeuvring Stability (Stick Force/G)
2C-25
(a) Longitudinal Manoeuvring Stability
N & NN Cruise
(b) Longitudinal Manoeuvring Stability
N & NN Approach
(c) Longitudinal Manoeuvring Stability
N & NN Landing
(7)
Longitudinal Static Stability
(8)
Stall Characteristics
(a) Stall Characteristics
(b) Stall Characteristics
N or NN Approach 2C-29
2C-32
N & NN Second
Segment Climb
N & NN Approach or
Landing
(9)
Phugoid Dynamics
NN
Cruise
2C-36
(10)
Short Period Dynamics
N & NN Cruise
2C-40
lxi
SEC. TEST TITLE
#
2d
LATERAL DIRECTIONAL
(1)
Minimum Control Speed, Air (Vmc or Vmcl)
(2)
Roll Response (Rate)
(a) Roll Response (Rate)
(b) Roll Response (Rate)
(3)
Step Input of Cockpit Roll Controller
(4)
Spiral Stability
(a) Spiral Stability
(b) Spiral Stability
(5)
(6)
(8)
lxii
PAGE
2D-1
N or NN Takeoff
2D-2
or Landing
2D-7
N
N
Cruise
Approach or
Landing
N & NN Approach 2D-10
or Landing
2D-13
NN
NN
Cruise
Approach or
Landing
Engine Inoperative Trim
(a) Engine Inoperative Trim
N
(b) Engine Inoperative Trim
N
Rudder Response
(a) Stability Augmentation Systems ON
2D-19
N & NN Approach or
Landing
N & NN Approach or
Landing
(b) Stability Augmentation Systems OFF
(7)
C.C.A. FLIGHT
STATE COND.
2D-16
Second
Segment Climb
Approach or
Landing
Dutch Roll (Yaw Damper OFF)
(a) Dutch Roll
(b) Dutch Roll
2D-23
NN
NN
Cruise
Approach or
Landing
Steady State Sideslip
N
Approach 2D-27
or Landing
SEC. TEST TITLE
#
C.C.A. FLIGHT
STATE COND.
PAGE
2e
LANDINGS
2E-1
(1)
Normal Landing
(a) Flap Position #1
(b) Flap Position #2
2E-2
N & NN Landing
N & NN Landing
(2)
Minimum/No Flap Landing (Max Weight)
N
Minimum 2E-6
Certificated
Landing Flap
(3)
Crosswind Landing
N
Landing
2E-9
(4)
One Engine Inoperative Landing
N
Landing
2E-13
(5)
Autoland Landing
N
Landing
2E-17
(6)
All Engine Autopilot Go Around
N or NN Per AFM
2E-20
(7)
One Engine Inoperative Go Around
NN
Per AFM
2E-23
(8)
Directional Control (Rudder Effectiveness) N
with Reverse Thrust (Symmetric)
Landing
2E-26
(9)
Directional Control (Rudder Effectiveness) N
with Reverse Thrust (Asymmetric)
Landing
2E-30
2f
GROUND EFFECT
(1)
A test to demonstrate ground effect
2g
WINDSHEAR
2G-1
(1)
Windshear
(a) Takeoff (with/without windshear)
(b) Landing (with/without windshear)
2G-10
2F-1
N or NN Landing
2F-3
Takeoff
Landing
lxiii
SEC. TEST TITLE
C.C.A. FLIGHT
#
STATE COND.
2h
ENVELOPE PROTECTION FUNCTIONS
(Applicable to Computer Controlled Aeroplanes only)
PAGE
(1)
Overspeed
2H-3
(2)
Minimum Speed
(a) Takeoff
(b) Cruise
(c) Approach
(3)
(4)
Load Factor
(a) Takeoff
(b) Cruise
N & NN* Cruise
2H-1
2H-6
N & NN* Takeoff
N & NN* Cruise
N & NN* Approach or
Landing
2H-9
N & NN* Takeoff
N & NN* Cruise
Pitch Angle
(a) Cruise
(b) Go-around
N & NN* Cruise
N & NN* G/A
(5)
Bank Angle
N & NN* Approach 2H-16
(6)
Angle of Attack
(a) Second Segment
(b) Approach or Landing
2H-12
2H-19
N & NN* Second
Segment Climb
N & NN* Approach or
Landing
* all tests should be run in both normal and non-normal control states where
the function is different.
lxiv
SEC. TEST TITLE
#
FLIGHT
COND.
PAGE
3
MOTION SYSTEM
3-1
3a
FREQUENCY RESPONSE
N/A
3A-1
3b
LEG BALANCE
N/A
3B-1
3c
TURN AROUND CHECK
N/A
3C-1
3d
MOTION EFFECTS
3D-1
3e
MOTION SYSTEM REPEATABILITY
3E-1
3f
MOTION CUEING PERFORMANCE SIGNATURE
Ground
3F-1
and Flight
3g
CHARACTERISTIC MOTION VIBRATIONS
(1)
(2)
(3)
(4)
(5)
(6)
(7)
Thrust Effects with Brakes Set
Landing Gear Extended Buffet
Flaps Extended buffet
Speedbrake Deployed buffet
Approach to Stall Buffet
High Speed or Mach Buffet
In Flight Vibrations
Ground
3G-1
and Flight
Ground
Flight
Flight
Flight
Flight
Flight
Flight (Clean)
lxv
SEC. TEST TITLE
#
FLIGHT
COND.
PAGE
4
VISUAL SYSTEM
4a
SYSTEM RESPONSE TIME
4b
VISUAL SCENE QUALITY
(1)
(2)
(3)
(4)
(5)
(6)
(7)
Field of View
System Geometry
Surface Contrast Ratio
Highlight Brightness
Vernier Resolution
Lightpoint Size
Lightpoint Contrast Ratio
N/A
N/A
N/A
N/A
N/A
N/A
N/A
4B-2
4B-4
4B-6
4B-8
4B-10
4B-12
4B-14
4c
VISUAL GROUND SEGMENT
Landing
4C-1
lxvi
4-1
N/A
4A-1
4B-1
SEC. TEST TITLE
#
FLIGHT
COND.
PAGE
5
SOUND SYSTEM
5-1
5a
TURBOJET AEROPLANES
5A-1
(1)
(2)
(3)
(4)
(5)
(6)
(7)
(8)
Ready for Engine Start
All Engines at Idle
All Engines at Maximum Allowable Thrust with
Brakes Set
Climb
Cruise
Speedbrake/Spoilers Extended
Initial Approach
Final Approach
5b
PROPELLER AEROPLANES
(1)
(2)
(3)
(4)
(5)
(6)
(7)
(8)
(9)
Ready for Engine Start
All Propellers Feathered
Ground Idle or Equivalent
Flight Idle or Equivalent
All Engines at Maximum Allowable Power
Climb
Cruise
Initial Approach
Final Approach
5c
SPECIAL CASES
5C-1
5d
FLIGHT SIMULATOR BACKGROUND NOISE
5D-1
5e
FREQUENCY RESPONSE
5E-1
Ground
Ground
Ground
Enroute Climb
Cruise
Cruise
Approach
Landing
5B-1
Ground
Ground
Ground
Ground
Ground
Enroute Climb
Cruise
Approach
Landing
lxvii
12.0
EVALUATION NOTES
12.1
GENERAL
The remainder of this document is dedicated to providing information on
the detail of the validation tests required for a Qualification Test Guide.
Each section has some introductory notes, followed by the requisite
information on Test Objective, Evaluation Criteria, Tolerances, Suggested
Plot Parameters and of course the Evaluation Notes themselves. Some
suggestions for the Manual Testing procedure are also included.
There is nothing in these notes which is intended to represent absolute
policy or to set out methods from which there can be no deviation. They
have been written, by a variety of sources, in order to provide assistance
to all sections of the flight simulator industry in determining what the
individual tests are all about.
12.2
OTHER COMMENTS
The Evaluation Criteria for a particular test within a simulator QTG may
well be very specific to that test or that aeroplane, and take account of
anomalies in the data, etc. To use the word 'Criteria' in a document which
by its very nature can only be generic may be misleading.
The information given in this Handbook under the 'Evaluation Notes' is
really only a collation of some of the experiences of several well qualified
simulator engineers and evaluators and as such does not form 'Criteria'
in the sense stated in the ICAO Manual. The intent in this Handbook is to
give assistance to a potential evaluator by providing some insight into the
possible peculiarities of each test and also some general remarks. The
information is given in 'good faith', but should certainly not be taken as
reason to lower the criteria set by the ICAO Manual in any way, nor should
it be used to excuse any particular shortcoming in a given flight simulator.
Finally, many sections contain example test results, and for many of these
it was decided to choose marginal failure cases, as there seemed little
merit in producing a Handbook in which all cases met the criteria perfectly,
since to do so would not be representative. The nature of the engineering
task associated with generating a QTG is such that it would be unheard
of to produce a document that invites no comments whatsoever, but it is
not the intent of this Handbook to suggest that QTG’s should contain
results that fail so that the regulators can find and comment on them. The
examples given are in some cases illustrations of things that can go wrong
during test development, and in others merely a comment on the lack of
lxviii
perfection which will, for the foreseeable future, always be inherent in the
flight simulator evaluation process. That, after all, is the reason for this
Handbook.
Note that any set of results shown by example in this Handbook can only
be a sub-set of the full results for that test, as space does not permit full
sets of results to be included. Also, the comments accompanying these
plots may in many cases be applied to other tests as well - the intention
is to give a fair cross-section of possible problems which may arise during
QTG development in the hope of promoting increased understanding of
the task and thereby a greater ability to evaluate.
lxix
SECTION 1
PERFORMANCE
1a
TAXI
1b
TAKEOFF
1c
CLIMB
1d
CRUISE/DESCENT
1e
STOPPING
1f
ENGINES
1-1
1.0
PERFORMANCE TESTS - GENERAL
The correct behaviour of the simulated aeroplane is clearly a major factor
in transfer of training from the simulator to the real world. The simulator
will obviously represent as accurately as possible the configuration of the
actual aeroplane, especially in regard to the airframe type and engine
variant, and it is primarily these two parameters, in combination, which
are being checked in the Performance Tests section of the QTG.
Concerning the engine variant, it is often the case that there are little or
no flight test data available (for simulator use) from the aeroplane
manufacturer with the airframe/engine combination operated by the
purchaser of the simulator. Under these circumstances it may be
necessary to present the engine information in terms of thrust (or torque,
etc) but this does not remove any need to also show the engine
parameter displayed to the flight crew so that it can be proven that the
QTG test is being run with the engines (and other) mathematical model
included as comprehensively as possible.
Some of the tests required in this section could be construed to be
combinations of both performance and handling qualities tests (e.g.
Crosswind Takeoff), but rather than create a separate section for this
category, thereby increasing the complexity of the documentation, it has
been deemed appropriate to include such tests in the area of prime
importance.
1-2
SECTION 1a
TAXI
1a(1)
Minimum Radius Turn
1a(2)
Rate of Turn vs. Nosewheel Steering Angle (NWA)
1A-1
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
1a(1) - MINIMUM RADIUS TURN
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR
MINIMUM RADIUS TURN ON THE GROUND
CONFORMS TO THE AEROPLANE
DEMONSTRATION
Taxi the aeroplane at a slow speed (10 knots or less)
along the runway. Apply maximum tiller and turn
through a heading change of at least 180 degrees.
Do not use wheel brakes.
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
GROUND
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
NOSEWHEEL STEERING CONTROLLER
POSITION
NOSEWHEEL ANGLE
GROUND SPEED
ENGINE KEY PARAMETERS
YAW RATE
HEADING ANGLE
TURN RADIUS
C.G. DISTANCE ALONG RUNWAY
C.G. DISTANCE ACROSS RUNWAY
NOSEWHEEL DISTANCE ALONG RUNWAY
NOSEWHEEL DISTANCE ACROSS RUNWAY
MAIN GEAR DISTANCE ALONG RUNWAY
MAIN GEAR DISTANCE ACROSS RUNWAY
EVALUATION NOTES
In the past, Operations Manual data were often used
for this check, since these data were not typically
recorded during an aeroplane flight test program.
More recently however Ops Manual data are not
used, being replaced by time histories of the relevant
ground steering information, including the locus of
the cg calculated from flight data and of the paths of
nosewheel and each main gear, which are plotted
1A-2
from engineering simulator data. See Figures 1a-1a
and 1a-1b for examples. The loci of the simulated
aeroplane centre of gravity and of each landing gear
strut should be plotted. The results should be in the
form of circular plots of longitudinal distance along
the runway versus lateral distance as well as time
histories showing all relevant parameters. The use of
engineering judgement tends to be fairly limited for
this test.
TOLERANCES
AEROPLANE TURN
RADIUS
±0.9 m (3 Ft) or ±20%
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
The simulated aeroplane should be aligned with either left or right main landing
gear nearest to the runway edge to ensure enough runway width for the
manoeuvre. The aeroplane speed should be kept constant (using engine
power as necessary) at a value just sufficient for the manoeuvre (typically less
than 10 knots, so that the tyre slip angle is kept to a minimum) and the tiller
should be at its maximum deflection.
EXAMPLE
Figure 1a-1 shows some of the time history plots for an older simulator. The
results look very good versus the aeroplane data, but the calculation of the
turning radius should be shown along with these results so that the simulator
can be compared with the aeroplane data.
1A-3
Figure 1a1-1
Example of Simulator Test Results for Minimum Radius Turn
1A-4
Figure 1a1-2
Example of OPS Manual Data for Minimum Turn Radius
1A-5
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
1a(2) - RATE OF TURN VS. NOSEWHEEL
STEERING ANGLE (NWA)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
HEADING RATES OF CHANGE DURING GROUND
STEERING MANOEUVRES CONFORM TO THE
AEROPLANE.
DEMONSTRATION
While taxiing at a ground speed of approximately 5
knots, turn the aeroplane in a step-wise fashion at
various nosewheel steering angles. The steering
angle should be increased slowly, then held constant
at each position until a constant yaw rate is
achieved. Do not use rudder control or wheel brakes.
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
GROUND
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
NOSEWHEEL STEERING CONTROLLER
POSITION
NOSEWHEEL ANGLE
GROUND SPEED
ENGINE KEY PARAMETERS
YAW RATE (TURN RATE)
HEADING ANGLE
EVALUATION NOTES
Compare the simulator yaw rate with that of the
aeroplane for the various combinations of nosewheel
angle and ground speeds provided. The simulator
ground speed is typically driven to the flight test
values to ensure the test conditions are correct. In
some older data packages, it may have been the
case that neither the engine power settings (except
perhaps thrust) nor runway slope and condition were
specified in the flight test data. Nevertheless the
evaluator should be satisfied that no external
influences such as crosswinds, asymmetric engine
1A-6
thrust or runway slope affects are employed unless
explicitly specified in the aeroplane data source.
The results may be presented either dynamically or
in tabular form. In either case several (two as a bare
minimum) conditions should be tested, preferably at
speeds differing by at least 5 knots ground speed. As
always with any dynamic (time history) data, it is
quite possible that the simulator value will
occasionally exceed the stated tolerance for a short
duration. Judgement must be exercised as to
whether the amount of deviation in a given set of
results is or is not excessive.
TOLERANCES
TURN RATE
±2o/Sec or ±10%
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
The method of running this test manually is theoretically straightforward.
However, if the ground speed is not maintained at its correct value, then the
test results may differ substantially from the aeroplane data. Symmetrical
engine power may be used as required to hold the requisite speed and will not
affect the test results. Ensuring that the correct nosewheel position is used will
probably require the use of an engineering terminal or one of the screens on
the instructor station. The relationship between the nosewheel steering and the
tiller force and position is not the issue here, as it is checked independently in
test 2a(4). The test will be much easier to run if the simulator is positioned on
to a valid runway and the visual system is operational.
EXAMPLE
Figure 1a2-1 shows a continuous time history, but the ground speed only
varies from around 20 knots down to 18.5 knots, and so does not fulfil the
requirements, though the time histories shown do meet the tolerances.
1A-7
Figure 1a2-1
Example of Simulator Test Results for Rate of Turn versus Nosewheel Steering Angle
1A-8
SECTION 1b
TAKEOFF
1b(1)
Ground Acceleration Time and Distance
1b(2)
Minimum Control Speed, Ground (Vmcg)
1b(3)
Minimum Unstick Speed (Vmu)
1b(4)
Normal Takeoff
1b(5)
Critical Engine Failure on Takeoff
1b(6)
Crosswind Takeoff
1b(7)
Rejected Takeoff
1b(8)
Dynamic Engine Failure after Takeoff
1B-1
1B
TAKEOFF TESTS - GENERAL
Most flight test packages for simulator use provide data for more than
one normal takeoff test, though for some aeroplanes there is only one
test available for the engine inoperative and/or the crosswind takeoffs.
The regulatory authorities are concerned that a spread of test data is
used to validate the simulator which identifies the correct characteristics
across all commonly-used takeoff flap settings, so care must be taken
when choosing which set of data is to be used for the QTG for tests
1b(3), 1b(4), 1b(5) and 1b(6).
Often this choice will be limited to which ‘normal’ takeoff data to use,
but if, as is the case in at least one data package, the only available
engine inoperative and crosswind takeoff tests were performed at the
same flap setting, then it will be necessary to use separate flap settings
for the maximum and light weight scenarios required for test 1b(4).
However, the requirements do allow for use of the same test data for
both the Ground Acceleration test (1b(1)) and either the Normal Takeoff
test (1b(4)) or the Rejected Takeoff test (1b(7)), since the former test
must start at brake release and therefore includes the entire ground roll.
1B-2
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
1b(1) - GROUND ACCELERATION TIME AND
DISTANCE
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE THAT THE TIME AND
DISTANCE REQUIRED FOR THE SIMULATOR TO
PERFORM A TAKEOFF RUN CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Perform a normal takeoff ground roll, recording the
time and distance from brake release to rotation
speed (Vr). This test may be combined with the
Normal Takeoff (1b4) or Rejected Takeoff (1b7)
tests.
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
TAKEOFF
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
AIRSPEED
GROUND SPEED
PITCH CONTROLLER POSITION
ELEVATOR ANGLE
PITCH ANGLE
STABILISER ANGLE
WIND SPEED COMPONENTS
ENGINES KEY PARAMETERS
DISTANCE ALONG RUNWAY
EVALUATION NOTES
Compare simulation results and validation data for at
least 80% of the total distance and time to reach Vr
from brake release. The aeroplane distance
information will quite probably have been derived
from the integral of the ground speed during the
post-processing of the flight test data. Since this is
essentially the same method by which the simulated
aeroplane distance is derived, there should be good
correlation between these two sets of data if the
speed match is close. Care should be taken when
1B-3
examining the data because large scales may have
been used when plotting the large changes in both
aeroplane speed and distance. These can cause
difficulty when attempting to determine whether these
parameters are within tolerance. The most critical
set-up parameter is the engine power, so this should
be very carefully matched throughout the test.
Additionally, it should be determined that the
requisite thrust can actually be obtained from the
simulated engine (which may not be the same variant
as used in the flight test program) otherwise manual
testing will be made much more complex. Any
potential problems in this area should be clearly
stated in the QTG. A closed-loop controller may have
been used during automatic testing to maintain
runway centre-line via directional control, but this
method of artificial control should merely represent
the actions of a pilot if he were present and therefore
will have no bearing on the parameters (time and
distance) in question.
TOLERANCES
TIME
DISTANCE
±5%
±5% or ±61 m (200 Ft)
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MANUAL TESTING
The data will probably be the ground-roll portion of a normal takeoff
manoeuvre, so normal takeoff procedures will apply. The engine power MUST
be set accurately so as to achieve and maintain the aeroplane value
throughout the duration of the test. The runway centre-line should be
maintained during the test using directional control only. Small deviations
should not adversely affect the results but, as always, the more accurately the
simulator is flown relative to the aeroplane data, the better the results.
EXAMPLE
The test results shown in Figures 1b1-1 and 1b1-2 show good correlation
between the aeroplane and the simulator. The engine thrust could perhaps
have been slightly better matched by driving the power levers more judiciously
to give the desired thrust increase at approximately 15 seconds, but the total
distance for the simulator is a little lower than for the aeroplane and so
1B-4
correlates. The airspeed clearly shows that there was some wind present
during the manoeuvre.
Figure 1b1-1
Example of Simulator Test Results for Ground Acceleration Time & Distance (Part 1)
1B-5
Figure 1b1-2
Example of Simulator Test Results for Ground Acceleration Time & Distance (Part 2)
1B-6
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TITLE
1b(2) - MINIMUM CONTROL SPEED, GROUND
(VMCG) AERODYNAMIC CONTROLS ONLY, PER
APPLICABLE AIRWORTHINESS REQUIREMENT
OR ALTERNATIVE ENGINE INOPERATIVE TEST
TO DEMONSTRATE GROUND CONTROL
CHARACTERISTICS
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR LOW
SPEED, ENGINE INOPERATIVE GROUND
CONTROL CHARACTERISTICS CONFORM TO
THOSE OF THE AEROPLANE.
DEMONSTRATION
For the Vmcg test, perform a normal takeoff ground
roll. Fail an engine using a fuel cut at the critical
airspeed and attempt to correct the resultant
deviation using rudder control only. Maintain full
rudder until the aeroplane c.g. has started to return
towards the runway centerline. Wheel brakes should
not be used during this demonstration. Minimise pitch
and roll control inputs.
An acceptable alternative test to a Vmcg test is a
snap engine deceleration to idle power at a speed
between V1 and V1-10 knots, otherwise performed n
a similar manner to a Vmcg test. For either method,
the nosewheel steering should be disabled (allowed
to castor) to ensure aerodynamic control only. An
alternative to disabling the nose gear steering is to
hold the nose slightly off the ground during the
engine-failure recovery.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
AIRSPEED
LATERAL DEVIATION FROM RUNWAY
1B-7
CENTRELINE
RUDDER PEDAL POSITION
RUDDER ANGLE
NOSEWHEEL ANGLE
ROLL CONTROLLER POSITION
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
HEADING ANGLE
YAW RATE
BANK ANGLE
RUDDER PEDAL FORCE (IF REVERSIBLE
CONTROLS)
EVALUATION NOTES
The Minimum Control Speed on the ground is
defined as the Calibrated Airspeed at which the
critical engine (for example, the most outboard) is
suddenly cut and at which it is possible to recover
control of the aeroplane with use of primary
aero-dynamic controls alone and without exceeding
30 feet lateral deviation from the runway centreline.
This is typically how the flight test data is presented,
and this should be used in preference to any
"estimated" data being supplied for the actual speed
at which the rudder ceases to be sufficiently effective
for directional control.
The nosewheel hydraulic pressure is normally off in
this test so that the nosewheel is free to castor,
allowing a worst case scenario to be represented
and ensuring that the yawing moment due to thrust
asymmetry is countered only with aerodynamic
control. Alternatively, the flight test may have been
run with the nosewheel raised slightly off the ground
during the engine failure recovery.
An acceptable alternative test to the strict Vmcg test
is a sudden throttle chop at a speed between Vr and
Vr-10 knots. This test may be required if there are
restrictions on the aeroplane, such as on engine fuel
cuts, or if a Vmcg test is not performed for aeroplane
certification.
1B-8
This test may well use data which were produced on
an aeroplane which did not have the same engine (or
engine variant) as that modelled on the simulator. If
this is the case, two test cases should be run. The
first should use the thrust decay from the flight test
data as the driving parameter to show that the overall
aerodynamic effects are simulated correctly. The
other should be run by cutting the fuel to the actual
simulated engine so that it can be shown that the
engine being used during normal training sessions is
representative of the aeroplane and does not cause
the pilot to over- or under-control in these
circumstances. It is highly unlikely that the results
would be identical for the two cases, but they are
necessary to show that all relevant areas of the
mathematical model are representative in this area.
The engine failure/fuel cut speed during the test must
occur within ±1 knot of the aeroplane speed during
the flight test manoeuvre. It is also extremely critical
that the rudder input is timed correctly. There are
occasions where one tenth of a second delay may
cause (on the simulator) a different result.
For what is arguably a more representative
demonstration of engine inoperative takeoff
characteristics, see also Test 1b(5).
TOLERANCES
MAXIMUM AEROPLANE ±25% or ±1.5 m (5 Ft)
LATERAL DEVIATION
(Engine failure speed must be within ±1kt of the
aeroplane data)
And additionally for aeroplanes with reversible flight
controls:
RUDDER PEDAL FORCE ±10% or ±2.2daN (5
Lbs)
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MANUAL TESTING
The minimum control speed test (or equivalent) is very dependent on the timely
input of both the engine failure (fuel cut/throttle slam to idle, etc.) and the
1B-9
appropriate rudder input at the required speed to correct the resultant yaw.
Thus it will almost certainly be necessary for some form of assistance to be
obtained by most pilots when manually running this test on the simulator. The
normal procedure will be to commence a normal takeoff roll with the power set
accurately and allow the pilot flying to concentrate on maintaining control of the
simulator. At the requisite speed the engine should be cut such that the thrust
decay compares favourably with the data and, immediately upon recognition of
this failure, the pilot should input full rudder to arrest the yaw such that the
lateral deviation stops increasing (or even decreases) and then regain and
maintain directional control. It is not necessary to perform the takeoff itself, but
the test should be terminated once it is confirmed that directional control has
been regained.
EXAMPLE
The test result shown in Figure 1b2-1 shows a very poor match for lateral
distance. Examination of the pilot input traces for engine thrust and rudder
position clearly reveals that the engine thrust decay begins approximately onethird of a second late. The effect is that the rudder is driven correctly in relation
to time and airspeed but not in relation to the state of the engine, and the
simulated aeroplane responds to the rudder before it should. This is a classic
example of why the engine failure needs to be within ±1kt of the aeroplane
data, but there may be certain occasions and certain simulators for which the
timing of the engine failure is even more critical than this. Here, the flight test
shows the engine failure to be at approximately 94 knots, but the simulator
engine failure is around 95.5 knots, so the difference is only slightly greater
than the requisite 1 knot, yet the test still fails badly.
1B-10
Figure 1b2-1
Example of Simulator Test Results for Minimum Control Speed, Ground
1B-11
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TITLE
1b(3) - MINIMUM UNSTICK SPEED (VMU) OR
EQUIVALENT TEST TO DEMONSTRATE EARLY
ROTATION CHARACTERISTICS
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATION OF
LOW SPEED HIGH ANGLE OF ATTACK
AERODYNAMICS AND CONTROL POWER NEAR
THE GROUND CONFORMS TO THE AEROPLANE.
DEMONSTRATION
Ground-limited pitch attitude should be achieved well
before minimum lift-off speed is reached. Pitch
control and thrust should match that of the flight-test
aeroplane, along with the simulator acceleration, liftoff speed and pitch attitude. Data should be recorded
from at least 10 kts before start of rotation until at
least 5 seconds after main gear lift-off. If data for a
Vmu test are not available, acceptable alternative
tests include a constant high-attitude takeoff which is
continued through main gear liftoff, or a takeoff with a
rotation speed less than the prescribed Vr for the flap
and weight condition. It should be stated whether the
data was gathered with a tail rubbing strip fitted.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
1B-12
AIRSPEED/GROUND SPEED
MAIN GEAR HEIGHT ABOVE GROUND/RADIO
ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
GEAR STRUT VERTICAL LOADS OR
DEFLECTIONS
EVALUATION NOTES
The Minimum Unstick Speed (Vmu) is defined as that
speed at which the last main landing gear leaves the
ground. Comparisons of airspeed, elevator angle and
pitch angle between the simulator and the aeroplane
should be performed, with particular reference to the
point at which the main gear struts leave the ground.
It is also helpful to plot radio altitude, but this should
not be relied upon to provide the datum point for
Vmu. It is not necessary to plot the entire takeoff roll
as it is only the latter portion of the manoeuvre which
is under scrutiny. However, the plotting process
should commence at least 10 knots prior to the start
of rotation and continue through to 5 seconds after
liftoff so that the data can be clearly interpreted. The
engine power settings and the elevator angle are of
particular importance and must be accurate to
enable a comparison to be made, though the test
should be driven with column position and thus may
not produce a perfect elevator match.
Note that Vmu should be within 3 kt, not just that the
trace should be always within 3 kt, but with unstick
(for example) several seconds late. If the airspeed
match does not hold within 3 kt throughout the time
history (within ‘reason’), but unstick is on-speed and
on-pitch then the test is a pass.
Reference 11 gives further background information.
TOLERANCES
AIRSPEED
PITCH ANGLE
±3 Kts
±1.5o
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MANUAL TESTING
The easiest way to run this test manually is to begin the takeoff roll from the
static condition. This will enable the power to be set accurately well in advance
of the elevator being applied. The plotting need not begin until 10 knots before
start of rotation though. The pitch control input must match the aeroplane data
as precisely as possible or the test is unlikely to pass. This may or may not
1B-13
mean that full nose up elevator is applied - typically the elevator is relaxed by
the flight test pilot in order to gently achieve ground contact pitch attitude. The
test may be concluded no sooner than 5 seconds after the main gear has left
the ground.
EXAMPLE
The result in Figure 1b3-1
below was taken from a
development version of a
test, so should never make
its way into a final QTG!
The engine thrust is clearly
much too high, with the
obvious result that the
airspeed is also high and
the pitch angle out of
tolerance at main gear
liftoff. The data also shows
that the two main gear
struts do not have
aeroplane data
accompanying them, which
is not untypical of older data
packages.
1B-14
Figure 1b3-1
Example of Simulator Test Results for Minimum Unstick
Speed
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TITLE
1b(4) - NORMAL TAKEOFF
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OBJECTIVE
NORMAL TAKEOFF CHARACTERISTICS
CONFORM TO THE AEROPLANE.
DEMONSTRATION
Perform a normal takeoff starting at brake release
and continuing through at least 61m (200 ft) altitude.
At least two tests must be shown: one at near
maximum certificated takeoff weight at a mid c.g.,
and one at a light takeoff weight at an aft c.g. If the
aeroplane has more than one certificated takeoff flap
position, a different flap setting should be used for
each test.
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FLIGHT CONDITION
a) TAKEOFF (NEAR MAXIMUM CERTIFICATED
TAKEOFF WEIGHT)
b) TAKEOFF (LIGHT WEIGHT WITH AFT CENTRE
OF GRAVITY)
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RECORDED
PARAMETERS
AIRSPEED
PITCH CONTROLLER FORCE (IF REVERSIBLE
CONTROLS)
ELEVATOR ANGLE
ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
RATE OF CLIMB
FLIGHT PATH ANGLE
LANDING GEAR POSITION
STABILISER ANGLE
BANK ANGLE
1B-15
AILERON ANGLE(S)
SPOILER ANGLES
EVALUATION NOTES
Whilst it must be borne in mind that this is an
important area for flight crew training, the exactness
of the match between the simulator and the flight test
aeroplane will probably be very difficult to achieve
continuously throughout the duration of the test.
Many factors (such as wind gusting, runway
unevenness and even slightly misrigged controls)
can have a bearing on the final results. The most
important aspect of this test is that the parameters
recorded on the simulator should not be significantly
different from the aeroplane for any prolonged
period. In other words it may be acceptable for very
short duration excursions out of tolerance where
these deviations may be explained by some transient
effect. Most simulator manufacturers will have
provided a test which will be controlled in a "closedloop" fashion. This method uses as its main
command parameters items such as pitch angle and
bank angle (from the flight test data) and drives the
appropriate controls to achieve the correct aeroplane
displacements. Thus a very good match of these
angles is not necessarily an indication that the
simulation has been correctly or adequately
modelled, as the degree to which the control
functions (column, elevator, wheel, etc.) differ from
the flight test data must also be taken into
consideration. See Appendix E for an in-depth
discussion of such methodology.
Because of concerns about aft-cg takeoff
characteristics, including a tendency for some
simulators to auto-rotate prior to Vr,, a new
requirement for a light-weight takeoff at aft cg is now
added. As noted above, it is desired to see test data
at more than one takeoff flap if the aeroplane has
more than one certificated takeoff flap position. If the
aeroplane is equipped with reversible flight controls,
comparisons of simulator results for control column
force with flight test data are also required.
1B-16
TOLERANCES
PITCH ANGLE
ANGLE OF ATTACK
CALIBRATED AIRSPEED
ALTITUDE
±1.5o
±1.5o
±3.0 Kts
±6 m (20 Ft)
controls:
COLUMN FORCE
±10% or ±2.2daN (5
Lbs)
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MANUAL TESTING
The typical form of the aeroplane flight test data will require the pilot to perform
a normal takeoff with the requisite preset conditions. The main issues in this
test are controllability on the runway, rotation forces at Vr and ground effects.
Careful reference to the flight test data may reveal some abnormality in the
method used during the flight test itself and hence the "feel" of this particular
test should not be used solely for any criticism of the simulated takeoff
performance. Ideally the data should call for the full regime from brake release
through at least 61m (200ft) above ground level, but if this is not the case the
manual test will probably be easier to execute when carried out in this fashion,
rather than beginning the test at a preset speed.
Obtaining a good match of airspeed, pitch angle and angle of attack may take
several attempts as it will be very difficult to follow all of the flight test pilot's
inputs simultaneously. Comparison of the results should therefore take this into
account.
EXAMPLE
The results shown in Figure 1b4-1 are generally quite good, but also illustrate
the principal mentioned above that it can be difficult to precisely match all
parameters within tolerance continuously for the entire duration of the test.
Ideally for a normal takeoff, there would be no wind, but in this test there is at
least one wind component which exhibits gusting of several knots, and it may
be this that accounts for the larger-than-desired bank angle excursion on liftoff.
Bank angle is, strictly speaking, not a toleranced parameter for a normal
takeoff test, but the regulatory authorities may choose not to ignore it as being
relevant if its value differs significantly from the aeroplane data. Another
feature of these plots is the scale used for the radio altitude trace, which has
been poorly chosen such that it is difficult to determine compliance with the
±20ft tolerance.
1B-17
Figure 1b4-1
Example of Simulator Test Results for Normal Takeoff
1B-18
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TITLE
1b(5) - CRITICAL ENGINE FAILURE ON TAKEOFF
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OBJECTIVE
THE TAKEOFF WITH THE FAILURE OF THE
CRITICAL ENGINE CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Perform an engine out takeoff to at least 61m (200 ft)
altitude. The engine cut should occur at
approximately V1 speed, and is usually simulated by
a throttle slam from the normal takeoff setting to the
idle position. Test at near maximum certificated
takeoff weight.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
LATERAL CONTROLLER FORCE (IF REVERSIBLE
CONTROLS)
BANK ANGLE
ROLL RATE
CONTROLS)
RUDDER ANGLE
ELEVATOR ANGLE
LONGITUDINAL CONTROL FORCE (IF
REVERSIBLE CONTROLS)
STABILISER ANGLE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ALTITUDE
AIRSPEED
1B-19
YAW RATE
SIDESLIP ANGLE
HEADING ANGLE
EVALUATION NOTES
For evaluation purposes, the engine failure should
occur within 3 knots of the speed demonstrated by
the aeroplane. Obviously the main area under
examination here is the simulated aeroplane
response to the engine failure. Again the data should
be carefully examined for any abnormalities, along
with the pilot inputs and responses used to
implement any "closed-loop control" (see the notes
for "Normal Takeoff"). Of particular interest is the
amount of rudder and wheel required to contain the
engine failure situation and whilst it may be the case
that there are short-term deviations in these
parameters between the simulator and the
aeroplane, the trends should nevertheless be very
close. Do not ignore the longitudinal parameters as,
for example, the required stick force may not
correspond directly with that of the Normal Takeoff.
When comparing this test with the Vmcg test (Test
1b(2)), tiny changes in the simulator start conditions
for that test can affect the result. This test is much
less sensitive to engine cut-conditions, and gives a
much more reliable indication of the simulators
dynamic response at engine failure.
If the aeroplane is equipped with reversible flight
controls, comparisons of simulator results for control
column force, wheel force and rudder pedal force
with flight test data are also required.
TOLERANCES
1B-20
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
ALTITUDE
BANK ANGLE
SIDESLIP ANGLE
HEADING
±3.0 Kts
±1.5o
±1.5o
±6 m (20 Ft)
±2.0o
±2.0o
±3.0o
and additionally for aeroplanes with reversible flight
controls:
COLUMN FORCE
±10% or ±2.2 daN (5
Lbs)
WHEEL FORCE
±10% or ±1.3 daN (3
Lbs)
RUDDER PEDAL FORCE
±10% or ±2.2 daN (5
Lbs)
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MANUAL TESTING
The data will usually provide for the use of normal takeoff procedures up to the
point at which the engine is failed. The timing of the failure is obviously critical,
and whilst the ICAO Manual allows the engine failure speed to be within ±3
knots of the flight test speed, the reality is that a speed much more accurate
than this is needed in order to facilitate a reasonable comparison. Reasons for
this are explained in the comments for test 1b(2). After the engine has been
failed, the pilot should continue with the takeoff manoeuvre while maintaining
and assessing both bank and heading control. The entire takeoff profile should
be recorded from brake release through 61m (200ft) above ground level. See
notes for the Normal Takeoff test regarding achievement of a good match.
EXAMPLE
The results shown in Figure 1b5-1 again illustrate the point that it can be
problematic to fully comply with the tolerances for the entire test duration. The
sideslip angle deviates from the requisite ±2 degrees, but this could probably
be improved by a better match of bank angle (and therefore heading, not
shown). Interestingly, however, the captain’s wheel position is already
significantly greater than the aeroplane data, and using wheel to reduce the
bank angle would make this even worse. The methods used to improve such
results as these can be very complex, but usually entail fine adjustments of
several parameters - particularly the airspeed and the liftoff point in this
particular case.
1B-21
Figure 1b5-1
Example of Simulator Test Results for Engine Inoperative Takeoff
1B-22
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TITLE
1b(6) - CROSSWIND TAKEOFF
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OBJECTIVE
CROSSWIND TAKEOFF CHARACTERISTICS
DEMONSTRATION
Perform a crosswind takeoff from brake release to at
least 61m (200ft) altitude. Set takeoff thrust prior to
brake release in order to assess the aeroplane
response at very low speed (less than 40 kts ground
speed). The crosswind component should be at least
60% of the AFM value.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
AIRSPEED
CONTROL WHEEL FORCE (IF REVERSIBLE
CONTROLS)
BANK ANGLE
ROLL RATE
AILERON ANGLE(S)
SPOILER ANGLES
CONTROLS)
RUDDER ANGLE
ELEVATOR ANGLE
LONGITUDINAL CONTROL FORCE (IF
REVERSIBLE CONTROLS)
STABILISER ANGLE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
1B-23
ALTITUDE
NOSEWHEEL ANGLE
SIDESLIP ANGLE
HEADING ANGLE
EVALUATION NOTES
As with the Critical Engine Failure on Takeoff test, it
is primarily (though by no means exclusively) the
lateral and directional parameters which are of
greatest concern here. However, the main feature
which should distinguish these results is that there is
no portion of this test which is free of these effects,
since the crosswind is present from the beginning of
the takeoff roll. Note that the simulator must exhibit
the correct trends for rudder/pedal and heading for
speeds up to 40 knots. The initial flight test value of
sideslip (i.e. at very low speeds) should not be
regarded with too much credence as the theoretical
value at zero forward speed is mathematically
undefined. Its value becomes much more important
at and above the point at which the aeroplane
forward speed equals the value of the crosswind
speed. Care should be taken to look for a tendency
to “anti-weathercock” or to yaw away from the
direction of the crosswind, at very low speed and with
high power setting, exhibited by some aeroplanes.
Because this test will almost certainly be run "closedloop", the value of the rudder required to maintain
heading should be carefully scrutinised. Again, small
short-term excursions outside of the allowable
tolerance may be explained by wind gusting
(perhaps more so in this test than in nearly all others)
but a constant offset of more than a degree or so
may well indicate that the rudder power is incorrect.
The roll control input on liftoff should also be
compared. If the aeroplane is equipped with
reversible flight controls, comparisons of simulator
versus aeroplane data must be performed for control
column force, wheel force and rudder pedal force.
TOLERANCES
CALIBRATED AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
1B-24
±3.0 Kts
±1.5o
±1.5o
BANK ANGLE
SIDESLIP ANGLE
ALTITUDE
HEADING
±2.0o
±2.0o
± 6 m (20 Ft)
±3.0o
controls:
COLUMN FORCE
±10% or ±2.2 daN (5
Lbs)
WHEEL FORCE
±10% or ±1.3 daN (3
Lbs)
RUDDER PEDAL FORCE
±10% or ±2.2 daN (5
Lbs)
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MANUAL TESTING
Perform takeoff manoeuvre with the preset conditions as per the flight test
data. The entire ground run should be flown, from brake release through at
least 61m (200ft) above ground level while maintaining heading control. The
simulation of this test will require the use of special test data for the wind profile
rather than the simple insertion of a generalised crosswind from the instructor
station.
Pitch angle, angle of attack, airspeed, height above ground, bank and sideslip
angles should be compared versus actual aeroplane data, but see the notes
from the Normal Takeoff test concerning the achievement of an exact match.
EXAMPLE
The example test results in Figure 1b6-1 show a generally good match
between the simulator and the aeroplane data for those parameters shown.
However, the plots begin when the airspeed reaches approximately 120 knots,
whereas the requirements call for the entire ground roll to be compared. Some
aeroplane data for this test has in the past been presented in two sections,
brake release to liftoff and liftoff to 200 ft above ground. If the data is presented
in this way it can mean that merging the two sets of flight test/proof of match
can present problems when attempting to use that data directly in the
simulation. The engine thrust match indicates that the engines were correctly
driven using power lever angles (or equivalent), but there is a significant
reduction in thrust - present in the aeroplane data, and matched by the
simulator - which may warrant explanation from the data provider.
1B-25
1B-26
Figure 1b6-1
Example of Simulator Test Results for Crosswind Takeoff
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TITLE
1b(7) - REJECTED TAKEOFF
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR TIME
AND DISTANCE TO STOP AFTER A REJECTED
TAKEOFF CONFORM TO THE AEROPLANE.
DEMONSTRATION
Perform a rejected takeoff starting from brake
release to a full stop using maximum braking effort,
where the speed of reject is at least 80% of V1. Test
at near maximum certificated takeoff weight. Use
maximum wheel brakes (autobrakes if available) and
ground spoilers as appropriate.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
STABILISER ANGLE
HEADING ANGLE
SPOILER ANGLES
SPEED BRAKE POSITION
BRAKE PEDAL POSITION
HYDRAULIC BRAKE PRESSURES
BRAKE TEMPERATURE
ELEVATOR ANGLE
PITCH ANGLE
EVALUATION NOTES
Of prime consideration here is the braking
performance. However, the close synchronisation of
all inputs, including rapid engine power reduction,
ground spoiler deployment and wheel brake
application will have an effect on the results. The test
should consist of the full takeoff run, beginning from
the static position, and it is this which will distinguish
1B-27
it from the stopping tests in Section 1e. Directional
control should at all times be maintained. The time
and distance are for brake release to a full stop.
TOLERANCES
TIME
DISTANCE
±5% or ±1.5 Sec
±7.5% or ±76 m (250
Ft)
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MANUAL TESTING
The usual procedure calls for a full takeoff roll with the preset conditions as per
the validation data. The manoeuvre should be executed from brake release
and the aborted takeoff initiated at the appropriate speed. Maximum braking
effort (auto or manual) should be applied, and braking continued until the
aeroplane comes to a complete stop. Autobrakes should be used if they are
available.
This is usually a fairly simple test and so a good match can often be achieved
versus the flight test data. It is necessary though to properly coordinate the
actions required to initiate the abortion of the takeoff.
EXAMPLE
The example in Figure 1b7-1 provides for generally good matches for each of
the parameters shown, though the initial value of engine thrust could perhaps
be increased slightly to remove the offset. However, the aeroplane data begins
at approximately 150 knots - only 3 seconds or so prior to the abort manoeuvre
is initiated, rather than showing the entire ground roll from brake release as per
the requirements. Little can be done by the simulator manufacturer if the data
is presented in this way, especially if there is wind profile data present, since all
parameters have indeterminate values below the initial speed provided in the
data. If there is a choice, then selection of a different set of data may be the
answer. In this particular case, there was no alternative.
1B-28
Figure 1b7-1
Example of Simulator Test Results for Rejected Takeoff
1B-29
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TITLE
1b(8) - DYNAMIC ENGINE FAILURE AFTER
TAKEOFF
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OBJECTIVE
TO DEMONSTRATE THAT THE SHORT-TERM
FREE RESPONSE OF THE SIMULATOR TO AN
ENGINE FAILURE AFTER TAKEOFF CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Perform a simulated takeoff at a safe altitude out of
ground effect, then fail an engine and allow the
aeroplane to respond freely for 5 seconds, or until
the bank angle reaches 30 degrees, whichever
occurs first, then initiate recovery. The engine failure
may be simulated by a snap deceleration to idle
power.
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FLIGHT CONDITION
TAKEOFF
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RECORDED
PARAMETERS
1B-30
BANK ANGLE
RUDDER ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
PITCH ANGLE
ALTITUDE
AIRSPEED
SIDESLIP ANGLE
HEADING ANGLE
PITCH RATE
YAW RATE
ROLL RATE
AILERON ANGLE
SPOILER ANGLES
EVALUATION NOTES
When comparing the aeroplane time history with the
simulator test results, it should be borne in mind that
the parameters in question for this test are the
simulated aeroplane body rates (roll, pitch and yaw)
which are generated by the engine failure with the
aeroplane in takeoff configuration. These in turn will
be highly dependent on the rate at which the failed
engine thrust decays and also on the thrust levels at
both takeoff and windmill (or idle) power, as well as
on the simulated aerodynamics themselves.
Therefore great care should be taken to ensure that
these parameters above all others are being
accurately reproduced during the test. This is
especially so, bearing in mind the very short duration
of the test after the engine has been failed (5
seconds or less). Further changes in the aeroplane
state beyond this period should not be expected to
correspond well with the aeroplane data, especially if
there is pilot activity for this portion of the
manoeuvre. The engine failure is usually replicated
by a snap deceleration to the idle position rather then
a fuel cut. The speed at which the engine failure is
introduced must be within ±3 knots of the aeroplane
data.
Note that for Computer Controlled Aeroplanes there
should be two tests, one each for the normal and
non-normal configurations.
The 'Body Angular Rates' in the Tolerances section
below refer to Pitch Rate, Roll Rate and Yaw Rate in the same axis system used in the flight test data.
TOLERANCES
BODY ANGULAR RATES
±20% or ±2o/sec
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MANUAL TESTING
After takeoff, fail the critical engine as per flight test data. The test parameters
should be recorded from 5 seconds before engine failure (which must be
1B-31
actuated as close as possible to the speed at which the engine was failed
during the flight test, and certainly within ±3 knots) to 5 seconds after engine
failure or when the aeroplane reaches a bank angle of 30 deg, whichever
occurs first. Naturally, the simulated aeroplane should not be allowed to fly out
of control beyond the above time or bank angle, but in any case the test will
usually be terminated as soon as it is feasible to do so, which may be prior to
wings level recovery.
The tolerances apply to pitch rate, roll rate and yaw rate, but for manual testing
it may be more useful to base the procedure on the appropriate angles
themselves rather than on the angular rates, which can be awkward to
manually quantify to the extent required.
For computer controlled aeroplanes, this test must be performed in both
normal and non-normal control states so as to ascertain that both the
electronic flight control system and the natural aeroplane aerodynamics are
being correctly modelled.
EXAMPLE
The example result in Figure 1b8-1 appears to exhibit generally good matches
for the three angular rates (pitch rate, roll rate and yaw rate), but closer
examination is likely to reveal that the ‘aeroplane’ data used here is actually
engineering simulation data, not true flight test. On this basis, the parameters
in question must follow the data more closely than for flight test data, and the
result shown would not adequately meet the tolerances.
1B-32
Figure 1b8-1
Example of Simulator Test Results for Dynamic Engine Failure After Takeoff
1B-33
1B-34
SECTION 1c
CLIMB
1c(1)
Normal Climb All Engines Operating
1c(2)
One Engine Inoperative Second Segment Climb
1c(3)
One Engine Inoperative Enroute Climb
1c(4)
One Engine Inoperative Approach Climb
1C-1
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TITLE
1c(1) - NORMAL CLIMB ALL ENGINES OPERATING
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OBJECTIVE
ENGINE THRUST, AERODYNAMIC DRAG AND
ATMOSPHERE IN A STEADY STATE NORMAL
CLIMB CONDITION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Establish a steady climb at nominal climb power with
flaps and landing gear retracted over an altitude
interval of at least 300 m (1000 ft) at mid initial climb
altitude.
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FLIGHT CONDITION
CLEAN
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
PITCH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
RATE OF CLIMB
FLIGHT PATH ANGLE
EVALUATION NOTES
The prime consideration for this test is whether the
recorded rate of climb matches that of the aeroplane.
However, rate of climb time histories can exhibit
sensitivities which are difficult to follow in the
simulator. Therefore, an equivalent method of
measuring the overall climb rate during the test is to
measure the time taken to climb from one specific
altitude to another at the airspeed recorded during the
flight test program. The altitude interval must be at
1C-2
least 300m (1000ft). Power settings are of particular
importance, especially over a relatively long duration,
and notwithstanding any transient inconsistencies, the
actual measured rate of climb should not be ignored.
Typically this test will be run with an automatic trimmer
on the horizontal stabiliser. The engine settings should
follow the flight test data very closely.
Manufacturer's performance manual data may be
used instead of actual flight test data, but note that
snapshot data is not acceptable.
TOLERANCES
RATE OF CLIMB
AIRSPEED
±5% or ±0.5 m/sec (100 Ft/Min)
±3 Kts
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MANUAL TESTING
Typically, the method for this test will require the pilot to climb at constant
calibrated airspeed for at least 60 seconds using stabiliser trim as required. The
average rate of climb for the whole manoeuvre is checked by measuring the
altitude change over 60 seconds, provided that the altitude change during this
period exceeds 300m (1000ft). It will be useful to the pilot flying this test if the
altitude is initially set to 1000 feet or so below that at which the recording/plotting
needs to commence, so as to allow him to stabilise the aeroplane before the
tolerances are applied. Also, it may be feasible to make use of the autopilot.
EXAMPLE
In any QTG test of reasonably long duration which is run in an open-loop
manner, there may be a tendency for certain parameters to ‘drift’. This may also
be true using closed-loop controllers if they are not set up very carefully. In
Figure 1c1-1 the airspeed is gradually reducing such that by the end of the test
it is getting close to the 3 knot tolerance. This is not a problem as such with the
result shown, but the operator would be prudent to check carefully all future
results for this tendency to ensure that the test does not fail and/or to adjust the
controller gains to give a more consistent result.
1C-3
Figure 1c1-1
Example of Simulator Test Results for Climb in Clean Configuration
1C-4
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TITLE
1c(2) - ONE ENGINE INOPERATIVE SECOND
SEGMENT CLIMB
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OBJECTIVE
ATMOSPHERE IN AN ENGINE OUT SECOND
SEGMENT CLIMB CONDITION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Establish a steady climb with one engine inoperative
and takeoff power on the operating engine(s) at
takeoff flaps and landing gear up over an interval of at
least 300 m (1000 ft)
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FLIGHT CONDITION
SECOND SEGMENT CLIMB
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
PITCH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
RATE OF CLIMB
FLIGHT PATH ANGLE
AILERON ANGLE
SPOILER ANGLES
YAW CONTROLLER POSITION
RUDDER ANGLE
SIDESLIP ANGLE
EVALUATION NOTES
recorded steady state rate of climb matches that of the
1C-5
aeroplane for the engine inoperative second segment
climb condition. However, rate of climb time histories
can exhibit sensitivities which are difficult to follow in
the simulator. Therefore, an equivalent method of
measuring the overall climb rate during the test is to
measure the time taken to climb from one specific
altitude to another at the airspeed recorded during the
flight test program. The altitude interval must be at
least 300 m (1000 ft). Power settings are of particular
importance, especially over a relatively long duration,
and notwithstanding any transient inconsistencies, the
actual measured rate of climb should not be ignored.
Typically this test will be run with an automatic trimmer
on the horizontal stabiliser. The operating engine
settings should follow the flight test data very closely.
Flight data for this test are typically from performance
check climbs for which the inoperative engine is shut
down (windmilling) and the remaining engines are set
at takeoff power.
Rudder closed-loop control may be used to balance
the asymmetric thrust and minimise sideslip so as to
maintain heading, but care should be taken to ensure
the rudder excursions are not excessive when
compared to flight test data.
Manufacturer's aeroplane performance manual data
may be used instead of flight test data, but note that
snapshot data is not acceptable. The reason that the
rate of climb must not be less than Approved Flight
Manual values is because AFM values are
conservative - usually based on a minimum thrust
engine.
TOLERANCES
RATE OF CLIMB
AIRSPEED
±5% or ±0.5 m/Sec (100 Ft/Min),
but not less than the Approved
Flight Manual rate of climb
±3 Kts
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MANUAL TESTING
The ICAO Manual specifies that the climb be maintained over an altitude interval
1C-6
of at least 300 m (1000 ft). Typically, stabiliser trim is used as required. Rudder
trim should be used to balance the asymmetric thrust and minimise sideslip so
as to maintain heading. Maintaining the aeroplane data airspeed is of particular
importance if good results are to be obtained. Rudder trim should be used to
balance the asymmetric thrust and minimise sideslip so as to maintain heading,
and the bank angle should be determined from the data and followed as closely
as possible. It will be useful to the pilot flying this test if the altitude is initially set
to 1000 feet or so below that at which the recording/plotting needs to commence,
so as to allow him to stabilise the aeroplane before the tolerances are applied.
If there are two pilots, one can fly the manoeuvre and the other check and correct
engine power settings.
EXAMPLE
Figure 1c2-1 is a clear illustration of a set of aircraft data that were inadequate
(no wind speeds were offered by the data provider), but which makes very little
difference to the overall test result. Under these circumstances it is doubtful that
it would be necessary to follow the wheel position accurately throughout the test,
and use of a closed-loop controller has been used to maintain bank angle (for
which again, no data was provided, but a mean value of -2 degrees or so would
presumably have been used). The rate of climb and rudder angle are clearly
within acceptable limits for this very long duration climb test.
1C-7
Figure 1c2-1
Example of Simulator Test Results for Engine Inoperative Climb, Second Segment
1C-8
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TITLE
1c(3) - ONE ENGINE INOPERATIVE ENROUTE
CLIMB
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OBJECTIVE
ATMOSPHERE IN AN ENGINE OUT ENROUTE
AEROPLANE.
DEMONSTRATION
Establish a steady climb with one engine out with
nominal climb power on the operating engine(s) and
with flaps and landing gear retracted over an interval
of at least 1550 m (5000 ft).
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FLIGHT CONDITION
CLEAN
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RECORDED
PARAMETERS
CALIBRATED AIRSPEED OR MACH NUMBER
PRESSURE ALTITUDE
RATE OF CLIMB
FLIGHT PATH ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
CONTROL WHEEL/LATERAL CONTROLLER
POSITION
AILERON ANGLE
SPOILER ANGLES
RUDDER ANGLE
FUEL FLOW OR FUEL QUANTITY
WIND COMPONENTS
EVALUATION NOTES
recorded time to climb, distance travelled and fuel
used match that of the aeroplane data for the engine
inoperative climb condition. The altitude interval must
1C-9
be at least 1550m (5000ft). Power settings are of
particular importance, especially over a relatively long
duration, and of course will especially affect the fuel
used value if they are not set correctly. Fuel used can
be based on a measurement of fuel quantity at the
beginning and at the end of the time segment for the
climb, or by integrating fuel flow over the same
interval. Typically this test will be run with an
automatic trimmer on the horizontal stabiliser.
Rudder closed-loop control may be used to balance
the asymmetric thrust and minimise sideslip so as to
maintain heading, but care should be taken to ensure
the rudder excursions are not excessive when
compared to flight test data.
Manufacturer's aeroplane performance manual data
may be used instead of flight test data, but note that
snapshot data is not acceptable.
TOLERANCES
TIME
DISTANCE
FUEL USED
±10%
±10%
±10%
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MANUAL TESTING
Typically, the method for this test will require the pilot to climb at constant
calibrated airspeed for several minutes (corresponding to at least 1500 m / 5000
ft) using stabiliser trim as required. Rudder trim should be used to balance the
asymmetric thrust and minimise sideslip so as to maintain heading. Maintaining
the aeroplane data airspeed is of particular importance if good results are to be
obtained. The time and perhaps the fuel used may be recorded by the pilot in the
cockpit, but the distance travelled for the whole manoeuvre will be checked by
examination of the computed values of these parameters.
1C-10
EXAMPLE
The example shown in
Figure 1c3-1 shows the
conventional parameters for
a climb test, rather than
those specified for this
particular test (i.e. time,
distance, fuel used).
However, what it illustrates is
the result of a poor
lateral/directional trim
immediately prior to the
beginning of the test. The
bank angle (not shown) has
rolled off and the pitch angle,
airspeed and altitude follow
accordingly. Obviously, with
a result such as this it is
impossible to accurately
assess the required
parameters. The test as
shown fails, but was
corrected by altering the way
in which the engine
inoperative trim was
performed.
Figure 1c3-1
Inoperative Enroute Climb
1C-11
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TITLE
1c(4) - ONE ENGINE INOPERATIVE APPROACH
CLIMB FOR AEROPLANES WITH ICING
ACCOUNTABILITY IF REQUIRED BY THE AFM FOR
THIS PHASE OF FLIGHT
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OBJECTIVE
ATMOSPHERE IN AN ENGINE OUT APPROACH
AEROPLANE FOR AN AEROPLANE WITH ICING
ACCOUNTABILITY.
DEMONSTRATION
Establish a steady climb with one engine out and goaround power on the operating engine(s) with
approach or go-around flaps and landing gear
retracted over an altitude interval of at least 300 m
(1000 ft). All anti-ice or de-icing systems should be
operating normally. It is not intended that ice
accumulation be present on the lifting surfaces.
Operational considerations for approach in icing, such
as adjustment to airspeed and weight limit, should be
in effect.
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FLIGHT CONDITION
APPROACH
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RECORDED
PARAMETERS
1C-12
PRESSURE ALTITUDE
AIRSPEED
PITCH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
RATE OF CLIMB
FLIGHT PATH ANGLE
AILERON ANGLE
SPOILER ANGLES
YAW CONTROLLER POSITION
RUDDER ANGLE
SIDESLIP ANGLE
EVALUATION NOTES
See notes for Test 1c(2). This test will almost certainly
use a very similar, if not identical technique. This test
is only required for those aeroplanes whose Approved
Flight Manuals require "Icing Accountability". In other
words, it is intended to apply only to aeroplanes for
which there is an operational requirement for flight in
icing. If the AFM does not state any operational
limitations (such as approach speed increment or
modified flap setting) in icing conditions then the
aircraft does not have icing accountability. - Most jet
transport aircraft do not, many turbo-prop aircraft do.
Approach climb means go-around climb condition with
one engine inoperative. It does not require testing of
the effects of ice accumulation; just the effects of
systems (anti-ice on engine bleeds, etc.) and
operational limitations (weight limit, addition to speed,
etc.).
Note that snapshot data is not acceptable.
TOLERANCES
RATE OF CLIMB
AIRSPEED
±5% or ±0.5 m/Sec (100 Ft/Min)
but not less than the Aeroplane
Flight Manual rate of climb
±3 Kts
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MANUAL TESTING
See notes for Test 1c(2). This test will almost certainly use a very similar, if not
identical technique, the exception being that the aeroplane should be configured
with all anti-ice and de-ice systems operating normally, gear up and go-around
flap. All icing accountability considerations, in accordance with the AFM for an
approach in icing conditions, should be applied.
EXAMPLE
1C-13
In Figure 1c4-1 the idle thrust (on the #2 engine) is lower by approximately 100
lbs than the aircraft data suggests. While this is small compared with the
combined total net thrust, nevertheless an attempt has been made to offset this
small inconsistency by applying an equivalent extra amount on the #1 engine.
The difference in the rate of climb that would result if this were not applied would
most likely be negligible, but the modification was at least made for a logical
reason. Few regulators would be concerned with this result, but it may still be
worth an explanatory note in the QTG.
Figure 1c4-1
Example of Simulator Test Results for Engine Inoperative Climb, Approach
1C-14
SECTION 1d
CRUISE/DESCENT
1d(1)
Level Flight Acceleration
1d(2)
Level Flight Deceleration
1d(3)
Cruise Performance
1d(4)
Idle Descent
1d(5)
Emergency Descent
1D-1
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TITLE
1d(1) - LEVEL FLIGHT ACCELERATION
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OBJECTIVE
ENGINE POWER VERSUS AERODYNAMIC DRAG,
IN THE CRUISE CONFIGURATION, CONFORMS TO
THE AEROPLANE.
DEMONSTRATION
After establishing a level-flight trim condition in the
clean configuration at cruise altitude, apply maximum
continuous power (or equivalent) and perform an
acceleration of at least 50 knots airspeed at constant
altitude. The test may be performed manually or with
autopilot / altitude hold function operating.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED OR MACH NUMBER
PITCH ANGLE
ANGLE OF ATTACK
RATE OF CLIMB
FLIGHT PATH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
EVALUATION NOTES
The test is usually conducted in a straightforward
manner by trimming for level flight at a particular
condition and then increasing the power to a
predetermined level at a predetermined rate and
allowing the airspeed to increase whilst maintaining
constant altitude. Obviously retrimming will be
necessary as the test progresses, either with pitch
1D-2
control or stabiliser trim. The prime parameter in
question is the time taken to increase the airspeed to
a given value, but it should be ensured that the other
relevant parameters such as engine power (or thrust)
and aeroplane gross weight are as per the flight test
data.
TOLERANCES
TIME
±5%
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MANUAL TESTING
The aeroplane should be trimmed for steady level flight in the cruise configuration
and ideally allowed to remain in this state for at least 5 seconds whilst recording
takes place prior to the initiation of the manoeuvre.
The engine throttle levers should be steadily pushed forward so that the engine
power increases to the requisite constant level, which should ideally be exactly
the same for each engine, and the test commenced only after the engine power
has stabilised. This may require the initial airspeed to be set lower then that at
which recording begins. The autopilot may be used to maintain altitude (or the
pilot may elect to do this manually) until the speed has increased by at least 50
knots from the initial value.
EXAMPLE
Referring to Figure 1d1-1, the aeroplane data (dotted line) indicates that the total
time taken to increase speed from 200 knots to 253 knots is 140 seconds.
Applying the 5% tolerance to this value gives a maximum time of 147 seconds,
whereas the simulator takes around 149 seconds. Hence the test is a marginal
failure. However, the thrust (at least on no. 1 engine) is a few hundred pounds
low. Increasing this to the correct value would probably result in the test result
just coming within the tolerance band, albeit still slightly on the high side of the
aeroplane value.
1D-3
1D-4
Figure 1d1-1
Example of Simulator Test Results for Level Flight Acceleration
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TITLE
1d(2) - LEVEL FLIGHT DECELERATION
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR LEVEL
FLIGHT DECELERATION PERFORMANCE
CONFORMS TO THE AEROPLANE.
DEMONSTRATION
After establishing a level-flight trim condition in the
clean configuration at cruise altitude, reduce power to
the idle setting and perform a deceleration of at least
50 knots airspeed at constant altitude. The test may
be performed manually or with autopilot / altitude hold
function operating.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
PITCH ANGLE
ANGLE OF ATTACK
RATE OF CLIMB
FLIGHT PATH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
SPEEDBRAKE POSITION
SPOILER ANGLES
EVALUATION NOTES
The test is usually conducted in a straightforward
manner by trimming for level flight at a particular
condition and then reducing the power to a given
value (usually idle) at a predetermined rate and
allowing the airspeed to decrease whilst maintaining
constant altitude. Obviously retrimming will be
1D-5
necessary as the test progresses, either with pitch
control or stabiliser trim. The prime parameter in
question is the time taken for the airspeed to decrease
to a given value of at least 50 knots less than the
initial value, but it should be ensured that the other
relevant parameters, especially engine power (or
thrust) and aeroplane gross weight, are as per the
flight test data.
TOLERANCES
TIME
±5%
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MANUAL TESTING
The aeroplane should be trimmed for steady level flight in the cruise configuration
and ideally allowed to remain in this state for at least 5 seconds whilst recording
takes place prior to the initiation of the manoeuvre.
The engine throttle levers should be steadily brought back to the flight test
(usually idle) position. The autopilot may be used to maintain altitude (or the pilot
may elect to do this manually) until the speed has decreased by at least 50 knots
from the trim value.
EXAMPLE
The result in Figure 1d2-1 is fairly straightforward to interpret, and seems to
match the aeroplane data well. Note though the unreliability of using rate of climb
as a tolerance parameter during a time history test, and this is why it has
generally been avoided in the ICAO Manual. Note that the full set of results would
obviously have included engine parameters, as well as other items necessary for
the interpretation of the results.
1D-6
Figure 1d2-1
Example of Simulator Test Results for Level Flight Deceleration
1D-7
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TITLE
1d(3) - CRUISE PERFORMANCE
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OBJECTIVE
ENGINE PERFORMANCE IN THE CRUISE
CONFIGURATION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Fly the aeroplane for at least 3 minutes in a level flight
trimmed state in the clean configuration at cruise
altitude. Do not alter the power setting, pitch control or
stabiliser trim position.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
PITCH ANGLE
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
TOTAL FUEL WEIGHT OR FUEL FLOW
EVALUATION NOTES
The prime purpose of this test is to ascertain that the
engine parameters are consistent with one another
and with the aeroplane during a steady state cruise
situation. Hence the test does not need to be plotted
as a time history, especially since all that would be
seen on the plots is a series of virtually straight lines.
Instead, two separate sets of the same parameters
should be recorded - one at the beginning of the 3
minute period and the second at the end, to check that
these items correspond well with the aeroplane data.
Alternatively, a single snapshot may be presented
showing instantaneous fuel flow. Also, the total fuel
1D-8
weight should be checked for consistency with the
recorded fuel flow. The simulated aeroplane will
probably be held in an altitude hold condition, perhaps
using the autopilot, but the autothrottle should not be
used.
TOLERANCES
EPR
N1
TORQUE
FUEL FLOW
±0.05 or
±5% or
±5%
±5%
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MANUAL TESTING
With the simulator trimmed at the preset conditions for level flight, continue to fly
with wings level and at constant altitude for a period of at least 3 minutes. Do not
alter the throttle settings, unless such alterations are evident from the flight test
data recordings. Record either EPR, N1 or Engine Torque and also Fuel Flow
and after the test is complete compare them with the aeroplane data. It may
assist the pilot if an instructor station maintenance page is displayed which gives
values of, for example, engine thrusts.
EXAMPLE
There are two sets of results shown in Figure 1d3-1, the first at time=0 seconds
and then the second 300 seconds later. This duration was formerly specified in
the ICAO Manual 2nd Edition, but is now reduced to 3 minutes (180 seconds) as
a minimum.
1D-9
Figure 1d3-1
Example of Simulator Test Results for Cruise Performance
1D-10
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
1d(4) - IDLE DESCENT
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE THAT THE AEROPLANE
PERFORMANCE DURING A NORMAL DESCENT
DEMONSTRATION
Perform a normal descent using idle engine power in
the clean configuration at a mid altitude over an
interval of at least 300 m (1000 ft).
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FLIGHT CONDITION
CLEAN
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
PRESSURE ALTITUDE
PITCH ANGLE
ANGLE OF ATTACK
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
RATE OF CLIMB
FLIGHT PATH ANGLE
SPEEDBRAKE POSITION
SPOILER ANGLES
EVALUATION NOTES
achieved rate of descent corresponds well with the
aeroplane during a descent with engines idle. The test
should be recorded over an altitude interval of at least
300m (1000ft). If the aircraft validation data were
gathered using an engine variant that is not present
on the simulator, a second test should be run using
engine thrusts rather than pilot controls as the driving
input to show that the simulator gives the same
1D-11
response as the aeroplane under similar conditions.
TOLERANCES
AIRSPEED
RATE OF DESCENT
±3 Kts
±5% or ±1.0 m/Sec (200
Ft/Min)
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MANUAL TESTING
With the simulator trimmed at the preset conditions for the descent, continue to
fly in a stabilized descent at the prescribed airspeed for an altitude interval of at
least 300m (1000ft). Note the rate of descent for comparison with the aeroplane
data.
EXAMPLE
The way in which a test is run can have a significant effect on the plotted results,
as Figure 1d4-1 illustrates. The test has clearly been run using automatic drivers,
probably attempting to maintain airspeed using the pitch controller. The result as
shown does not indicate any particular problem with the simulation itself, but the
automatic driver gains have been set too high, or are otherwise incorrectly
programmed such that the result is unlikely to be acceptable by the authorities.
1D-12
Figure 1d4-1
Example of Simulator Test Results for Idle Descent
1D-13
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TITLE
1d(5) - EMERGENCY DESCENT
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OBJECTIVE
TO DEMONSTRATE THAT THE AEROPLANE
PERFORMANCE DURING AN EMERGENCY
DESCENT CONFORMS TO THE AEROPLANE.
DEMONSTRATION
Perform an emergency descent at mid altitude over an
interval of at least 900 m (3000 ft). Use idle power,
speedbrakes extended near Vmo speed, or according
to emergency descent procedures.
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FLIGHT CONDITION
AS PER AFM
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
PITCH ANGLE
ANGLE OF ATTACK
BANK ANGLE
ELEVATOR ANGLE
STABILISER ANGLE
RATE OF CLIMB
FLIGHT PATH ANGLE
SPEEDBRAKE POSITION
SPOILER ANGLES
EVALUATION NOTES
achieved rate of descent corresponds well with the
aeroplane during an emergency descent. The test
should be recorded over an altitude interval of at least
900m (3000ft). If the aircraft validation data were
gathered using an engine variant that is not present
on the simulator, a second test should be run using
engine thrusts rather than pilot controls as the driving
1D-14
input to show that the simulator gives the same
response as the aeroplane under similar conditions.
TOLERANCES
AIRSPEED
RATE OF DESCENT
±5 Kts
±5% or ±1.5 m/Sec (300
Ft/Min)
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
With the simulator stabilized at mid-altitude and at or near Vmo for the descent,
continue to fly in a stabilized descent at constant airspeed for an altitude interval
of at least 900m (3000ft). Note the rate of descent for comparison with the
aeroplane data, and perform a brief check of spoiler blowdown.
EXAMPLE
Running each QTG test manually has always been important to the regulators,
and Figure 1d5-1 is an example of a manually run emergency descent. The
duration is considerably longer than required by the ICAO Manual, but controlling
the simulated aeroplane for long periods should not be problematic for relatively
steady state tests such as this. The initial deviation in rate of descent may be
because the trimmed state is slightly incorrect, but it should be borne in mind that
the value yielded during a snapshot check of descent rate - especially at high
altitude/mach number combinations - will not be the same as when the test is run
as a time history
1D-15
Figure 1d5-1
Example of Simulator Test Results for Emergency Descent (Manual)
1D-16
SECTION 1e
STOPPING
1e(1)
Deceleration Time and Distance, Manual Wheel
Brakes, Dry Runway, No Reverse Thrust
1e(2)
Deceleration Time and Distance, Reverse Thrust, No
Wheel Brakes, Dry Runway
1e(3)
Stopping Distance, Wheel Brakes, Wet Runway
1e(4)
Stopping Distance, Wheel Brakes, Icy Runway
1E-1
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TITLE
1e(1) - DECELERATION TIME & DISTANCE,
MANUAL WHEEL BRAKES, DRY RUNWAY, NO
REVERSE THRUST
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OBJECTIVE
LANDING PERFORMANCE USING MANUAL
WHEEL BRAKES ONLY ON A DRY RUNWAY
DEMONSTRATION
Complete a normal landing on a dry runway, then
apply wheel brakes only, using maximum braking
pressure until reaching a full stop. No other
deceleration devices should be used.
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FLIGHT CONDITION
a) LANDING (MEDIUM WEIGHT)
b) LANDING (NEAR MAXIMUM LANDING WEIGHT)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
SPEEDBRAKE POSITION
SPOILER ANGLES
BRAKE PRESSURES
PITCH ANGLE
SPEEDBRAKE HANDLE POSITION
SPOILER ANGLES
EVALUATION NOTES
The test will begin with the simulated aeroplane set
up on the runway at the prescribed speed and
runway reference heading, usually with speed and
position frozen. It is not necessary, and may even
confuse the results, for a complete landing or
rejected takeoff manoeuvre to be executed.
1E-2
Aeroplane manufacturer's engineering data may be
used for the medium weight condition. Ground
speed, if available, should be used in preference to
airspeed.
During this test maximum brake effort should be
used continuously. Brake system pressure should,
however, be recorded. Time and distance data
should be recorded for at least 80% of the total time
from touchdown to a full stop. However, occasionally
during the flight test the pilot may have partially
released the brakes prior to coming to a full stop and
this may cause difficulties trying to fully repeat the
pilot actions during the simulator test.
TOLERANCES
TIME
DISTANCE
±5%
±61 m (200 Ft) or ±10%,
whichever is the smaller, for
distances up to 1220 m (4000
Ft).
±5% for distances greater than
1220 m (4000 Ft) .
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
This test is fairly simple to execute as it merely involves setting the simulated
aeroplane up on the runway at the prescribed configuration, speed and runway
reference heading, allowing it to stabilize with speed and position frozen, and
then releasing both freezes with the brakes fully applied. Since the runway is
dry and the braking symmetrical, no significant steering inputs should be
necessary.
Manual braking (not auto) is to be used for this test, unless the data specify
otherwise. Ideally, ground spoilers will not be used, but this will be dependent
on the aeroplane data.
EXAMPLE
In Figure 1e1-1 the result is just out of tolerance with the airspeed taking too
little time to reduce relative to the aircraft. This is almost certainly because the
heading deviation has caused the aircraft to yaw which would have the effect
of adding a slight extra stopping force. However, the other item of note, which
1E-3
may ultimately have a greater effect on the result, is the discontinuity in
airspeed after approximately 10.5 seconds. This may have been due to a wind
gust (which is why ground speed should also be shown), or the pilot may have
reduced the braking effort slightly. Either way a note should be added to the
QTG to explain the inconsistency.
Figure 1e1-1
Example of Simulator Test Results for Stopping Time & Distance, Dry Runway
1E-4
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TITLE
1e(2) - DECELERATION TIME & DISTANCE,
REVERSE THRUST, NO WHEEL BRAKES, DRY
RUNWAY
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OBJECTIVE
LANDING PERFORMANCE USING REVERSE
THRUST ONLY ON A DRY RUNWAY CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Complete a normal landing on a dry runway, then
apply maximum reverse thrust until reaching the full
thrust reverser aeroplane minimum operating speed.
Other than ground spoilers, no other deceleration
devices should be used.
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FLIGHT CONDITION
a) LANDING (MEDIUM WEIGHT)
b) LANDING (NEAR MAXIMUM LANDING WEIGHT)
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RECORDED
PARAMETERS
SPEEDBRAKE POSITION
SPOILER ANGLES
BRAKE PRESSURES
PITCH ANGLE
SPOILER ANGLES
EVALUATION NOTES
The test will begin with the simulated aeroplane set
up on the runway at the prescribed speed and
runway reference heading, usually with speed and
position frozen. It is not necessary, and may even
confuse the results, for a complete landing or
rejected takeoff manoeuvre to be executed.
1E-5
Aeroplane manufacturer's engineering data may be
used for the medium weight condition. Ground
speed, if available, should be used in preference to
airspeed.
During this test the engine instruments should all
function normally, but when run automatically the
simulator power levers will not physically move from
the idle position into reverse thrust as with most
aeroplanes it is physically impossible (or extremely
impractical) to do this. This does not affect the
validity of the results, providing the appropriate
aeroplane engine parameters (EPR, N1, Thrust, etc.)
are closely matched.
Time and distance data should be recorded for at
least 80% of the total time from touchdown to full
thrust reverser minimum operating speed.
TOLERANCES
TIME
DISTANCE
±5%
The smaller of ±10% or ±61 m
(200 Ft)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
This test is not as simple to execute as Test 1e(1) because whilst it also
involves setting the simulated aeroplane up on the runway at the prescribed
configuration, speed and runway reference heading, following the flight test
pilot's actions for the correct movement of the power levers to achieve the flight
test values of EPR, N1 or Thrust can be quite difficult. Thus it may require
some very well-worded manual test procedures to obtain a good match. Once
again though, since the runway is dry and the reverse thrust (usually)
symmetrical, no significant steering inputs should be necessary. Ideally,
spoilers will not be used so that just the effects of reverse thrust can be
ascertained, but this will be dependent on the aeroplane data.
EXAMPLE
Two (partial) sets of contrasting results are displayed in Figures 1e2-1 and
1e2-2 below. The first produces an exact match for both distance and ground
speed, whereas the second gives a result which is slightly out of tolerance for
both parameters. The differences were eventually resolved, but the ‘first
1E-6
passes’ at each test shown here reveal how much easier it usually is to match
engineering simulation data, which was the source for the first of the two sets
of results.
Figure 1e2-1
Example of Simulator Test Results for Reverse Thrust Stopping Time & Distance (1)
1E-7
1E-8
Figure 1e2-2
Example of Simulator Test Results for Reverse Thrust Stopping Time &
Distance (2)
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TITLE
1e(3) - STOPPING DISTANCE, WHEEL BRAKES,
WET RUNWAY
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OBJECTIVE
LANDING PERFORMANCE USING WHEEL
BRAKES ONLY ON A WET RUNWAY CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Complete a normal landing on a wet runway, then
apply wheel brakes only, using maximum brake
pressure until reaching a full stop. Other than ground
spoilers, no other deceleration devices should be
used.
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FLIGHT CONDITION
LANDING
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
SPEEDBRAKE POSITION
SPOILER ANGLES
BRAKE PRESSURES
PITCH ANGLE
SPOILER ANGLES
EVALUATION NOTES
See the notes for test 1e(1). Either flight test or
manufacturer’s performance manual data must be
used where available. An acceptable alternative is to
use engineering data based on dry runway flight-test
stopping distance and the effects of wet runway
braking coefficient. Clearly, there should be an
increase in the time and distance to stop over that
achieved during the dry runway test, although
1E-9
because of training considerations it is the distance
rather than the time which is in question here. The
test should clearly show that the simulated aeroplane
should still be able to stop within the confines of the
runway.
TOLERANCES
DISTANCE
±10% or ±61 m (200 Ft)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
See the notes for test 1e(1). The test is typically run in a very similar, if not
identical fashion.
EXAMPLE
The diagram below, Figure 1e3-1, contains a partial set of plots and partial
printed pass/fail data for a wet runway stopping distance test. The asterisk by
the check points on the printout indicate that the test has failed, however a
close look at the plotted data shows that these parameters - including both
ground sped and ground distance - are virtually exact overlays of the
aeroplane data. The obvious conclusion to draw is that the ‘aeroplane’ values
against which the simulator is being checked have been incorrectly specified
by the simulator engineer in the test information. This situation is surprisingly
common, but also very easy to rectify.
1E-10
Figure 1e3-1
Example of Simulator Test Results for Stopping Time & Distance, Wet Runway
1E-11
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TITLE
1e(4) - STOPPING DISTANCE, WHEEL BRAKES,
ICY RUNWAY
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OBJECTIVE
LANDING PERFORMANCE USING WHEEL
BRAKES ONLY ON AN ICY RUNWAY CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Complete a normal landing on an icy runway, then
apply wheel brakes only, using maximum brake
pressure until reaching a full stop. Other than ground
spoilers, no other deceleration devices should be
used.
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
LANDING
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
SPEEDBRAKE POSITION
SPOILER ANGLES
BRAKE PRESSURES
PITCH ANGLE
SPOILER ANGLES
EVALUATION NOTES
See the notes for test 1e(1). Either flight-test or
manufacturer’s performance manual data should be
used, though flight test data are not often available
for this test. An acceptable alternative is to use
engineering data based on dry runway flight-test
stopping distance and the effects of icy runway
braking coefficient. Clearly, there should be an
increase in the time and distance to stop over that
1E-12
achieved during the wet runway test, although
because of training considerations it is the distance
rather than the time which is in question here. For
this runway condition the simulated aeroplane may
not be able to stop within the confines of the runway.
TOLERANCES
DISTANCE
±10% or ±61 m (200 Ft)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
See the notes for test 1e(1). The test is typically run in a very similar, if not
identical fashion. For icy runways, it may be expected that the simulated
aeroplane will overshoot the end of the runway, but this may be dependent on
the selected configuration.
EXAMPLE
The results shown in Figure 1e4-1 below are for a footprint test, indicating that
no aeroplane manufacturer’s data was available for this condition. The most
up-to-date packages will contain engineering simulator data, but that was not
the case for this aircraft type. The only potentially confusing item on the plots is
the reference to ‘Flight Test Data’ in the bottom left corner, which may lead an
evaluator to the erroneous conclusion that the simulator matches the
aeroplane so perfectly that the plots are indistinguishable.
1E-13
Figure 1e4-1
Example of Simulator Test Results for Stopping Time & Distance with Icy Runway
1E-14
SECTION 1f
ENGINES
1f(1)
Acceleration
1f(2)
Deceleration
1F-1
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TITLE
1f(1) - ENGINE ACCELERATION
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATED
ENGINE ACCELERATION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Starting from a stabilised condition, rapidly advance
the throttles from idle power to the go-around thrust
setting.
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FLIGHT CONDITION
APPROACH OR LANDING
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
For each engine:
POWER LEVER ANGLE (or equivalent)
NET THRUST
EGT
FUEL FLOW
ENGINE PRESSURE RATIO (EPR) or N1 & N2
PRESSURE ALTITUDE
MACH NUMBER
AMBIENT TEMPERATURE
EVALUATION NOTES
The items of importance for this test are the prime
(key) engines parameters presented in the data, and
in particular the time taken to achieve the values.
The definitions of Ti and Tt are given in the ICAO
Manual but are repeated below for clarity. The actual
aeroplane response is not the issue for this test, but
obviously the test conditions should be accurately
represented so that a fair comparison can be made.
TOLERANCES
TIME (Ti)
±10% or ±0.25 Sec
(Where Ti is the total time from initial throttle
movement until a 10% response of a critical engine
parameter)
TIME (Tt)
±10%
1F-2
(Where Tt is the total time from initial throttle
movement to 90% of go-around power)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
Typical procedures call for the simulator and engines to be set for approach in
a trimmed condition. The throttles are then rapidly advanced to Go-Around
power. Record critical engine performance parameters (Power Lever Angle,
Net Thrust per engine, N1, N2, EGT, Fuel Flow and EPR) and compare versus
aeroplane data. Following flight test aeroplane parameters such as airspeed
and altitude is desirable if they are available, but small deviations in these
values should not adversely affect the results.
EXAMPLE
A good match has been achieved in the simulator when compared with the
manufacturer’s proof of match (see Figures 1f1-1 and 1f1-2), but the plot scale
chosen does not properly allow an evaluator to determine whether or not the
test passes or fails. This is one example where use of the manufacturer’s
original plot scale can be bettered when designing the simulator QTG.
1F-3
Figure 1f1-1
Example of Aeroplane Manufacturer's Proof of Match Data
(Engine Acceleration)
Figure 1f1-2
Example of Simulator Test Results for Engine Acceleration
1F-4
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
1f(2) - ENGINE DECELERATION
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATED
ENGINE DECELERATION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Starting from a stable on-ground condition with the
engines at maximum takeoff power and stabilised,
rapidly retard the throttles to the idle power position.
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
GROUND
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
For each engine:
POWER LEVER ANGLE (or equivalent)
NET THRUST
EGT
FUEL FLOW
PRESSURE ALTITUDE
ATMOSPHERIC PRESSURE
MACH NUMBER
AMBIENT TEMPERATURE
EVALUATION NOTES
The items of importance for this test are the prime
(key) engines parameters presented in the data, and
in particular the time taken to achieve the values,
usually idle thrust. The definitions of Ti and Tt are
given in the ICAO Manual but are repeated below for
clarity. The actual aeroplane response is not the
issue for this test, especially as it is to be performed
on ground, but obviously the test conditions should
be accurately represented so that a fair comparison
can be made. Note that the final time to achieve the
same value of (idle) thrust is not the foremost issue,
as it is recognised that the simulator and aeroplane
times need not be the same if the thrust is way below
1F-5
a level of significance.
TOLERANCES
TIME (Ti)
±10% or ±0.25 Sec
(Where Ti is the total time from initial throttle
movement until a 10% response of a critical engine
parameter)
TIME (Tt)
±10%
(Where Tt is the total time from initial throttle
movement to 90% decay of maximum takeoff power)
)))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
Typical procedures call for the simulator and engines to be set on ground static
in a stabilised condition with the throttles set for Takeoff power. The throttle
levers should then be rapidly retarded to the idle position whilst the critical
engine performance parameters (Power Lever Angle, Net Thrust per engine,
N1, N2, EGT, Fuel Flow and EPR) are recorded and compared versus
aeroplane data. The test is usually best performed with the parking brake on,
but note should be taken of the method used during the acquisition of the flight
test data.
EXAMPLE
The example in Figure 1f2-1 shows that, while it is possible to get a match
which looks good when plotted, the nature of the tolerances for this test is such
that the test still technically fails in the acceleration phase though not in the
deceleration phase. Other parameters, if available in the data, should also be
shown, and the power lever angle (or equivalent) should be driven exactly as in
the data.
1F-6
Figure 1f2-1
Example of Simulator Test Results for Engine Acceleration & Deceleration (Combined)
1F-7
1F-8
SECTION 2
HANDLING QUALITIES
2a
2b
2c
LONGITUDINAL
2d
LATERAL DIRECTIONAL
2e
LANDINGS
2f
GROUND EFFECT
2g
WINDSHEAR
2h
FLIGHT AND MANOEUVRE ENVELOPE
PROTECTION FUNCTIONS
2-1
2.0
HANDLING QUALITIES - GENERAL
The purpose of this section of the QTG is to provide the evaluator of the
simulator with adequate, objective evidence that the handling qualities of
the simulator correspond within reasonable limits to those of the aeroplane
being simulated. The validation tests which assist in this function have
been chosen to ensure that the stability and control characteristics of the
simulator are satisfactory relative to the actual aeroplane throughout its
speed, altitude, weight and centre of gravity envelope. Whilst much of the
training of jet transport flight crew is carried out at low altitude and in the
vicinity of the airfield, this does not remove the necessity to prove the
capability of the simulator as an effective training tool in cruise conditions
as well. It is important to pilot training and certification (licensing) that the
simulator handling qualities closely match those of the respective
aeroplane. These essential characteristics of simulators, therefore, must
be demonstrated and must be repeatable. Repeatability must not, however,
be so designed into the testing system that the objective to demonstrate
proper handling qualities is diminished.
As with the Performance Tests, the most effective way of carrying out these
tests to give the desired accuracy is by using an automatic test system.
However, confirming the automatic test result with a selection of tests
which have been flown manually by a suitably qualified pilot is perhaps of
even greater importance when assessing handling qualities than it is when
evaluating the simulated aeroplane performance.
2.1
CONTROL CHECKS
The static and dynamic force "feel" characteristics of an aeroplane control
system form an important feed-back to the pilot when flying the aeroplane.
In the following sections, tests are specified to evaluate the static and
dynamic "feel" characteristics and position calibration of the simulator
control systems compared to the aeroplane control systems.
Due to the nature of control systems, and the very limited possibilities to
measure actual pilot control force without affecting the control system
characteristics, the following considerations should be taken into account
when comparing measured control system characteristics of an aeroplane
and a simulator.
For any control system, especially if the applied control forces are
important, the exact system configuration during the test must be
documented (e.g. yaw damper on/off, control wheel steering on/off,
hydraulics on/off, feel pressure). Furthermore, it must be verified that the
2-2
controls are free to move, and not obstructed by any crew member or
equipment (e.g. co-pilot's knee obstructing wheel movement, manuals on
co-pilot's seat obstructing column movement).
If special equipment is attached to the flight controls to measure pilot
applied force and control deflection, the inertia and unbalance of the
equipment may affect the measured characteristics. Direct comparison
with aeroplane data is only possible if the same equipment and equipment
configuration is used for both the aeroplane and simulator measurements,
under the same conditions.
A full specification of which equipment was used for the aeroplane
measurements, and a drawing showing how the equipment was attached
to the flight controls, is necessary in order to be able to reproduce the
results in the simulator.
In general, flight test recorded control forces and control deflections are not
measured at the point of pilot force application. This implies that if data
from an aeroplane installed data acquisition systems is used, the exact
location in the control system where the signal is measured, and the
applied conversions to obtain equivalent pilot control force and control
deflection must be specified (e.g. measured in the control cables,
measured at the aft quadrant, etc).
Due to inertial effects and the filtering of strain gauge signals, the
measured control force will inevitably be different from the theoretical pilot
control force, for example if the pilot releases the control, the control force
will by definition instantaneously become zero, but recorded data will
always show a less abrupt change in force due to inertial and damping of
the controls.
Also, some contributions to the pilot control force, such as friction and
unbalance in the pilot controls, may not be reflected in the forces as
measured by the data acquisition system. Known differences must be
documented for each test by the data supplier, in order to be able to judge
the acceptability of observed differences.
When using data from a prototype aeroplane as reference, it should be
recognised that prototype aeroplanes often do not exactly represent
production aeroplane standards, due to installed instrumentation or
development process. For example, controller-to-control-surface gearing,
the position of stops, friction levels, inertia of controls and characteristics
of feel springs may be slightly different from production aeroplanes.
2-3
Since a training simulator must represent production aeroplane
characteristics rather than prototype aeroplane characteristics, data
measured at the pilot controls (using Control Force Measuring equipment)
of a production aeroplane must be used to tune the characteristics of the
control loading system, by comparison with equivalent recordings of the
simulator control characteristics. The ICAO Manual should be consulted for
further explanation.
When an evaluator is examining a simulator ne/she has not seen before,
the instructor maintenance pages should be used to check travel limits and
rudder neutral with trim at zero. If a force balance gauge is available,
check a couple of forces at the pilots point of application. There is no check
of mass unbalance in the regulatory requirements, so the stick forces to
rotate should be briefly checked with and without motion.
Finally, note that force versus position testing in several of the tests
contained in sections 2a (Static Controls Checks) and 2b (Dynamic
Controls Checks) is not required if an actual aeroplane hardware controller
is employed in the simulator. This typically applies to certain Computer
Controlled Aircraft.
2-4
SECTION 2a
2a(1)
Pitch Controller Position vs. Force and Surface
2a(2)
Roll Controller Position vs. Force and Surface
2a(3)
Rudder Pedal Position vs. Force and Surface
2a(4)
Nosewheel Steering Controller Force and Position
Calibration
2a(5)
Rudder Pedal Steering Calibration
2a(6)
Pitch Trim Indicator vs. Surface Position Calibration
2a(7)
Pitch Trim Rate
2a(8)
Alignment of Cockpit Throttle Lever vs. Selected
Engine Parameter
2a(9)
Brake Pedal Position vs. Force and Brake System
Pressure Calibration
2A-1
2A.0 STATIC CONTROL CHECKS
The purpose of the static control tests is to verify the simulated
quasi-stationary control system force characteristics, and the relation
between pilot control position and surface position.
In order to exclude dynamic effects, the tests must be performed using very
small control deflection rates. Also, the control deflection rate should be as
constant as possible.
The tests are performed by very slowly moving the pilot control over its full
range: from neutral to the stop, then to the opposite stop, then back to the
neutral position (full sweep). The exceptions here are the pitch trim tests
and the throttle lever test, which can be accomplished by spot checking
rates and positions as appropriate, and therefore do not need a full sweep
in both directions.
Except as noted for the pitch trim and throttle lever, tolerances for these
tests are on pilot control force and surface position (or brake system
pressure in test 2a(9)). Compliance should be shown by comparison of
cross-plots of control force versus pilot control position and surface position
versus pilot control position rather than time histories, except for the pitch
trim rate test.
The controller position versus force shall be measured at the pilot control.
An alternative method would be to instrument the simulator in an equivalent
manner to the flight test aeroplane. The force and position data from this
instrumentation can be directly recorded and matched to the aeroplane
data. Prior to an initial evaluation, the regulatory authorities require that a
physical calibration is performed using a control force measuring (CFM)
system on the primary controls (the so-called ‘Fokker’ tests, though some
simulator manufacturers use methods other than that which was devised
by Fokker themselves). These tests are usually time-consuming and
labourious to perform, but are important in that they are designed to give
confidence that the computed values - as displayed on the IOS for example
- are a good match for the true values of force and position so that the
computed values may be used for recurrent and any subsequent tests. In
general, the regulatory authorities do not ask for the calibration tests using
CFM equipment to be re-run on a recurrent basis, but they are entitled to
request them, and do so occasionally if there appears to be good reason.
Note that for some aeroplanes with reversible flight controls, the tests will
need to be run at a suitable airspeed condition.
2A-2
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TITLE
2a(1) - PITCH CONTROLLER POSITION vs. FORCE
AND SURFACE POSITION CALIBRATION
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR PITCH
CONTROLLER POSITION vs. PITCH CONTROLLER
FORCE AND THE PITCH CONTROLLER POSITION
vs. ELEVATOR POSITION CHARACTERISTICS
DEMONSTRATION
Starting from the neutral position, move the pitch
controller at a very slow rate over its full range to the
aft or forward limit, then back through neutral to the
opposite limit, then back to neutral again.
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
PITCH CONTROLLER FORCE
ELEVATOR ANGLE
EVALUATION NOTES
After confirming that the pitch controller is in the
neutral position, the test is run by slowly driving the
controller to either the forward or the aft stop, then
slowly driving it back through neutral to the other stop,
then finally driving it back again to the neutral position.
The results should be overplotted with those obtained
on the aeroplane to enable an effective comparison to
be made. For initial evaluations, or at the request of
the authorities, repeat the test with a control force
measurement (cfm) system fitted. It is not necessary
to run this test provided the aeroplane cockpit
controller unit has been employed in the simulator and
it has not been modified from its status in the
aeroplane. The longitudinal control system
characteristics for all aeroplane types are further
validated by the tests included in other sections, such
2A-3
as Longitudinal Static Stability (section 2c(7)) and Stall
Characteristics (section 2c(8)). Note that it may be
necessary to apply the same control feel pressure to
the simulation as was in effect for the aeroplane data.
TOLERANCES
BREAKOUT FORCE
FORCE
ELEVATOR ANGLE
±0.9 daN (2 Lbs)
±2.2 daN (5 Lbs) or ±10%
±2o
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MANUAL TESTING
Slowly move the pitch controller such that approximately 100 seconds are
required to achieve a full sweep. A full sweep is defined as movement of the
controller from neutral to the stop (forward or aft), then to the opposite stop, then
back to the neutral position. Before performing the test, verify that the pitch
controller movement is not obstructed by any crew member or equipment (e.g.
manuals on co-pilot's seat obstructing control column movement) as this can
cause severe deformations in the measured control characteristics, especially
near the stops. Note that the cockpit controller position versus force
measurement is not required if a self contained aeroplane controller (i.e.
aeroplane part(s)) which has integrated force and damping systems is used in the
simulator. Be aware that on some aeroplanes feel forces vary with stabiliser trim
position.
EXAMPLE
Figure 2a1-1a shows a good result for force versus position, however the
‘aeroplane data’ used for the overplot is almost certainly not taken directly from
an aeroplane using CFM equipment - the data trends are too ‘smooth’ and the
end-stops do not represent what actually occurs when using CFM equipment in
the aeroplane. This does not render the test results invalid, as there may be good
reasons why such data has been used (for example, this is what the aeroplane
manufacturer provided in their proof-of-match document), though this should be
properly explained in the QTG.
2A-4
Figure 2a1-1a
Example of Simulator Test Results for Pitch Controller Force
versus Position Calibration
2A-5
Figure 2a1-1b
Example of Simulator Test Results for Elevator versus Pitch Controller Position
Calibration
2A-6
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TITLE
2a(2) - ROLL CONTROLLER POSITION vs. FORCE
AND SURFACE POSITION CALIBRATION
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR ROLL
CONTROLLER POSITION vs ROLL CONTROLLER
FORCE AND ROLL CONTROLLER POSITION vs
AILERON
AND
SPOILER
ANGLE
CHARACTERISTICS CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Starting from the neutral position, move the roll
controller at a very slow rate over its full range to the
left or right limit, then back through neutral to the
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
ROLL CONTROLLER FORCE
AILERON AND SPOILER ANGLES
EVALUATION NOTES
After confirming that the roll controller is in the neutral
position, the test is run by slowly driving the controller
to either the left or the right stop, then slowly driving it
back through neutral to the other stop, then finally
driving it back again to the neutral position. The results
should be overplotted with those obtained on the
aeroplane to enable an effective comparison to be
made. For initial evaluations, or at the request of the
authorities, repeat the test with a control force
measurement (cfm) system fitted. It is not necessary
to run this test provided the aeroplane cockpit
controller unit has been employed in the simulator and
it has not been modified from its status in the
aeroplane. The lateral control system characteristics
for all aeroplane types are further validated by the
2A-7
tests included in other sections, such as Engine
Inoperative Trims (section 2d(5)) and Steady Sideslip
(section 2d(8)). Note that it may be necessary to apply
the same control feel pressure to the simulation as
was in effect for the aeroplane data.
TOLERANCES
BREAKOUT FORCE
FORCE
AILERON ANGLE
SPOILER ANGLES
±0.9 daN (2 Lbs)
±1.3 daN (3 Lbs) or ±10%
±2o
±3o
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MANUAL TESTING
Slowly move the roll controller such that approximately 100 seconds are required
to achieve a full sweep. A full sweep is defined as movement of the controller
from neutral to the stop (right or left), then to the opposite stop, then back to the
neutral position. Before performing the test, verify that the roll controller
movement is not obstructed by any crew member or equipment (e.g. manuals on
co-pilot's seat obstructing control wheel movement). Note that the cockpit
controller position versus force measurement is not required if a self contained
aeroplane controller (i.e. aeroplane part(s)) which has integrated force and
damping systems is used in the simulator.
EXAMPLE
The lack of an aeroplane data overplot for the example in Figure 2a2-1a shows
the general format and shape of a typical roll controller force (in this case wheel
force) calibration test for a conventionally-controlled (i.e. non-computer
controlled) aeroplane. Note that the breakout force, at a little under 6 lbs, is a
fairly large percentage of the maximum effort required to apply full wheel. This
plot was run on an older simulator which did not use overplots for the control
force calibration tests, relying instead on transparency copies of the aeroplane
data to facilitate the comparison.
2A-8
Figure 2a2-1a
Example of Simulator Test Results for Roll Controller Force versus Position Calibration
Figure 2a2-1b below shows the roll controller (wheel) versus surface position
plots for the same simulator, illustrating both the small breakout value of the
aileron surface and also the amount of wheel required before spoiler movement
is initiated.
2A-9
Figure 2a2-1b
2A-10 Example of Simulator Test Results for Aileron & Spoiler versus Roll Controller
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TITLE
2a(3) - RUDDER PEDAL POSITION vs. FORCE AND
SURFACE POSITION CALIBRATION
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OBJECTIVE
RUDDER PEDAL STATIC CHARACTERISTICS
CONFORM TO THE AEROPLANE
DEMONSTRATION
Starting from the neutral position, move the rudder
pedals at a very slow rate over their full range to the
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
RUDDER PEDAL FORCE
RUDDER ANGLE
EVALUATION NOTES
After confirming that the rudder pedals are in the
neutral position, the test is run by slowly driving the
pedals to either the left or the right stop, then slowly
driving them back through neutral to the other stop,
then finally driving them back again to the neutral
position. The results should be overplotted with those
obtained on the aeroplane to enable an effective
comparison to be made. For initial evaluations, or at
the request of the authorities, repeat the test with a
control force measurement (cfm) system fitted. The
directional control system characteristics are further
validated by the tests included in Sections 1b, 2d and
2e.
TOLERANCES
BREAKOUT FORCE
FORCE
RUDDER ANGLE
±2.2 daN (5 Lbs)
±2.2 daN (5 Lbs) or ±10 %
±2o
2A-11
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MANUAL TESTING
Slowly move the rudder pedals such that approximately 100 seconds are
required to achieve a full sweep. A full sweep is defined as movement of the
controller from neutral to the stop, usually the right stop, then to the opposite
stop, then back to the neutral position. Before performing the test, verify that the
rudder pedal movement is not obstructed by any crew member or equipment.
EXAMPLE
What at first looks like a major problem with the simulated pedal force in Figure
2a3-1 (the simulator is the full line, the aeroplane data is the dotted line) turned
out to be nothing more than a problem related to the rate at which the automatic
test system was driving the rudder pedals (see the upper plot). This result was
taken from a simulator that was in service at the time, and some engineering
effort was obviously required to find and eliminate the cause of the anomaly
during the autotest, but running the test manually (as illustrated in the lower plot)
revealed no actual problem with the pedals themselves.
2A-12
Figure 2a3-1
Example of Simulator Test Results for Rudder Pedal Force versus
2A-13
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TITLE
2a(4) - NOSEWHEEL STEERING CONTROLLER
FORCE AND POSITION CALIBRATION
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OBJECTIVE
NOSEWHEEL STEERING CONTROL STATIC
CHARACTERISTICS
CONFORM TO THE
AEROPLANE
DEMONSTRATION
Starting from the neutral position, move the nosewheel
steering controller at a very slow rate over its full
range to the left or right limit, then back through
neutral to the opposite limit, then back to neutral
again.
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
NOSEWHEEL STEERING CONTROLLER FORCE
NOSEWHEEL STEERING CONTROLLER POSITION
NOSEWHEEL ANGLE
MAIN GEAR ANGLE (if applicable)
EVALUATION NOTES
After confirming that the tiller is in the neutral position,
the test is run by slowly driving the nosewheel steering
controller to either the left or the right stop, then slowly
driving it back through neutral to the other stop, then
finally driving it back again to the neutral position. The
results should be overplotted with those obtained on
the aeroplane to enable an effective comparison to be
made. The nosewheel steering system characteristics
are further validated by the tests included in Section
1a.
TOLERANCES
BREAKOUT FORCE ±0.9 daN (2 Lbs)
FORCE
±1.3 daN (3 Lbs) or ±10%
NOSEWHEEL ANGLE ±2o
2A-14
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MANUAL TESTING
Slowly move the tiller such that approximately 100 seconds are required to
achieve a full sweep. A full sweep is defined as movement of the controller from
neutral to the stop, either right or left, then to the opposite stop, then back to the
neutral position. Before performing the test, verify that the tiller movement is not
obstructed by any crew member or equipment.
EXAMPLE
Figure 2a4-1 again shows a typical plot for a nosewheel steering controller force
versus position test. As is also a typical feature of the roll controller calibration
test, the breakout is a very large percentage of the maximum force required for
full deflection.
2A-15
Figure 2a4-1
Example of Simulator Test Results for Nosewheel Steering Controller Force versus Position
2A-16
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TITLE
2a(5) - RUDDER PEDAL STEERING CALIBRATION
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OBJECTIVE
RUDDER PEDAL STEERING CHARACTERISTICS
CONFORM TO THE AEROPLANE
DEMONSTRATION
Starting from the neutral position, move the rudder
pedals at a very slow rate over their full range to the
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
NOSEWHEEL ANGLE
EVALUATION NOTES
The rudder pedal steering force characteristics are
usually the same as the rudder control forces. The
criteria of note here is that the nosewheel angle only
travels through the range that it should - much smaller
than when driven by the nosewheel controller - within
the prescribed tolerances.
TOLERANCES
NOSEWHEEL ANGLE ±2o
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MANUAL TESTING
The rudder pedal steering force characteristics are usually the same as the
rudder control forces. Therefore, the check may be executed as part of the rudder
calibration test. While not typically necessary, it may better represent the
aeroplane data if the simulated runway friction is set to a very low value while
running this test so as to better simulate the ‘greasy plate’ sometimes employed
by aircraft manufacturers when conducting this test.
2A-17
EXAMPLE
The rudder pedal steering check, which is often done in conjunction with the
rudder pedal position versus force calibration, is exemplified by Figure 2a5-1. The
small discontinuity in the plot has little overall significance in terms of the test
evaluation, but could probably be eliminated by running the test for a little longer
or by nudging the pedals back past the neutral position prior to actually ending
the test.
Figure 2a5-1
Example of Simulator Test Results for Rudder Pedal Steering Calibration
2A-18
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TITLE
2a(6) - PITCH TRIM INDICATOR vs. SURFACE
POSITION CALIBRATION
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OBJECTIVE
TRIM VISUAL COCKPIT INDICATIONS ARE
SATISFACTORILY CALIBRATED RELATIVE TO THE
COMPUTED VALUE AND CONFORM TO THE
AEROPLANE DESIGN DATA.
DEMONSTRATION
The pitch trim is manually commanded (using manual
switches or trimwheel as appropriate) to the nose-up
and nose-down limits. The stabiliser angle and
computed trim value are then checked.
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
INDICATED PITCH TRIM POSITION
COMPUTED TRIM POSITION
STABILISER ANGLE
EVALUATION NOTES
This test is not practical to run fully automated,
because the requirement is to check that the readout
of stabiliser position or angle as perceived by the pilot
in the cockpit corresponds closely with the value
computed and available either on an engineering
terminal or workstation or at the instructor's screen.
Hence the typical method is to set the simulator up at
the required condition with all integrators frozen whilst
the autotest system drives the stabiliser to specified
positions and waits for the person running the test to
confirm that the two values correspond. Several
points should be checked over the range, but special
attention should be paid to the values at either end.
The purpose of this test is to compare the simulator
trim indicator value against aeroplane design data or
2A-19
equivalent, so flight-test data are not required.
TOLERANCES
TRIM ANGLE
±0.5o
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MANUAL TESTING
See the 'EVALUATION NOTES' section. Because this test requires manual
confirmation that the computed and indicated values agree within the tolerance,
the manual test will typically only differ from the automatic one in that the
stabiliser trim switch is used to set the indicated stabiliser position instead of the
autotest system performing this function.
EXAMPLE
A plotted example would be of little benefit for this test, since it is essentially just
a check that the indicated position is correctly aligned with the computed value,
available on the IOS and/or on an engineering terminal. Clearly, the value
indicated to the pilot in the cockpit must be confirmed by a pilot seated in the
correct position.
2A-20
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TITLE
2a(7) - PITCH TRIM RATE
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OBJECTIVE
TRIM RATE CONFORMS TO THE AEROPLANE.
DEMONSTRATION
Command pitch trim (using manual switches or trim
wheel as appropriate) to the nose-up and nose-down
limits. The stabiliser angle and computed trim value
and trim rate (especially for the Go-Around case) is
then checked.
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FLIGHT CONDITION
i) GROUND
ii) APPROACH
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RECORDED
PARAMETERS
COMPUTED TRIM POSITION / INDICATED PITCH
TRIM POSITION
STABILISER ANGLE
TRIM RATE
TRIMMED SURFACE ANGLE RATE
PILOT PRIMARY TRIM SWITCH POSITION
AUTOPILOT TRIM SIGNAL (FOR GO-AROUND
CASE)
EVALUATION NOTES
This test should be run after the calibration test 2a(6)
has confirmed that the actual and indicated trim
positions are closely aligned. These tests can be run
fully automated, but should also be run manually on a
recurrent basis to check that the trim rate as perceived
by the pilot in the cockpit corresponds closely with the
computed value. The trim rate should first be checked
on ground static, then in a go-around configuration.
The on-ground test requires the trim rate to be
checked only with manual (primary) trim. The
Approach/Go-Around case requires the trim rate to be
checked either using manual trim (i.e. the trim switch),
2A-21
or using autopilot trim. The commanded trim rate for
each configuration should be compared with the
achieved surface angle rate and the aeroplane trim
rate.
TOLERANCES
TRIM RATE (o/Sec)
±10%
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MANUAL TESTING
This test requires firstly that the simulator be set on ground in a static condition,
then the manual trim switch(es) used to drive the stabiliser from near one
extreme to near the opposite extreme, noting the time taken to travel between the
two points. The trim rate can then be cross-checked against the plotted value of
stabiliser. The second case entails setting up an approach condition, then either
repeating the method used on ground or using autopilot trim for a go-around
manoeuvre. Clearly, if large excursions in pitch angle are to be avoided, the test
at the approach condition must be of short duration, just sufficient to determine
the trim rate, and this may be achieved using either the manual switches or the
autopilot.
EXAMPLE
A plotted example of this test is shown in Figure 2a7-1. Aside from the slightly
odd values on the time axis, the test shows very well the co-ordination between
pitch trim rate and position, and even directly plots the value for trim rate which
can easily be used to compare with the aeroplane data.
2A-22
Figure 2a7-1
Example of Simulator Test Results for Pitch Trim Rate Test
2A-23
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TITLE
2a(8) - ALIGNMENT OF COCKPIT THROTTLE
LEVER vs. SELECTED ENGINE PARAMETER
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OBJECTIVE
TO DEMONSTRATE THAT THE ALIGNMENT OF
THE THROTTLE LEVERS IN THE SIMULATOR WITH
REFERENCE TO THE RESULTANT KEY ENGINE
PARAMETERS CONFORMS TO THE AEROPLANE.
DEMONSTRATION
The throttle levers are moved to several specified
positions and once the engines have stabilised the
engine key parameters (EPR, N1, torque, as
appropriate) are recorded.
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
For each engine, as appropriate to the engine type:
THROTTLE LEVER ANGLE
TORQUE (Turboprop only)
EVALUATION NOTES
The typical method is to set the simulator up at the
required on ground static condition, with the engines
idle and parking brake on. The autotest system then
slowly drives the throttle levers over the entire range
of movement. At specified values of engine parameter
it may be possible to automatically confirm that the
simulator power lever angles correspond with those of
the aeroplane, but the engines should have been
allowed to stabilize at each position first. Several
points should be checked over the range, and special
attention should be paid to the values at each end of
travel. The throttle levers themselves may be backdriven, and so it may be that a slight delay can be
expected between the engines apparently achieving
their stabilized condition and the cessation of throttle
lever movement, but this is unlikely to materially affect
2A-24
the results.
Note that this test can be conducted either way, set
the throttle levers and read N1/EPR/Torque, or set
N1/EPR/Torque and read off throttle lever angle. This
is the reason for the ‘or’ in the tolerances, it is not
acceptable use both the TLA tolerance and
N1/EPR/Torque tolerance in combination.
Note that for propeller powered aeroplanes, if an
additional lever, usually referred to as the propeller
lever, is present, it must also be checked. The
tolerance for throttle lever angle applies for the
simulator against aeroplane data and also to the
throttle levers relative to each other - see the ICAO
Manual for additional information. If more than one
engine variant or model is being simulated on the
same simulator, this test should be run for each
variant.
TOLERANCES
THROTTLE LEVER ANGLE
±5.0o or
N1
±3% or
EPR
±0.03 or
TORQUE
±3%
PROPELLER LEVER TRAVEL
±2 cm (0.8 in)
(Used for propellor-driven aeroplanes where such
levers do not have angular travel)
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MANUAL TESTING
The method for this test typically requires the person conducting the test to first
set the throttles to idle, then to each of several specified throttle lever angles, and
then wait for the engines to stabilise before noting the value of each of the prime
engine parameters at that point. This procedure is then repeated several times
over the range of throttle movement. It may be possible to move the throttle
levers slowly enough and at a sufficiently constant speed to obtain a smooth set
of time history plots, but it is more usual to plot the engine parameters in question
versus the throttle lever angle. Note that the cancellation of any configuration
warning should not affect the outcome of the test.
EXAMPLE
2A-25
There may be occasions when a set of plots is especially useful, but the
important aspect of this test is that the cockpit throttle levers are properly aligned,
both with each other and with the appropriate engine indications for each throttle
position. The intent is of course to ensure that the pilot does not learn to position
the levers in different positions on the simulator compared to the aeroplane for
the same thrust levels. Whilst not specifically stated in the ICAO Manual, the test
condition must include the correct pressure altitude, static air temperature and
Mach number for a proper comparison of the results with the aeroplane data.
Figure 2a8-1
Example of Simulator Test Results for Cockpit Throttle Lever versus EPR
2A-26
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TITLE
2a(9) - BRAKE PEDAL POSITION vs. FORCE AND
BRAKE SYSTEM PRESSURE CALIBRATION
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR BRAKE
PEDAL CHARACTERISTICS ARE CALIBRATED AND
DEMONSTRATION
Depress the brake pedal very slowly until its full range
has been achieved, then slowly release the pedals
until they return to neutral.
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FLIGHT CONDITION
GROUND
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RECORDED
PARAMETERS
BRAKE PEDAL FORCE (LEFT & RIGHT)
BRAKE PEDAL POSITION (LEFT & RIGHT)
BRAKE HYDRAULIC PRESSURE (LEFT & RIGHT)
BRAKE SYSTEM HYDRAULIC PRESSURE(S)
EVALUATION NOTES
The test should be run in the on-ground static
condition and the results overplotted with those
obtained on the aeroplane to enable an effective
comparison to be made, though the simulator
computer output results may be used to show
compliance. Compare pedal force and hydraulic
system pressure with the aeroplane values for given
pedal positions. The wheel braking system
characteristics are further validated by the tests
included in the Rejected Takeoff (1b(7)) and Stopping
Time & Distance (1e(1)) sections.
TOLERANCES
FORCE
BRAKE SYSTEM
PRESSURE
±2.2daN (5 Lbs) or ±10%
±1.0MPa (150 psi) or ±10%
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MANUAL TESTING
2A-27
Slowly move the brake pedals such that at least 30 seconds are required to
achieve a full sweep. Before performing the test, verify that the brake movement
is not obstructed by any crew member or equipment. It is worth noting however,
that the method of checking brake pedal forces manually can be difficult,
especially on older simulators.
EXAMPLE
In most modern transport aeroplanes, brake pedal loads result from pressure
feedback and deflection of the normal and alternate brake metering valve internal
return springs from the active brake hydraulic system. Pedal load characteristics
taken from a simulator test are shown in Figure 2a9-1 and tend to be
representative of both the normal and alternate brake systems. These
characteristics will change only if both active hydraulic systems are lost and the
accumulator is depressurized. Brake pedal deflections are due to stretch in the
brake cables and deflection of the metering valve return springs. Metered
pressures as a function of pedal position are shown in the lower plot. A small
initial deflection of the brake pedals is necessary before the valves begin
metering pressure to the brakes, and this should be represented in the simulator
within the prescribed tolerances.
No aeroplane data has been overplotted in Figure 2a9-1, but the plots are fairly
typical.
2A-28
Figure 2a9-1
Example of Simulator Test Results for Brake Pedal Calibration
2A-29
2A-30
SECTION 2b
2b(1)
Pitch Control
2b(2)
Roll Control
2b(3)
Yaw Control
2b(4a/b)
Small Control Inputs - Pitch (Forward/Aft)
2b(5)
Small Control Inputs - Roll
2b(6)
Small Control Inputs - Yaw
2B-1
2B
CONTROL DYNAMICS
2B.1
GENERAL
The characteristics of an aeroplane flight control system have a major
effect on handling qualities. A significant consideration in pilot
acceptability of an aeroplane is the "feel" provided through the cockpit
controls. Considerable effort is expended on aeroplane feel system
design in order to deliver a system with which pilots will be comfortable
and consider the aeroplane desirable to fly. In order for a simulator to be
representative, it too must present the pilot with the proper feel; that of
the respective aeroplane. Compliance with this requirement shall be
determined by comparing a recording of the control feel dynamics of the
simulator to aeroplane measurements in the takeoff, cruise and landing
configurations.
Recordings such as free response to an impulse or step function are
classically used to estimate the dynamic properties of electromechanical
systems. In any case, it is only possible to estimate the dynamic
properties as a result of only being able to estimate true inputs and
responses. Therefore, it is imperative that the best possible data be
collected since close matching of the simulator control loading system to
the aeroplane systems is essential.
For initial and upgrade evaluations, it is required that control dynamics
characteristics be measured at and recorded directly from the cockpit
controls. This procedure is usually accomplished by measuring the free
response of the controls using a step input or pulse input to excite the
system. The procedure must be accomplished at conditions that
represent flight at takeoff, cruise and landing.
Most large jet transport aeroplanes have fully powered, irreversible
controls with artificially provided feel forces and a balanced control
column. Variations in control release dynamics are due to variations in
feel force gradient, which is usually programmed as a function of
stabiliser position and impact pressure. Therefore, ground test conditions
may be used to represent flight conditions if the appropriate impact
pressure and stabiliser positions are used to provide the correct feel
gradients.
Because the alternate method (“rate method”) of demonstrating control
system dynamics (discussed below) is the preferred method for some
aeroplane manufacturers, they may sometimes not present data for
control releases.
2B-2
Note that the characteristics of the controller position time history are
influenced significantly by how "cleanly" the controller is released. To
provide the best matching of controller position time histories, it is
recommended that the initial portions of the force release time histories
which may be shown in the aeroplane data be approximated in the
simulator and emphasis placed on the precise positioning of the
controller. The forces of interest begin just before controller release and
continue until controller force initially approaches zero. Subsequent
forces that may be present in the aeroplane data are almost always
noise produced by aeroplane instrumentation and should be ignored.
2B.1.1
Irreversible Control Systems
For aeroplanes with irreversible control systems, measurements may be
obtained on the ground if proper pitot-static inputs are provided to
represent airspeeds typical of those encountered in flight. Likewise, it
may be shown that for some cases, takeoff, cruise and landing
configurations have like effects. Thus, one may suffice for another. If
either or both considerations apply, engineering validation or aeroplane
manufacturer rationale must be submitted as justification for ground tests
or for eliminating a configuration. For simulators requiring static and
dynamic tests at the controls, special test fixtures will not be required
during initial and upgrade evaluations if the operator's QTG shows both
test fixture results and the results of an alternate approach, such as
computer plots which were produced concurrently and show satisfactory
agreement. Repeat of the alternate method during the initial evaluation
would then satisfy this test requirement.
It should be kept in mind that, especially in overdamped systems, scatter
of up to 100% of amplitude may be observed between results of different
tests in the same configuration, due to effects of friction and unbalance.
In general the deflection (amplitude) at which the control is released is
of much more importance than the force with which it is released.
2B.1.2
Reversible Control Systems
For aeroplanes with reversible controls, the aero force gradient has an
overwhelming influence on the dynamic response of the control.
Therefore, a different dynamic response of the simulator as compared to
the aeroplane may spoil the dynamic response of the simulator control
completely due to its effect on the actual simulated aero force gradient.
Note that a relatively small offset in the surface deflection may cause a
markedly different response of the simulated aeroplane. In some cases
it may be necessary to force the simulated aeroplane to follow the
2B-3
measured dynamic response of the aeroplane in terms of attitudes, in
order to obtain a fair and independent comparison of the control
dynamics.
2B.2
CONTROL DYNAMICS EVALUATION
The dynamic properties of control systems are often stated in terms of
frequency, damping and a number of other classical measurements
which can be found in texts on control systems.
In order to establish a consistent means of validating test results for
simulator control loading, criteria are needed that will clearly define the
interpretation of the measurements and the tolerances to be applied.
Criteria are needed for both underdamped and critically and overdamped
systems. In the case of an underdamped system with very light
damping, the system may be quantified in terms of frequency and
damping. In critically damped or overdamped systems, the frequency
and damping are not readily measured from a response time history.
Therefore, some other measurement must be used.
Tests to verify that control feel dynamics represent the aeroplane must
show that the dynamic damping cycles (free response of the controls)
match that of the aeroplane within specified tolerances. The method of
evaluating the response and the tolerance to be applied is described in
the next two subparagraphs for the underdamped and critically damped
cases.
2B.2.1
Underdamped Response
Two measurements are required for the period, the time to first zero
crossing (in case a rate limit is present) and the subsequent frequency
of oscillation. It is necessary to measure cycles on an individual basis
in case there are non-uniform periods in the response. Each period will
be independently compared to the respective period of the aeroplane
control system and, consequently, will enjoy the full tolerance specified
for that period.
The damping tolerance should be applied to overshoots on an individual
basis. Care should be taken when applying the tolerance to small
overshoots since the significance of such overshoots becomes
questionable. Only those overshoots larger than 5% of the total initial
displacement should be considered. The residual band, labelled T(Ad)
on Figure 2b-1 is ±5% of the initial displacement amplitude Ad from the
steady state value of the oscillation. Oscillations within the residual band
2B-4
are considered insignificant. When comparing simulator data to
aeroplane data, the process should begin by overlaying or aligning the
simulator and aeroplane steady state values and then comparing
amplitudes of oscillation peaks, the time of the first zero crossing, and
individual periods of oscillation. The simulator should show the same
number of significant overshoots to within 1 when compared against the
aeroplane data. This procedure for evaluating the response is illustrated
in Figure 2b-1.
Figure 2b-1
Underdamped Step Response
2B.3
TOLERANCES
The following table summarises the tolerances, T. Note that the
tolerance on P0 applies to underdamped as well as critically damped
(see Figure 2b-2 below for an example) and overdamped systems. The
remaining tolerances apply only to underdamped systems. See Figures
2b-1 and 2b-2 for an illustration of the referenced measurements.
T(P0)
T(P1)
T(P2)
T(Pn)
T(An)
±10% of P0
±20% of P1
±30% of P2
±10(n+1)% of Pn
±10% of A1
2B-5
T(Ad)
Significant Overshoots
±5% of Ad = Residual Band
First Overshoot and ±1 Subsequent
Overshoots
Tolerances apply against the absolute values of each period
(considered independently).
Figure 2b-2
Critically Damped Step Response
2B.4 ALTERNATE METHOD FOR CONTROL DYNAMICS
One aeroplane manufacturer has proposed, and his regulatory authority
has accepted, an alternate means for dealing with control dynamics. The
method applies to aeroplanes with hydraulically powered flight controls and
artificial feel systems. Instead of free response measurements, the system
would be validated by measurements of control force and rate of
movement.
For each axis of pitch, roll and yaw, the control shall be forced to its
maximum extreme position for the following distinct rates. These tests
shall be conducted at typical taxi, takeoff, cruise and landing conditions.
2B.4.1
Static Test
Slowly move the control such that approximately 100 seconds are required
2B-6
to achieve a full sweep. A full sweep is defined as movement of the
controller from neutral to the stop, usually aft or right stop, then to the
opposite stop, then to the neutral position.
2B.4.2
Slow Dynamic Test
Achieve a full sweep in approximately 10 seconds.
2B.4.3
Fast Dynamic Test
Achieve a full sweep in approximately 4 seconds.
NOTE: Dynamic sweeps may be limited to forces not exceeding 44.5 daN
(100 lb).
2B.4.4
Tolerances
a) Static Test - see Tests 2a(1), 2a(2) and 2a(3), Section 2a.
b) Dynamic Test - ±0.9 daN (2 lb) or ±10% on dynamic increment above
static test.
The authorities are open to alternative means such as the one described
above. Such alternatives must, however, be justified and appropriate to the
application. For example, the method described here may not apply to all
manufacturers' systems and certainly not to aeroplanes with reversible
control systems. Hence, each case must be considered on its own merit on
an ad-hoc basis. Should the authority find that alternative methods do not
result in satisfactory performance, then more conventionally accepted
methods must be used.
2B-7
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TITLE
2b(1) - PITCH CONTROL DYNAMICS
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OBJECTIVE
CONTROL DYNAMIC CHARACTERISTICS
CONFORM TO THE AEROPLANE CONTROL
RESPONSE.
DEMONSTRATION
The longitudinal controller is moved to a specified
initial amplitude of approximately 25% to 50% of
maximum available travel and then abruptly released.
Alternatively, the pitch controller may be moved
through a full sweep at a slow, moderate, and fast rate
(approximately 100, 10, and 4 seconds, respectively).
For aeroplanes with irreversible control systems,
measurements may be obtained on the ground if
proper pitot-static inputs to an artificial feel system are
provided to represent airspeeds typical of those
encountered in flight.
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FLIGHT CONDITION
a) TAKEOFF
b) CRUISE
c) LANDING
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RECORDED
PARAMETERS
PITCH CONTROLLER FORCE
EVALUATION NOTES
The oscillatory characteristics of the control response
are evaluated in terms of the period, magnitude and a
number of the overshoots on the plot. These
parameters are then compared with the equivalent
terms from the aeroplane data plots to check for good
correlation. The reason for the increasing tolerance on
each successive oscillation period is to account for
any accumulative errors which tend to build up whilst
the controls are responding. It is not necessary to
include the elevator angle in the assessment of the
2B-8
response, but it is very important to ascertain that the
correct feel pressure is used during the test. Note that
this test does not need to be supplied if the dynamic
response is generated solely by use of aeroplane
hardware in the simulator. Data should be for normal
control displacements in both directions
(approximately 25% to 50% full throw or approximately
25% to 50% of maximum allowable controller
deflection for flight conditions limited by the
manoeuvring load envelope - particularly for tab-driven
control systems).
TOLERANCES
For underdamped response:
Time from 90% of initial
±10%
displacement (Ad) to first
zero crossing
Time for nth period
±10(n+1)% of period
thereafter
(where n = the sequential period of a full oscillation)
Amplitude of all overshoots
±10% amplitude of
greater than 5% of initial
first overshoot
displacement (Ad))
Number of significant
±1
overshoots
(first significant overshoot should be matched)
For overdamped systems:
±10% of time from 90% of initial displacement (Ad) to
10% of initial displacement (0.1 Ad)
For the alternate method (slow, moderate, rapid
control sweeps):
(The slow sweep is equivalent to the static test 2a(1))
For the moderate and rapid sweeps:
Dynamic increment above the ±0.9 daN (2 Lb) or
static force
±10%
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MANUAL TESTING
This test typically is carried out with the simulator carefully trimmed in the correct
configuration and at the appropriate condition. The pitch controller is displaced
to the value specified in the flight test data, held briefly and then released and
2B-9
allowed to freely respond for a few seconds whilst the controller position is
plotted. For the alternative method using full control sweeps, the slow test is
equivalent to the static test contained in test 2a(1), and the medium and fast
sweeps should be carried out in a similar manner, but at faster rates.
EXAMPLE
A typical result for conventional pitch control dynamics is shown in Figure 2b1-1,
clearly illustrating a classic example of an underdamped response, and is almost
always more difficult to replicate in the simulator than a response which is much
more damped. The result below may not meet the criteria within the required
tolerance for the amplitude of the first overshoot.
Figure 2b1-1
Example of Simulator Test Results for Pitch Control Dynamics
2B-10
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TITLE
2b(2) - ROLL CONTROL DYNAMICS
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR ROLL
CONTROL
DYNAMIC CHARACTERISTICS
RESPONSE.
DEMONSTRATION
The roll controller is moved to a specified initial
amplitude of approximately 25% to 50% of maximum
available travel and then abruptly released.
Alternatively, the roll controller may be moved through
a full sweep at a slow, moderate, and fast rate
It may be shown that for aeroplanes with irreversible
control systems, the takeoff, cruise and landing
configurations have like effects on lateral control
system dynamic characteristics. Thus one may suffice
for another.
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FLIGHT CONDITION
a) TAKEOFF
b) CRUISE
c) LANDING
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RECORDED
PARAMETERS
ROLL CONTROLLER FORCE
EVALUATION NOTES
include the aileron or spoiler angles in the assessment
2B-11
of the response, but it is very important to ascertain
that the correct feel pressure is used during the test if
an artificial feel system is used in the lateral axis. Data
should be for normal control displacements in both
directions (approximately 25% to 50% full throw or
approximately 25% to 50% of maximum allowable
controller deflection for flight conditions limited by the
manoeuvring load envelope - particularly for tab-driven
control systems).
Note that this test does not need to be supplied if the
dynamic response is generated solely by use of
aeroplane hardware in the simulator.
TOLERANCES
±10%
zero crossing
Time for nth period
thereafter
±10% amplitude of
first overshoot
displacement (Ad))
±1
overshoots
control sweeps):
static force
±10%
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MANUAL TESTING
This test typically is carried out with the simulator carefully trimmed in the correct
2B-12
configuration and at the appropriate condition. The roll controller is displaced to
the value specified in the flight test data, held briefly and then released and
allowed to freely respond for a few seconds whilst the controller position is
plotted. For the alternative method using full control sweeps, the slow test is
equivalent to the static test contained in test 2a(2), and the medium and fast
sweeps should be carried out in a similar manner, but at faster rates.
EXAMPLE
Figure 2b2-1 gives a typical result for a control wheel dynamics test. There are
usually only a small number of overshoots, and in this certainly no more than two
that ‘count’ (i.e. are $5% of the initial displacement). Whilst the time for the zero
crossing is probably just about within the 10% tolerance, the simulator result may
be problematic in that the amplitude of the first overshoot is too low. Strictly
speaking the number of overshoots is within the tolerance of ±1, but some
regulatory authorities might consider it worth further scrutiny.
Figure 2b2-1
Example of Simulator Test Result for Roll Control Dynamics
2B-13
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TITLE
2b(3) - YAW CONTROL DYNAMICS
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR YAW
CONTROL DYNAMIC CHARACTERISTICS
RESPONSE.
DEMONSTRATION
The rudder pedals are moved to a specified initial
amplitude of approximately 25% to 50% of maximum
available travel and then abruptly released.
Alternatively, the rudder pedals may be moved
through a full sweep at a slow, moderate, and fast rate
It may be shown that for aeroplanes with irreversible
control systems, the takeoff, cruise and landing
configurations have like effects on yaw control system
dynamic characteristics. Thus one may suffice for
another.
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FLIGHT CONDITION
a) TAKEOFF
b) CRUISE
c) LANDING
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RECORDED
PARAMETERS
RUDDER PEDAL FORCE
EVALUATION NOTES
include the rudder angle in the assessment of the
2B-14
response, but it is very important to ascertain that the
correct feel pressure is used during the test, if an
artificial feel system is used in the directional axis.
Data should be for normal control displacement
(approximately 25% to 50% of full throw). It may be
shown that for aeroplanes with irreversible control
systems the takeoff, cruise and landing configurations
have like effects on lateral control system dynamic
characteristics. Thus, one may suffice for another.
Note that this test does not need to be supplied if the
dynamic response is generated solely by use of
aeroplane hardware in the simulator.
TOLERANCES
±10%
zero crossing
Time for nth period
thereafter
±10% amplitude of
first overshoot
displacement (Ad))
±1
overshoots
control sweeps):
static force
±10%
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MANUAL TESTING
This test is typically carried out with the simulator carefully trimmed in the correct
2B-15
configuration and at the appropriate condition. The captain’s rudder pedal is
displaced to the value specified in the flight test data, held briefly and then
released and allowed to freely respond for a few seconds whilst the pedal
position is plotted. For the alternative method using full control sweeps, the slow
test is equivalent to the static test contained in test 2a(3), and the medium and
fast sweeps should be carried out in a similar manner, but at faster rates.
EXAMPLE
In general, obtaining a really good result for pedal position dynamics seems to
cause more problems than the other two axes. Figure 2b3-1 shows a surprisingly
good match, as long as one takes into account the fact that the third overshoot
does not need to be counted.
Figure 2b3-1
Example of Simulator Test Results for Yaw Control Dynamics
2B-16
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TITLE
2b(4) - SMALL CONTROL INPUTS - PITCH
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OBJECTIVE
RESPONSE TO SMALL PITCH CONTROL INPUTS
DEMONSTRATION
During trimmed flight make small pitch control inputs
typical of minor corrections made while established on
an ILS approach. The pitch controller should be
moved in both directions and result in a pitch rate of
approximately 0.5 to 2 deg/sec. The control input
should be large enough to overcome the breakout
force, but usually not more than 5% of the total control
travel from neutral to one stop.
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FLIGHT CONDITION
a) APPROACH or LANDING, FORWARD
b) APPROACH or LANDING, AFT
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
CALIBRATED AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
PITCH RATE
ELEVATOR ANGLE
EVALUATION NOTES
The purpose of this test is to determine that the
longitudinal pilot control "feel" for small pitch control
inputs affects the simulated aeroplane in the same
way as an equivalent control force and movement
would do in the real aeroplane. The plot scales for the
pitch rate, elevator angle and pitch controller position
must all be carefully chosen so as to facilitate proper
analysis of the results.
The control input in the aeroplane data acquisition
2B-17
should be large enough to overcome the breakout
force, but usually not more than 5% of the total control
travel from neutral to one stop. The test may be of
short duration, but pitch control inputs should be made
in both the nose up and nose down directions, thereby
forming a single test which encompasses both
directions. Data should be plotted from 5 seconds
before until at least 5 seconds after initiation of the
control input.
CCA: Test in normal and non-normal control state.
TOLERANCES
BODY PITCH RATE
±0.15o/Sec or
±20% of peak body pitch rate
applied throughout the time
history
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MANUAL TESTING
The basic method is as stated in the 'EVALUATION NOTES' section above. It is
extremely important that a very stable trim condition is achieved prior to any
movement of the pitch controller. Notwithstanding the presentation of the data
from the aeroplane, the actual longitudinal control input should be carried out
slowly and from a very steady trim condition, so that the resultant simulator pitch
response can be clearly seen as soon as it is caused by the elevator
displacement.
EXAMPLE
By no means a terrible result, but the plots shown in Figure 2b4-1 could still be
improved upon. The initial speed is slightly off, plus the combination angle of
attack and pitch angle differences may mean that the initial rate of climb was
incorrect. However, the movement of the control column (not shown) gives the
correct change in elevator angle and the pitch responds quite well.
2B-18
Figure 2b4-1
Example of Simulator Test Results for Small Control Inputs,
Pitch
2B-19
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TITLE
2b(5) - SMALL CONTROL INPUTS - ROLL
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OBJECTIVE
RESPONSE TO SMALL ROLL CONTROL INPUTS
DEMONSTRATION
During trimmed flight make small roll control inputs
an ILS approach. The roll controller should be moved
in one direction, or if the aeroplane exhibits nonsymmetrical behaviour in the lateral axis, the control
inputs should be made in both directions. The control
inputs should result in a roll rate of approximately 0.5
to 2 deg/sec. The control input should be large enough
to overcome the breakout force, but usually not more
than 5% of the total control travel from neutral to one
stop.
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FLIGHT CONDITION
APPROACH or LANDING
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
CALIBRATED AIRSPEED
BANK ANGLE
ROLL RATE
AILERON ANGLE(S)
SPOILER ANGLES
SIDESLIP ANGLE
HEADING ANGLE
RUDDER ANGLE
YAW RATE
EVALUATION NOTES
The purpose of this test is to determine that the area
of pilot lateral control "feel" for small roll control inputs
2B-20
affects the simulated aeroplane in the same way as an
equivalent control force and movement would do in
the real aeroplane. The plot scales for the roll rate,
aileron and spoiler angles and roll controller position
must all be carefully chosen so as to facilitate proper
analysis of the results. The test may be of short
duration, and data should be plotted from 5 seconds
before until at least 5 seconds after initiation of the
control input.
TOLERANCES
BODY ROLL RATE
±0.15o/Sec or
±20% of peak body roll rate
history
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MANUAL TESTING
movement of the roll controller. Notwithstanding the presentation of the data
from the aeroplane, the actual lateral control input should be carried out slowly
and from a very steady trim condition, so that the resultant simulator roll response
can be clearly seen as soon as it is caused by the lateral control surface
displacement.
EXAMPLE
Something is clearly amiss in the result shown in Figure 2b5-1. There appears
to be a roll response even before the aileron is displaced. The most likely
explanation is that the initialisation procedure failed to complete properly; the
control displacements used for the trimming process would then be incorrectly
positioned for the start of the test. One other possibility is that the controls were
positioned correctly at the completion of the trim phase, but then were released
prematurely prior to the proper beginning of the test.
2B-21
Figure 2b5-1
Example of Simulator Test Results for Small Control Inputs, Roll
2B-22
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TITLE
2b(6) - SMALL CONTROL INPUTS - YAW
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OBJECTIVE
RESPONSE TO SMALL YAW CONTROL INPUTS
DEMONSTRATION
During trimmed flight make small yaw control inputs
an ILS approach. The rudder pedals should be moved
in one direction, or if the aeroplane exhibits nonsymmetrical behaviour in the yaw axis, the control
inputs should be made in both directions. The control
inputs should result in a yaw rate of approximately 0.5
to 2 deg/sec. The control input should be large enough
to overcome the breakout force, but usually not more
than 5% of the total control travel from neutral to one
stop.
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FLIGHT CONDITION
APPROACH or LANDING
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
CALIBRATED AIRSPEED
BANK ANGLE
HEADING ANGLE
YAW RATE
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE
SPOILER ANGLES
ROLL RATE
EVALUATION NOTES
The purpose of this test is to determine that the area
of directional pilot control "feel" for small control inputs
2B-23
affects the simulated aeroplane in the same way as an
equivalent control force and movement would do in
the real aeroplane. The plot scales for the yaw rate,
rudder angles, and rudder position must all be
carefully chosen so as to facilitate proper analysis of
the results. The test may be of short duration, and
data should be plotted from 5 seconds before until at
least 5 seconds after initiation of the control input.
TOLERANCES
BODY YAW RATE
±0.15o/Sec or
±20% of peak body yaw rate
history
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MANUAL TESTING
movement of the rudder pedals. Notwithstanding the presentation of the data
from the aeroplane, the actual directional control input should be carried out
slowly and from a very steady trim condition, so that the resultant simulator yaw
response can be clearly seen as soon as it is caused by the rudder displacement.
EXAMPLE
The result shown in Figure 2b6-1 is at best marginal, probably because the
rudder pedal input (not shown) was insufficient to give the required rudder
displacement. It is to be expected that using the correct amount of rudder and/or
following the change in engine thrust would significantly improve the match.
2B-24
Figure 2b6-1
Example of Simulator Test Results for Small Control Inputs, Yaw
2B-25
2B-26
SECTION 2c
LONGITUDINAL
2c(1)
2c(2)
2c(3)
Spoiler/Speedbrake Change Dynamics
2c(4)
2c(5)
Longitudinal Trim
2c(6)
Longitudinal Manoeuvring Stability (Stick Force/g)
2c(7)
Longitudinal Static Stability
2c(8)
Stall Characteristics
2c(9)
Phugoid Dynamics
2c(10)
Short Period Dynamics
2C-1
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TITLE
2c(1) - POWER CHANGE DYNAMICS
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OBJECTIVE
RESPONSE TO AN ENGINE POWER CHANGE
DEMONSTRATION
Starting from the prescribed trimmed condition,
increase engine power setting to match the desired
power setting and allow the aeroplane to respond
freely. The initial condition should be thrust for
approach or level flight followed by a sudden
commanded power increase to maximum continuous
or go-around power. The test should be conducted
from at least 5 seconds before the initiation of the
power change until at least 15 seconds after the thrust
has reached a stabilised value.
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FLIGHT CONDITION
APPROACH
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RECORDED
PARAMETERS
AIRSPEED
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
RATE OF CLIMB
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
EVALUATION NOTES
For this test, it is clearly of fundamental importance
that the throttle lever position for each engine matches
the validation data throughout the duration of the test.
To obtain the closest correlation, it will be necessary
2C-2
that the flight test engine thrust is followed as closely
as possible. For some simulators however, this thrust
may be difficult to achieve on the simulator, perhaps
because the flight test engine differs from that
represented on the simulator, or else because the
thrust itself is not included in the validation data.
Whichever engine parameter is plotted, further
information relating to the thrust should be supplied
with the simulator test results which enables a clear
understanding to be gained. Any extra pilot activity
towards the end of the flight test time history which
renders the last few seconds as 'hands-on' rather than
a free response can usually be ignored, but the
simulator test should still be run for the full duration as
in the ICAO Manual. Some residual pitch rate may be
present and discernable from the aeroplane time
history. If this is the case, its inclusion in the simulator
test may be necessary for the test to pass within the
required tolerances. The engines should be
manipulated using the throttle levers, thus exercising
the propulsion system model as well as the
aerodynamics and equations of motion. Small
perturbations of the elevator surface (e.g. less than 1o)
may mean the difference between the test passing
and failing, but nevertheless this test should be run as
a free response.
Note that two tests are required for computer
controlled aeroplanes, one for the normal control state
and the other for a non-normal control state. Also,
stability augmentation systems must be configured as
the aeroplane system was configured during the data
acquisition.
TOLERANCES
PITCH ANGLE
AIRSPEED
ALTITUDE
±1.5o or ±20%
±3 Kts
±30 m (100 Ft)
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MANUAL TESTING
Above all else, it is important that the simulated aeroplane be accurately trimmed
in accordance with the validation data before commencing this test. After the
2C-3
prescribed interval, which should ideally be of several seconds (at least 5) to
check the trim, the throttles are advanced to achieve the required engines
parameters (EPR, N1, Torque, etc.) at the appropriate rate. The simulator should
then be left to respond freely for at least 15 seconds after the engines have
reached their final stable state. If the aeroplane data indicate that the pilot did not
fly the manoeuvre "hands-off" then attempts to duplicate his actions should be
made. However, the manoeuvre should really be flown without pilot input and if
it is not certain whether the flight test was flown in this manner, then it is easiest
to assess the results if the flight controls are left alone. If there is a residual pitch
rate present and discernable from the aeroplane time history, its inclusion in the
simulator test may be necessary for the test to pass within the required
tolerances, but this will render the manoeuvre difficult to fly manually. Under such
circumstances it is better to fly the manoeuvre from a stable, trimmed condition
and account for the differences in the results obtained by use of engineering
judgement.
EXAMPLE
In the plot of Figure
2c1-1, the pitch angle
deviates outside the
specified tolerance
after approximately
36 seconds. An
evaluator would have
to examine the other
parameters plotted for
this test, but to see a
result deviate in this
way towards the end
of a manoeuvre is not
uncommon. It usually
signifies that the pilot
began correcting the
aeroplane response
during the flight test
by applying control
i n p u t s ,
b u t
occasionally it may be
caused by an
a t m o s p h e r i c
disturbance.
Figure 2c1-1
Example of Simulator Test Results for Power Change Dynamics
2C-4
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TITLE
2c(2a) - FLAP CHANGE DYNAMICS (Retraction)
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OBJECTIVE
RESPONSE TO A FLAP RETRACTION IN THE
SECOND TO THIRD SEGMENT CLIMB CONDITION
DEMONSTRATION
Starting from the prescribed trimmed condition, set the
flap handle to retract the flaps from the takeoff position
to the initial flap retracted and allow the aeroplane to
respond freely. The test should be conducted from at
least 5 seconds before the initiation of the flap change
until at least 15 seconds after the flap motion has
ended.
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FLIGHT CONDITION
TAKEOFF THROUGH INITIAL FLAP RETRACTION
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
STABILISER ANGLE
ELEVATOR ANGLE
FLAP LEVER POSITION/FLAP SURFACE ANGLE(S)
INDICATED FLAP ANGLE (to demonstrate timing)
PRESSURE ALTITUDE
RATE OF CLIMB
BANK ANGLE
EVALUATION NOTES
The timing of the flap/slat movement during the
retraction must match the validation data. However,
the precise profile of the flap and slat movement may
not be exactly synchronised on the simulator when
2C-5
compared to the flight test aeroplane. A mean value of
flap/slat time and position is therefore usually
acceptable, makes little difference to the test outcome
and will produce more consistent results for recurrent
evaluations. Any extra pilot activity towards the end of
the flight test time history which renders the last few
seconds as 'hands-on' rather than a free response can
usually be ignored, but the simulator test should still
be run for the full duration stated in the ICAO Manual..
The flaps should be retracted using the lever, thus
exercising the high-lift device actuation model as well
as the aerodynamics and the equations of motion.
See notes for test 2c(1) concerning residual pitch rate
and small elevator perturbations.
acquisition.
TOLERANCES
PITCH ANGLE
AIRSPEED
ALTITUDE
±1.5o or ±20%
±3 Kts
±30 m (100 Ft)
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MANUAL TESTING
Above all else, it is important that the aeroplane be accurately trimmed in
accordance with the validation data before commencing this test. After the
check the trim, the flap lever is moved to the appropriate detent position as
specified in the data. The simulator should then be left to respond freely for at
least 15 seconds after both the flaps and slats have reached their final stable
state. If the aeroplane data indicates that the pilot did not fly the manoeuvre
"hands-off" then attempts to duplicate his actions should be made. However, the
manoeuvre should really be flown without pilot input and if it is not certain
whether the flight test was flown in this manner, then it is easiest to assess the
results if the flight controls are left alone. If there is a residual pitch rate present
and discernable from the aeroplane time history, its inclusion in the simulator test
may be necessary for the test to pass within the required tolerances, but this will
render the manoeuvre difficult to fly manually. Under such circumstances it is
2C-6
better to fly the manoeuvre from a stable, trimmed condition and account for the
differences in the results obtained by use of engineering judgement.
EXAMPLE
There is nothing particularly untoward about the plots shown in Figure 2c2-1, but
it can clearly be seen that the aeroplane flight test data (the dotted line) ceased
to be hand-free after approximately 30 seconds. However, the flap motion ended
at around 7 seconds, so in any case the requirement of the ICAO Manual was
satisfied by 22 seconds (7 seconds + 15 seconds). There has been no attempt
to replicate the pilot’s manipulation of the elevator for the last two seconds, nor
is there any necessity to do so.
Figure 2c2-1
Example of Simulator Test Results for Flap Change Dynamics,
Retraction
2C-7
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TITLE
2c(2b) - FLAP CHANGE DYNAMICS (Extension)
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OBJECTIVE
RESPONSE TO A FLAP EXTENSION IN THE
APPROACH TO LANDING CONDITION CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Starting from the prescribed trimmed condition, set the
flap handle to extend the flaps from the approach
position to the landing flap position, and allow the
aeroplane to respond freely. The test should be
conducted from at least 5 seconds before the initiation
of the flap change until at least 15 seconds after the
flap motion has ended.
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FLIGHT CONDITION
APPROACH TO LANDING
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
STABILISER ANGLE
ELEVATOR ANGLE
FLAP LEVER POSITION/FLAP SURFACE ANGLE(S)
INDICATED FLAP ANGLE (to demonstrate timing)
PRESSURE ALTITUDE
RATE OF CLIMB
BANK ANGLE
EVALUATION NOTES
See notes for test 2c(2a). The difference is that in this
test the flaps are to be extended rather than retracted.
TOLERANCES
PITCH ANGLE
2C-8
±1.5o or ±20%
AIRSPEED
ALTITUDE
±3 Kts
±30 m (100 Ft)
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MANUAL TESTING
See notes for test 2c(2a).
EXAMPLE
The result in Figure 2c2-2 is just about in tolerance, with the possible exception
of the airspeed at 65 seconds, though this may be beyond the 15 second period
after the flaps and slats (not shown) have achieved the commanded value. Even
if it is within the 15 seconds, a tiny excursion outside the tolerance band such as
this is unlikely, as an isolated case, to cause the regulators to take drastic action
against the qualification of the simulator. The result should be improved if it is
possible to do so, and careful use may need to be made of the aeroplane proofof-match data. This is an example of a test which is slightly affected by an
atmospheric disturbance (i.e. wind/turbulence, visible in the airspeed trace), but
this can be easily explained in a note accompanying the test result.
2C-9
Figure 2c2-2
Example of Simulator Test Results for Flap Change Dynamics, Extension
2C-10
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TITLE
2c(3a) - SPOILER/SPEEDBRAKE
DYNAMICS (Extension)
CHANGE
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OBJECTIVE
RESPONSE TO A SPEEDBRAKE EXTENSION IN
THE CRUISE CONDITION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Starting from the prescribed trimmed condition with
the speedbrake handle fully retracted, set the
speedbrake handle to near the inflight detent position,
and allow the aeroplane to respond freely. The test
should be conducted from at least 5 seconds before
the initiation of the speedbrake change until at least
15 seconds after the speedbrake motion has ended.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
SPOILER ANGLES
PRESSURE ALTITUDE
RATE OF CLIMB
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
EVALUATION NOTES
As with all the configuration change dynamics tests, it
is obviously important that the way in which the
spoilers move during the extension closely match the
2C-11
aeroplane data throughout the duration of the test,
though because of the much higher speed at which
the spoilers move the effects of a slight difference are
likely to be less obvious in the results. Any extra pilot
activity towards the end of the flight test time history
which renders the last few seconds as 'hands-on'
rather than a free response can usually be ignored,
but the simulator test should still be run for the full
duration stated in the ICAO Manual. The speedbrake
should be extended using the lever, thus exercising
the speedbrake actuation system model as well as
the aerodynamics and equations of motion. See notes
for test 2c(1) concerning residual pitch rate and small
elevator perturbations.
acquisition.
TOLERANCES
PITCH ANGLE
ALTITUDE
AIRSPEED
±1.5o or ±20%
±30 m (100 Ft)
±3 Kts
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MANUAL TESTING
accordance with the flight test data before commencing this test. After the
check the trim, the speedbrake lever is moved to the appropriate inflight detent
position as specified in the data. The simulator should then be left to respond
freely for at least 15 seconds after the spoilers have reached their final stable
state. If the aeroplane data indicates that the pilot did not fly the manoeuvre
"hands-off" then attempts to duplicate his actions should be made. However, the
manoeuvre should really be flown without pilot input and if it is not certain
whether the flight test was flown in this manner, then it is easiest to assess the
results if the flight controls are left alone. If there is a residual pitch rate present
and discernable from the aeroplane time history, its inclusion in the simulator test
may be necessary for the test to pass within the required tolerances, but this will
render the manoeuvre difficult to fly manually. Under such circumstances it is
2C-12
better to fly the manoeuvre from a stable, trimmed condition and account for the
differences in the results obtained by use of engineering judgement.
EXAMPLE
Figure 2c3-1 is another example of a test where the result(s) ends up being
slightly out of tolerance at the very end. This is not uncommon for the
uncontrolled free-response tests such as this one, but there are usually ways to
make slight improvements that will ensure the test remains in tolerance to the
end, though sometimes it may mean that an intermediate portion of the time
history (which is usually well within tolerance) is slightly worse.
Figure 2c3-1
Example of Simulator Test Results for Speedbrake Change
Dynamics, Extension
2C-13
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TITLE
2c(3b) - SPOILER/SPEEDBRAKE
DYNAMICS (Retraction)
CHANGE
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OBJECTIVE
RESPONSE TO A SPEEDBRAKE RETRACTION IN
THE CRUISE CONDITION CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
From the prescribed trimmed condition, retract the
speedbrakes and allow the simulator to respond
freely. Compare the resulting response with the
aeroplane data.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
SPOILER ANGLES
PRESSURE ALTITUDE
RATE OF CLIMB
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
EVALUATION NOTES
test the speedbrake is to be retracted rather than
extended.
TOLERANCES
PITCH ANGLE
ALTITUDE
2C-14
±1.5o or ±20%
±30 m (100 Ft)
AIRSPEED
±3 Kts
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MANUAL TESTING
See notes for test 2c(3a).
EXAMPLE
A reasonable result for speedbrake retraction dynamics is shown in Figure 2c3-2.
The altitude is barely in tolerance by the end of the time history though. These
minor items are often overlooked and can usually be corrected quite easily, in
this case most probably by slight adjustment of the trim pitch angle and/or rate
of climb.
2C-15
2C-16
Figure 2c3-2
Example of Simulator Test Results for Speedbrake Change Dynamics, Retraction
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TITLE
2c(4a) - GEAR CHANGE DYNAMICS (Retraction)
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OBJECTIVE
TO DEMONSTRATE THAT THE DYNAMIC
RESPONSE TO A LANDING GEAR RETRACTION
DEMONSTRATION
Starting from the prescribed trimmed first segment
climb condition with the gear handle in the extended
position, set the gear handle to the retracted position
the initiation of the gear change until at least 15
seconds after the landing gear motion has ended.
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FLIGHT CONDITION
1ST TO 2ND SEGMENT CLIMB
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
LANDING GEAR HANDLE POSITION
LANDING GEAR INDIVIDUAL POSITIONS
GEAR POSITION INDICATION/LIGHTS STATUS (to
demonstrate timing)
PRESSURE ALTITUDE
RATE OF CLIMB
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
EVALUATION NOTES
Once again, as this is a configuration change
dynamics test, it is obviously important that the way in
which each landing gear moves during the retraction
2C-17
closely matches the aeroplane data throughout the
duration of the test. Any extra pilot activity towards the
end of the flight test time history which renders the last
few seconds as 'hands-on' rather than a free response
can usually be ignored, but the simulator test should
still be run for the full duration stated in the ICAO
Manual. The landing gear should be retracted using
the lever, thus exercising the gear actuation system
model as well as the aerodynamics and equations of
motion. See notes for test 2c(1) concerning residual
pitch rate and small elevator perturbations.
acquisition.
TOLERANCES
PITCH ANGLE
AIRSPEED
ALTITUDE
±1.5o or ±20%
±3 Kts
±30 m (100 Ft)
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MANUAL TESTING
accordance with the validation data before commencing this test. After the
check the trim, the landing gear lever is retracted. The simulator should then be
left to respond freely for at least 15 seconds after the landing gear lights on the
flight deck have extinguished. If the aeroplane data indicates that the pilot did not
fly the manoeuvre "hands-off" then attempts to duplicate his actions should be
made. However, the manoeuvre should really be flown without pilot input and if
it is not certain whether the flight test was flown in this manner, then it is easiest
to assess the results if the flight controls are left alone. If there is a residual pitch
rate present and discernable from the aeroplane time history, its inclusion in the
simulator test may be necessary for the test to pass within the required
tolerances, but this will render the manoeuvre difficult to fly manually. Under such
circumstances it is better to fly the manoeuvre from a stable, trimmed condition
and account for the differences in the results obtained by use of engineering
judgement.
2C-18
EXAMPLE
Figure 2c4-1 is a typical result for this test. In most modern jet transport
aeroplanes retracting the gear does not usually have a very large effect on
flightpath. If the test conditions have been correctly initialised and the trimmed
state defined in the validation data properly replicated, in most cases little can go
wrong.
Figure 2c4-1
Example of Simulator Test Results for Gear Change Dynamics, Retraction
2C-19
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TITLE
2c(4b) - GEAR CHANGE DYNAMICS (Extension)
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OBJECTIVE
TO DEMONSTRATE THAT THE DYNAMIC
RESPONSE TO A LANDING GEAR EXTENSION
DEMONSTRATION
Starting from the prescribed trimmed approach
condition with the gear handle in the retracted
position, set the gear handle to the extended position
the initiation of the gear change until at least 15
seconds after the landing gear motion has ended.
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FLIGHT CONDITION
APPROACH TO LANDING
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
LANDING GEAR HANDLE POSITION
LANDING GEAR INDIVIDUAL POSITIONS
GEAR POSITION INDICATION/LIGHTS STATUS (to
demonstrate timing)
PRESSURE ALTITUDE
RATE OF CLIMB
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
EVALUATION NOTES
test the landing gear is to be extended rather than
retracted.
2C-20
TOLERANCES
PITCH ANGLE
AIRSPEED
ALTITUDE
±1.5o or ±20%
±3 Kts
±30 m (100 Ft)
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MANUAL TESTING
See notes for test 2c(4a). The difference is that in this test the landing gear is to
be extended rather than retracted.
EXAMPLE
A similar situation exists for the gear extension test, as illustrated in Figure 2c4-2,
as for the gear retraction test 2c(4a). In addition, both these tests are usually
quite straightforward to run manually and achieve a pass.
Figure 2c4-2
Example of Simulator Test Results for Gear Change Dynamics, Extension
2C-21
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TITLE
2c(5) - LONGITUDINAL TRIM
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OBJECTIVE
INTER-RELATIONSHIPS OF LIFT, DRAG, THRUST
AND LONGITUDINAL TRIM CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Establish a steady state wings-level constant altitude
flight condition, setting thrust as required to achieve
the target speed. Data may be presented as a series
of snapshots.
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FLIGHT CONDITION
a) CRUISE
b) APPROACH
c) LANDING
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
PRESSURE ALTITUDE
RATE OF CLIMB
STABILISER ANGLE
ELEVATOR ANGLE
LINEAR ACCELERATIONS (Longitudinal, Lateral,
Vertical)
EVALUATION NOTES
This test is generally straightforward in that the only
requirement is to check certain longitudinal
parameters in a steady state, wings level trimmed
condition. The simulator automatic test system will
usually be used to quickly perform any trim test. The
longitudinal control system should be exercised fully,
including any aerodynamic hinge moment
calculations, especially for aeroplanes in which the
2C-22
elevator neutral position does not necessarily
correspond to zero degrees. The engines should be
driven from the throttle levers. There is not a
requirement for a time history to be plotted.
Note that only one test is required for computer
controlled aeroplanes for each of the three flight
conditions, which may be for either the normal control
state or a non-normal control state.
TOLERANCES
ELEVATOR ANGLE
STABILISER ANGLE
PITCH ANGLE
NET THRUST or equivalent
±1o
±0.5o
±1o
±5%
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MANUAL TESTING
These cases are to verify the recorded parameters in a stable condition at the
weight and configuration specified in the data. Sufficient time should be spent in
ensuring that the trim is accurate.
EXAMPLE
The cruise case result in Figure 2c5-1 proves well that the results as given are
within the tolerances, but elevator angle, which is a toleranced parameter, has
not been printed nor is it demonstrated that the condition being checked is
actually a trim. Linear (and preferably rotational as well) accelerations should be
printed to enable an evaluator to be sure that the result is as it should be. Also,
the use of the term ‘glideslope angle’ for this printout is confusing. It can be
assumed that the actual parameter referred to is flightpath angle, but the result
should state this clearly.
2C-23
Figure 2c5-1
Example of Simulator Test Results for Longitudinal Trim
2C-24
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TITLE
2c(6) - LONGITUDINAL MANOEUVRING STABILITY
(STICK FORCE/G)
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OBJECTIVE
MANOEUVRING STABILITY, NORMALLY
MEASURED AS STICK FORCE PER 'G',
DEMONSTRATION
Trim wings level at the prescribed conditions. The test
is performed either by steadily increasing bank angle
until reaching the prescribed maximum angle, or by
establishing a steady-state condition at several
intermediate bank angle up to and including the
maximum angle. Use longitudinal control force to
maintain the trim speed
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FLIGHT CONDITION
a) CRUISE
b) APPROACH
c) LANDING
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RECORDED
PARAMETERS
AIRSPEED/MACH NUMBER
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
NORMAL ACCELERATION (or NORMAL LOAD
FACTOR)
PITCH CONTROLLER FORCE & POSITION
EVALUATION NOTES
A critical factor for this test is to obtain an accurate
trim before banking the simulated aeroplane to the
2C-25
requisite angles. Results may be shown either as a
time history or a series of snapshots.
If results are shown by a series of snapshots, the
steady state bank angles should be approximately 20
and 30 degrees for the approach and landing
configurations and 20, 30 and 45 degrees for the
cruise configuration. If conducting the test using
discrete bank angle snapshots, it may be necessary
for the pitch controller to be displaced in one direction
only from the neutral (trim) position so as to avoid
hysteresis effects which would almost certainly give
erroneous results. The results should be recorded
only once the aeroplane is stable at each new bank
angle.
Showing the test as a time history can provide
assistance to the evaluator by clearly showing the
effects of control system hysteresis and breakout
force. On no account must the stabiliser, flap, landing
gear or throttle levers be moved from the trim position.
The alternative method, with tolerances only on
elevator angle, applies to those aeroplanes using
control augmentation to compensate for the normal
stick-force-per-g characteristics. For some
aeroplanes, this may be partially the case up to
certain bank angles. Note that the force tolerance is
not applicable if the forces are generated solely by the
use of actual aeroplane hardware in the simulator.
controlled aeroplanes for each of the three flight
conditions, one for the normal control state and the
other for a non-normal control state.
TOLERANCES
PITCH CONTROLLER FORCE ±2.2 daN (5 Lbs) or
±10%
Alternative Method:
(applies to aeroplanes which do not exhibit stick force
per g characteristics)
ELEVATOR
±1o or ±10%
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2C-26
MANUAL TESTING
The principal purpose of this test is to determine the pitch controller force
required to maintain speed at a specified bank angle. Note that longitudinal
stability augmentation systems must be as the validation data and that no trim
change should be used during the banking manoeuvre or at the new bank angle.
The test may consist of a continuous time history with slowing increasing bank
angle, or a series of steady-state conditions at stabilised bank angles. The
results may be "snapshot" once the aeroplane has been stabilized at the required
bank angle and at the trim airspeed, or the data may call for a time history to be
run, in which case careful study should be made of the way that the flight test
data was gathered before attempting to replicate it in the simulator. Selecting
altitude freeze will help to provide sufficient time to set up each case. The engine
power setting must not be altered.
Note that with some computer-controlled aircraft the bank angle protection
functions can make it difficult to maintain some of the higher bank angles.
EXAMPLE
Figure 2c6-1 shows a time history for a longitudinal manoeuvrability (stick
force per g) test. The method is similar to that using ‘snapshots’, but the test is
run in a continuous manner instead of maintaining constant speed at two or
three discrete bank angles. Using a time history in this manner eliminates the
need to take special account of pitch controller hysteresis, but will be more
difficult to achieve an accurate result when flying the test manually.
2C-27
Figure 2c6-1
Example of Simulator Test Results for Longitudinal Manoeuvring Stability
2C-28
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TITLE
2c(7) - LONGITUDINAL STATIC STABILITY
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OBJECTIVE
STATIC LONGITUDINAL STABILITY
AEROPLANE.
DEMONSTRATION
From the prescribed trim conditions apply a
longitudinal control command to achieve a deviation
from the trimmed airspeed whilst maintaining wings
level. Use longitudinal control force to maintain a
steady-state condition at each of at least two speeds
above and two speeds below the initial trim speed.
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FLIGHT CONDITION
APPROACH
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RECORDED
PARAMETERS
AIRSPEED
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
FACTOR)
PITCH CONTROLLER FORCE & POSITION
EVALUATION NOTES
A critical factor for this test is once again to obtain an
accurate trim before displacing the longitudinal
controller either forward or aft to achieve the requisite
airspeeds. If conducting the test using discrete
airspeed snapshots, it may be expedient to attempt to
displace the pitch controller in one direction only from
2C-29
the neutral (trim) position so as to avoid hysteresis
effects which would almost certainly give erroneous
results. The results should be recorded only once the
aeroplane is stable at each new airspeed.
The stabiliser, flap, landing gear and throttle levers
should not be moved from the trim position. The
alternative method, with tolerances only on elevator
angle, applies to those aeroplanes using control
augmentation to compensate for the normal speed
stability characteristics.
The force tolerance is not applicable if the forces are
generated solely by the use of actual aeroplane
hardware in the simulator.
controlled aeroplanes, which may be for either the
normal control state or a non-normal control state.
TOLERANCES
PITCH CONTROLLER FORCE ±2.2 daN (5 Lbs) or
±10%
Alternative Method:
(applies to aeroplanes which do not exhibit speed
stability characteristics)
ELEVATOR
±1o or ±10%
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MANUAL TESTING
The principal purpose of this test is to determine the pitch controller force
required to maintain specific airspeeds without re-trimming. Each case starts
from a stable trimmed condition, then control column or pitch controller force is
applied to achieve and stabilise at the desired airspeeds. It is a consequence of
this type of manoeuvre that a rate of climb or descent will develop, so it is usual
to perform the test with the altitude frozen at the value specified in the data, and
this will not appreciably affect the results. Note that the longitudinal stability
augmentation systems must be as stated in the validation data and if the results
of the aeroplane flight test data are "snapshots", then a time history plot is not
necessary. The engine power and stabiliser settings must not be altered from
trim.
EXAMPLE
2C-30
An example of a more recent longitudinal static stability test is shown in Figure
2c7-1. In the past, this test has typically been accomplished using several
‘snapshots’ at speeds above and below the trim speed. Whilst this method may
still be valid, a time history such as the one below may better exemplify the
normal flying characteristics in this regime.
Figure 2c7-1
Example of Simulator Test Results for Longitudinal Static Stability
2C-31
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TITLE
2c(8) - STALL CHARACTERISTICS
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OBJECTIVE
LONGITUDINAL CONTROL POWER, LIFT,
PITCHING MOMENT AND STALL WARNING
INDICATIONS CONFORM TO THE AEROPLANE.
DEMONSTRATION
Conduct a wings-level stall entry at or near idle power
using longitudinal control to achieve a steady rate of
deceleration until reaching the minimum speed before
initiating recovery.
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FLIGHT CONDITION
a) SECOND SEGMENT CLIMB
b) APPROACH (or LANDING)
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RECORDED
PARAMETERS
AIRSPEED
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
FACTOR)
PITCH CONTROLLER FORCE (if reversible controls)
EVALUATION NOTES
The aeroplane flight test may or may not have been
performed using a consistent 1 kt/second deceleration
rate. This may not be critical, but in any case attempts
should be made to match the actual data rather than
perform the manoeuvre using classical techniques.
Nevertheless, the main criteria are whether the
2C-32
speeds for initial buffet, stick shaker and minimum
stall are accurate, though these parameters cannot be
taken in isolation from the technique used. If the
aeroplane demonstrates an abrupt nose down
tendency, or “g-break”, after full stall this characteristic
should be represented.
The modelling of stall characteristics through the stall
recovery requires incremental changes to the basic lift
and pitching moment data and the aerodynamic tail
efficiency. Wing flow separation occurs differently
during stall entry than does flow re-attachment during
stall recovery, and simulation of this effect requires the
addition of hysteresis increments to the basic data. To
model the hysteresis effects associated with
reattaching the separated flow following a stall, a
typical simulation model uses a simple first-order lag
filter function to lag angle of attack. As angle of attack
increases above the initial buffet angle of attack, an
initial buffet flag is triggered. As the angle of attack
decreases, the simulation begins calculation of the
lagged body angle of attack. The initial buffet flag
remains on until the lagged angle of attack decreases
below the angle of attack trip point for initial buffet (this
lagged angle of attack is only used in the calculation
of the initial buffet flag). The stall hysteresis
increments are a function of the difference between
aeroplane angle-of-attack and the re-attachment
angle-of-attack, and they return to zero when the wing
flow re-attaches. Thus the simulation of stall behaviour
uses special aaerodynamic models which are not
exercised in other flight regimes.
controlled aeroplanes for each of the two flight
conditions, one for the normal control state and the
other for a non-normal control state.
TOLERANCES
AIRSPEED
±3 Kts
(For Initial Buffet, Stall Warning and Stall Speeds)
controls:
PITCH CONTROLLER FORCE ±10% or ±2.2 daN
(5Lbs)
2C-33
(Prior to g-break only)
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MANUAL TESTING
Accurate stall warning test results are obtained by using classical aeroplane
'flight test' techniques. Thus ensure that the entry rate to the stall is close to 1
kt/sec, but take into account any significant deviations from this value visible on
the flight test results. The stick shaker speed may be defined from the flight test
data or it may have been obtained from another approved data source. The
aeroplane manufacturers may also provide the buffet and minimum stall speeds
in tabular form, though the test should be run as a time history. For aeroplanes
with stall protection systems, special care needs to be taken to ensure that the
published stall speeds are not those at which the stick pusher or other such
device operates, but the stall speed itself. Note that the stability augmentation
systems must be as stated in the validation data, and whilst the bank angle
tolerance has now been removed from the requirements, it will the results if the
bank angle differs greatly from flight test data.
EXAMPLE
The reason why the initial airspeed trace in Figure 2c8-1 suddenly ‘jumps’ from
130 knots down to 119 knots is because the wind speeds were inadvertently
omitted from the initialisation process This needs to be corrected in this test,
along with the noticeable difference in initial altitude. The overall effect of both
these errors is very little however, proving that, at least for simulation testing
purposes, the precise speed at which the aeroplane is trimmed ready to begin
a stall manoeuvre does not necessarily have much bearing on the outcome of the
test. The assumption is made with this result that the stall warning, initial buffet
and minimum speeds are all recorded elsewhere, otherwise the time histories
would benefit from markers to show the values.
2C-34
Figure 2c8-1
Example of Simulator Test Results for Stall Characteristics
2C-35
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TITLE
2c(9) - PHUGOID DYNAMICS
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OBJECTIVE
PHUGOID DYNAMIC CHARACTERISTICS
DEMONSTRATION
From the prescribed trimmed condition, excite the
phugoid mode by applying longitudinal control in one
direction in order to change airspeed by approximately
10 kt and then releasing. Record the pertinent
longitudinal parameters to enable a mathematical
analysis of the oscillations to be carried out for
comparison with aeroplane data. The test should
proceed hands-off for three full cycles or long enough
to determine the time to one-half or double amplitude.
It is extremely important that the aeroplane test be
conducted with little or no atmospheric turbulence.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
AIRSPEED
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
FACTOR)
EVALUATION NOTES
The purpose of this test is not to attempt to obtain a
perfect match of all plotted parameters for the entire
2C-36
length of the manoeuvre. For any long-duration
hands-off test, the small errors which are inevitably
present between the flight test data and the simulator
dynamics will almost certainly accumulate and cause
some of the parameters (e.g. altitude) to be noticeably
different by the time the test is complete. These types
of errors can usually be minimised by careful use of
trim parameters such as pitch rate, rate of climb, etc.,
but the simulator should never be artificially controlled
during this (or any hands-off) manoeuvre. The items
to be checked are the periodic time of the phugoid
oscillation and also the time to half or double
amplitude. For most modern jet transport aeroplanes
the oscillation will be damped, but there may be
exceptions to this. The period itself will typically be of
the order of several tens of seconds. The
mathematical analysis will probably be carried out
within the simulator computer, but it may be
worthwhile performing a cursory check of the value so
obtained. Due to the long-term accumulation of small
errors discussed above it is quite likely that any two
sets of calculations of period and damping for this test
will differ by a small amount.
controlled aeroplanes - for a non-normal control state.
TOLERANCES
PERIOD
and either
TIME TO HALF OR DOUBLE
AMPLITUDE
or
DAMPING RATIO
±10%
±10%
±0.02
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MANUAL TESTING
Typical procedures for this test call for the aeroplane to be set up in a level flight
trimmed condition, and then for the pilot to either pull or push the pitch controller
to a given value for a specific period (usually a few seconds) to reduce or
increase the airspeed. Once the time has expired and the target airspeed
achieved (ideally simultaneously) the pitch controller is released to the neutral
position and the remainder of the test duration is ‘hands-off’. A minimum of three
2C-37
full cycles is typically needed to determine the time to half (or double amplitude).
Parameters are recorded as above and the calculations and analysis will
probably be carried out within the simulator computer. The exact duplication of
the flight test control inputs should not be strictly necessary, though as always
the closer they can be matched the easier it may be to interpret the results. It
may be that some small lateral adjustments using the roll controller are needed
during the test in order to maintain a wings level configuration. Alternatively some
pilots have found it useful to use slight pressure on the rudder pedals to achieve
the same end. Ensure that the simulator stability augmentation systems are
configured as stated in the validation data.
EXAMPLE
The deviation in the pressure altitude value in Figure 2c9-1 is not necessarily
significant from the point of view of the analysis, but it would be of very doubtful
use in order to ascertain the period and time to half amplitude. Typically,
simulator autotest systems make allowances for the data, and permit choice of
one of several longitudinal parameters. Here, the pitch angle looks the most
promising parameter to compare with the aeroplane data. The overall impression
of the result is not of a high standard, but in this particular case there had been
many industry comments from several simulator manufacturers who had been
struggling to replicate the validation data results.
2C-38
Figure 2c9-1
Example of Simulator Test Results for Phugoid DYnamics
2C-39
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TITLE
2c(10) - SHORT PERIOD DYNAMICS
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OBJECTIVE
SHORT PERIOD DYNAMIC CHARACTERISTICS
DEMONSTRATION
From the prescribed trim condition, excite the short
period mode by applying a brief (one second or less)
longitudinal control input in one direction then allowing
the aeroplane to freely respond.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
AIRSPEED
PRESSURE ALTITUDE
PITCH ANGLE
PITCH RATE
ANGLE OF ATTACK
ELEVATOR ANGLE
STABILISER ANGLE
BANK ANGLE
FACTOR)
EVALUATION NOTES
Because of the relatively short duration of this test it is
not practical to perform a mathematical analysis of the
oscillatory nature of the short period pitch response.
In any case, for large transport aeroplanes the
oscillation tends to be very highly damped, making
analysis even more difficult. However, the fact that the
test is of short duration (typically around 20 seconds
or so) often makes it simpler to obtain a good match
of the parameters plotted in the time history.
2C-40
Duplication of the control inputs is important for this
test, but due to the rapid nature of these inputs it may
be difficult to discern from the plotted data whether the
elevator position has been stimulated accurately.
and the other for a non-normal control state.
TOLERANCES
PITCH ANGLE
(or PITCH RATE
NORMAL ACCELERATION
±1.5o
±2o/Sec)
±0.1 g
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MANUAL TESTING
The short period mode is typically excited with a pitch controller pulse or double
input. The test is usually begun with the aeroplane perfectly trimmed in level flight
(though the data should be carefully checked for any slight pitch rate, climb rate,
etc.). The control column or longitudinal controller is then displaced minimally but
rapidly to induce the short period pitching oscillation whilst the relevant
parameters are being recorded. Because of the very small movement required,
which must be tailored to achieve the desired result, this case should be
practised beforehand whilst the control parameters are being monitored at an
engineering terminal or workstation. The simulated stability augmentation
systems must be configured as stated in the validation data.
EXAMPLE
Figure 2c10-1 illustrates the point that, being of short duration, it is often easy to
obtain a good result for this test. The plots also serve to illustrate quite well the
type of oscillation to be expected from a jet transport when the short period has
been excited. This test does not suffer from the susceptibility to accumulative
errors that are inherent in, for example, the phugoid test.
2C-41
Figure 2c10-1
Example of Simulator Test Results for Short Period Dynamics
2C-42
SECTION 2d
LATERAL DIRECTIONAL
2d(1)
Minimum Control Speed, Air (Vmc or Vmcl), per
Applicable Airworthiness Standard - or - Low Speed
Engine Inoperative Handling Characteristics in the Air
2d(2)
Roll Response (Rate)
2d(3)
Step Input of Cockpit Roll Controller
2d(4)
Spiral Stability
2d(5)
Engine Inoperative Trim
2d(6)
Rudder Response
2d(7)
Dutch Roll (Yaw Damper OFF)
2d(8)
Steady State Sideslip
2D-1
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TITLE
2d(1) - MINIMUM CONTROL SPEED, AIR (VMCA OR
VMCL), PER APPLICABLE AIRWORTHINESS
STANDARD OR LOW SPEED ENGINE
INOPERATIVE HANDLING CHARACTERISTICS IN
THE AIR
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OBJECTIVE
DIRECTIONAL CONTROL AT MAXIMUM
ASYMMETRIC THRUST AT THE LOW SPEED IN
THE AIR CONFORMS TO THE AEROPLANE.
DEMONSTRATION
For a time-history demonstration, start from the
prescribed initial condition, and slowly decelerate with
idle power on one engine and maximum takeoff power
on the other engine(s) using full rudder control and
lateral control as necessary to maintain heading until
reaching the minimum speed. The Vmca test is an
aeroplane certification requirement to determine the
minimum airspeed at which heading can be
maintained with not more than 5 degrees of bank with
a critical engine inoperative and with maximum takeoff
power on the remaining engine(s). When
demonstrating this test it is acceptable to use idle
power to simulate the inoperative engine. During
actual flight testing, heavy buffet may be encountered
before reaching a bank angle of 5 degrees requiring
that the test be terminated. The data may be shown
as a time history or as a series of snapshot tests.
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FLIGHT CONDITION
TAKEOFF or LANDING (whichever is most critical in
the aeroplane)
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RECORDED
PARAMETERS
2D-2
PRESSURE ALTITUDE
AIRSPEED
PITCH ANGLE
PITCH RATE
ELEVATOR ANGLE
BANK ANGLE
RUDDER ANGLE
HEADING ANGLE
SIDESLIP ANGLE
AILERON ANGLE
SPOILER ANGLES
YAW RATE
ROLL RATE
EVALUATION NOTES
Ideally, the regulatory authorities tend to prefer this
test to be run dynamically rather than using trim points
at progressively lower speeds. Depending on whether
the validation data are available as a time history or a
declaration of the VMCA, it may be best to run a
classical VMCA case by trimming with maximum thrust
asymmetry at a speed a few knots in excess of the
declared VMCA and then using all primary controls to
reduce the speed whilst maintaining heading at a bank
angle of no more than five degrees (failed engine
high). The Air Minimum Control Speed is then the
speed at which heading can no longer be maintained
using full rudder control. However, the validation data
may have been obtained from a test where heavy
buffet or stall speed was reached during the
deceleration before a steady-state bank angle of five
degrees could be achieved. In other cases, the
validation data are presented in the form of a time
history, though it may not be for a classical VMCA test,
and this is acceptable provided the thrust asymmetry
is sufficient to determine that the simulator handling
qualities are equivalent to the aeroplane at low
speeds. Some data is presented as a snapshot at a
defined point close to (or at the declared) VMCA.
Obviously the simulator needs to represent the
aeroplane data as published, in whatever form that is.
Note that for computer controlled aeroplanes this test
may be run in either the normal or non-normal control
2D-3
state.
TOLERANCES
AIRSPEED
±3 Kts
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MANUAL TESTING
Unless the pilot flying the test (in either the aeroplane or the simulator) is very
experienced and well practised at performing VMCA manoeuvres, it is likely that
accurate flying of a classical VMCA test will take several attempts to get right. The
test should ideally start with the simulated aeroplane trimmed for level flight at
around 1.3Vstall. The appropriate engine(s) can then be set to idle thrust or shut
down (depending on the method used to acquire the validation data) and the
thrust on the operating engine(s) increased to takeoff level whilst heading is
maintained and the bank angle kept at or below five degrees. The airspeed will
then need to be reduced using elevator control only. It is quite possible that the
aeroplane VMCA is below the stall speed and if this is so, the simulator should
reflect this also. If the data is presented in the form of a snapshot, the simulated
aeroplane should be trimmed at this point and the relevant parameters recorded.
It still may be difficult to fly, but the results, once obtained, are usually easier to
interpret.
EXAMPLE
The result shown in Figure 2d1-1 is one section of a snapshot version of this test.
Obviously there have been three previous sections, each demonstrating a trim
condition at progressively lower airspeed, to show the increase in rudder required
to maintain heading.
2D-4
Figure 2d1-1
Example of Simulator Test Results for Minimum Control Speed, Air
Figure 2d1-2 is an example of a time history implementation of this test. To fully
comply with the requirement the plots shown below would need to be
supplemented by others, especially rudder angle and the key engine parameters,
but the roll rate conveys how difficult it can be to fly this test manually in either the
aeroplane and by definition therefore the simulator. There are some jet transport
aeroplanes for which Vmca may be at or below the stall speed for a given
condition, so other means may need to be investigated in the data if the
requirement is to be fulfilled.
2D-5
Figure 2d1-2
Example of Simulator Test Results for Minimum Control Speed, Air - Time History
2D-6
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TITLE
2d(2) - ROLL RESPONSE (RATE)
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR ROLLRATE RESPONSE TO ROLL CONTROL INPUT
DEMONSTRATION
From a wings-level trim condition, execute a roll
manoeuvre using lateral control. The lateral control
input should consist of a rapid displacement to
approximately one-third of maximum roll controller
travel, holding this value at least until a steady roll rate
is established. This test may be combined with the
Step Input of Cockpit Roll Controller test, 2d(3).
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FLIGHT CONDITION
a) CRUISE
b) APPROACH or LANDING
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
BANK ANGLE
PITCH ANGLE
ROLL CONTROLLER FORCE (if reversible controls)
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER ANGLE
ROLL RATE
ELEVATOR ANGLE
EVALUATION NOTES
This is usually a relatively simple test to accomplish,
but it should be borne in mind that a fairly small
2D-7
deviation in the roll rate may result in a large
discrepancy between the bank angles at the end of
the manoeuvre, even though the test could still be well
within tolerance. The test should be driven from the
cockpit roll controller and not by directly driving the
ailerons and/or spoilers, though use of the latter
method as well may in certain cases be valid backup.
Longitudinal parameters such as airspeed and pitch
angle should be maintained close to the validation
data values, but this can be achieved automatically
using closed-loop drivers or by manipulating the pitch
controller as needed when running manually. Sources
of asymmetry other than roll control displacement,
such as the rudder or engines, should be checked that
they represent the aeroplane as it was during the flight
test manoeuvre, though the test will probably have
been carried out with the yaw damper on. If the test is
run for both right and left wing down cases, these
possible asymmetries will be better quantified.
Occasionally an aeroplane may exhibit slight intrinsic
asymmetry, (e.g., left and right flap positions which are
slightly different), but it is not expected that the
simulator should replicate this type of asymmetry.
Since the initial portion of the test is identical to the
method for performing test 2d(2), the test for 2d(2) and
2d(3) may be performed as a single manoeuvre,
though two separate tests should still be present in the
QTG.
TOLERANCES
ROLL RATE
±10% or ±2o/Sec
controls:
ROLL CONTROLLER FORCE ±10% or ±1.3 daN (3
Lbs)
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MANUAL TESTING
The normal procedure for this test calls for the pilot to establish the simulator in
a trimmed condition with symmetrical engine power and at an initial bank angle
appropriate to the test aeroplane, which may be wings level or it may be at an
initial bank angle of say, 30 degrees. After a few seconds has been allowed to
2D-8
confirm stability, apply a roll control input to match that of the aeroplane. Use
longitudinal control as necessary to maintain a pitch angle as closely as possible
to that of the aeroplane. When the required final bank angle has been achieved,
return the roll controller to neutral. Ideally, the roll controller deflection will be
around one third of maximum, but as always this will be dependent on the
aeroplane data.
EXAMPLE
The first plot
shown in Figure
2d2-1 shows the
peak roll rate out
of tolerance at 43
seconds, even
though the wheel
position and
airspeed closely
follow
the
aeroplane data.
The reason, in
this case, was
that the aeroplane
wheel/aileron/spo
iler relationship
was slightly misrigged during the
flight test data
gathering. The
only solution for
the simulator was
to run a second
test,
using
ailerons and
spoilers as the
d r i v i n g
parameters, and
present both sets
of results in the
QTG.
Figure 2d2-1
Example of Simulator Test Results for Roll Response (Cruise
Condition)
2D-9
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TITLE
2d(3) - STEP
CONTROLLER
INPUT
OF
COCKPIT
ROLL
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OBJECTIVE
ROLL CHARACTERISTICS AFTER A STEP ROLL
CONTROL INPUT HAS BEEN REMOVED
DEMONSTRATION
After establishing a constant roll rate as described for
the Roll Response (Rate) test 2d(2), at a bank angle
of about 20 to 30 degrees, abruptly return the roll
controller to neutral, and then allow at least 10
seconds of free response. This test may be combined
with the Roll Response (Rate) Approach or Landing
test, 2d(2b).
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FLIGHT CONDITION
APPROACH or LANDING
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
ROLL CONTROLLER FORCE (if reversible controls)
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER ANGLE
ROLL RATE
BANK ANGLE
PITCH ANGLE
ELEVATOR ANGLE
EVALUATION NOTES
Whilst this test, like test 2d(2), will give a good
2D-10
indication of the roll response, its main purpose is to
determine that the simulator is like the aeroplane
during the free response which follows the removal of
the lateral control input, especially for any tendency for
roll overshoot immediately after the controller is
released. Thus it tends to be performed in a similar
manner to the Phugoid test, 2c(9) or the Dutch Roll
tests, 2d(7) in that a primary control is displaced briefly
and then returned to its neutral position to allow the
simulator free response to be recorded and the degree
to which the roll angle is damped to be examined.
The flight test pitch angle should be maintained fairly
closely so that the airspeed does not deviate
significantly, and this again may be achieved by use of
an automatic closed-loop controller or by manipulating
the pitch controller as needed when running manually.
The yaw damper must be as the aeroplane. Since the
initial portion of the test is identical to the method for
performing test 2d(2b), the test for 2d(2b) and 2d(3)
may be performed as a single manoeuvre.
Note that for computer controlled aeroplanes there
must be a test for both the normal and non-normal
control states.
TOLERANCES
BANK ANGLE
±10% or ±2o
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MANUAL TESTING
Usual procedures call for the simulator to be established in a trimmed flight
condition with symmetrical engine power and zero bank angle. A roll control input
to match that of the aeroplane is then rapidly applied and removed such that the
control is returned to its neutral position. The free response of the simulated
aeroplane is then observed and plotted for a period of typically 15 to 20 seconds
to obtain the bank angle characteristics.
EXAMPLE
Figure 2d3-1 shows a slight roll overshoot, as the roll controller input was
returned fully to neutral between 14 and 15 seconds. The plots are another
example of the use of tolerance bands and show the test to pass, except that
2D-11
they have been applied to roll rate, which was the parameter required in the
previous version of the ICAO Manual. Under the new requirements the same test
would fail, as it is not within the requisite 2o or 10% of bank angle. Nor does it
allow for at least 10 seconds of free response after the control input has been
removed.
2D-12
Figure 2d3-1
Example of Simulator Test Results for Step Input of Cockpit Roll
Controller
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TITLE
2d(4) - SPIRAL STABILITY
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OBJECTIVE
THE DYNAMIC LATERAL/DIRECTIONAL
CHARACTERISTICS IN THE SPIRAL MODE
DEMONSTRATION
From an established bank angle of approximately 30
degrees, release the lateral controller and allow the
aeroplane to freely respond until the bank angle
reaches approximately 10 degrees if decreasing, or
approximately 45 degrees if increasing. In either case
the time interval for the free response need not
exceed one minute. The test should be conducted in
both directions. As an alternative, the test may be
conducted by establishing a steady bank angle of
approximately 30 degrees, then applying the lateral
control required to maintain that bank angle. The
intent of this test is to establish the spiral mode
characteristics of the unaugmented aeroplane, so the
yaw damper should be turned OFF.
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FLIGHT CONDITION
a) CRUISE
(Tests must be in both directions)
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
BANK ANGLE
PITCH ANGLE
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
2D-13
RUDDER ANGLE
ROLL RATE
ELEVATOR ANGLE
EVALUATION NOTES
Like test 2d(3), the spiral stability test has as its main
purpose the determination that the simulator is like
the aeroplane during the free response which follows
the removal of a lateral control input. For the spiral
mode, however, the control is removed slowly once
the aeroplane is stable at a given bank angle and the
tendency (if any) of the simulated aeroplane to either
continue banking (unstable) or to return to wings level
(stable) is recorded. The flight test pitch angle should
be maintained fairly closely so that the airspeed does
not deviate significantly, and this again may be
achieved by use of an automatic closed-loop controller
or by manipulating the pitch controller as needed
when running manually. The yaw damper must be as
the aeroplane, but will usually be off. Bearing in mind
that most modern jet transport aeroplanes tend to
exhibit a relatively neutral spiral mode, it is very
important that any slight asymmetries present in the
flight test data (such as rudder or engine thrust
differences) are properly recognised during analysis of
the spiral mode test. Therefore this test should be
performed in both directions. The duration of the test
should be at least 20 seconds after the free response
period has commenced.
As an alternative, the test may be performed by
establishing a steady bank angle of about 30 degrees,
then simply maintaining the bank with a constant roll
control input. For the tolerance on bank angle, “correct
trend” means that the simulator should exhibit the
same tendency as the aeroplane to either increase or
decrease the bank angle during a free response, or
require the same lateral controller direction for the
alternative method.
Note that for computer controlled aeroplanes the test
must be for the non-normal control state.
TOLERANCES
2D-14
BANK ANGLE
Correct Trend and
±2o or ±10% bank in
20 sec
If alternative test is used:
AILERON ANGLE
Correct Trend and ±2o
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MANUAL TESTING
The pilot is usually required to establish the simulator in a trimmed flight condition
with symmetrical engine power and zero bank angle. Particular care should be
exercised to precisely trim the simulator in wings level, stable flight with
symmetrical power, since initial trim strongly affects the test result. Smoothly roll
the simulator to the bank angle of the test aeroplane, and stabilise airspeed and
bank angle. Return the roll control slowly to neutral and allow a free response of
the simulator in roll. Pitch control may be applied to match the pitch angle of the
aeroplane. If the aeroplane test had the yaw damper on, verify that the rudder in
the simulator is very close to the aeroplane.
For the alternative test method, bank the aeroplane smoothly to the requisite
angle (should be about 30 degrees) and maintain that angle using roll control.
Airspeed must also be maintained using pitch control.
EXAMPLE
The result in Figure 2d4-1 is essentially a good one, and again uses tolerance
banding to prove the simulator meets the requirements. The absence of the initial
oscillation in roll which is present in the aeroplane data has no significant effect
on the outcome of the test.
Figure 2d4-1
Example of Simulator Test Results for Spiral Stability
2D-15
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TITLE
2d(5) - ENGINE INOPERATIVE TRIM
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OBJECTIVE
THE EFFECTS OF DIRECTIONAL TRIM DURING AN
ENGINE INOPERATIVE MANOEUVRE CONFORMS
TO THE AEROPLANE.
DEMONSTRATION
Establish a steady-state engine-out trim condition
using techniques similar to that for which a pilot is
trained to trim an engine failure condition, typically
using little or no lateral trim or maintaining near wingslevel flight. For the climb condition, the thrust should
be set to takeoff power. For the approach or landing
test, the power should be set for thrust for level flight
with the engine inoperative. The inoperative engine
may be simulated by setting idle thrust. The data may
be shown as a series of snapshot tests.
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FLIGHT CONDITION
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RECORDED
PARAMETERS
2D-16
PRESSURE ALTITUDE
AIRSPEED
BANK ANGLE
PITCH ANGLE
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER TRIM POSITION
RUDDER ANGLE
ROLL RATE
ELEVATOR ANGLE
EVALUATION NOTES
These two tests are designed to ascertain that the
simulated rudder allows the same degree of control
power as the real rudder does in the aeroplane. They
are akin to the longitudinal trim tests found in 2c(5)
and as such tend to be quite straight forward to
perform. Whilst the main tolerance parameters do not
include items such as airspeed, pitch angle or rate of
climb, it is worth checking these values as well so as
to be sure that the flight condition is correct according
to the aeroplane data. Slightly different results can be
obtained depending on the engine-out trim technique
used, so the data should be carefully studied so that
the flight test manoeuvre is repeated accurately and
preferably in a manner similar to that for which a pilot
is trained to trim an engine failure condition. The tests
should be run dynamically, but the results can be
confirmed using a snapshot technique once the
simulated aeroplane has been established in a steady
state condition for a few seconds. It is important to
match the aeroplane bank angle accurately as there
can be a large change of rudder and sideslip with
bank angle for a stabilised constant heading condition.
The ICAO Manual states that for the approach or
landing test the power should be set to thrust for level
flight (i.e. level flight for the engine inoperative
condition).
TOLERANCES
RUDDER ANGLE
(or TAB ANGLE
or EQUIVALENT PEDAL)
SIDESLIP ANGLE
±1o
±1o
±2o
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MANUAL TESTING
It is usually a requirement to establish the simulator in a constant heading
stabilised flight condition at constant airspeed using pedal (or rudder trim) and roll
control as necessary to maintain the aeroplane recorded trimmed bank angle,
with the engines set to the aeroplane conditions. The inoperative engine may be
shutdown or set to idle power, and the ‘live’ engine set to a high thrust level for
the second segment condition to ensure that sufficient thrust is used to require
2D-17
substantial yaw control. For the approach condition, thrust will probably have
been set for level flight, but the validation data should be followed whichever
technique is used. Use lateral and directional trim to minimise the pilot control
forces as necessary to match the aeroplane conditions. Maintain this stabilised
flight condition for several seconds before verifying the results.
If the test involves a climb or descent, it may be helpful for the manual test to
start at a lower or higher altitude, respectively, so as to be stabilised at the
required pressure altitude of the aeroplane data.
EXAMPLE
The result shown in Figure 2d5-1 is another example of a snapshot test, showing
all the required parameters, and including a pass/fail assessment on engine
power lever angle as well as the requisite rudder deflection and sideslip angle.
While it is sometimes useful to mention items such as ‘Parameters with no a/c
data’, some explanation as to the meaning of this statement should be included
in the QTG.
Figure 2d5-1
Example of Simulator Test Results for Engine Inoperative Trim
2D-18
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TITLE
2d(6) - RUDDER RESPONSE
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OBJECTIVE
DIRECTIONAL RESPONSE FROM RUDDER
CONTROL MOVEMENTS CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Starting from a wings-level trimmed condition, initiate
a rapid rudder input to approximately 25 per cent of
full rudder pedal travel. One method which can
achieve a clean rapid rudder input is to use the rudder
trim system to command the desired rudder position
while using rudder pedal to counter the trim command.
Then abruptly release the pedal and allow the
aeroplane to freely respond. The response may be of
short duration, usually not to exceed 15 seconds or 30
degrees of bank. The test should be performed both
with stability augmentation ON and OFF.
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FLIGHT CONDITION
a)
APPROACH
or
LANDING, STABILITY
AUGMENTATION ON
b) APPROACH or LANDING, STABILITY
AUGMENTATION OFF
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
BANK ANGLE
PITCH ANGLE
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER ANGLE
ROLL RATE
2D-19
ELEVATOR ANGLE
EVALUATION NOTES
Of primary consideration in this test is the yaw rate
reponse associated with the initial pedal movement.
With the yaw damper disengaged, the simulator (like
the aeroplane) will tend to enter a dutch roll oscillation
after a few seconds, but it is not the dutch roll
characteristics which are under scrutiny, hence the
test duration need not exceed about 15 seconds after
completion of the rudder movement, or when the bank
angle reaches 30 degrees. The test should ideally be
driven through the rudder pedals, so as to confirm the
relationship between control position and control
surface position. The amount of rudder deflection
need not be excessive, but should typically be limited
to around 25% of full rudder pedal throw.
For computer controlled aeroplanes, a test should be
provided in both the normal and the non-normal
control state. Also, even for conventionally controlled
aeroplanes, a separate test should be supplied for
stability augmentation (usually a yaw damper and/or
turn coordinator) ON as well as with the system turned
OFF.
TOLERANCES
YAW RATE
±2o/Sec or ±10%
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MANUAL TESTING
The data will most likely require the pilot to establish the simulator in a steady
level flight condition with symmetric engine power in the configuration specified.
Apply a rapid rudder pedal movement to match that of the aeroplane, but other
lateral/directional control inputs should not be used, nor should there be any
change away from the initial trim or power settings. Ideally there should follow a
free response of the simulator, but if the data deems it necessary the pilot should
maintain the rudder pedal (and thus the rudder) as close to the aeroplane
recorded time history as possible. Typical test duration will be around 20
seconds. The test technique is similar for both settings of the stability
augmentation (e.g. yaw damper) system.
2D-20
EXAMPLE
Figure 2d6-1 is included because it was taken from the beginning of a Dutch roll
test. The duration however, is too short to determine the difference in
characteristics between the response with yaw damper off and on. Figure 2d6-2
has the yaw damper engaged. This test was driven with the rudder pedals (not
shown) and so allows the yaw damper simulation to be properly exercised.
Figure 2d6-1
Example of Simulator Test Results for Rudder Response (Yaw Damper
Off)
2D-21
Figure 2d6-2
Example of Simulator Test Results for Rudder Response (Yaw Damper On)
2D-22
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TITLE
2d(7) - DUTCH ROLL (YAW DAMPER OFF)
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OBJECTIVE
TO
DEMONSTRATE
THAT
THE
LATERAL/DIRECTIONAL DYNAMIC STABILITY
CHARACTERISTICS OF THE SIMULATOR IN THE
DUTCH ROLL MODE CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Starting from a wings-level trimmed condition, initiate
a 'dutch roll' oscillation using a rapid pedal input in
both directions, then allowing the aeroplane to
respond freely for at least six cycles of the oscillation.
Stability augmentation for the lateral-directional axes
should be turned off. Record the pertinent lateraldirectional parameters to enable a mathematical
analysis of the oscillations to be carried out for
comparison with aeroplane data.
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FLIGHT CONDITION
a) CRUISE
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RECORDED
PARAMETERS
BANK ANGLE
PITCH ANGLE
PRESSURE ALTITUDE
AIRSPEED
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER ANGLE
ROLL RATE
ELEVATOR ANGLE
2D-23
EVALUATION NOTES
Of primary consideration in this test are the dutch roll
characteristics excited by the rudder pedal
movements, which will usually be a 'doublet', which
requires the pedals to be depressed the same amount
in each direction before releasing. This technique
should then result in the oscillations being roughly
symmetrical about zero bank angle, making the results
easier to assess. With the rudder pedals returned to
the neutral position, the oscillations will be free,
allowing mathematical analysis after about 6 cycles,
with a typical dutch roll period in the order of 5 to 10
seconds. The test should ideally be driven through the
rudder pedals, so as to confirm the relationship
between control position and control surface position,
but any automatic rudder pedal driver should be
released after the pedals have been returned to
neutral. The amount of rudder deflection need not be
excessive, and like test 2d(6) should typically be
limited to around 25% of full rudder pedal throw.
For computer controlled aeroplanes, this test should
be provided in the non-normal control state.
Some notes on the analysis of dynamic time histories
are given in Appendix B.
TOLERANCES
PERIOD
TIME DIFFERENCE
BETWEEN PEAKS OF
BANK AND SIDESLIP
EITHER
TIME TO HALF OR DOUBLE
AMPLITUDE
OR
DAMPING RATIO
±0.5 Sec or ±10%
±20% or ±1 Sec
±10%
±0.02
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MANUAL TESTING
The validation data will most likely require the pilot to establish the simulator in
a steady level flight condition with symmetric engine power in the configuration
specified and with the yaw damper switched off. Apply a rapid rudder pedal
movement (or doublet) to match that of the aeroplane, but other lateral/directional
2D-24
control inputs should not be used, nor should there be any change away from the
initial trim or power settings. There should follow a free response of the simulator
for approximately 60 seconds (at least 6 Dutch Roll cycles) with no control inputs
except as necessary to maintain the approximate flight test pitch angle.
It is desirable, but may not be necessary, to get an exact match of the aeroplane
rudder time history, since the purpose here is to check the dutch roll dynamic
characteristics - period, damping and bank angle to sideslip data.
It will be necessary to have a rudder input that results in the same general bank
trend as the aeroplane, following the rudder pedal release. The rudder pedal
input to excite the dutch roll is normally a doublet; one pedal in for a period of
time, then release, followed by the other pedal in for the same period of time,
then release. The general bank trend - either an oscillation with the average
drifting right wing down, an oscillation with the average drifting left wing down, or
an oscillation with the average near zero - can be controlled by the symmetry of
the doublet. A bank angle trend that clearly differs from the data may well result
in airspeed and angle of attack deviations, both of which can significantly affect
dutch roll dynamic characteristics.
EXAMPLE
The criteria for passing the dutch roll test does not have to include the precise
matching of all parameters such as roll rate and bank angle, etc., as the result in
Figure 2d7-1 illustrates. The results as shown passes, but this may not be
immediately obvious until the mathematical analysis is carried out. See Appendix
B for more details about such analysis.
2D-25
Figure 2d7-1
Example of Simulator Test Results for Dutch Roll
2D-26
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TITLE
2d(8) - STEADY STATE SIDESLIP
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OBJECTIVE
EXHIBITS THE CORRECT INTER-RELATIONSHIP
OF STEADY STATE LATERAL/DIRECTIONAL
FLIGHT CHARACTERISTICS IN CONFORMANCE
WITH THE AEROPLANE.
DEMONSTRATION
Establish a steady-state sideslip condition using a
constant rudder pedal position and sufficient lateral
control to maintain a constant heading. The steadystate sideslip condition should be performed for at
least two rudder pedal displacements, one of which
should be near the maximum available rudder
position. For propeller-driven aeroplanes, the tests
should be conducted in each direction. The data may
be shown as a time history or a series of snapshot
tests.
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FLIGHT CONDITION
APPROACH or LANDING
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RECORDED
PARAMETERS
PRESSURE ALTITUDE
AIRSPEED
BANK ANGLE
PITCH ANGLE
AILERON ANGLE
SPOILER ANGLES
SIDESLIP ANGLE
YAW RATE
RUDDER ANGLE
ROLL RATE
ELEVATOR ANGLE
2D-27
ROLL CONTROLLER FORCE (If reversible controls)
RUDDER PEDAL FORCE (If reversible controls)
EVALUATION NOTES
The usual way in which this test is run is to begin by
configuring the aeroplane for trimmed level flight, then
applying a specified rudder pedal position and holding
it steady whilst maintaining heading using the control
wheel or roll controller. The sideslip angle thus
developed is probably the most critical parameter,
since a small deviation in this value can result in a
significant difference in bank angle and/or roll
controller angle. Several rudder deflections should be
used (at least two, preferably three or four), one of
which should be near the maximum available rudder
deflection. For propellor driven aeroplanes the test
should be performed with the deflections both to the
left and to the right. Some flight test data may be
presented such that there are inconsistencies in the
aeroplane results. If this is the case then it may be
necessary to take an average value or to carefully
scrutinise the values so that the best data set is
utilised. During these tests in the aeroplane, it is
possible for the wing fuel to move (usually inboard on
the upper wing), which gives a lateral centre of gravity
movement and a roll moment. This affects the aileron
angle required for trim. Care must be taken to verify
this in the aeroplane data and subsequently in the
simulator.
It is important to have a very well stabilised flight state
for each of the rudder angles specified. This can be
verified by checking that the rate terms are very small
for each of the snapshots. If they are not the test must
be repeated for the appropriate condition. If sufficient
data are available, a plot of the aeroplane parameters
versus rudder angle may be included. This allows
overplotting of the simulator results and allows use of
manual test results which may not have the exact
specified rudder.
TOLERANCES
2D-28
For a given rudder position:
BANK ANGLE
SIDESLIP ANGLE
AILERON
±2o
±1o
±10% or ±2o
SPOILER or EQUIVALENT
ROLL CONTROLLER
POSITION or FORCE
±10% or ±5o
controls:
ROLL CONTROLLER FORCE ±10% or ±1.3daN (3
Lbs)
RUDDER PEDAL FORCE
±10% or ±2.2daN (5
Lbs)
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MANUAL TESTING
The simulator should first be established in steady level flight with symmetric
engine power in the configuration and flight condition specified by the flight test
data. Whilst maintaining constant airspeed, apply a rudder deflection appropriate
to the validation data (using rudder pedal or rudder trim - whichever was
employed in the aeroplane) and use roll control to stabilise the simulator at a
bank angle required to hold a constant heading.
It may be possible or helpful to make use of the autopilot, in airspeed and
heading hold, to assist in this manoeuvre, though usually these tests are not
difficult to fly, and are of relatively short duration.
Once the steady sideslip has been achieved, acquire a snapshot of the stabilised
condition. Repeat for at least two rudder angles, ensuring that one of these is
near the maximum allowable rudder, as there is significant interest in
characteristics at large sideslip angles.
EXAMPLE
The result shown in Figure 2d8-1 shows a magnificent array of passes against
all the required parameters, but this is not strictly necessary since this condition
is clearly for the initial wings level (symmetrical) trim (with zero rudder deflection).
Other subsequent conditions no doubt show the true status of the test results.
2D-29
Figure 2d8-1
Example of Simulator Test Results for Steady State Sideslip
2D-30
SECTION 2e
LANDINGS
2e(1)
Normal Landing
2e(2)
Minimum Flap Landing
2e(3)
Crosswind Landing
2e(4)
2e(5)
Autopilot Landing (if applicable)
2e(6)
All Engine Autopilot Go Around
2e(7)
One-Engine-Inoperative Go-Around
2e(8)
Directional Control (Rudder Effectiveness) with
Reverse Thrust (Symmetric)
2e(9)
Directional Control (Rudder Effectiveness) with
Reverse Thrust (Asymmetric)
2E-1
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TITLE
2e(1) - NORMAL LANDING
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OBJECTIVE
NORMAL LANDING CHARACTERISTICS
DEMONSTRATION
Perform a normal manual or automatic approach and
landing. Record data from at least 61 m (200 ft)
altitude through nose gear touchdown. For aeroplanes
with more than one certified landing flap, two tests
must be shown, each at a different flap position. One
test must be near the maximum landing weight, and
one at a light or medium weight.
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FLIGHT CONDITION
a) LANDING, near maximum landing weight
b) LANDING, light or medium landing weight
(The tests should include two normal landing flaps, if
applicable)
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RECORDED
PARAMETERS
2E-2
ALTITUDE
AIRSPEED
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
EVALUATION NOTES
This test is designed to show that the simulator overall
normal landing characteristics, primarily in the
longitudinal axis, are sufficiently like those of the
aeroplane to allow pilot training for landing
manoeuvres. The test should be commenced at a
radio altitude of not less than 61 metres (200 feet), so
that all ground effects can be examined as the
simulated aeroplane descends. It is not necessary to
show the entire landing ground roll, but the time
history must include details of the nose gear
touchdown. The important parameters are many,
since this is such a critical area for pilot training, but it
may be unrealistic to expect all parameters to be in
tolerance all of the time due to the complex nature of
pilot activity that is usually present during the
aeroplane flight test. When the test is run
automatically it will typically be controlled by closedloop drivers on, for example, pitch angle (driven by
pitch controller, or elevator as a last resort) and
possibly bank angle (with roll controller), though the
latter should not in theory be significant for this
particular test. Flare characteristics should be
examined carefully to ensure that over- or underrotation has not occurred or that the elevator used to
perform the flare does not deviate inexplicably from
the validation data.
Note that for computer controlled aeroplanes there
must be two tests, one for the normal control state and
the other for the non-normal state.
TOLERANCES
HEIGHT
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
±3 m (10 Ft) or ±10%
±3 Kts
±1.5o
±1.5o
controls:
PITCH CONTROLLER
±10% or ±2.2 daN (5
FORCE
Lbs)
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MANUAL TESTING
2E-3
The simulator will probably be automatically trimmed in the correct configuration
with an appropriate descent rate to allow the pilot to easily fly down and complete
the landing manoeuvre. The important points to assess are firstly that the
simulation is synchronised with the aeroplane data, especially for the flare and
touchdown portions, and secondly that the control positions and thrusts are the
same. The data should be carefully studied before beginning the manoeuvre so
that the pilot is able to reproduce as accurately as possible the control inputs
used during the flight test. The most critical part of the manoeuvre is the flare
from 50 ft to touchdown. The threshold speed should be as in the data. It should
not be expected to achieve perfect matches of all parameters and it will quite
probably be necessary for several attempts to be made before even a reasonably
acceptable result is obtained due to the complexity of coordinating and repeating
several simultaneous pilot inputs. If the aeroplane has autoland capability it
should only be used if the flight test also employed this method of achieving the
landing.
EXAMPLE
Figure 2e1-1 shows the difference that incorrect thrust is likely to make to a
landing test. The average thrust here is too high, but this has little effect on the
airspeed over the first 15 seconds. Where the problem manifests itself is in the
poor match with angle of attack (and also pitch angle, not shown). This test was
easily corrected by setting the correct thrusts prior to the commencement of the
test execution phase.
2E-4
Figure 2e1-1
Example of Simulator Test Results for Normal Landing
2E-5
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TITLE
2e(2) - MINIMUM FLAP LANDING
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OBJECTIVE
CHARACTERISTICS DURING EITHER A MINIMUM
OR NO FLAP LANDING CONFORM TO THE
AEROPLANE.
DEMONSTRATION
Carry out either an automatic or a manual normal
landing
to nosewheel touchdown as per the
prescribed aeroplane data at the minimum flap setting
or with flaps retracted. Record the data and compare
results.
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FLIGHT CONDITION
MINIMUM
CERTIFIED
CONFIGURATION
LANDING FLAP
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
EVALUATION NOTES
This test is designed to show that the simulator
characteristics for a minimum flap landing, primarily in
the longitudinal axes, are sufficiently like those of the
2E-6
aeroplane to allow pilot training for landing
manoeuvres. The test should be commenced at a
radio altitude of not less than 61 metres (200 feet), so
that all ground effects can be examined as the
simulated aeroplane descends. It is not necessary to
show the entire landing ground roll, but the time
history must include details of the nose gear
touchdown. The important parameters are many,
since this is such a critical area for pilot training, but it
aeroplane flight test. When the test is run
automatically it will typically be controlled by closedloop drivers on, for example, pitch angle (driven by
pitch controller, or elevator as a last resort) and
possibly bank angle (with roll controller), though the
latter should not in theory be significant for this
particular test. Flare characteristics should be
examined carefully to ensure that over- or underrotation has not occurred or that the elevator used to
perform the flare does not deviate inexplicably from
the validation data.
TOLERANCES
HEIGHT
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
±3 m (10 Ft) or ±10%
±3 Kts
±1.5o
±1.5o
controls:
PITCH CONTROLLER
±10% or ±2.2 daN (5
FORCE
Lbs)
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MANUAL TESTING
The simulator will probably be automatically trimmed in the correct configuration
with an appropriate descent rate to allow the pilot to easily fly down and complete
the landing manoeuvre. The important points to assess are firstly that the
simulation is synchronised with the aeroplane data, especially for the flare and
same. The data should be carefully studied before beginning the manoeuvre so
2E-7
that the pilot is able to reproduce as accurately as possible the control inputs
used during the flight test. The most critical part of the manoeuvre is the flare
from 50 ft to touchdown, along with the possibility of scraping the tail. The
threshold speed should be as in the data. It should not be expected to achieve
perfect matches of all parameters and it will quite probably be necessary for
several attempts to be made before even a reasonably acceptable result is
obtained due to the complexity of coordinating and repeating several
simultaneous pilot inputs. If the aeroplane has autoland capability it should only
be used if the flight test also employed this method of achieving the landing.
EXAMPLE
A poor result for
this test is shown
in Figure 2e2-1.
The radio altitude
remains well
within tolerance
right down to the
touchdown point,
but the simulator
then ‘bounces’ 20
feet
before
returning to the
ground. The
solution though
proved to be
simple - namely a
very
small
reduction in the
rate of descent
during
the
trimming phase,
demonstrating
just how sensitive
landing tests are
to the finer
details.
Figure 2e2-1
Example of Simulator Test Results for Minimum Flap Landing
2E-8
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TITLE
2e(3) - CROSSWIND LANDING
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OBJECTIVE
CROSSWIND LANDING CHARACTERISTICS
DEMONSTRATION
Perform an approach and landing in a crosswind
condition. Record data from at least 61 m (200 ft)
altitude through at least a 50% decrease in main
landing gear touchdown speed. The magnitude of the
crosswind component should be at least 60% of the
maximum demonstrated value provided in the AFM.
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE(S)
SPOILER ANGLES
EVALUATION NOTES
This test is designed to show that the simulator
characteristics for a crosswind landing are sufficiently
2E-9
like those of the aeroplane to fulfil the pilot training
requirements for landing manoeuvres. As for the other
landing tests, the manoeuvre should be commenced
at a radio altitude of not less than 61 metres (200
feet), so that all ground effects, both lateral and
longitudinal, can be examined as the simulated
aeroplane descends. It is not necessary to show the
entire landing ground roll, but the time history must
include details of the nose gear touchdown, followed
by a speed decrease to 50% of the main landing gear
touchdown speed. The important parameters are as
for test 2e(1) plus roll controller position, rudder angle,
bank angle and sideslip/heading angle. Once again it
aeroplane flight test, even more so in this test, where
the crosswind (at least 60% of the AFM value
measured at a height of 10 m (30 ft) above the
runway) was unlikely to have been steady. For this
reason the wind profile speeds in all three linear axes
should be provided as part of the data package. When
the test is run automatically it will typically be
controlled by closed-loop drivers on, for example,
pitch angle (driven by pitch controller, or elevator as a
last resort), bank angle (with roll controller) and
heading or yaw angle (with rudder or rudder pedal
position). Flare and de-crab characteristics should be
examined carefully to ensure that over- or undercontrol has not occurred or that the control surfaces
used to perform the flare and de-crab do not deviate
unduly from the validation data.
TOLERANCES
HEIGHT
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
BANK ANGLE
SIDESLIP ANGLE
HEADING
±3 m (10 Ft) or ±10%
±3 Kts
±1.5o
±1.5o
±2.0o
±2.0o
±3.0o
controls:
PITCH CONTROLLER
±10% or ±2.2 daN (5
2E-10
FORCE
ROLL CONTROLLER
FORCE
RUDDER PEDAL FORCE
Lbs)
±10% or ±1.3 daN (3
Lbs)
±10% or ±2.2 daN (5
Lbs)
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MANUAL TESTING
The simulator will probably be automatically trimmed (with the crosswind active)
in the correct configuration with an appropriate descent rate to allow the pilot to
easily fly down and complete the landing manoeuvre. The important points to
assess are firstly that the simulation is synchronised with the aeroplane data,
especially for the de-crab, flare and touchdown portions, and secondly that the
control positions and thrusts are the same or very similar. The data should be
carefully studied before beginning the manoeuvre so that the pilot is able to
reproduce as accurately as possible all the control inputs used during the flight
test. The most critical part of the manoeuvre is the de-crab and flare from 50 ft
to touchdown and then the speed decrease during the ground roll. The threshold
speed should be as in the data. It should not be expected to achieve perfect
matches of all parameters and it will quite probably be necessary for several
attempts to be made before even a reasonably acceptable result is obtained due
to the complexity of coordinating and repeating several simultaneous pilot inputs.
If the aeroplane has autoland capability it should only be used if the flight test
also employed this method of achieving the landing. The technique used can
vary between the wing down or crabbed approach and kick-off drift prior to
touchdown.
EXAMPLE
Figure 2e3-1 on the next page is a development version of one such test. The
plots appear to show that the ground effect is at fault, when in fact the solution
was to slightly increase the rate of descent during the trim phase. This had the
effect of causing the simulated aeroplane to touchdown slightly earlier and
thereby avoiding the erroneous pitch attitude which ensued when the simulated
aeroplane should already have been in ground contact. Slightly more difficult to
reconcile was the rate of airspeed decrease after touchdown. The validation data
clearly shows an anomaly here at around 105 knots and unless a wind speed
component is available but has not been properly implemented then further
information may be sought from the data provider.
2E-11
2E-12
Figure 2e3-1
Example of Simulator Test Results for Crosswind Landing
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TITLE
2e(4) - ONE ENGINE INOPERATIVE LANDING
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OBJECTIVE
ENGINE-OUT LANDING CHARACTERISTICS
DEMONSTRATION
Perform an approach and landing with a single engine
inoperative at the appropriate landing flap for an
engine-out approach. The inoperative engine may be
represented by setting thrust to idle. Record data from
at least 61 m (200 ft) altitude through at least a 50%
decrease in main landing gear touchdown speed.
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE(S)
SPOILER ANGLES
EVALUATION NOTES
The purpose of this test is to show that the simulator
characteristics for a landing with one engine
2E-13
inoperative are sufficiently like those of the aeroplane
to fulfil the pilot training requirements for landing
manoeuvres. As usual for the landing tests, the
manoeuvre should be commenced at a radio altitude
of not less than 61 metres (200 feet), so that all
ground effects, both lateral and longitudinal, can be
examined as the simulated aeroplane descends. It is
not necessary to show the entire landing ground roll,
but the time history must include details of the nose
gear touchdown, followed by a speed decrease to
50% of the main landing gear touchdown speed. The
important items are as for the crosswind landing test
plus the engines key parameters. Again it will be
unrealistic to expect all parameters to be in tolerance
all of the time due to the complex nature of pilot
activity usually present during the aeroplane flight test.
When the test is run automatically it will typically be
controlled by closed-loop drivers on, for example,
pitch angle (driven by pitch controller or elevator as a
last resort), bank angle (with roll controller) and
heading or yaw angle (with rudder or rudder pedal
position). Flare and de-crab characteristics should be
examined carefully to ensure that over- or undercontrol has not occurred or that the control surfaces
used to perform the flare and de-crab do not deviate
unduly from the aeroplane data. It may be helpful to
show the de-rotation as a separate segment from the
time of main landing gear touchdown, but this will
depend on the way in which the data is presented.
TOLERANCES
HEIGHT
AIRSPEED
PITCH
ANGLE OF ATTACK
BANK ANGLE
SIDESLIP ANGLE
HEADING
±3 m (10 Ft) or ±10%
±3 Kts
±1.5o
±1.5o
±2.0o
±2.0o
±3.0o
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MANUAL TESTING
The simulator will probably be automatically trimmed (with the appropriate engine
inoperative) in the correct configuration with a descent rate corresponding to the
2E-14
flight test data to allow the pilot to easily fly down and complete the landing
manoeuvre. The important points to assess are firstly that the simulation is
synchronised with the aeroplane data, especially for the de-crab, flare and
same or very similar. The data should be carefully studied before beginning the
manoeuvre so that the pilot is able to reproduce as accurately as possible all the
control inputs used during the flight test. The most critical part of the manoeuvre
is the de-crab and flare from 50 ft to touchdown and then the speed decrease
during the ground roll. The threshold speed should be as in the data. It should not
be expected to achieve perfect matches of all parameters and it will quite
probably be necessary for several attempts to be made before even a reasonably
acceptable result is obtained due to the complexity of coordinating and repeating
several simultaneous pilot inputs. If the aeroplane has autoland capability it
should only be used if the flight test also employed this method of achieving the
landing. When conducting the test manually, ensure the simulator is controllable
with reverse thrust on the unaffected engine(s) down to the speed that reverse
thrust is disengaged.
EXAMPLE
In Figure 2e4-1 there are problems with both the de-rotation characteristics and
with the speed degree during the ground roll-out, as well as the fact that the
aeroplane data does not decrease sufficiently to match the requirement of 50%
of touchdown speed, though it is usually possible to run the test for longer to
show the simulator characteristics at lower speeds.
Figure 2e4-1
Example of Simulator Test Results for Engine Inoperative Landing
2E-15
A second engine inoperative landing example is shown in Figure 2e4-2 below,
and serves to illustrate the kind of different thrust response which may be seen
when the simulator is ‘fitted’ with an engine model that is different from that used
to gather the aeroplane data. In the plot for No.1 Engine Thrust, the engine is
retarded to idle at approximately the correct time, but the simulated engine (the
unbroken line) takes longer to reach idle thrust. As it is, this result could stand
improvement anyway, but there may be no getting around the way the thrust
decays differently other then by running a second, separate test which overwrites
thrust so that the correct implementation of the aeroplane manufacturer’s model
can be proven.
2E-16
Figure 2e4-2
Example of Simulator Test Results for Engine Inoperative Landing,
Alternate Engine Fit
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TITLE
2e(5) - AUTOPILOT LANDING (if applicable)
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OBJECTIVE
AUTOMATIC
LANDING CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Perform a normal autoland approach and landing.
Record data from a least 61 m (200 ft) altitude to 50%
of main gear touchdown speed (if the autoland system
includes rollout guidance).
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
RATE OF CLIMB
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE(S)
SPOILER ANGLES
LATERAL DISPLACEMENT FROM RUNWAY
CENTRELINE
FLARE ENGAGE DISCRETE
WEIGHT ON WHEELS/GEAR CONTACT FLAG
2E-17
EVALUATION NOTES
The autoland test is designed to ascertain that the
auto-approach and autoland system as installed
and/or programmed in the simulator, is capable of
producing the same landing performance in the
simulator as the real system does in the aeroplane.
The test assumes therefore that the aeroplane
handling qualities are correct, having been checked
under various conditions in tests 2e(1), 2e(2), 2e(3)
and 2e(4). The test is required whether the simulator
autopilot system utilises the actual aeroplane part
number or it is software-simulated. The lateral
displacement specified above in the list of recorded
parameters should be plotted from touchdown to the
point at which the autopilot was disconnected, or at
least until a 50% decrease in main landing gear
touchdown speed. Since the test is usually run from a
stabilised condition, achieving an accurate rate of
descent does not cause too many problems. However,
if the flare height and time deviate from the data the
problem could lie with the aerodynamic ground effects
though possibly with the autoland system simulation
itself. Note that, since the purpose of the test is to
check the simulated system, it is the system itself
which should be used for the test. Overwriting elevator
angle or thrust, for example, is not acceptable, and
thus running the test ‘automatically’ (other than the
setup of initial conditions on approach) is not
appropriate.
TOLERANCES
FLARE HEIGHT
DURATION OF FLARE (Tf)
RATE OF DESCENT AT
TOUCHDOWN
LATERAL DEVIATION
DURING ROLLOUT
(Time of autopilot flare mode
touchdown should be noted)
±1.5 m (5 Ft)
±0.5 Sec or ±10%
±0.7 m/Sec (140
Ft/Min)
±3 m (10 Ft)
engage and main gear
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MANUAL TESTING
Usual procedures call for the simulated aeroplane to be trimmed such that the
autoland system is able to cope with the aeroplane dynamics when the reposition
2E-18
is complete. The pilot intervention in this test should be minimal, since the
intention is to check out the automatic landing system. Thus the test is in itself
manual rather than ‘automatic’ in the conventional (i.e. autotest) sense and the
inputs from the autotest system are typically confined to wind speed components
and terrain height. Where the aeroplane will provide lateral control to a full stop,
the simulator must also.
It may be that the results are sensitive to pilot inputs; on certain computercontrolled aeroplanes the flare time is dependant of the exact moment of throttle
back during the flare.
EXAMPLE
An ‘interesting’ result is shown in Figure 2e5-1, but it is one that arguably looks
worse than it actually is. The overall synchronisation of the manoeuvre is not
correct, as exemplified by the radio altitude trace passing 56 feet approximately
0.5 second late. The flare begins over a second late as a result and is slightly too
long. Correct setup of the initial conditions cured the problem, as is often the
case.
Figure 2e5-1
Example of Simulator Results for Autopilot Landing
2E-19
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TITLE
2e(6) - ALL-ENGINE-AUTOPILOT GO-AROUND
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OBJECTIVE
AUTOPILOT GO AROUND WITH ALL ENGINES
OPERATING CONFORMS TO THE AEROPLANE.
DEMONSTRATION
Perform a normal approach with autopilot ON, then
conduct a go-around at the appropriate decision
height according to the AFM with all engines
operating.
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FLIGHT CONDITION
AS PER AFM (medium weight)
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
RATE OF CLIMB
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE(S)
SPOILER ANGLES
FLAP POSITION
GEAR POSITION
EVALUATION NOTES
The intention behind this test is to determine that the
simulator autopilot exhibits the correct characteristics
2E-20
(primarily longitudinal) when subjected to a pilot
decision to go around with all engines operative. The
test must be conducted with the autopilot itself, and
not by using substitute inputs such as elevator or roll
controller, etc. This therefore makes it an easy test to
accomplish, but care should be taken to select goaround operation at precisely the correct moment. The
preference is to use the usual, physical means of
selection rather than a programmed test input to make
the selection. This has the benefit of alleviating the
requirement for a separate means of performing the
test manually.
Note that for computer controlled aeroplanes two tests
are required, one for the normal flight control state and
the other for a non-normal state.
TOLERANCES
PITCH ANGLE
AIRSPEED
ANGLE OF ATTACK
±1.5o
± 3 Kts
±1.5o
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MANUAL TESTING
The test will normally commence with the simulated aeroplane trimmed in the
appropriate configuration for a descent down the glideslope. Once the
manoeuvre has begun it is important to ensure that the autopilot behaviour is
synchronised in accordance with the aeroplane data, this will confirm that all
associated longitudinal, (and lateral and directional, if appropriate) trim changes
are correctly reflected.
EXAMPLE
An example of an all engine go-around is shown in Figure 2e6-1. The results
seem to be within tolerance, but the data provided was not specifically for a
manoeuvre using the autopilot, or if it was then the data provider did not make
that fact clear in the validation document.
2E-21
2E-22
Figure 2e6-1
Example of Simulator Test Results for Go-Around, All Engines Operating
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TITLE
2e(7) - ONE-ENGINE-INOPERATIVE GO-AROUND
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR ONE
ENGINE INOPERATIVE GO-AROUND
AEROPLANE.
DEMONSTRATION
Perform an approach with one engine inoperative at
the appropriate landing flap setting for an engine-out
approach and conduct a go around at the appropriate
decision height according to the AFM
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FLIGHT CONDITION
a) AS PER AFM (with autopilot, if applicable)
b) AS PER AFM (manual)
(Both tests to be at near maximum certificated landing
weight)
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RECORDED
PARAMETERS
ALTITUDE
AIRSPEED
RATE OF CLIMB
PITCH ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
BANK ANGLE
HEADING ANGLE
ANGLE OF ATTACK
RUDDER ANGLE
SIDESLIP ANGLE
AILERON ANGLE(S)
SPOILER ANGLES
EVALUATION NOTES
The intention behind these tests is to determine that
2E-23
the simulator exhibits the correct longitudinal, lateral
and directional characteristics when subjected to a
pilot decision to go-around when one (critical) engine
is inoperative, both with and without the autopilot
engaged. For the ‘manual’ test case running under
automatic test system control, the driving inputs used
should be only those which were used by the pilot
during the flight test. The response is very unlikely to
be free at any time during the time history, especially
once the engine power has been applied and so there
may be a case for using closed-loop controllers during
the non-autopilot simulator test. If this is so, there
must be good correlation between simulator and
aeroplane control surface positions and the scales
chosen for the plotted results should enable an easy
comparison to be made.
For the autopilot case, see notes for test 2e(6a).
Note that for computer controlled aeroplanes the nonautopilot test is to be conducted in a non-normal
mode.
TOLERANCES
PITCH ANGLE
AIRSPEED
ANGLE OF ATTACK
BANK ANGLE
SIDESLIP ANGLE
±1.5o
±3 Kts
±1.5o
±2o
±2o
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MANUAL TESTING
The test will normally commence with the simulated aeroplane trimmed in the
appropriate configuration for a descent down the glideslope. Once the
manoeuvre has begun it is important to ensure that the power increase, flap
selection and gear selection are synchronised in accordance with the aeroplane
data, this will confirm that all associated longitudinal, lateral and directional trim
changes are correctly reflected.
For the autopilot case, the autopilot should be used with normal approach
procedures for one engine inoperative.
EXAMPLE
2E-24
The captain’s wheel position plot in Figure 2e7-1 indicates that a closed-loop
controller was used to assist the simulator in maintaining the bank angle
specified (not shown) for the duration of the test. Probably the gains used for
the closed-loop controller driver are too high though, as it is somewhat
unlikely that the wheel would have been used to the extent indicated by the
plot. Unfortunately, the wheel angle was not supplied by the data provider, so
there is no way of knowing. Running the test manually would serve as a good
comparison in this case.
Figure 2e7-1
Example of Simulator Test Results for One Engine Inoperative Go-Around
2E-25
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TITLE
2e(8) - DIRECTIONAL CONTROL (RUDDER
EFFECTIVENESS) WITH REVERSE THRUST
(SYMMETRIC)
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OBJECTIVE
DIRECTIONAL CONTROL (RUDDER
EFFECTIVENESS) ON GROUND WITH SYMMETRIC
REVERSE THRUST CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
Starting at a speed near normal touchdown speed on
the runway, apply rudder pedal input in both directions
using full reverse thrust until reaching full thrust
reverser minimum operating speed. The nose gear
steering should be disabled to isolate the effects of
rudder control. Delay the use of wheel brakes as long
as possible.
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
AIRSPEED
RUDDER ANGLE
NOSEWHEEL STEERING ANGLE
HEADING ANGLE
SIDESLIP ANGLE
YAW RATE
YAW ACCELERATION
LATERAL DEVIATION FROM RUNWAY CENTRE
LINE
EVALUATION NOTES
This test was originally formulated with particular
reference to aeroplanes with tail-pod mounted
engines, where the jet blast with the reverser doors
2E-26
open could impinge directly on to the vertical tail,
causing difficulty with directional control at medium to
low speeds. However, the effect can also be
significant even with wing-mounted engines, so the
simulator test must be generated for all aeroplane
types. The test should be conducted using full reverse
thrust from a speed near a normal touchdown speed
to the minimum operating speed for maximum reverse
thrust by applying full rudder pedal input in both
directions.
Of special importance is the airspeed/reverse
thrust/rudder deflection combination, all three of which
must be properly synchronised if the effects are to be
correctly reproduced. The tolerance is applied to the
minimum rudder effectiveness speed and also to the
yaw rate. The aeroplane manufacturer's data must be
explicit in its definition of rudder effectiveness speed
to enable a clear application of the standard. Without
this specific flight datum, subjective comparison of
simulator data with flight data is required. If no
aeroplane test data is available, then the aeroplane
manufacturer's engineering simulator data may be
used for reference data.
TOLERANCES
AIRSPEED
YAW RATE
± 5 Kts
± 2o/SEC
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MANUAL TESTING
Typical manual testing procedures call for the simulated aeroplane to be
positioned on the runway in the landing configuration at an approximate
touchdown speed. The thrust reversers are deployed to give maximum
permissible reverse thrust whilst the rudder pedals are used (with the nosewheel
castoring) in both directions to deliberately attempt to deviate from the runway
centreline. If time permits, it may be useful to confirm the results by running a
second test which utilises a lower value of reverse thrust. The speed at which
directional control is lost should be lower than in the first instance.
EXAMPLE
Figures 2e8-1 and 2e8-2 below illustrate one of the reasons why obtaining
integrated validation data is important. Figure 2e8-1 has been run by driving both
2E-27
the rudder surface position and the nosewheel angle independently and
overwriting the software locations directly. The yaw rate and speed are not
perfect matches, but they are very close and it may be that the reverse thrust
was not initialised quite correctly.
2E-28
Figure 2e8-1
Example of Simulator Test Results for Directional Control with Symmetric Reverse
Thrust (1)
Figure 2e8-2 was run by driving rudder pedal position and allowing the
nosewheel to castor as it would when ‘flying’ the simulated aircraft normally. The
rudder pedal position had to be derived for the purposes of the test, based on the
rudder angle, as it was not provided as part of the validation data package. The
obvious difference between this result and that shown on Figure 2e8-1 is that the
nosewheel response is totally different - and in fact much more logical in this
result than in the former. This has resulted in a somewhat different yaw rate
profile, and an airspeed which is not within tolerance beyond approximately 12
seconds.
Figure 2e8-2
Example of Simulator Test Results for Directional Control with Symmetric
Reverse Thrust (2)
2E-29
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TITLE
2e(9) - DIRECTIONAL CONTROL (RUDDER
EFFECTIVENESS) WITH REVERSE THRUST
(ASYMMETRIC)
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OBJECTIVE
DIRECTIONAL CONTROL (RUDDER
EFFECTIVENESS) ON GROUND WITH
ASYMMETRIC REVERSE THRUST CONFORMS TO
THE AEROPLANE.
DEMONSTRATION
Starting at a speed near normal touchdown speed on
the runway, with the thrust reverser inoperative on one
engine and full reverse thrust on the remaining
engine(s), apply sufficient rudder control to maintain a
constant heading along the runway. Continue until
heading cannot be maintained with full rudder control
or minimum speed for thrust reverser operation is
reached. The nose gear steering should be disabled
to isolate the effects of rudder control. The inoperative
thrust reverser may be represented by a setting of
forward idle thrust.
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
2E-30
AIRSPEED
RUDDER ANGLE
NOSEWHEEL STEERING ANGLE
HEADING ANGLE
SIDESLIP ANGLE
YAW RATE
YAW ACCELERATION
LATERAL DEVIATION FROM RUNWAY CENTRE
LINE
EVALUATION NOTES
The Evaluation Notes for test 2e(8) above apply to this
test, except that the test should be performed by using
a steadily increasing pilot rudder control input in the
direction to counter the yawing moment due to the
asymmetric thrust. The tolerance on airspeed applies
throughout the test, but it is especially important to
note the speed at the point where heading can no
longer be maintained (or at the declared minimum
speed for thrust reverser operation) to check that the
simulator maximum rudder effectiveness in the
presence of asymmetric reverse thrust matches the
validation data. The speed profile should be
maintained, along with the pedal steering inputs.
Typically a slow decrease in speed is combined with
a slow increase in pedal position to counteract the
yaw moment from the engine asymmetry. In some
flight tests, a nosewheel angle is present, even though
the nosegear steering system may be disconnected,
and the use or non-use of this nosewheel value this
may have a significant influence on the test results.
TOLERANCES
AIRSPEED
HEADING ANGLE
± 5 Kts
± 3o
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MANUAL TESTING
Typical manual testing procedures call for the simulated aeroplane to be
positioned on the runway in the landing configuration at an approximate
touchdown speed. One engine is set to forward idle thrust, while the remaining
engine(s) are set to full reverse thrust. Rudder control is applied to counter the
yawing moment due to the asymmetric reverse thrust usually until full rudder
pedal position is reached and heading on the runway can no longer be
maintained at which point the test may be terminated. If time permits, it may be
useful to confirm the results by running a second test which utilises a lower value
of reverse thrust. The speed at which directional control is lost should be lower
than in the first instance.
EXAMPLE
The result shown in Figure 2e9-1 illustrate how easily the lateral distance can
deviate even when the yaw rate is easily within tolerance. There may be reasons
why the lateral deviation plot should be omitted, but the regulatory authorities
2E-31
tend to request its inclusion. Obviously, the interactions between the
aerodynamic calculations and the ground handling are tested to the extreme
here.
Figure 2e9-1
Example of Simulator Test Results for Directional Control with Asymmetric Reverse
2E-32
Thrust
SECTION 2f
GROUND EFFECT
2f(1)
A test to demonstrate ground effect
2F-1
2F
GROUND EFFECT
For a simulator to be used for takeoff and, in particular, landing credit, it
must faithfully reproduce the aerodynamic changes which occur in ground
effect. The parameters chosen for simulator validation must obviously be
indicative of these changes.
A dedicated test should be provided which will validate the aerodynamic
ground effect characteristics.
The selection of the test method and procedures to validate ground effect
is at the option of the organisation performing the flight tests; however, the
flight test should be performed with enough duration near the ground to
sufficiently validate the ground-effect model.
Acceptable tests for validation of ground effect include:
a)
Level fly-bys. The level fly-bys should be conducted at a minimum
of three altitudes within the ground effect, including one at no more
than 10% of the wingspan above the ground, one each at
approximately 30% and 50% of the wingspan where height refers to
main gear tyre above the ground. In addition, one level-flight trim
condition should be conducted out of ground effect, e.g., at 150% of
wingspan.
b)
Shallow approach landing. The shallow approach landing should be
performed at a glide slope of approximately one degree with
negligible pilot activity until flare.
If other methods are proposed, rationale shall be provided to conclude that
the tests performed do validate the ground-effect model.
The lateral-directional characteristics are also altered by ground effect. For
example, because of changes in lift, roll damping is affected. The change
in roll damping will affect other dynamic modes usually evaluated for flight
simulator validation. In fact, Dutch roll dynamics, spiral stability and roll-rate
for a given lateral control input are altered by ground effect. Steady
heading side-slips will also be affected. These effects shall be accounted
for in the simulator modelling. Several tests such as "crosswind landing",
"one engine inoperative landing" and "engine failure on take-off" serve to
validate lateral-directional ground effect since portions of them are
accomplished whilst transiting heights at which ground effect is an
important factor.
2F-2
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TITLE
2f(1) - GROUND EFFECT DEMONSTRATION
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OBJECTIVE
GROUND
EFFECT
AERODYNAMIC
CHARACTERISTICS CONFORMS TO THE
AEROPLANE.
DEMONSTRATION
For level fly-bys, trim the aeroplane at a constant
height above a runway. The level flybys should be
performed at least four altitudes: one that is out of
ground effect (e.g., 150% of the wing span), plus three
more at approximately 10%, 30% and 50% of
wingspan. The test out of ground effect should be
conducted first. Note the stabiliser position, and use
that same setting for the tests in ground effect.
Maintain constant altitude and airspeed using pitch
control and thrust control, respectively. For the
shallow approach technique, set up a landing
approach well above ground effect at a glide slope of
approximately -1 degree. Continue the approach with
a minimum of control activity until or just prior to main
gear touchdown, reducing power, as required, during
the flare. For good results, it is essential that ground
effects testing be conducted in nearly calm air, i.e.,
with little or no atmospheric turbulence
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FLIGHT CONDITION
LANDING
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RECORDED
PARAMETERS
HEIGHT ABOVE GROUND/RADIO ALTITUDE
AIRSPEED
STABILISER ANGLE
PITCH ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
2F-3
BANK ANGLE (to show wings level)
EVALUATION NOTES
The purpose of this test is to show that the simulator
aerodynamic model includes terms to adequately
represent the longitudinal ground effect on the
aeroplane. As discussed in Paragraph 2F above,
there are at least two acceptable means to
demonstrate that the longitudinal ground effect
characteristics match the validation data. For the level
fly-by method, the aeroplane will have been trimmed
for level flight at three or more heights both within and
just above ground effect at approximately the same
airspeed and configuration. The results should
illustrate the different longitudinal control required at
each height with the same trim position that was
established for the free-air fly-by condition, and the
thrust change required to maintain a constant
airspeed as the height decreases.
Because of the difficulty in acquiring this type of data,
it is possible that the trim conditions given in an
aeroplane manufacturer's time history will not be very
stable. If this is the case, then engineering judgement
should be used to determine that the simulator
appears to conform to the general trend of the
aeroplane data within the tolerances given below. The
shallow approach technique allows a relatively gradual
and continuous descent through the ground effect to
a height corresponding to main gear touchdown. For
an approach glide slope of about one degree, the
aeroplane tends to flare automatically, so a well
stabilised approach in calm air should require very
little pilot pitch control activity to just prior to main gear
touchdown, enabling a nearly controls-free ground
effect evaluation.
TOLERANCES
2F-4
STABILIZER ANGLE
ELEVATOR ANGLE
PITCH ANGLE
ANGLE OF ATTACK
NET THRUST or equivalent
AIRSPEED
HEIGHT
±0.5o
±1o
±1o
±1o
±5%
±3 Kts
±1.5 m (5Ft) OR ±10%
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MANUAL TESTING
The fly-by test will usually consist of a series of longitudinal trims at several
heights above the ground with all except one in ground effect. The purpose is to
demonstrate that the trim conditions, either with regard to stabiliser position for
no pitch control input or pitch control for a constant stabiliser position, are
different for each height. In addition, small throttle adjustments will be required
to maintain a constant airspeed at each height. The differences however are
likely to be small, so it is of the utmost importance that all attempts are made to
obtain accurate steady state trim conditions so that this difference is easy to
distinguish. The results will be snapshots and as such do not require or benefit
from being plotted as a time history, and it may be reasonable to set and lock the
simulated aeroplane at the correct radio altitude to make the test points easier
to fly.
EXAMPLE
Referring to the example results in Figure 2f1-1, at first sight the result may
indicate that the ground effect is in error - obviously fundamental if the simulator
is to be used for landing manoeuvres. However, examination of the setup
conditions in this case revealed that the gross weight had not initialised correctly,
hence the large discrepancy between the aeroplane and simulator net thrust
values. The remedy was simply to re-run the test and make sure that the weight
was correct the second time. One-off anomalous test runs such as this are not
as common as they used to be, but can still sometimes happen.
2F-5
Figure 2f1-1
Example of Simulator Test Results for Ground Effect Demonstration (Snapshots)
2F-6
SECTION 2g
WINDSHEAR
2g(1)
A test to demonstrate windshear models
2G-1
2G
WINDSHEAR
2G.1 GENERAL
The fidelity of windshear modelling has developed significantly in the last
few years such that simulators are very effective tools for training flight
crews in the techniques necessary to combat these phenomena. Various
studies have been undertaken by several research bodies, mainly in the
United States and United Kingdom, and from these studies wind models
have been derived which can be successfully correlated to measurements
taken during actual windshear encounters.
2G.2 REQUIREMENTS
For the purposes of flight crew training, there are four critical phases of flight
for which wind models should be available in the simulator:
1)
2)
3)
4)
Prior to takeoff rotation
At liftoff
During initial climb
Short final approach
The most obvious acceptable means of complying with these requirements
is by making use of the information contained in the "Windshear Training
Aid" (Reference 17). The data for the four models contained in the
document are presented in Figures 2g-1 through 2g-4. Figure 2g-5 is a
graph of the wind factor which must be applied, depending on the aircraft
type, to wind models 1 and 2 to enhance the training value. References 13
through 17 give a great deal of information on the simulation of wind-related
effects such as windshear and turbulence.
A diagrammatic representation of the microburst model generated by the
UK Royal Aerospace Establishment is shown in Figure 2g-6. The
mathematical data tables for this model are complex and difficult to
represent in any standard plot format. The reader is referred to Reference
14 for more specific information.
2G.3 OTHER DATA SOURCES
Whilst the Windshear Training Aid profiles provide one solution to the ICAO
Manual requirement, other sources of wind model data are also acceptable,
provided they are from a recognised source such as the UK Royal
Aerospace Establishment, Bedford or the Joint Airport Weather Studies
Project. See References 13 and 14 respectively for further details.
2G-2
Whatever the source of data used, it must be properly supported and
referenced in the QTG. Use of alternate data must be coordinated with the
regulatory authorities prior to submission of the QTG for approval.
2G-3
FAA Wind Training Aid Profile #1
Longitudinal Wind Velocity
0
U.Wind (kts)
-10
-20
-30
-40
-50
0
2000
4000
6000
8000
Distance Travelled (ft)
10000
12000
10000
12000
10000
12000
Lateral Wind Velocity
1
V.Wind (kts)
0.8
0.6
0.4
0.2
0
0
2000
4000
6000
8000
Vertical Wind Velocity
1
W.Wind (kts)
0.8
0.6
0.4
0.2
0
0
2000
4000
6000
8000
Figure 2g-1
2G-4
0
U.Wind (kts)
-10
-20
-30
-40
-50
-60
0
5000
10000
15000
20000
15000
20000
1
V.Wind (kts)
0.8
0.6
0.4
0.2
0
0
5000
10000
W.Wind (kts)
0
-2
-4
-6
-8
-10
0
5000
10000
15000
20000
Figure 2g-2
2G-5
0
U.Wind (kts)
-10
-20
-30
-40
-50
-60
0
2000
4000
6000
8000
10000
12000
14000
12000
14000
12000
14000
20
V.Wind (kts)
15
10
5
0
-5
-10
-15
0
2000
4000
6000
8000
10000
20
W.Wind (kts)
10
0
-10
-20
-30
0
2000
4000
6000
8000
10000
Figure 2g-3
Wind training Aid Model #3
2G-6
U.Wind (kts)
30
20
10
0
-10
-20
-30
-40
0
5000
10000
15000
Distance Travelled
20000
25000
20000
25000
20000
25000
V.Wind (kts)
20
15
10
5
0
-5
-10
-15
0
5000
10000
15000
Distance Travelled
20
W.Wind (kts)
10
0
-10
-20
-30
0
5000
10000
15000
Distance Travelled
Figure 2g-4
2G-7
NOTE: THIS CHART SHOULD BE USED DIRECTLY FOR ALL
WINDSHEAR EXERCISES EXCEPT THE FOLLOWING:
1) FOR THE ‘ON GROUND PRIOR TO VR’ EXERCISE,
REDUCE THE WIND FACTOR BY 0.1
2) FOR REFERENCE WIND MODEL NO. 2, ‘DURING INITIAL
CLIMB EXERCISE’, USE THE LINE LABELLED ‘3 & 4 ENGINE
AIRPLANES’ FOR ALL CASES
TWIN ENGINE
AIRPLANES
USE THIS DATA
2
1.9
3 & 4 ENGINE
AIRPLANES
USE THIS DATA
1.8
1.7
WIND FACTOR
1.6
1.5
1.4
1.3
1.2
1.1
1
0.9
0.8
0.16 0.18 0.2 0.22 0.24 0.26 0.28 0.3 0.32
2G-8
Figure 2g-5
Wind Training Aid Wind Factor Chart
Figure 2g-6
United Kingdom Royal Aerospace Establishment (now Qinetiq) Microburst Vortex Ring Air
Flow Model
2G-9
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TITLE
2g(1) - A TEST TO DEMONSTRATE WINDSHEAR
MODELS
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OBJECTIVE
TO DEMONSTRATE THAT THE SIMULATOR IS
PROVISIONED WITH WINDSHEAR MODELS
CAPABLE OF PROVIDING POSITIVE TRAINING
DURING WINDSHEAR ENCOUNTERS.
DEMONSTRATION
Perform the specified manoeuvres (i.e. takeoff and
approach/go-around) with and without windshear
selected and record the effects of the windshear
models on handling and performance.
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FLIGHT CONDITION
a) TAKEOFF
b) LANDING
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RECORDED
PARAMETERS
HEIGHT ABOVE GROUND/RADIO ALTITUDE
AIRSPEED
RATE OF CLIMB
PITCH ANGLE
ANGLE OF ATTACK
PRIMARY CONTROL POSITIONS & SURFACES
BANK ANGLE
HEADING ANGLE
STABILISER ANGLE
FLAP POSITION(S)
EVALUATION NOTES
Each windshear profile must be implemented as per
the nominated model so that the simulator is assured
the use of data which have been properly
researched, rather than arbitrarily decided upon. The
model(s) nominated must provide, through the
simulation, adequate recognition cues and the
capability to execute recovery manoeuvres. Usually,
the requirement for these tests amounts to a total of
2G-10
four manoeuvres - two for a normal takeoff (one with
and the other without windshear) and two for an
approach to go-around case (again, one with and
the other without windshear). The reason for this is
to illustrate the difference that the windshear makes
and to ease the task of the evaluator when trying to
determine the effectiveness of the models used for
each of the two flight conditions. For repeatability
during recurrent evaluations, automatic drivers may
be used to reproduce correct pilot techniques and
thus prevent the simulated aeroplane from crashing.
Time history plots of relevant parameters are also
helpful and should be provided.
TOLERANCES
None
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MANUAL TESTING
This test should be performed after it has been determined that the nominated
model has been implemented correctly by comparing simulator data
(windshear profiles) with the flight (model) data. The effect of windshear in
each of the two flight conditions should be demonstrated firstly by the pilot
performing the appropriate manoeuvre (i.e. takeoff and approach/go around)
with no windshear inserted and then with the windshear, using a corresponding
model for each which should be provided (and clearly labelled as either takeoff
or approach/landing) on the instructor's station. Recording and plotting the
various parameters of interest will assist greatly in the subsequent evaluation
of the results. It may enhance the effectiveness of the test if the pilot flies
through each of the models without being briefed on which profile to expect. If
the profile selected is recognised by the flight crew as a windshear problem,
then the demonstration is successful.
EXAMPLE
Figure 2g1-1 below shows a partial set of equivalent results for both a
simulated takeoff with and without windshear present. There is no aeroplane
data for comparison, nor does there need to be, as the intent is to show the
effects of the windshear on the simulation. These tests should be backed up by
subjective evaluation.
2G-11
Figure 2g1-1
Example of Simulator Test Results for Takeoff Windshear Demonstration
2G-12
SECTION 2h
FLIGHT AND MANOEUVRE ENVELOPE
PROTECTION FUNCTIONS
(This Section is only applicable to Computer Controlled Aeroplanes)
2h(1)
Overspeed
2h(2)
Minimum Speed
2h(3)
Load Factor
2h(4)
Pitch Angle
2h(5)
Bank Angle
2h(6)
Angle of Attack
2H-1
2H
FLIGHT AND MANOEUVRE ENVELOPE PROTECTION FUNCTIONS
As a general note, all of the tests in this section are applicable only to
Computer Controlled Aeroplanes and should show time history results of
the response to control inputs during entry into each envelope protection
function. The requirements of the ICAO Manual state that all these tests
must be run in both normal and degraded control states if the function is
different. However, it is evident that for envelope limiting functions, there
is little to be gained by running a test where that function is inactive,
hence the requirement really refers to testing for the most degraded
states where the function is still active.
For each test in this section an example procedure is given. If the
aeroplane manufacturer provides information on the test procedure used
in the aeroplane, then the same procedure should be used in the
simulator. The aeroplane manufacturer's procedure would normally
include information such as trim speed for a specific test, rate of speed
change, rate of pitch angle change, aeroplane configuration and other
factors which may be important to successful demonstration of simulator
modelling.
In some cases, where the original aircraft hardware is used to
implement the computer control functions, it may be permissible to omit
some of these tests from the QTG. In these cases the QTG must include
the full rationale to cover such omissions.
2H-2
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TITLE
2h(1) - OVERSPEED
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OBJECTIVE
OVERSPEED PROTECTION FUNCTION
OPERATES AS IN THE AEROPLANE.
DEMONSTRATION
From a stabilised flight condition near Vmo/Mmo,
push forward smoothly on the control
column/longitudinal controller and hold until the
aeroplane speed begins to decrease or stabilises at
its limited value.
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FLIGHT CONDITION
CRUISE
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
PRESSURE ALTITUDE
STABILISER ANGLE
ENGINES KEY PARAMETERS (including net thrust)
EVALUATION NOTES
The test should first be run with all aeroplane
systems functional so that the normal operation of
the overspeed protection system can be tested, this
may have been achieved either by the utilisation of
the actual flight warning unit or else by software
simulation. Running the test automatically will
probably involve the driving of the pitch controller
position, so that the test does not bypass or ignore
any interaction between the electronic flight control
system and the overspeed protection system. The
method of detecting the onset of the overspeed
protection should be ascertained, this can be
typically identified by changes in the control surface
2H-3
angles which do not correspond to the inputs at the
controller position. The aeroplane maximum
allowable speed may be either airspeed or mach
number defined, and may therefore change with
altitude. The overspeed protection system may also
be dependant on the rate at which the limit is being
approached, affecting both the protection onset
speed and the maximum speed reached during the
recovery. Note that the test must be run for both
normal and non-normal flight control system states
(see note on page 2H-2).
TOLERANCES
AIRSPEED
±5 Kts
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MANUAL TESTING
This test should be continued until the airspeed is stable for the current
altitude. The entry into the overspeed function should be smooth, with a
constant control position, which may be stick forward or stick neutral, as
required for accelerating flight in the specific aeroplane configuration.
EXAMPLE
See next page.
2H-4
Figure 2h1-1
Example of Simulator Test Results for Overspeed Protection Function
2H-5
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TITLE
2h(2) - MINIMUM SPEED
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OBJECTIVE
MINIMUM SPEED PROTECTION FUNCTION
DEMONSTRATION
From a stabilised flight condition near the minimum
speed for the configuration, pull back smoothly on
the control column/longitudinal controller and hold
until the aeroplane speed stabilises at its limited
value.
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FLIGHT CONDITION
a) TAKEOFF
b) CRUISE
c) APPROACH or LANDING
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
PRESSURE ALTITUDE
STABILISER ANGLE
EVALUATION NOTES
the minimum speed system can be tested. Running
the test automatically will probably involve the
driving of the elevator surface position, but this
should not adversely affect the test results providing
such methods do not bypass or ignore any
interaction between the electronic flight control
system and the minimum speed warning system.
The method of detecting the minimum speed
warning should also be ascertained, so that the
2H-6
evaluator is assured of obtaining results which truly
test the functionality of the simulated aeroplane
system, which may be achieved either by the
utilisation of the actual aeroplane flight warning unit
or else by software simulation. The intention is to
demonstrate limitation at the lowest permitted
operating speed for the configuration, and should
continue until the speed is stable. Note that the test
must be run for both normal and non-normal flight
control system states (see note on page 2H-2).
TOLERANCES
AIRSPEED
±3 Kts
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MANUAL TESTING
The entry into the speed limiting function should be smooth, with a constant
control position. This may be stick aft or stick neutral, as required for
decelerating flight in the specific aeroplane configuration. Any other incidence
limiting function, such as an alpha floor function on the engine power, should
be inactive during this test (these functions will be tested in test 2h(6)).
EXAMPLE
In the example in Figure 2h2-1, although the result exceeds tolerance at the
end of the time history, the result is within tolerance during the minimum
speed portion of the test, and therefore it is acceptable.
2H-7
2H-8
Figure 2h2-1
Example of Simulator Test Results for Minimum Speed Protection Function
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TITLE
2h(3) - LOAD FACTOR
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OBJECTIVE
LOAD FACTOR PROTECTION FUNCTION
DEMONSTRATION
From a stabilised flight condition, roll the aeroplane
into a turn, progressively increasing the bank angle
until the load factor envelope protection function
operates and has the time to stabilise the bank
angle or load factor.
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FLIGHT CONDITION
a) TAKEOFF
b) CRUISE
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RECORDED
PARAMETERS
LATERAL CONTROLLER POSITION
AIRSPEED
PITCH ANGLE
NORMAL LOAD FACTOR
BANK ANGLE
STABILISER ANGLE
ELEVATOR ANGLE
EVALUATION NOTES
the load factor protection system can be tested, this
the actual aeroplane flight warning unit or else by
software simulation. Running the test automatically
will probably involve the driving of the pitch and roll
controller positions, so that the test does not bypass
or ignore any interaction between the electronic
flight control system and the load factor protection
system. The method of detecting the onset of the
2H-9
load factor protection should be ascertained, this
can be typically identified by changes in the control
surface angles which do not correspond to the
inputs at the controller position. The intention is to
demonstrate limitation of the load factor at its
maximum permitted value. Note that the test must
be run for both normal and non-normal flight control
system states (see note on page 2H-2).
TOLERANCES
NORMAL ACCELERATION ±0.1 g
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MANUAL TESTING
The test should commence when trimmed for steady level flight, after which
the pilot needs to bank the using the control wheel/lateral controller. The
entry into the load factor limiting function should be smooth, and the
aeroplane should be held against its load factor limit for long enough to
establish the stabilised value. Systems designed to limit bank angle above a
certain load factor should also be demonstrated by this test, which should be
of sufficient length for the aeroplane to obtain a stabilised bank angle in these
conditions. . Care must be taken that the test demonstrates the correct
protection, as even slight mishandling can cause other protections, such as
the AOA, alpha floor or pitch angle protections to become active also. It may
be appropriate to disable, if possible, other such limiting functions, in order
that the load factor protection can be demonstrated in isolation.
2H-10
THIS PAGE LEFT INTENTIONALLY BLANK
2H-11
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TITLE
2h(4) - PITCH ANGLE
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OBJECTIVE
PITCH ANGLE PROTECTION FUNCTION
DEMONSTRATION
From a stabilised flight condition, pull back smoothly
on the control column/longitudinal controller until the
pitch angle protection function operates. Maintain
the input until the pitch angle stabilises.
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FLIGHT CONDITION
a) CRUISE
b) APPROACH
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
PRESSURE ALTITUDE
STABILISER ANGLE
EVALUATION NOTES
the pitch angle protection system can be tested, this
system and the pitch angle protection system. The
method of detecting the onset of the protection
should be ascertained, which is typically identified
by changes in the control surface angles that do not
2H-12
correspond to the inputs at the controller position.
The intention is to demonstrate limitation of the pitch
angle at its maximum permitted value. Note that the
test must be run for both normal and non-normal
flight control system states (see note on page 2H-2).
TOLERANCES
PITCH ANGLE
±1.5o
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MANUAL TESTING
With the aeroplane stabilised in steady level flight, the entry into the pitch
angle limiting function should be made using the control column or pitch
controller and should be smooth, with the input being held for long enough to
establish the stabilised value. Care must be taken that the test demonstrates
the correct protection, as even slight mishandling can cause other protections,
such as the AOA, normal load factor or minimum speed protections to become
active also. It may be appropriate to disable, if possible, other such limiting
functions, in order that the load factor protection can be demonstrated in
isolation.
EXAMPLE
In this example (Figure 2h4-1a & 1b) the pitch angle is stabilised following a
small overshoot at 25 deg by the protection, but before the end of the time
history the minimum speed protection becomes active and starts to reduce the
pitch angle still further without any input from the pilot.
2H-13
Figure 2h4-1a and 1b (below)
Example of Simulator Test Results for Pitch Angle Protection Function
2H-14
2H-15
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TITLE
2h(5) - BANK ANGLE
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OBJECTIVE
BANK ANGLE PROTECTION FUNCTION
DEMONSTRATION
From a stabilised flight condition, roll the aeroplane
smoothly until the bank angle protection function
operates. Maintain the input until the bank angle
stabilises.
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FLIGHT CONDITION
APPROACH
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RECORDED
PARAMETERS
AIRSPEED
BANK ANGLE
AILERON ANGLE
SPOILER ANGLES
PITCH ANGLE
PRESSURE ALTITUDE
STABILISER ANGLE
ELEVATOR ANGLE
EVALUATION NOTES
the pitch angle protection system can be tested, this
probably involve the driving of the roll controller
system and the bank angle protection system. The
method of detecting the onset of the protection
should be ascertained, which is typically identified
2H-16
by changes in the control surface angles that do not
correspond to the inputs at the controller position.
The intention is to demonstrate limitation of the bank
angle at its maximum permitted value. Note that the
test must be run for both normal and non-normal
flight control system states (see note on page 2H-2).
TOLERANCES
BANK ANGLE
±2o OR ±10%
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MANUAL TESTING
With the aeroplane stabilised in steady level flight, the entry into the bank
angle limiting function should be made using the control wheel or lateral
controller and should be smooth, with the input being held for long enough to
establish the stabilised value. Care must be taken that the test demonstrates
the correct protection, as even slight mishandling can cause other protections,
such as the AOA or normal load factor protections to become active as well.
EXAMPLE
See next page.
2H-17
2H-18
Figure 2h5-1
Example of Simulator Test Results for Bank Angle Protection Function
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TITLE
2h(6) - ANGLE OF ATTACK
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OBJECTIVE
ANGLE OF ATTACK PROTECTION FUNCTION
DEMONSTRATION
From a stabilised flight condition above the
minimum speed for the configuration, pull back
smoothly on the control column/longitudinal
controller and hold until the angle of attack
protection function operates.
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FLIGHT CONDITION
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RECORDED
PARAMETERS
AIRSPEED
PITCH ANGLE
ANGLE OF ATTACK
ELEVATOR ANGLE
PRESSURE ALTITUDE
STABILISER ANGLE
EVALUATION NOTES
the angle of attack protection system can be tested,
this may have been achieved either by the utilisation
of the actual flight warning unit or else by software
system and the bank angle protection system. The
intention is to demonstrate protection of the
aeroplane against excessive angles of attack. Note
2H-19
that the test must be run for both normal and
non-normal flight control system states (see note on
page 2H-2).
TOLERANCES
ANGLE OF ATTACK
±1.5o
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MANUAL TESTING
Having ascertained that the simulated aeroplane is in trim for steady level
flight, the entry into the angle of attack limiting function should be smooth, with
a constant longitudinal control position. This may be stick aft or stick neutral,
as required for decelerating flight in the specific aeroplane configuration. Care
must be taken that the test demonstrates the correct protection, as even slight
mishandling can cause other protections, such as the normal load factor or
minimum speed protections to become active also. It may be appropriate to
disable, if possible, other such limiting functions, in order that the angle of
attack protection can be demonstrated in isolation.
The test should continue for long enough to demonstrate recovery from, or
stabilisation at, the high incidence condition.
EXAMPLE
Figure 2h6-1 shows an example of the use of tolerance banding on the plots.
Not considered universally useful, they are appropriate in this test because of
the need to check the simulated system against the aeroplane limitation.
2H-20
Figure 2h6-1
Example of Simulator Test Results for Angle of Attack Protection Function
2H-21
2H-22
SECTION 3
MOTION SYSTEM
3a
FREQUENCY RESPONSE
3b
LEG BALANCE
3c
TURN-AROUND CHECK
3d
MOTION EFFECTS
3e
3f
MOTION CUEING PERFORMANCE SIGNATURE
3g
CHARACTERISTIC BUFFET MOTIONS
3-1
3.0
MOTION SYSTEMS - GENERAL
3.1
INTRODUCTION
The motion envelope of a real aeroplane, both in terms of linear
displacements and angular rotations, is virtually unlimited. It allows for
altitudes up to 50000 feet, displacements of up to 12000 nautical miles and
attitudes of up to ±90 degrees. On the other hand, Figure 3-1 below is an
example of a full flight simulator six-axis synergistic motion system, for
which each actuator (‘jack’) typically has a maximum stroke of between 60
to 72 inches (1.52 to 1.83 metres).
Figure 3-1
Flight Simulator Six-Axis Synergistic Motion System
During the flight of a real aeroplane, the movement of the aeroplane gives
rise to perceivable stimuli to the pilot's sensory organs. These stimuli,
referred to as motion cues, form an important source of information to the
pilot when performing the task of controlling the aeroplane. It is important
therefore to produce the relevant motion cues in a simulator to obtain the
overall fidelity necessary for using it as an adequate training tool.
This fact is recognised by the regulatory authorities, leading to the
requirement that a zero flight-time simulator should be equipped with a
six-degree-of-freedom motion system, producing motion cues
3-2
representative of the aeroplane motions. However, the precise details of
how the motion system hardware is to be checked for its performance (as
a stand alone system) has always been left to the operator of the device.
Any ground-based flight simulator with motion capabilities inevitably has
severe limitations with respect to motion generation. These limitations
apply to linear displacements, velocities and accelerations as well as
angular rotations, rotation rates and rotational accelerations.
As a consequence, modifications have to be made to the signals derived
from the aeroplane's states to keep the motion system from running into
its limits and giving erroneous motion perception to the pilot. Such
modifications mean that resulting cues generated through the motion
system will inevitably deviate from those experienced in the real aeroplane
and so sophisticated use of the available motion capability is required to
avoid adverse effects. Figure 3-2 below illustrates the process by which
calculated values in the simulation modelling are translated into outputs to
the motion system actuators.
Figure 3-2
Simulator Motion System Drive Block Diagram (Simplified)
3-3
3.2
ACTUATOR STROKE REQUIREMENTS
The translational operational excursion envelope (surge, sway and heave)
decrease proportionally with the reduction of the actuator stroke. For a
reduction from a commonly used 60" stroke (1.52m) to, for example, a 24"
stroke (0.61m), the excursion envelope decreases to approximately 40%
of the 60" system envelope, assuming synergistic lay-out. The exact
reduction also depends on detailed motion system geometry. The angular
operational excursion envelope (roll, pitch, yaw) decreases to
approximately 75% (depending on motion system geometry).
In most cases, the available attitudes from a standard 60" stroke
synergistic motion system are sufficient for an acceptable level of
sustained specific force generation.
Reduction of the excursion envelope directly results in an increase of the
false cues during the transition from onset (linear acceleration) cue to
sustained (simulator tilt) cue when simulating sustained specific force (for
example, during take-off, climb-out, braking).
Generally speaking, if insufficient excursion is available, the transition from
onset to sustained specific force cannot be achieved without severe
deformation in both magnitude and direction of the resulting simulated
specific force.
In a well designed motion law scheme it may still be necessary to perform
some input limiting to prevent the motion system from running into its limits
in certain motion critical manoeuvres, such as an emergency stop.
3.3
MOTION SYSTEM TESTS
3.3.1 Historical Background
The frequency response, leg balance and turn-around bump tests have
been the traditional motion tests that have been in existence for many
years. More precise definitions of these items and the way in which they
are applied to flight simulators are given in the succeeding pages.
As always, care should be taken when running these tests, particularly if
they are run in the motion system maintenance mode, which may require
personnel to be in the vicinity of the jacks.
Usually a special program is used to run these tests which bypasses the
main motion drive software normally used for training purposes. This is
3-4
because the tests are not checking the types of cues to be sensed on the
flight deck (these are done in the functional and subjective test section of
the QTG) but the performance of the motion system itself (with its
"payload").
In 2001, a Motion Working Group met to discuss other ways in which a
flight simulator motion system should be tested. The general conclusion of
the Motion Working Group was that the frequency response, leg balance
and turn-around bump tests represented tests of the hardware set-up,
calibration and wear and were normally run in "maintenance mode". This
meant the motion cueing software drives were not being checked in any
way.
The regulators in the working group wanted to insert an additional objective
test that would include the operation of the motion cueing and filtering
software as part of the drive for the test. This resulted in the creation of the
Motion System Repeatability test.
3.3.2 Motion System Repeatability Testing
The Motion System Repeatability test would be a laboratory test of the
motion system's reaction to an injected laboratory input to the motion
cueing software and would be independent of the aircraft characteristics.
The aircraft characteristics would be isolated from the motion test by
injecting predefined test acceleration and rate profiles into the motion
cueing programs in place of the aircraft accelerations and rates which
would normally be produced by the Equations of Motion program. This
injection of laboratory accelerations and rates would take place at a
convenient point between the Equations of Motion software and the Motion
cueing software where the aircraft centre of gravity accelerations and
velocities would normally be transformed into pilot reference point
accelerations and velocities prior to entering the motion cueing program.
The Motion System Repeatability test was therefore described as a
"diagnostic test" in the ICAO document as opposed to the frequency
response, leg balance and turn-around bump tests which assumed the
unfortunate label of being "robotic" tests.
To ensure that very small test input amplitudes could not be used,
guidance was provided in the form of suggested minimums for the test
inputs. Since this was a new concept of test, there was no experience to
draw upon to define the thresholds that were set for these minimums. An
initial guess was proposed with the view that this could be adjusted as
experience was gained. The intent was that these thresholds should
3-5
appear in guidance material rather than in the regulations, but the way the
document evolved precluded this as the guidance material became
absorbed into the section entitled Test Requirements.
Isolating the aircraft characteristics from the motion system ensured that
there was no need for aircraft data to be used. This made the test generic
and independent of the aircraft simulated. Test results across platforms
could differ because of the gains set in the motion cueing software.
It was recognised that software cueing gains could be different for onground and in-air cases to optimise the motion performance. For this
reason, two tests conditions were specified, an on-ground test and an inair test. The shape and form of the injected test input was left to the
discretion of the operator, hence the wording "One test case on-ground: to
be determined by the Operator" and " One test case in-air: to be
determined by the Operator"
The concept of the test was to generate a footprint test during the initial
evaluation, which would produce the Master QTG result. Subsequent
recurrent evaluations would be compared with the MQTG result to highlight
any changes in both the hardware and software performance of the motion
cueing system. Specifically, the amplitude during recurrent testing must
remain within ±0.05g relative to that measured during the initial
qualification.
The second aspect of Motion System Repeatability requires a
‘Performance Signature’ to be taken by running several manoeuvres (both
on ground and in flight) and recording the resultant motion system
accelerations and positions. Examples may include such manoeuvres as
Normal and Engine Inoperative Takeoffs, Rejected Takeoff, Normal and
Engine Inoperative Landings, Speedbrake Deployment, Fast Roll
Response, etc. No tolerances or specific flight conditions are stated in the
ICAO Manual, so the method by which this criteria is fulfilled has been left
up to the simulator manufacturers and operators. However, the
manoeuvres must be repeatable during recurrent testing and must provide
sufficient data for an evaluator to determine that the motion system cueing
performance has not degraded over time.
3-6
SECTION 3a
FREQUENCY RESPONSE
3a
FREQUENCY RESPONSE
3A-1
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TITLE
3a - FREQUENCY RESPONSE
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OBJECTIVE
TO DEMONSTRATE THE FREQUENCY RESPONSE
OF THE MOTION SYSTEM WHEN SUBJECTED TO
AN OSCILLATORY INPUT.
DEMONSTRATION
Using the appropriate motion test facilities, drive each
actuator independently with a sinusoidal signal.
Record and analyse actuator feedback signals.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
REFERENCE (DRIVING) INPUT SIGNAL
ACTUATOR POSITION FEEDBACK SIGNAL
EVALUATION NOTES
The purpose of this test is to determine the phase lag
and the attenuation experienced by the motion system
when subjected to an oscillatory input. It is normally
run at two separate frequencies, a slow frequency
(typically 0.1 Hz) and a higher frequency (typically 0.5
Hz). At 0.1 Hz, the phase lag between the reference
input and the actuator position feedback will usually be
of the order of less than 10 degrees, whereas at 0.5
Hz the phase lag will be more in the region of 30
degrees, though some variation can be expected for
different motion systems. The attenuation will usually
be greater than -1.0 dB for both cases, though
naturally both phase and gain are frequency
dependent.
This check should be run on all 6 actuators
independently, with the results shown on a print out,
which may be from the simulator line printer or a
separate multi-channel recording device. The
tolerances have been left to the discretion of the
simulator operator, but results should be submitted as
3A-2
part of the initial QTG and should be available for
inspection at each evaluation.
TOLERANCES
AS SPECIFIED BY THE
SIMULATOR ACCEPTANCE.
OPERATOR
FOR
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MANUAL TESTING
It is not recommended that this test is performed manually because of the special
driving signals needed to allow the test to be run correctly so as to permit
analysis of the results. In any case it is not the motion cues experienced by the
flight crew which are under scrutiny, but the physical response of the motion
system itself.
EXAMPLE
Figures 3a-1 and 3a-2 on the next page shows two differing sets of simulator test
results for Motion System Frequency Response. The top set is in the form of a
table which gives the phase and gain values for each jack after running the test
at a particular frequency. This table will be accompanied by plots of the response
of each of the actuators driven simultaneously by a sinusoidal reference input.
The lower set is in the more conventional Bode Plot format, which shows the
phase and gain over a range of frequencies against a logarithmic frequency axis
(1Hz, 10Hz, 100Hz in this case). Note that the response falls away rapidly above
approximately 10 Hz.
3A-3
Figure 3a-1
3A-4
Figure 3a-2a
3A-5
Figure 3a-2b
3A-6
SECTION 3b
LEG BALANCE
3b
LEG BALANCE
3B-1
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TITLE
3b - LEG BALANCE
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OBJECTIVE
TO
DEMONSTRATE
THE
DYNAMIC
PERFORMANCE OF THE MOTION SYSTEM.
DEMONSTRATION
Using the appropriate motion system test facilities,
drive each actuator independently and examine the
results to determine the effects on the undriven
actuators.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
REFERENCE (DRIVING) INPUT SIGNAL
ACTUATOR POSITION FEEDBACK SIGNAL
EVALUATION NOTES
"Leg Balance", or more correctly "Dynamic Stiffness",
is tested to ensure the simulator motion actuators are
not unduly affected by external forces applied to them.
Since the majority of actuators employed in six-axis
motion systems are hydrostatic, it follows that their
friction is very small and that the inherent damping of
the oil column itself is also very small. To provide the
necessary damping for stability of the servos, a signal,
proportional or approximately proportional to the
acceleration of the actuator, is generated. The most
common method of producing this signal is to
determine the actuator force by use of pressure
transducers or load cells. The actuator force signal is
applied to a high pass filter to remove the d.c. and low
frequency components. When an external force is
applied to the actuator, i.e. the reaction force
generated as a result of the displacement of another
actuator, the computed force signal generated as a
result of the reaction force in the 'static' actuator
causes this actuator to displace. The purpose of this
test is to establish that this displacement is maintained
3B-2
within given limits following the application of an
external force to the actuator.
This check is run on all 6 actuators independently,
with the results shown as plots. There will typically be
two frequencies used, examples being 0.5 Hz and 3.0
Hz.
The plots in Figures 3b-1 and 3b-2 show an example
set of results for one jack being driven at 0.5 Hz and
at 3 Hz. For a six-axis synergistic motion system there
will always be some movement of the other 5 jacks,
but the movement of the 5 undriven actuators should
normally be less than 5% fraction of full scale (peak to
peak) of the drive signal amplitude.
TOLERANCES
AS SPECIFIED BY THE
OPERATOR
FOR
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MANUAL TESTING
system itself.
EXAMPLE
See Figures 3b-1 and 3b-2.
3B-3
Figure 3b-1
Example of Simulator Test Results for Motion System Cross-Drive (Leg Balance) at 0.5 Hz
3B-4
Figure 3b-2
Example of Simulator Test Results for Motion System Cross-Drive (Leg Balance) at 3 Hz
3B-5
3B-6
SECTION 3c
TURN AROUND
3c
TURN AROUND
3C-1
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TITLE
3c - TURN AROUND
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OBJECTIVE
TO DEMONSTRATE THAT THE MOTION SYSTEM
RESPONSE DURING OSCILLATORY
MANOEUVRES DOES NOT EXHIBIT EXCESSIVE
NOISE WHEN THE DIRECTION OF THE DRIVING
SIGNAL IS BEING REVERSED.
DEMONSTRATION
Using the appropriate motion test facilities, drive and
plot each actuator independently to determine the
degree to which noise is present in the actuator
feedback signal.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
REFERENCE (DRIVING) INPUT
PLATFORM ACCELERATIONS
EVALUATION NOTES
With any servo driven system - including a flight
simulator motion system - noise will be present which,
if excessive, may be perceived by the flight crew as an
extra vibration or turbulence effect which should not
actually be there. To conduct this test the
accelerometer produces a sine wave in phase with the
reference input signal. The peak noise spikes must be
imperceptible to the flight crew on the linear portions
of the graph. Typically, this will mean less than 0.02g.
This check should be run on all 6 actuators
independently, with the results shown as plots on the
print out.
See Figure 3c-1 for example plot of one jack (actuator)
versus the drive reference demand.
TOLERANCES
3C-2
AS
SPECIFIED
BY
THE
OPERATOR
FOR
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MANUAL TESTING
system itself.
EXAMPLE
The example below (Figure 3c-1) shows the response of two motion system
actuators (‘jacks’) when subjected to a sinusoidal reference drive input at 0.5 Hz.
Both plots exhibit reversal characteristics at peak amplitude, i.e. there are visible
non-linearities just as the direction of travel is beginning to reverse, though the
nature of the non-linearities is slightly different between the two actuators. The
deviation from the ‘standard’ sinusoidal path can be clearly seen in both these
plots, however all such deviations are within the ±0.01g tolerance limit set by the
simulator manufacturer.
3C-3
Figure 3c-1
Example of Simulator Test Results for Motion System Turn Around/Smoothness
(Actuator #1, upper plot & Actuator #5, lower plot)
(Platform heave motion based on a reference drive demand at 0.5 Hz)
3C-4
SECTION 3d
MOTION EFFECTS
(These requirements are stated as being Validation Tests, but are specified
in the Functions and Subjective Testing section of the ICAO Manual)
3d(1)
Effects of Runway Rumble, Oleo Deflections, Ground
Speed, Uneven Runway, Runway Centreline Lights
and Taxiway Characteristics
3d(2)
Buffets on the Ground Due to Spoiler/Speedbrake
Extension and Thrust
3d(3)
Bumps Associated with the Landing Gear
3d(4)
Buffet During Extension and Retraction of Landing
Gear
3d(5)
Buffet in the Air Due to Flap and Spoiler/Speedbrake
Extension
3d(6)
Approach to Stall Buffet
3d(7)
Touchdown Cues for Main and Nose Gear
3d(8)
Nosewheel Scuffing
3d(9)
Thrust Effects with Brakes Set
3d(10)
Mach and Manoeuvre Buffet
3d(11)
Tyre Failure Dynamics
3d(12)
Engine Malfunction and Engine Damage
3d(13)
Tail Strikes and Pod Strikes
3D-1
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TITLE
3d - MOTION EFFECTS
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OBJECTIVE
TO CONFIRM THAT THE SIMULATOR MOTION
BUFFETS EXPERIENCED DURING VARIOUS
FLIGHT CONDITIONS ARE QUALITATIVELY LIKE
THE AEROPLANE.
DEMONSTRATION
Taxi and fly the simulated aeroplane at various speeds
and flight conditions and note the onset, amplitude,
frequency and general quality of the simulator buffet.
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FLIGHT CONDITION
ALL (see title page)
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RECORDED
PARAMETERS
QUALITATIVE ASSESSMENT ONLY IS REQUIRED
EVALUATION NOTES
This test is only required to check the motion buffets,
but the general ‘feel’ of the main motion system
simulation can also be subjectively examined at the
same time. Whilst some of the flight conditions and
aeroplane configurations may be set up using the
simulator autotest system, because of there being no
requirement to meet tolerances, along with the
necessary pilot input which is required during this test,
automatic running and checking against tolerances is
not possible. The reader is referred to Volume II of this
Handbook, as these tests essentially fall into the
category of ‘Functions & Subjective Tests’.
TOLERANCES
NONE
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MANUAL TESTING
The tests require the simulator to be flown in all the specified regimes.
3D-2
SECTION 3e
3e
Motion System Repeatability
3E-1
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TITLE
3e - MOTION SYSTEM REPEATABILITY
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OBJECTIVE
SYSTEM CONTINUES TO PERFORM AS
ORIGINALLY QUALIFIED
DEMONSTRATION
Drive the motion system and record the response in
such a way as to be able to determine that the
actuators are maintaining the driven amplitudes within
the prescribed tolerances
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FLIGHT CONDITION
NONE
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RECORDED
PARAMETERS
TIME
MOTION LINEAR ACCELERATION DEMANDS
MOTION ROTATIONAL ACCELERATION
DEMANDS
MOTION ROTATIONAL VELOCITY DEMANDS
MOTION LINEAR ACCELEROMETER - X, Y, Z
MOTION JACK POSITIONS
EVALUATION NOTES
This test is required to make sure that the motion
system is being properly maintained over the life of
the simulator. It is a test of both the motion
hardware and the motion cueing and filtering
software to monitor change. The test represents the
motion system's reaction to a series of demands to
the motion cueing software and is totally
independent of aircraft data. Pre-defined demands
are injected at the point where the aircraft centre of
gravity accelerations and velocities would be
transformed into the pilot reference point
accelerations and velocities prior to entering the
motion cueing software.
3E-2
Instrumentation requirement for this test will be linear
accelerometers. There is no requirement for angular
accelerometers. The pre-defined driving signal has
not been specified, allowing individuals to produce
their own, but the intent of the test is to prove motion
responses to both linear and rotational acceleration
stimulation. Two tests are to be run, one in an
on-ground state and the other in a in-air state. This is
to cater for possible differences in motion system
gains for the on-ground and in-air conditions.
For the initial qualification, the test would be run to
create a master footprint which would not have any
criteria for comparison. For recurrent qualifications,
the amplitudes of the accelerations achieved should
remain within ±0.05g of the original MQTG response.
TOLERANCES
ACTUAL PLATFORM ±0.05g
LINEAR ACCELERATIONS
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MANUAL TESTING
Manual testing is not required to fulfil this requirement.
EXAMPLE
Figure 3e-1 (six diagrams on the following pages) shows the Master footprint
result for the Motion Repeatability Test - On Ground. Note the linear
acceleration inputs have been provided for X, Y and Z as independent inputs,
with responses both in terms of accelerations measured by accelerometer and
also with motion jack positions. Angular acceleration/rates inputs in roll, pitch
and yaw are then provided as independent inputs.
3E-3
Figure 3e-1a (and 1b through 1f following pages)
Example of Motion System Repeatability Test Results
3E-4
Figure 3e-1b
3E-5
Figure 3e-1c
3E-6
Figure 3e-1d
3E-7
Figure 3e-1e
3E-8
Figure 3e-1f
3E-9
3E-10
SECTION 3f
MOTION CUEING PERFORMANCE
SIGNATURE
3f(1)
Normal Takeoff Signature
3f(2)
Engine Inoperative Takeoff Signature
3f(3)
Power Change Dynamics Signature
3f(4)
Flap Change Dynamics Signature
3f(5)
Gear Change Dynamics Signature
3f(6)
Normal Landing Signature
3f(7)
All Engine Autopilot Go-Around Signature
3F-1
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TITLE
3f - MOTION CUEING PERFORMANCE
SIGNATURE
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OBJECTIVE
SYSTEM CONTINUES TO GIVE ADEQUATE
CUEING AS ORIGINALLY QUALIFIED
DEMONSTRATION
Perform several manoeuvres on ground and in flight
and record the motion platform responses in such a
way as to be able to determine that the motion
system cueing is consistent over the life of the
simulator
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FLIGHT CONDITION
GROUND AND FLIGHT
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RECORDED
PARAMETERS
TIME
USUAL PARAMETERS FROM THE RECORDED
PARAMETER LIST FOR THE ORIGINAL SOURCE
TEST
LINEAR ACCELERATIONS AT PILOT
REFERENCE POINT
ANGULAR ACCELERATIONS AT PILOT
REFERENCE POINT
ANGULAR RATES AT PILOT REFERENCE POINT
MOTION ACTUATOR POSITIONS
MOTION PLATFORM LINEAR DISPLACEMENT
AND ANGULAR POSITION
MOTION LINEAR ACCELEROMETER - X, Y, Z
EVALUATION NOTES
This test is required to make sure that the motion
system cues in various flight regimes remain
consistent and also that this cueing is being properly
maintained over the life of the simulator. No
tolerances are prescribed, but the tests are to check
the ability of the motion system to give repeatable
cues and so that the response should not markedly
3F-2
deviate from the initial qualification performance
when undergoing recurrent testing. This will demand
that a method of checking this repeatability is used,
such as overplotting the initial response with the
recurrent results.
The proposed method of demonstrating this
requirement is to run several tests from the
Performance and Handling Qualities sections of the
ICAO Manual, but providing motion system cueing
plots in addition to the usual performance
parameters. These will typically include plots of
motion platform accelerations in all six axes, but
there is no requirement to compare the simulator
accelerations with those of the aircraft. The tests
which typically may be selected are as follows:
Source Test
Number
Test Title
1b(4)
Normal Takeoff
1b(5)
Engine Inoperative Takeoff
2c(1)
2c(2)
2c(4)
2e(1)
Normal Landing
2e(6)
All Engine Autopilot Go-Around
It was specifically stated by the ICAO working group
that for this type of test the motion platform
performance should in no way be compared with the
aircraft performance. These tests were to be run as
footprints only for the initial evaluation and were not
necessarily to be used for recurrent checks. The
only time they would be used was if the regulator
wanted to do a comparison if perceived changes
had occurred to the Motion System Repeatability
test. This would provide some additional indication
of the possible impact on the training value by the
degradation of the motion system.
3F-3
It is also the case that if the simulator motion system
has undergone some modification (agreed with the
regulator) then the footprint Motion Cueing
performance signature must be re-run to form an
updated master record of the motion performance.
This creates a somewhat unsatisfactory situation in
that a simple comparison and tick cannot
necessarily now be done for example on a monthly
or quarterly basis, but tolerances were thought to be
inappropriate and so no tolerances were set.
TOLERANCES
NONE
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
Manual testing is not recommended to fulfil this requirement, as the main
criteria by which the results are judged are that the cues are totally repeatable.
EXAMPLE
The plots of Figure 3f-1 shows 1b(4) Normal Takeoff with the additional
motion related plots. This is done to provide a general example of a Motion
Cueing Performance Signature Test. Please note that the full complement of
plots for the original 1b(4) source test have not been provided, but this
example places emphasis on the additional motion related plots.
3F-4
Figure 3f-1a
Motion Cueing Performance Signature - Normal Takeoff
3F-5
Figure 3f-1b
3F-6
Figure 3f-1c
3F-7
Figure 3f-1d
3F-8
Figure 3f-1e
3F-9
Figure 3f-1f
3F-10
Figure 3f-1g
3F-11
Figure 3f-1h
3F-12
Figure 3f-1i
Motion Cueing Performance Signature - Norm al Takeoff
3F-13
Figure 3f-1j
3F-14
SECTION 3g
CHARACTERISTIC BUFFET MOTIONS
3g
A test with recorded results and a Statement of
Compliance are required for characteristic buffet
motions which can be sensed at the flight deck.
The motion buffets required are:
3g(1)
THRUST EFFECTS WITH BRAKES SET
3g(2)
LANDING GEAR EXTENDED BUFFET
3g(3)
FLAPS EXTENDED BUFFET
3g(4)
SPEEDBRAKE DEPLOYED BUFFET
3g(5)
APPROACH-TO-STALL BUFFET
3g(6)
HIGH SPEED OR MACH BUFFET
3g(7)
IN-FLIGHT VIBRATIONS
3G-1
3G.0 MOTION SYSTEM VIBRATION TESTS
3G.1 INTRODUCTION
The simulator vibration system is used to provide motion effects for
buffets and vibrations (eg runway rumble), which are not simulated in
the aeroplane dynamics, and so do not automatically follow through with
the motion system drive signals from the main aeroplane dynamics. The
system of representing these types of cues tends to use a combination
of sine wave frequency drives, random noise sources and step
displacements.
Vibration data recorded on a flight test aeroplane will have been
analysed by the simulator manufacturer and the simulator hardware
and/or software adjusted to approximate the responses of the
aeroplane. Acceleration against time plots and spectral decomposition
are used to analyse the aeroplane data, and the same measurement
techniques are used on the simulator for comparison.
Vibrations experienced on the flight deck are generated in the simulator
hardware and software and added into the main motion drives.
Vibrations simulated must include buffets due to gear, flaps, airbrake,
engine vibration, runway rumble, high speed buffet, stall buffet and
turbulence buffet.
The vibration levels on jet transport aeroplanes during normal operating
conditions are usually fairly low. The main exceptions are high
speed/stall buffet and runway roughness which must be simulated to an
adequate level for training purposes.
3G.2 VIBRATION RECORDING
Recorded data showing the frequency, magnitude, and orientation of
the vibrations and similar effects that can be felt or observed in the
cockpit during ground and flight operations are presented and analyzed
in the simulator data document(s) which define the aeroplane vibration
characteristics. Such data is intended to provide simulator
manufacturers and simulator operators with environmental information
for use in approximating the vibratory sensations associated with the
handling qualities of the aircraft.
The dynamic response of a conventional aircraft to external
3G-2
disturbances can be considered in terms of independent vertical
(normal), longitudinal, and lateral rigid body movements, and oscillatory
modes associated with structural flexibility.
The aeroplane response in the rigid body modes is at low frequencies
and these movements can be detected visually by the pilot through
observations of the relative motion of the flight deck compared to a
stationary external reference point such as the horizon, runway
markers, or a specific landmark. This motion is also displayed to the
pilot by the instruments and may be observed as relatively slow
fluctuations in quantities such as airspeed and flight path deviations.
Although rigid body motions are most certainly felt by the flight crew, the
most apparent physiological impression is visual.
The oscillatory modes associated with structural flexibility are referred to
as the aircraft structural modes. These modes significantly affect the
vibration environment of the flight deck since the effect of structural
flexibility is to increase the accelerations to which the flight crew is
exposed. The situation is aggravated by the location of the flight deck at
the fuselage extremity for all jet transport aeroplanes. For modes
involving fuselage motion the amplitude of the response can be a
maximum at this location, and the aircraft response to these modes is
perceived by the flight crew through subjective impressions and from
motion cues.
In practical flight situations the exact frequency values associated with
these modes are variable and dependent upon the distribution of
passengers and/or cargo in the fuselage, the fuel loading configuration,
flap setting and the airspeed at which the aircraft is being flown. As the
aircraft experiences various flight maneuvers and ground operations,
energy is imparted to the structure from external forces acting on the
vehicle. The aircraft reacts to this energy input by responding in a
combination of rigid body and structural modes. The particular
combination of rigid response and structural response is dependent
upon the nature, orientation, and duration of the external disturbance.
Normal, longitudinal, and lateral accelerations may have been recorded
during the aeroplane flight test program, during test flights intended to
record engineering data relative to aircraft stability, control, handling
qualities. However these recordings are usually at low sample rate and
made from measurements close to the aircraft center of mass, and
therefore only provide information regarding the rigid body movement of
the airframe.
3G-3
For aircraft structural mode vibration analysis specific recordings are
required, taken from three accelerometers installed at the base of the
pilot's (captain's) seat. Vertical (normal), lateral and longitudinal
accelerations are recorded in flight and processed post-flight to obtain
"quick-look" time history information. Time histories are assembled and
categorized according to the external vibration source and type of
maneuver performed, then scrutinized and certain traces identified
which display typical flight deck vibration responses to the various
categories of vibration. These typical traces can then be further
processed to obtain final data suitable for use in simulator design and
qualification.
Simulated vibration is measured using an accelerometer package fitted
under the simulator. The exact position of the accelerometer package
may vary from one simulator manufacturer to another, though it is
typically placed at or near the centre of the simulator motion platform
frame. This is sufficiently close to the pilots' seats in most, if not all,
simulators to render the results more than adequate for comparison with
the flight test data from the aeroplane.
3G.3 PERCEPTION OF VIBRATION
The human pilot is sensitive to both the amplitude and the frequency of
vibration. Research has determined that humans are particularly
sensitive to vibratory accelerations occurring in the frequency range
from 1-10 Hz. Human physiological sensitivity to vibration is dependent
on body position, method of support relative to the direction of
oscillation, and what portions of the body resonate. The whole body
natural frequency for a human when seated occurs at 3-5 Hz, and
disorientation (vertigo) can occur at about 1 Hz. The maximum motion
of the head relative to the seat occurs at 3-6 Hz. If vibration levels
become severe, breathing becomes difficult for structural responses
within the range 1-4 Hz, and chest pains result from 3-10 Hz
oscillations.
Because of the characteristics of the human body, flight deck
accelerations occurring in the 1-10 Hz frequency range are considered
significant for the purposes of qualification testing, although data is
usually presented up to about 20hz, which is where the human hearing
system becomes active.
3G.4 THE CONCEPT OF POWER SPECTRAL DENSITY
3G-4
It is virtually impossible to identify the dominant frequencies present in a
recording by visual inspection of acceleration against time plot alone. To
accurately identify the frequency content of an acceleration recording,
the signal needs to be processed to produce a plot of amplitude against
frequency; this conversion is performed via a mathematical process
called Fourier Transformation, named after Jean Baptiste Joseph
Fourier (1768-1830), a French mathematician and physicist.
The processed signal is normally presented using units of “power
density”, rather than simple acceleration, on the amplitude axis. Power
density is an additive quantity that gives an indication of the energy and
thus perceived magnitude of the vibration at a particular frequency. The
resulting graphs are commonly called Acceleration Power Spectral
Density (APSD) plots. The human pilot can perceive vibrations over an
energy range of several orders of magnitude, and so the power
spectrum plots are normally rectilinear displays of log-magnitude versus
frequency.
"Peaks" in the power spectrum correspond to the resonant frequencies
associated with the vibration modes of the aircraft. By observing the
character of the power spectrum, the particular response modes and the
relative energy levels of each mode can be discerned for the various
sources of vibration under scrutiny. Due to the short duration of many
test conditions in the data test acquisition programs, data below 1 Hertz
is not usually considered.
3G.5 CONTROL OF THE DISCRETE FOURIER TRANSFORMATION
The Fourier transformation, as originally described in 1807 is a
mathematical ideal, which holds true for perfectly repeating signals,
recorded over infinite periods, and then considered using infinitely
narrow frequency bands. In practice of course real life limitations
degrade the quality of the transformation, and the resulting APSD is
only an approximation to the true content of the signal. These limitations
to the quality of the APSD apply equally to analysis conducted using
purpose made equipment or via the simulator testing software. The
analyser may provide some control over certain parameters used during
the transformation, of which Windowing, Bandwidth, Averaging, Overlap
and Smoothing are the most important. It is preferable to analyse both
the aircraft data and the simulator result using the same equipment and
the same settings, then direct comparison can be conducted between
the two APSD plots. But where this is not possible, the control
parameters should be similar for the processing of the aircraft data and
3G-5
the simulator data, and allowances made when comparing the resulting
APSD. The following sub-sections provide a brief introduction into how
the major control parameters affect the resulting APSD.
3G.5.1 Windowing
The finite sampling period, of typically a few seconds, will introduce
erroneous frequency content on the APSD due to the sudden start and
stop of the signal. Fading the signal in and out, at the beginning and end
of the recording minimizes this effect. This is called windowing, and the
shape of the fade function used is called the windowing function. The
disadvantage to using a windowing function is that the fade function will
inevitably reduce the overall magnitude of the signal being analysed. If a
rapid the fade is achieved the overall magnitude is affected less but
more erroneous frequency content is added, which can swamp the
underlying signal at low frequencies. In the plots in Figure 3g-1 the
same signal has been analysed using three of the most common
windowing functions.
3G-6
Figure 3g-1
Vibration Analysis Windowing Functions
3G.5.2 Bandwidth
Both purpose made equipment and the simulator testing software
operate by sampling the input signal, and then processing these
discrete samples using a Discrete Fourier Transformation (DFT). The
sample rate and number of samples used during the DFT dictate the
bandwidth of the resulting APSD, a parameter that critically affects the
character and magnitude of the resulting trace.
The following plots are for exactly the same simulator recording
processed so as to generate APSD plots with different bandwidths.
Figure 3g-2a
APSD Plot, Processed using 0.25 hz Bandwidth
Figure 3g-2b
Clearly the increased bandwidth has smoothed the peaks of the trace
and reduced the apparent magnitude, if direct comparison is to be
conducted between the aircraft trace (dotted) and the simulator result
then both should be processed to the same bandwidth.
3G.5.3 Averaging and Overlap
If the signal being analysed contains a significant random element, then
a more repeatable result will be obtained if the sampling and Fourier
3G-7
transformation stages are repeated for several times and the results
averaged. The aircraft trace will usually contain more random variation
than the simulator result, and therefore benefit more from averaging.
Averaging is usually used in conjunction with windowing, where each
successive ‘average’ overlaps the former, so as to ensure that parts of
the signal which may have been masked by the window fade function
are analysed at full magnitude by the following average.
Using a single pass analysis, when the windowing function gain (Figure
3g-3a, dotted) is multiplied with the input signal, a large percentage of
the signal is processed at less than full gain.
Figure 3g-3a
Single Pass Analysis
By averaging the results taken from multiple pass analysis, a greater
percentage of the signal is processed at high gain, and sporadic bursts
of activity are less likely to be masked by the windowing function.
Figure 3g-3b
Multiple Pass Analysis
However, if too many averages are performed upon a short duration
signal, then the benefits of the windowing function are lost, as the fade
in and out of the signal by the windowing function becomes too rapid.
3G-8
3G.5.4 Smoothing
Some data vendors further process the APSD produced by Fourier
transformation by simply smoothing the resulting trace. The effect is to
reduce the peak, and increase the trough values, and can be easily
accounted for if the analyser does not support post-analysis smoothing.
3G.6 ANALYSIS OF VIBRATION
Usually, two types of final data plots are prepared for most of the
operating conditions:
1. Time histories of the normal, lateral, and longitudinal accelerations
are constructed and annotated with sufficient supplementary aircraft
parameters to adequately define aeroplane configuration and flight
condition.
2. Power spectral density plots are created for each axis of acceleration.
These plots display frequency from 1.0 to 20.0 Hz as the abscissa and
acceleration power spectral density, in G2/Hz units, as the ordinate.
For each flight case checked, a set of time histories for X, Y and Z axes
should be provided from the simulator which allows side-by-side viewing
with similar plots from the aeroplane.
Comparisons of spectral analysis plots are provided for the X, Y and Z
axes. Some spectral densities may be very small and therefore difficult
to drive with any accuracy, so under these circumstances plots may not
be provided for comparison. For the spectral analysis plots provided, the
simulator vibration levels should have been adjusted to match the
spectral analysis data of the aeroplane within reasonable bounds. No
formal tolerances exist for these tests and the results may be assessed
using engineering judgement based on the general guidelines given in
this document.
For each axis, the simulator vibration system will usually provide more
than one noise source of slightly different frequency content and several
periodic frequency drives. As a minimum, these will probably be
low-frequency drive (0-4 Hz), mid-frequency drive (4-10 Hz) and
high-frequency drive (10-30 Hz), though there may be more than these
for some simulators or from some simulator manufacturers. The
predominant frequency in each of these frequency ranges will have
been chosen from the flight test data and programmed into the vibration
3G-9
software. Occasionally there may be two predominant frequencies close
to each other. In this case the simulator may have been programmed
with the frequency that 'looks' most predominant on the time history plot
and/or feels most predominant to the pilot, or else the simulator
manufacturer's engineers may have decided to represent both
frequencies with a single one placed equidistant between them. Having
adjusted the vibration levels to match the test data, the time history is
checked and compared against the equivalent aeroplane time history.
Comparison of simulator to aeroplane data can only be approximate
and the simulator manufacturer should have made use of pilot
assessment during tuning of the vibration.
The characteristics of the Motion system may well be such that 'cross
coupling' of the effects are evident, particularly at medium to high
frequencies. The effect is such that, when driving high axis vibrations in
one axis (e.g. the lateral axis, 'Y'), a moderate amount of vibration in
another axis (e.g. the vertical axis, 'Z') will be induced. This will be
noticed predominantly on the spectral plots. Given the nature of a flight
simulator and its motion system this may be unavoidable and should
therefore be taken into consideration when reviewing the results.
3G.7 SIMULATED VIBRATIONS
The ICAO Manual is now definitive with regard to exactly which
vibrations and buffets are required to be tested for the highest level
simulator and these are as follows:
1
2
3
4
5
6
7
3G-10
Thrust Effect with Brakes Set
(conducted on the ground at maximum possible thrust with
brakes set)
Landing Gear Extended Buffet
(conducted at a normal operational speed)
Flaps Extended Buffet
Speedbrake Deployed Buffet
Approach-to-Stall Buffet
(conducted only for approach-to-stall; post-stall characteristics
are not required)
High Speed or Mach Buffet
(conducted for high speed manoeuvre buffet/wind-up turn or
alternatively Mach buffet)
In-Flight Vibrations
(conducted to be representative of in-flight vibrations for
propellor-driven aeroplanes)
Note. For some aeroplanes, there may be no data available for some of
the conditions. Under these circumstances no objective comparison can
be made but simulator plots can still be useful in order to provide
objective results which can be used as a reference point for future
evaluations once the effect has been subjectively assessed as being
acceptable.
3G.8 SUBJECTIVE TUNING
There may well be occasions when pilots are not happy with the
amplitudes at a particular QTG check point. Assuming that the
conditions and scalings have been checked objectively with the time
histories and power spectral density plots, then the best that can be
done is to tune the drives high or low as far as can be permitted within
the limitations of achieving a reasonable match.
3G.8 REVIEW OF OBJECTIVE TESTS
Objective matching of motion vibration is required for the highest level of
simulator qualification. Generally, the airframe manufacturers produce
two types of data for these conditions: time histories and spectral
analysis, both of which are normally provided for the simulator. The time
histories show the acceleration (in g's) against time for the X, Y and Z
axes. The spectral analysis plots show the power spectral density
(G2/Hz) against frequency for X, Y and Z axes (see Figure 3e-1 for an
example), and are used to demonstrate that the simulator has a
frequency content very similar to that of the aeroplane in each of the
above conditions. Principally, the flight simulator results should exhibit
the overall appearance and trends of the aeroplane plots, with at least
some of the frequency ‘spikes’ being present within 1 or 2 Hz or the
aeroplane data.
The digitization process of sampling can also introduce spurious spikes
on the trace, by way of aliasing higher frequencies present at the
accelerometer, or even electrical noise introduced by the cabling. A well
designed vibration data acquisition system will include analogue
anti-aliasing filters to as to prevent these problems.
In some simulator installations the accelerometers may be sensitive to
3G-11
"rumble" sounds from the loudspeakers which may be capable of
masking the vibration signature on the time histories. The tests should
be run with the sound system switched off, but may also be run with the
sound system on for purposes of comparison. Differences in the plots
obtained can then usually be explained quite easily.
3G.9 TEST METHODOLOGY
There follows a single generic description of the procedure for running a
characteristic motion buffet test. For more details on the individual
scenarios under test the reader is referred to Section 3d and also the
ICAO Manual itself. The set-up for each condition will obviously be very
similar to that specified in Section 3d, the main difference being that the
simulated aeroplane must be held - against normal flying procedures if
necessary - within that buffet regime for a fixed period of time whilst the
measurements are being taken and analysed by the equipment. The
different requirements of Sections 3d and 3e therefore make it difficult to
run the two sets of tests concurrently.
3G-12
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TITLE
3g - CHARACTERISTIC BUFFET MOTIONS
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OBJECTIVE
BUFFETS ARE CHARACTERISTIC OF THE
AEROPLANE.
DEMONSTRATION
Position and maintain the simulated aeroplane in the
required configuration and record the simulator x-,
y-, and z-axis buffet responses.
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FLIGHT CONDITION
THRUST EFFECTS WITH BRAKES SET
LANDING GEAR EXTENDED BUFFET
FLAPS EXTENDED BUFFET
SPEEDBRAKE DEPLOYED BUFFET
APPROACH-TO-STALL BUFFET
HIGH SPEED OR MACH BUFFET
IN-FLIGHT VIBRATIONS
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RECORDED
PARAMETERS
VERTICAL POWER SPECTRAL DENSITY vs
FREQUENCY
LONGITUDINAL POWER SPECTRAL DENSITY vs
FREQUENCY
LATERAL POWER SPECTRAL DENSITY vs
FREQUENCY
VERTICAL ACCELERATION (G) vs TIME
LONGITUDINAL ACCELERATION (G) vs TIME
LATERAL ACCELERATION (G) vs TIME
EVALUATION NOTES
No particular guidance can really be given on the
evaluation of results, but the general points at the
beginning of this section should be borne in mind
when engineering judgement is exercised.
Principally, the simulator results should exhibit the
overall appearance and trends of the aeroplane
plots, with at least some of the frequency 'spikes'
being present within 1 or 2 Hz of the aeroplane data.
3G-13
The positioning of the accelerometer is important,
since the measured accelerations should be
representative at the same point on the simulator as
on the aeroplane - typically the data will have been
gathered at a point beneath the pilot's seat.
These tests are not required to be demonstrated
for a Level C or below simulator qualification.
TOLERANCES
NONE
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MANUAL TESTING
Position the simulated aeroplane at the prescribed point within the appropriate
motion buffet regime. A typical period for which this condition would have to be
maintained is 20 to 30 seconds. The motion system platform vibrations should
be plotted in all three linear axes and this will almost certainly entail the use of
a spectral analyser. Understanding exactly how the analyser functions is not
necessary for the evaluation of the test results. Results should be hard-copied
for comparison with aeroplane data. Ideally the scales should be the same for
both aeroplane and simulator data sets.
EXAMPLE
Several sets of test results for both time histories and power spectral density
(in two axes) is shown on the following pages in Figures 3g-4 to 3g-9
inclusive. Note that these are only representative, and therefore should not be
taken as being valid for all (or any) aeroplane, either for flap buffet or any
other flight condition. What they are intended to illustrate is that the evaluation
of such results must primarily look at the trends of the recordings, along with
certain of the ‘spikes’ that should be present at or near the frequencies at
which they may be seen on the aeroplane data.
3G-14
Figure 3g-4
Example of Simulator Test Results for Flap Buffet Amplitude Time History (Y- and ZAxes Only)
Figure 3g-5
Example of Aeroplane Manufacturer’s Data for Flap Buffet - Amplitude Time History
3G-15
Figure 3g-6
Example of Simulator Test Results for Flap Buffet - PSD Plots (Y- and Z-Axes Only)
3G-16
Figure 3g-7
Example of Aeroplane Manufacturer’s Data for Flap Buffet - PSD Plots
3G-17
Figure 3g-8
Example of PSD Plot Obtained from Stand-alone Test Equipment
Figure 3g-9
Example of Time History Plot Obtained from Stand-alone Test Equipment
3G-18
SECTION 4
VISUAL SYSTEM
4a
4b
DISPLAY SYSTEM TESTS
4c
4-1
4.0
VISUAL SYSTEMS - GENERAL
Most, if not all, advanced flight simulators are built with a visual system
which employs computer generated imagery. The testing of these systems
entails the use of methods and techniques with which a simulator training
captain or evaluation pilot may not be very familiar. In particular, technical
terms such as foot-lamberts (or candles per square metre) are not
encountered on an every day basis and as a result may well be alien to him
- at least initially. Nevertheless, an understanding of these and other terms
and the way in which they are applied to the evaluation of flight simulator
visual systems is fundamental if the evaluator is to perform his function
properly.
The requirements on modern visual systems, and especially those for which
qualification is being sought, are stringent and well defined. For a lower
level qualified device, the visual must provide dusk and night scenes,
whereas for the higher level daylight scenes are necessary. The overall
brightness capability of any system must be such that realistic simulation of
aeroplane landing lights and lights on the ground, as well as significant
topographical feature, is provided. Naturally, the field of view from each
pilot station and the portrayal of the general environment must be fully
compatible with the aeroplane being simulated, with special emphasis being
placed on the visible ground segment on approach.
Visual system computers, whilst powerful on their own, are normally run as
slave to the simulator host computer via some kind of electronic
communications link (usually an Ethernet). This hardware set-up works
well, but has the obvious limitation that a delay is introduced between the
response of the simulated aeroplane and the response of the visual system
image generator. Limits have been set on the maximum allowable delay,
but in all flight regimes this delay should be small enough to remain
unnoticed by the flight crew.
The above notes are intended only as an introduction to visual system
testing. The remainder of this chapter is dedicated to providing more details
information about each test as specified in the ICAO Manual requirements.
4-2
SECTION 4a
4a(1)
Transport Delay (or Latency)
4A-1
4A
SYSTEM RESPONSE (TRANSPORT DELAY OR LATENCY) TESTS
4A.1 INTRODUCTION
Latency, as applied to a flight simulator, is defined as the additional time
beyond that of the basic perceivable response time of the aeroplane due to
the response of the simulator. This additional time taken by the simulator is
as a result of the computer and hardware interfaces and the execution of
the software.
The test technique used should be designed to measure only those delays
introduced by the purely simulator aspects and not the response of the
aerodynamic or certain control features which would also be found on the
aeroplane itself. For example, there will always be a small time delay in the
aeroplane between the control column being displaced and the movement
of the elevator surface and also between the movement of the elevator
surface and a resultant change in pitch angle. It is not the purpose of the
latency requirement contained in the ICAO Manual to use complicated flight
test equipment to measure these delays on the aeroplane so that they can
be accurately reproduced on the simulator. The response of the simulator
model compared to the aeroplane is checked in the relevant QTG validation
tests.
4A.2 PREVIOUS METHOD OF LATENCY DEMONSTRATION
In the past latency was demonstrated by duplicating a test performed on the
aeroplane and subtracting the aeroplane responses in order to determine
the delay caused by the computing system. Whilst this method is still a
valid one, it was difficult to determine the onset of the cues as they were
masked by the lags in the aerodynamic and control models as described
above. The response of the aeroplane had to be measured very accurately
and subtracted from the result, and this assumed that the simulator
combined aerodynamic and control system response was exactly the same
as the aeroplane for the particular case tested. Figure 4a-1 gives a pictorial
representation of a typical set of results obtained using this method.
4A-2
Figure 4a-1
Example of Simulator System Response (Latency) Results
4A.3 TRANSPORT DELAY
The transport delay is defined as the total simulator system processing time
required for an input signal from a pilot primary flight control until motion
system, visual system or instrument response. It is the overall time delay
incurred from signal input until output response. It does not include the
characteristic delay of the aeroplane simulated.
The results obtained by this method of testing for simulator response times
are more clearly defined than when running conventional latency tests and
can be measured more easily with less chance of ambiguity. The method
gives a direct measurement of latency without having to subtract the
aeroplane response and as such no aeroplane flight test data is required.
The tests ensure that all the computing elements, in the critical path, are
executed in the optimum order.
The following method is typical of that used to perform the transport delay
checks:
With the software executing normally, a force demand is injected into the
primary control causing it to move. The control position signal passes along
4A-3
the normal controls path to the host computer where the change is detected.
This is used to initialise a discrete index counter which is incremented by
each of the software modules in turn as the critical path is executed. The
index in a given module can only be updated if it contains the index number
of the module designated to run immediately before it. This allows the
software path to be traced with a signal which is not degraded by the
simulated model and checks that each element of the critical path is
executed in the correct order. The time overhead introduced by the
incrementing of the discrete index is negligible.
In the appropriate software modules for instrumentation, visual and motion
systems, the index is used to input a step signal into the simulated
aeroplane pitch, roll and yaw moments (one test for each) as well as the
motion outputs at the appropriate point in the software path. This provides
a sharp, clear signal to the visual, motion and instrument drive software.
The signals then pass through the normal computing path to generate the
visual picture, motion deflection and instrumentation response. A recording
device is used to plot the deflections and the time delay between the onset
of the control deflection and the change in state of the visual, motion and
instruments gives a direct measurement of the transport delays in the
system.
Figure 4a-2
Example of Simulator System Response (Transport Delay) Results
4A-4
A typical example of a transport delay test is illustrated in Figure 4a-2
above. The point in time at which the instrument signal moves is not
important provided the movement is within the applicable tolerance for the
level of approval sought (i.e. within 150 milliseconds of the control
movement). The visual signal, however, must not change before the motion
signal in order to prevent the pilot perceiving strange cues when flying the
simulator in a training programme. Under no circumstances should the
movement of the instrument, motion or visual signals occur before the
controls signal moves as this would indicate severe problems with the
software or a faulty method of driving the transport delay checks.
It should be borne in mind that the example given has been "styled" to show
the type of results which would be expected. The exact appearance of the
results may well differ among the different simulator manufacturers. The
important aspect is the time delay between the movement of the flight
controls and the corresponding movement of the other signals. The
direction of movement displayed on this type of plot is not usually of any
significance and will depend on various factors such as the type of recording
device used for these tests and the sign conventions used for the plotted
parameters.
4A.4 PRACTICAL ASPECTS
The inputs must of necessity be abrupt, though not necessarily large in
amplitude, producing sharp accelerations on the motion platform. Hence for
safety reasons the tests are normally carried out automatically with all
personnel offboard and the motion engaged from a maintenance facility.
The types of recording device used vary from one simulator manufacturer
to another, and may include pen recorders, ultraviolet recorders, dynamic
analysers, PC-related hardware and software or computer line printers
(where the resolution of the printer is adequate for the task).
There are some inherent difficulties in the acquisition of aeroplane flight test
data for information which needs to be measured over a very short duration
(i.e. less than 1 second in this particular case). Therefore most modern
simulators use only the transport delay method for accomplishing these
tests. Note that the reason for there only being 3 tests required for the
transport delay method versus 9 for the conventional latency method is that
the differences in transport delay which might be said to occur between the
takeoff, cruise and approach or landing flight conditions are assumed to be
negligible. This assumption is broadly correct, in that the critical path
computations are essentially the same whichever axis is being tested.
4A-5
There has been a move away from use of external plotting equipment,
largely because of cost and availability/maintainability, but many older
simulators do still use such methods. They are however much more labour
intensive and present results that are generally more difficult to interpret
than those generated through the main automatic test systems. With
experience though, this does not present a problem.
4A.5 FURTHER INFORMATION
The above has been merely a cursory treatment of the subject of simulator
transport delay methodology. For further information the reader is referred
to Appendix 5 to ACJ No.1 to JAR-STD 1A.030 (part of Reference 20) which
covers what is and is not acceptable in considerably more detail.
4A-6
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TITLE
4a(1) - SYSTEM RESPONSE TIME
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OBJECTIVE
TO CONFIRM THAT THE CUE CORRELATION AND
RESPONSES OF THE MOTION, VISUAL AND
INSTRUMENT DRIVES ARE SUFFICIENT TO BE
REPRESENTATIVE OF THE CUES PERCEIVED IN
THE AEROPLANE.
DEMONSTRATION
A signal is driven through the control system and the
resultant effects on motion, visual and instruments are
monitored to ensure that there are no unacceptable
delays in the pilot-perceived cues of the simulation.
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FLIGHT CONDITION
EITHER for transport delay tests:
a) PITCH
b) ROLL
c) YAW
OR for conventional latency tests:
a) PITCH - TAKEOFF
b) PITCH - CRUISE
c) PITCH - APPROACH OR LANDING
d) ROLL - TAKEOFF
e) ROLL - CRUISE
f) ROLL - APPROACH OR LANDING
g) YAW - TAKEOFF
h) YAW - CRUISE
i) YAW - APPROACH OR LANDING
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RECORDED
PARAMETERS
Use a multi-track recorder to record the following:
CONTROL POSITION
Longitudinal, Lateral or
Directional as appropriate
MOTION SYSTEM
In the appropriate pitch,
ACCELERATION
roll or yaw axis
VISUAL SYSTEM SIGNAL 'X', 'Y' or 'Video' drive
4A-7
INSTRUMENT SIGNAL
from the image generator
Pitch angle and bank
angle from the attitude
direction indicator, yaw
signal from the simulated
slip bubble - or
corresponding electronic
flight instrumentation
system parameters
EVALUATION NOTES Generally, some experience is required in the
interpretation of the results seen on the multichannel
recorder, though better and clearer presentation is to
be encouraged such that the values required are
easily determined. Some of the more modern systems
actually measure the results and give graphical
pass/fail determination as well. It may be the case that
running the same test twice does not necessarily yield
exactly the same time delay. This is generally
because of the highly complex program execution
tasks which are a feature of modern multi-processor
simulator computer systems (there may have been
similar occurrences prior to multi-processor systems,
but the basic reason is the same - that the relative
point in time during software critical pathexecution at
which the inputs are injected and/or individual signals
start to be measured will probably not be the same
each time the test is run). Any such inconsistencies
may be allowed for at the discretion of the evaluation
team, but the general requirement of a maximum
delay of 150 milliseconds must still be met.
TOLERANCES
MOTION RESPONSE SHALL PRECEDE VISUAL
RESPONSE BUT NOT OCCUR LATER THAN 150
MILLISECONDS AFTER INITIAL CONTROL
DEFLECTION.
VISUAL RESPONSE SHALL FOLLOW MOTION
RESPONSE BUT NOT OCCUR LATER THAN 150
MILLISECONDS AFTER INITIAL CONTROL
DEFLECTION.
INSTRUMENT RESPONSE SHALL NOT OCCUR
LATER THAN 150 MILLISECONDS AFTER INITIAL
4A-8
AEROPLANE CONTROL DEFLECTION.
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MANUAL TESTING
Manual testing for this requirement is possible, but due to the necessity of
carefully coordinating the control input with the measurement of the simulator
responses, along with the fact that what occurs is almost always too fast for a
pilot to observe the effects of the simulated aeroplane, there is not usually
much to be gained from such an approach. Nevertheless, with good
coordination between the pilot operating the controls and the person or team
operating the recording equipment, reasonable results can be achieved,
though usually not at the first attempt.
EXAMPLE
The plots shown in Figure 4a-1a and 4a-1b, right, form half of a set of results
obtained using the transport delay method. These first two sets of data show
nothing untoward. The control movement was at approximately 1.02 seconds
on the timescale, and so to meet the tolerance the responses must all occur
before 1.17 seconds. The instrument and visual responses are both within this
limit, but reference to Figure 4a-1b below shows that the motion system did not
exhibit any movement at all. The reason for this could be obvious (e.g. the
motion platform was not engaged when the test was run!), or it may indicate a
transducer failure or other problem associated with the hardware or
electronics. It is unlikely to be a software fault, because the driving signal
clearly worked for the visual and instrument.
4A-9
Figure 4a-1a
Example of Simulator Test Results for Transport Delay (Yaw) Part 1
4A-10
Figure 4a-1b
Example of Simulator Test Results for Transport Delay (Yaw) Part 2
4A-11
4A-12
SECTION 4b
DISPLAY SYSTEM TESTS
4b(1)
Field of View
4b(2)
System Geometry
4b(3)
Surface Contrast Ratio
4b(4)
Highlight Brightness
4b(5)
Vernier Resolution
4b(6)
Lightpoint Size
4b(7)
Lightpoint Contrast Ratio
4B-1
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TITLE
4b(1) - FIELD OF VIEW
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OBJECTIVE
TO DEMONSTRATE THAT THE VISUAL SYSTEM
FIELD OF VIEW FROM THE COCKPIT WINDOW IS
SUFFICIENT FOR THE LEVEL OF SIMULATOR
QUALIFICATION SOUGHT
DEMONSTRATION
Evaluate the vertical and horizontal visual system
field of view as presented using a demonstration
model.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
SEE BELOW FOR ITEMS TO BE OBSERVED
DURING THIS TEST
EVALUATION NOTES
The method of verification that the visual system field
of view is adequate can be checked by use of a
theodolite. Set-up of such equipment can be
elaborate and generally will already have been
performed during initial system test. A visual check of
the capability can be performed more quickly.
Therefore see the 'Manual Testing' section below.
TOLERANCES
Continuous, cross-cockpit, minimum collimated
visual field of view providing each pilot with 180
degrees horizontal and 40 degrees vertical field of
view.
4B-2
Horizontal FOV:
Not less than a total of 176
measured degrees (including
not less than ±88 measured
degrees either side of the centre
of the design eye point).
Vertical FOV:
Not less than a total of 36
measured degrees from the
pilot’s and co-pilot’s eye point.
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MANUAL TESTING
A typical procedure for this test is to provide a grid pattern of lines or light
points that subtend 5 Degrees per line/point allowing the number of squares to
be counted to demonstrate the required field of view. The squares should
appear square, not rectangular. If there is doubt about the angle subtended a
theodolite should be used to prove the exact Field of View.
The adequacy of the system is determined by the evaluator(s) performing the
test.
EXAMPLE
See below.
Figure 4b1-1
Example of Spherical Grid Test Pattern
(Front Channel with Partial Side Channels)
4B-3
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TITLE
4b(2) - SYSTEM GEOMETRY
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OBJECTIVE
TO DEMONSTRATE ADEQUATE VISUAL DISPLAY
SYSTEM GEOMETRY
DEMONSTRATION
Using a test pattern which fills the entire visual scene
(all channels) with a matrix of black and white 5o
squares or a 5o grid, evaluate the display system
geometry as presented
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
The method of verification that the visual system
geometry is adequate cannot practically be done by
any other method than by use of measuring
equipment using a specially designed test pattern.
The operator should demonstrate that the angular
spacing of any chosen 5o square and the relative
spacing of adjacent squares are within the stated
tolerances. For example, if the angle from the edge
of one square to its other edge is between 4 and 6
deg it falls within tolerance. The angle to the edge of
the next square may be between 8.5 and 11.5
degrees. The intent of this test is to demonstrate
local linearity of the displayed image at either pilot
eyepoint. It is generally impractical to test every
square so a judgement should be made by eye to
determine which squares are to be tested on the
basis of which appears most in error.
TOLERANCES
5o angular spacing within ±1o as measured from
either pilot eyepoint, and within 1.5o for adjacent
squares
4B-4
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MANUAL TESTING
The procedure for this test is to display a 5 o grid to show the angular spacing.
Most systems are provided with a fixed grid projected from a slide or similar
device. If visual system image can be shown to align within the tolerances
specified it may be assumed that the requirements are met based on the initial
acceptance of the visual system. Records of these checks should be available.
In the event there is any doubt about the observations or original checks a
Theodolite should be used to demonstrate compliance.
EXAMPLE
See below.
o
5 +/- 1
o
10 +/- 1.5
o
o
Figure 4b2-1
Example of Spherical Grid Test Pattern with example angular measurement
(Front Channel with Partial Side Channels)
4B-5
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TITLE
4b(3) - SURFACE CONTRAST RATIO
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OBJECTIVE
SURFACE CONTRAST RATIO IS ADEQUATE FOR
THE LEVEL OF SIMULATOR QUALIFICATION
SOUGHT.
DEMONSTRATION
Evaluate the visual system picture contrast using a
demonstration pattern and a photometer.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
Using a demonstration model or test pattern (usually
designed by the visual system manufacturer to yield
the best brightness possible from the system), the
evaluator is usually required to place himself in the
pilot's seat with the appropriate instrumentation (i.e.
the photometer) so that the results are determined
objectively. There is not really any practical method
of performing this test without significant manual
intervention, therefore see the 'Manual Testing'
section below.
TOLERANCES
Not less than 5:1
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MANUAL TESTING
The evaluator should be situated in the cockpit with a 1o photometer and the
visual demonstration model or test pattern loaded. The test pattern may be
raster drawn and should fill the entire visual scene (at least three channels),
consisting of a matrix of 5o squares with a white square in the centre of each
4B-6
channel.
Measurement shall be made on the centre bright square for each channel
using the photometer. The minimum brightness of this square shall be 7cd/m2
(2 foot-lamberts). Any adjacent dark squares should then be measured and the
contrast ratio found by dividing the bright square value by the dark square
value, with the minimum acceptable ratio being 5:1.
EXAMPLE
An example of a test pattern used to determine surface contrast ratio is shown
below. The contrast checkerboard test pattern consists of a spherical pattern of
alternating black and white polygons covering the full horizontal and vertical
Field of View.
−35
0
35
20
0
−25
Figure 4b3-1
Example of Surface Contrast Checkerboard Pattern
4B-7
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TITLE
4b(4) - HIGHLIGHT BRIGHTNESS
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OBJECTIVE
HIGHLIGHT BRIGHTNESS IS ADEQUATE
DEMONSTRATION
Evaluate the visual system brightness using a
demonstration model and a photometer.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
The method of verification that the brightness of a
daylight visual system is adequate should not be by
a method that relies solely on the use of the human
eye. Instead, using a demonstration model or test
pattern (usually designed by the visual system
manufacturer to yield the best brightness possible
from the system), the evaluator is usually required to
place himself in the pilot's seat with the appropriate
instrumentation (i.e. the photometer) so that the
results are determined objectively. However, there is
not really any practical method of performing this test
without significant manual intervention, therefore see
the 'Manual Testing' section below.
TOLERANCES
Not less than 20cd/m2 (6 foot-lamberts) on the
display
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MANUAL TESTING
The evaluator should be situated in the cockpit with a 1 degree photometer and
the visual demonstration model or test pattern loaded. The test pattern may be
4B-8
raster drawn and should fill the entire visual scene (at least three channels),
consisting of a matrix of 5 o squares with a white square in the centre of each
channel. A highlight is superimposed on the centre white square of each
channel and the brightness measured using a 1 degree photometer. Note that
the use of lightpoints is not acceptable, but it is acceptable to use calligraphic
capabilities to enhance raster brightness. (The highlight square should appear
a reasonably even brightness). Verify the visual system display brightness as
measured by the photometer is 20cd/m2 (6 foot-lamberts) on the display and
17.5cd/m2 (5 foot-lamberts) at an approach plate positioned at the pilot's knee.
EXAMPLE
An example of a test pattern used to determine surface contrast ratio is shown
in Figure 4b4-1. The contrast checkerboard test pattern consists of a spherical
pattern of alternating black and white polygons covering the full horizontal and
vertical Field of View, with a white-highlight square subtending 5 o located in
the centre of each channel.
−35
0
35
20
0
−25
High brightness square
Figure 4b4-1
Example of Highlight Brightness Checkerboard Pattern
4B-9
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TITLE
4b(5) - SURFACE RESOLUTION
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OBJECTIVE
HAS ADEQUATE SURFACE RESOLUTION
CAPABILITY.
DEMONSTRATION
Evaluate the surface resolution using a test pattern
which consists of objects shown to occupy the
required visual angle in each visual display used on
a scene from the pilot’s eye-point..
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
The method of verification that the surface resolution
of a visual system is adequate should not be solely
by a method that relies on the use of the human eye.
Consequently, while this test will also employ a
demonstration model that can be visually assessed
by the evaluator, the results must be backed up by
calculations that should be presented as part of a
statement of compliance for this test and presented
in the QTG. The eye should be positioned on a 3degree glideslope 6876 feet slant range from the
centrally located threshold of a black runway surface
painted with white threshold bars that are 16 feet
wide with 4 feet gaps in-between. At this range the
gaps will subtend two arc minutes to the eyepoint.
TOLERANCES
Not greater than 2 arc minutes
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MANUAL TESTING
4B-10
The preferred method of demonstrating that the visual system meets this
requirement is described above. As always, the cockpit lighting should be
switched off and the eyes allowed to adjust before attempting to perform this
test.
EXAMPLE
An example of a surface resolution test pattern is shown in Figure 4b5-1. The
test pattern consists of a black runway surface 10,000 ft long and 200 ft wide,
the origin of which is located at the centre of the runway. The white threshold
bars are 16 feet wide with 4 feet gaps in-between.
Figure 4b5-1
Example of Surface Resolution Test Pattern
4B-11
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TITLE
4b(6) - LIGHTPOINT SIZE
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OBJECTIVE
TO DEMONSTRATE THAT THE MAXIMUM
LIGHTPOINT SIZE YIELDS SUFFICIENTLY HIGH
RESOLUTION DISPLAYS.
DEMONSTRATION
Evaluate a row of lightpoints on the demonstration
model where modulation is just discernible.
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
The usual method of demonstrating that the visual
system meets this requirement is to display a test
pattern consisting of a centrally located single row of
lightpoints which are reduced in length until
modulation is just discernible.
Note that modulation means an ability to determine
that there are variations in brightness along the row
such that it can be determined that there is a
difference between one lightpoint and the next. It
does NOT mean that the lightpoints have to be
separated by total blackness
By the very nature of the test though, there is no
practical way of determining this point accurately
using any objective measuring equipment, therefore
see the 'Manual Testing' section below.
TOLERANCES
Not greater than 5 arc minutes
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4B-12
MANUAL TESTING
Usually, a test pattern will be displayed consisting of a single row of lightpoints
reduced in length until modulation is just discernible by the pilot or other
evaluator. For example, a row of 48 lights will form an angle of 4 o or less if the
system meets the criteria (i.e. that the lightpoint size should not be greater than
5 arc minutes). Two ‘goalposts’ show a 4 o angle, such that for the test
requirements to be met the row of lightpoints will fall between the goalposts.
EXAMPLE
See below.
Green Goal Posts
Figure 4b6-1
Example of Lightpoint Size Test Pattern
(Note: For the sake of clarity, not all lights are shown)
4B-13
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TITLE
4b(7) - LIGHTPOINT CONTRAST RATIO
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OBJECTIVE
LIGHTPOINT CONTRAST RATIO IS ADEQUATE.
DEMONSTRATION
Using a test pattern filling an area greater than 1o by
1o filled with lightpoints, evaluate the contrast ratio
between the lightpoints and the background
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FLIGHT CONDITION
NOT APPLICABLE
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RECORDED
PARAMETERS
DURING THIS TEST
EVALUATION NOTES
The method of verification that the contrast ratio of
lightpoints is adequate should not be by a method
which relies solely on the use of the human eye.
Instead, using a demonstration model or test pattern
(usually designed by the visual system
manufacturer), the evaluator is usually required to
place himself in the pilot's seat with the appropriate
instrumentation (i.e. the photometer) so that the
results are determined objectively. However, there is
not really any practical method of performing this test
without significant manual intervention, therefore see
the 'Manual Testing' section below.
TOLERANCES
Not less than 25:1
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MANUAL TESTING
The evaluator should be situated in the cockpit with a 1o photometer and the
visual demonstration model or test pattern loaded. The test pattern should
4B-14
consist of an array of calligraphic lightpoints that touch one another to fill an
area of at least 1o . Measure the brightness using a 1o photometer and make a
note of the brightness. With the lightpoint array just outside the Field of View of
the photometer measure the brightness of the black background. The contrast
ratio is found by dividing the bright value by the background value, with the
minimum acceptable ratio being 25:1.
EXAMPLE
The light array measures 20 ft lamberts and the background measures 0.1 ft
lamberts, therefore the contrast ratio is 20/0.5 = 40:1
0
40 × 40 DOT ARRAY
BLACK SURFACE
Figure 4b7-1
Example of Lightpoint Array Test Pattern
4B-15
4B-16
SECTION 4c
4c(1)
Visual Ground Segment
4C-1
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TITLE
4c - VISUAL GROUND SEGMENT
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OBJECTIVE
GROUND SEGMENT VISIBLE TO THE PILOT
WHEN CONDUCTING A LANDING MANOEUVRE
IN LOW VISIBILITY CONFORMS TO THE
Aeroplane.
DEMONSTRATION
Visibly determine that the visual ground segment is
correct when trimmed at 100 feet radio altitude for
landing in low visibility.
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FLIGHT CONDITION
TRIMMED IN THE LANDING CONFIGURATION AT
30M (100FT) WHEEL HEIGHT ABOVE
TOUCHDOWN ZONE ELEVATION, ON GLIDE
SLOPE AT A RVR SETTING OF 300M (1000FT) OR
350M (1200FT)
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RECORDED
PARAMETERS
See the description of Manual Testing for full details.
EVALUATION NOTES
This test will typically be automatically set up with the
simulated aeroplane trimmed for landing at 100 feet
above the runway height and on the glideslope. The
manufacturer will have produced a chart showing
calculations which set out what runway lights are
visible from that point with a specified Runway Visual
Range (RVR) (either 1000 or 1200 feet). Clearly, the
aeroplane result is dependent on both the aeroplane
geometry and pitch angle when trimmed for a final
approach, and the pilot or other evaluator performing
this test must be properly seated in the cockpit. The
furthest lights may only just be visible, so the
simulator cab lighting should be switched off and
time taken to allow the eyes to adjust if necessary.
4C-2
TOLERANCES
Near End:
THE LIGHTS COMPUTED TO
BE VISIBLE SHOULD BE
VISIBLE IN THE FLIGHT
SIMULATOR
Far End:
± 20% OF THE COMPUTED
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MANUAL TESTING
Select a scene that has a fully marked and lighted runway and collect the
following data for that runway:
Glideslope Threshold Crossing Height (TCH)
ILS glideslope angle
In addition, use the aeroplane manufacturer data for the following parameters:
Pilot design eyepoint
Cockpit cutoff angle
Main landing gear location (bottom of wheels)
Approach speed
ILS glideslope antenna location on aeroplane
Typical weight of aeroplane on landing
Pitch angle of aeroplane at 100 feet wheel height on landing glideslope
Select the reduced visibility conditions, either 1000 or 1200ft RVR (depending
on the regulatory authority) and trim the aeroplane for landing on the glideslope
at a radio altitude of 100ft. Then use the formulas at the end of this section to
compute these positions:
NEAR END OF VISUAL GROUND SEGMENT (VGS)
FAR END OF VGS
PILOT EYEPOINT
Graphically show the aeroplane position, near and far ends of the VGS on a
map of the selected runway that shows the approach, centreline, edge and
runway threshold lights (see Figure 4c-3 for an example).
Either fly the simulated aeroplane on the glideslope and freeze at 100ft radio
altitude or select the pre-computed position. Verify that the simulated
aeroplane's pitch angle is correct since the near end of the VGS is very
4C-3
sensitive to small pitch changes.
From the pilot eyepoint observe the near and far ends of the VGS along the
runway extended centreline. Compare the VGS to the plot. The 20%
tolerance of the VGS is applied at the far end of the VGS. If the calculations
show that the runway threshold is visible when in the same position in the
aeroplane, then the threshold in the simulator visual system must also be
visible. Perform the test in the day, twilight and night modes as necessary to
confirm the results. No instrumentation is required to perform this test.
CALCULATIONS
Given the waterline and station of the Main Gear (WMG, SMG) and the Pilot's
Eyepoint (WEP, SEP) we can find the horizontal and vertical distances of the
Pilot's Eyepoint from the Main Gear when the aeroplane is level (EP_MG_Xo,
EP_MG_Zo):
EP_MG_Xo = SMG - SEP
EP_MG_Zo = WEP - WMG
Similarly the distances of the Glideslope Antenna from the Main Gear
(GA_MG_Xo, GA_MG_Zo) can be easily found.
In general, when a point in 2-dimensional space (Xo, Zo) is rotated through an
angle 2 about the origin, its new coordinates (X2, Z2) are given by:
X2 = Xo cos 2 - Zo sin 2
Z2 = Xo sin 2 + Zo cos 2
4C-4
Figure 4c-1
Visual Ground Segment Horizontal and Vertical Distances - Pilot/Glideslope
Antenna/Main Gear
Refer to Figure 4c-1. On approach, the aeroplane will be rotated through a
pitch angle 2, positive being pitch up, so that the new horizontal and vertical
distances of the Pilot's Eyepoint from the Main Gear will be:
EP_MG_X2 = EP_MG_Xo cos 2 - EP_MG_Zo sin 2
EP_MG_Z2 = EP_MG_Xo sin 2 + EP_MG_Zo cos 2
Again, the rotated distances for the Glideslope Antenna (GA_MG_X2 ,
GA_MG_Z2) are similarly found.
It is also easily seen from Figure 4c-1 that the horizontal and vertical distances
from the Pilot's Eyepoint to the Glideslope Antenna (EP_GA_X2, EP_GA_Z2)
are:
EP_GA_X2 = GA_MG_X2 -EP_MG_X2
EP_GA_Z2 = EP_MG_Z2 - GA_MG_Z2
Since the radar altimeter reading is the altitude of the Main Gear, AMG, then the
altitude of the Pilot's Eyepoint, AEP, and the Glideslope Antenna, AGA, are:
AEP = AMG + EP_MG_Z2
AGA = AMG + GA_MG_Z2
Figure 4c-2
Visual Ground Segment Horizontal and Vertical Distances - Aeroplane to Ground
4C-5
Refer to Figure 4c-2. With the aeroplane on a glideslope which has an angle of
*, the horizontal distance from the glideslope antenna to the point the
glideslope intersects the runway centreline, DXMIT_GA_HZ is:
DXMIT_GA_HZ = AGA/tan(*)
Using the published Threshold Crossing Height (TCH), the distance from the
glideslope transmitter to the threshold, DXMIT_TH is given by:
DXMIT_TH = TCH/tan(*)
With the transmitter at a distance DXMIT_TH from the threshold of the runway, the
horizontal distance from the Pilot's Eyepoint to the runway, DR is given by:
DR = DXMIT_GA_HZ - DXMIT_TH + EP_GA_X2
With the Runway Visual Range of RVR feet, the horizontal distance that will be
visible on approach is:
VDHZ = /[RVR2 - AEP2]
The distance down the runway that is visible for a given RVR is:
DV = VDHZ - DR
Given a cut-off angle of :, measured with a zero pitch, the closest distance that
the pilot can see, DMIN is:
DMIN = AEP COT(:-2)
The total visible ground segment then is:
GndSeg = VDHZ - DMIN
4C-6
Figure 4c-3
Visual Segment Diagram Example 1
4C-7
Edge light spacing = 200 ft
Visual Segment for a Slant Range Visibility of 1200 feet and a Radio Altitude of 100
feet. The visual segment is 875 ft long and starts 826 ft before the threshold.
Aircraft type:
Aircraft pitch:
Cockpit Cutoff Angle @ 100ft
Calibrated Airspeed:
Pilot's eye height above ground:
A797-83
2.35 deg
20.79 deg
128.0 kts
120.7 ft
Aircraft data when at zero pitch
Pilot's eye ahead of main gear:
Pilot's eye above main gear:
G/S antenna ahead of main gear:
G/S antenna above main gear:
72.8 ft
17.7 ft
80.1 ft
10.8 ft
G/S Txmr ref.point to threshold:
G/S Txmr offset from runway c/line:
Glide Slope Angle:
Example runway:
1000 ft
400 ft
3.00 deg
MIB_35R
Figure 4c-4
4C-8
SECTION 5
SOUND SYSTEMS
5a
JET AEROPLANES
5b
PROPELLOR AEROPLANES
5c
SPECIAL CASES
5d
5e
FREQUENCY RESPONSE
5-1
5.0
SOUND SYSTEM TESTS - GENERAL
The purpose of the simulator sound system tests is to confirm that what is
being heard by the flight crew in the cockpit correlates well with what they
would hear in the aeroplane. Clearly, the requirements here are such that
abnormal sounds (eg compressor stall) must be simulated as well as
normal ones such as engine whine and aerodynamic noise. A further
complication is that some of the sounds heard will be extremely transient
in nature (eg engine seizure or landing gear uplock) and will not always be
supported by quality data.
The definition of the requirement varies from a Level 1 qualified simulator
to a Level 2 device, and has changed since the original ICAO Manual was
issued in so far as the first two sections in the previous document really
dealt with subjective tests rather than objective measurement of sound
cues and levels. Thus this section of the Handbook is now able to be
confined to the realistic amplitude and frequency of cockpit sounds.
5.1
LEVEL 1 QUALIFICATION REQUIREMENTS
For this level of qualification, the requirement is that the sounds should be
demonstrated as being representative of those heard in the aeroplane.
Since objective data is not required, the sounds are subjectively evaluated
and accepted by experienced persons. See Volume 2 of this Handbook for
more information on Functions and Subjective Tests involving sound cues.
5.2
LEVEL 2 QUALIFICATION REQUIREMENTS
The Level 2 requirements are not intended to be viewed in isolation but as
extensions to those of Level 1. They state that the sounds must have
realistic amplitude and frequency content compared with the aeroplane and
that they should be coordinated with weather representations which are
required to be displayed on the visual scene.
There is, therefore, still some subjective content, in that it is not really
possible to devise an objective test to demonstrate the coordination aspect,
but the major workload with these tests consists of setting up and using a
Frequency Response Analyser (FRA) to determine the simulator
compliancy. All tests in this section should be presented using an
unweighted 1/3-octave band format from band 17 to 42 (50 Hz to 16 kHz).
A minimum 20 second average should be taken at the location
corresponding to the aeroplane data set. This is usually, though not
necessarily, close to the position of the captain’s right ear. Obviously, the
aeroplane and simulator results should be produced using comparable
5-2
data analysis techniques.
The ICAO Manual requires objective testing for the highest qualification
level to be carried out to confirm that the simulation of sounds in various
phases of ground and flight operations correspond well with the data
obtained during those manoeuvres in the aeroplane. The main item of test
equipment used for these tests is a sound analyzer, which typically will be
a hand-held device capable of various types of sound measurement, but
including 1/1- and 1/3-octave frequency analysis and broadband statistical
distributions. Often the data collected is then transferred to a PC where
data can be displayed using off-the-shelf spreadsheet software and hard
copies made for inclusion in the QTG.
5.2.1 Usage of a sound analyser
The unit has many modes of operation. For the purposes of obtaining
simulator sound test results the analyser should be used in an
instantaneous mode, which gives an immediate readout of the sound
pressure level in the simulator. This will give a good indication of the
stability of the sound pressure levels, in case of doubt.
The analyser is then switched to an averaging mode, whereby a sample is
taken over at least 20 seconds, at the end of which the analyser reverts to
the ‘paused’ mode. A store function is then usually used to save the
results.
Once all the readings have been recorded and stored they may be
transferred to a PC (using RS232, etc.) for display and hard copy.
5.2.2 Typical setup and test procedures
The way in which a particular piece of equipment is used for these test will
vary, but below is given some generic guidance, based on experience with
using such a device.
1.
Position the analyser such that the microphone can easily be
maintained close to the pilot’s right ear, as this is the position from
which the aircraft data was probably gathered. A tripod may be used
to mount the device, but in any case the operator should hold the
device away from himself as far as is possible so as to avoid picking
up extraneous sounds as much as possible.
2.
Ensure the crew seats are in a normal flying configuration, that all
doors to the flight deck and all air vents are closed.
5-3
3.
Set the sound level to maximum at the simulator IOS. Leave Flight
Freeze on for the moment (or a similar control which turns the
simulator sound off).
4.
The analyser should be calibrated using its own defined procedure
for doing so. If the reading is outside the acceptable value then the
microphone should be recalibrated using the standard procedure,
which will be described in the documentation for the device.
5.
With the analyser now calibrated, the background sound level in the
cockpit needs to be checked to make sure that it has no impact on
the on the sound levels when the simulator sound system is switched
on.
6.
Remove the acoustic calibrator and position the analyser at the
pilot’s right ear. Begin recording. Ensure that the level is below or
very close to a baseline minimum level (e.g. 60dBA). If the level is
above this, items such as the flight deck cooling air will need to be
checked and reduced accordingly until an acceptable baseline level
is achieved.
7.
The main sound tests can now be run. Set the simulated aircraft up
in each of the specified configurations and initiate the recordings.
During recordings, all flight deck occupants must maintain strict
silence.
8.
When the analyser has completed each set of readings, store the
data and proceed to the next test. Once all tests are complete,
transfer the sets of data to the PC and convert the data to
spreadsheet format and then use the appropriate method of plotting
the results on a printer connected to the PC.
5.2.3 Interpretation of the Sound Data Results
The evaluation of the results must compare simulator SPL results against
flight test data, which will have been supplied by the aeroplane
manufacturer or data provider and processed in the same way as the data
is gathered on the simulator.
The tolerances suggested in the ICAO Manual are based on the ability of
the human ear to perceive sounds - especially in the 1/3 Octave band that
is most prominent to human hearing, and this has attracted a tolerance
level of 5 decibels, based on the standard dBA correction curve that is
used to calculate human noise perception in any noisy environment.
5-4
In general, the comparison should encompass both the general trends of
the plots and also the presence of higher and lower variations in the SPL
in the vicinity of - but not necessarily at the exact frequency of - those
present on the flight test data.
On a final note, the sound measuring equipment can be sensitive to
'rumble' sounds from the motion system which sometimes tend to mask the
SPL signature. Therefore it may be expedient that the tests are run with
motion system switched off, but may also be run with the motion system on
for purposes of comparison.
5.2.4 Simulator Sound Results Rationale
Prior to collecting the data for the specified simulator tests, some
preliminary work is required to ascertain the ambient noise levels in the
simulator flight deck.
It is inherent with any flight simulator that a level of air
conditioning/equipment cooling is necessary even though none may be
required in the aircraft for an equivalent configuration. For example, if the
aircraft is on ground, static and with engines and other systems off, no
cooling fan noise will be heard. This is not true for the simulator and so
rigorous attempts should be made during testing of the simulator to
quantify this effect. Given below is the experience of one such attempt:
For readings taken in a quiet room with a low noise fan running in the
background, note the steady reduction in level between 1 kHz and 5 kHz
(Figure 5-1), the latter value being approximately 25.5 dB. In contrast, the
one example of aeroplane data gathered with the aeroplane parked and all
systems switched of has a broadly similar trend, but the reduction is much
greater, with the final 5 kHz value being approximately 11 dB, which is an
extremely low value for a reading taken with any kind of noise (such as a
fan) present. In the simulator, it is therefore to be expected that the higher
frequencies will show increasingly higher SPL values than does the aircraft
data.
Two tests were then conducted on the simulator itself, firstly with the
simulator air conditioning off, then with it on. Both tests were with the sound
system turned off, but with the simulation otherwise running normally (i.e.
avionics/electrical equipment internal cooling noises were present).
5-5
Figure 5-1
Example of Simulator Calibration Results - Quiet Room with Low Noise Fan
The results, given below, show that the ambient simulator state, even with
the air conditioning off, gives a greater SPL than is indicated by the
aeroplane data for an equivalent condition. It is clear therefore that some
allowance needs to be made to adjust the readings taken on the simulator
to enable a fair comparison to be made with the aircraft data.
Figure 5-2 helps to quantify the effect of the simulator air conditioning
system on the sounds heard in the cabin when the simulator sound system
is switched off. Figure 5-3 shows directly the effect of the simulator cab air
conditioning when applied to recorded results for the engines at idle with
the sound system switched on.
5-6
Figure 5-2
Example of Simulator Calibration Results - Air Conditioning On versus Off
5-7
Figure 5-3
Example of Simulator Calibration Results - Adjustment for Air Conditioning
The sets of calibration results lead to two main conclusions:
1.
That for most configurations the simulator results are always going
to be affected by the simulator cab air conditioning, and also that, at
higher frequencies, there will probably be a greater deviation.
2.
That some adjustment needs to be made in order to properly
compare the results - at least for the test results obtained in
configurations with low noise levels.
The key is therefore to make some kind of consistent adjustment
throughout the tests, whilst still attempting to obtain results that are
recognisable versus the aircraft data.
5-8
SECTION 5a
JET AEROPLANES
5a
JET AEROPLANES
5A-1
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
5a - JET AEROPLANES
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
TO DEMONSTRATE OBJECTIVELY THAT SOUNDS
HEARD IN THE COCKPIT UNDER VARIOUS
CONDITIONS CORRESPOND TO THOSE HEARD IN
THE AEROPLANE.
DEMONSTRATION
Initialise the simulated aeroplane in each of the
configurations specified below and use a sound
analyser to record and analyse the sound cues.
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
1)
2)
3)
4)
5)
6)
7)
8)
READY FOR ENGINE START
ALL ENGINES AT IDLE
ALL ENGINES AT MAXIMUM ALLOWABLE
THRUST WITH BRAKES SET
CLIMB
CRUISE
SPEEDBRAKE/SPOILERS EXTENDED (AS
APPROPRIATE)
INITIAL APPROACH
FINAL APPROACH
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
SOUND PRESSURE LEVELS
EVALUATION NOTES
In all cases the sound analyser should be positioned
close to the captain’s right ear - or wherever the
readings were taken during the flight test data
gathering. Exact matches will not be achieved in the
same sense as for other objective comparisons
between the simulator and the aeroplane, but the
results should nevertheless be examined to ensure
they are within tolerance. The effect of the simulator
air conditioning should be borne in mind, but in any
case it should not be allowed to drastically affect the
perception of sounds heard by the flight crew.
5A-2
TOLERANCES
±5 dB per 1/3 octave band
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
The simulator should be set up in the following
configurations, and allowed to stabilise before taking
the SPL readings:
1)
2)
3)
4)
5)
6)
7)
8)
READY FOR ENGINE START - Normal
condition prior to engine start, with the APU on
if appropriate
ALL ENGINES AT IDLE - Normal condition prior
to takeoff
THRUST WITH BRAKES SET - Normal
condition prior to takeoff
CLIMB - Medium altitude
CRUISE - Normal cruise configuration
SPEEDBRAKE/SPOILERS EXTENDED (AS
APPROPRIATE) - Normal and constant
speedbrake deflection for descent at a constant
airspeed and power setting
INITIAL APPROACH - Constant airspeed, gear
up, flaps and slats as appropriate
FINAL APPROACH - Constant airspeed, gear
down, full flaps
The above conditions may be set by use of an
automatic test driver system, but usually the results
will then be taken manually, using a sound analyser.
All personnel should be silent during the 30 seconds
or so that the readings are being taken.
EXAMPLE
Figure 5a-1 shows a test result for the gear down, landing flap condition which
has been transferred to a PC and plotted from within a spreadsheet program.
5A-3
Figure 5a-1
Example of Simulator Test Results for Landing Condition Sound Test
5A-4
SECTION 5b
5b
5B-1
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
5b - PROPELLOR AEROPLANES
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
THE AEROPLANE.
DEMONSTRATION
Initialise the simulated aeroplane in each of the
configurations specified below and use a sound
analyser to record and analyse the sound cues.
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
1)
2)
3)
4)
5)
6)
7)
8)
9)
READY FOR ENGINE START
ALL PROPELLORS FEATHERED
GROUND IDLE OR EQUIVALENT
FLIGHT IDLE OR EQUIVALENT
POWER WITH BRAKES SET
CLIMB
CRUISE
INITIAL APPROACH
FINAL APPROACH
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
EVALUATION NOTES
5B-2
TOLERANCES
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
The simulator should be set up in the following
configurations, and allowed to stabilise before taking
the SPL readings:
1)
2)
3)
4)
5)
6)
7)
8)
9)
READY FOR ENGINE START - Normal condition
prior to engine start, with the APU on if
appropriate
ALL PROPELLORS FEATHERED - Normal
GROUND IDLE OR EQUIVALENT - Normal
FLIGHT IDLE OR EQUIVALENT - Normal
POWER WITH BRAKES SET - Normal condition
prior to takeoff
CLIMB - Medium altitude
CRUISE - Normal cruise configuration
INITIAL APPROACH - Constant airspeed, gear
up, flaps extended as appropriate, RPM as per
operating manual
FINAL APPROACH - Constant airspeed, gear
down, full flaps, RPM as per operating manual
The above conditions may be set by use of an automatic test driver system, but
usually the results will then be taken manually, using a sound analyser. All
personnel should be silent during the 30 seconds or so that the readings are
being taken.
5B-3
5B-4
SECTION 5c
SPECIAL CASES
5c
SPECIAL CASES
5C-1
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
5c - SPECIAL CASES
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
THE AEROPLANE.
DEMONSTRATION
Initialise the simulated aeroplane in a configuration
which is identifiable as particularly significant to the
pilot for that aeroplane and use a sound analyser to
record and analyse the sound cues.
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
ANY WHICH IS/ARE PARTICULARLY SIGNIFICANT
TO THE PILOT, IMPORTANT IN TRAINING, OR
UNIQUE TO A SPECIFIC AEROPLANE TYPE OR
MODEL
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
EVALUATION NOTES
TOLERANCES
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
5C-2
MANUAL TESTING
The simulator should be set up in whatever
configuration is required to demonstrate the particular
sound, and allowed to stabilise before taking the SPL
readings:
The above condition may be set by use of an
automatic test driver system, but usually the results
will then be taken manually, using a sound analyser.
5C-3
5C-4
SECTION 5d
FLIGHT SIMULATOR BACKGROUND
NOISE
5d
5D-1
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
5d - FLIGHT SIMULATOR BACKGROUND NOISE
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
HEARD IN THE COCKPIT WHICH ARE NOT PART
OF THE SIMULATION MODELLING DO NOT
INTERFERE WITH TRAINING.
DEMONSTRATION
Initialise the simulator with the simulation running, the
sound system muted and a ‘dead’ cockpit and use a
sound analyser to record and analyse the sound
pressure levels.
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
GROUND (STATIC,
SWITCHED OFF)
WITH
ALL SYSTEMS
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
EVALUATION NOTES
Again, the sound analyser should be positioned close
to the captain’s right ear when taking the readings.
The object of this test is to ascertain the sound
pressure level inherent in the simulator flight deck
even before the sound system is switched on. The
cockpit instruments should not be powered, so as to
eliminate electronic and other noise (e.g. cooling)
associated with the aircraft systems. The test must be
performed for the Master QTG prior to the initial
evaluation and the results retained as a ‘baseline’
against which future recurrent tests can be compared,
when the tolerance below will be applied. No specific
objective for a tolerable noise level has been specified
in the ICAO Manual, but the background noise should
not interfere with training.
TOLERANCES
±3 dB per 1/3 octave band at recurrent evaluations,
compared to initial evaluation
5D-2
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
The simulator should be set up in an on-ground static
configuration, with the sound system and all aircraft
systems switched off, and allowed to stabilise before
taking the SPL readings. The results should show the
levels inherent in the simulator with just the cab air
conditioning running. Other noises in the simulator
building which are perceptible to the test
instrumentation will also form part of the recordings.
EXAMPLE
See the discussion in section 5.2 above. The ICAO Manual is the source for the
following diagram (Figure 5d-1). Figure 5-2 above allows a comparison to be
made from a real simulator test result for background noise..
Figure 5d-1
Recommended Maximum Simulator Background Noise
5D-3
5D-4
SECTION 5e
FREQUENCY RESPONSE
5e
FREQUENCY RESPONSE
5E-1
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
TITLE
5e - FREQUENCY RESPONSE
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
OBJECTIVE
HEARD IN THE COCKPIT DURING THE INITIAL
EVALUATION ARE MAINTAINED OVER THE LIFE
OF THE SIMULATOR.
DEMONSTRATION
Initialise the simulator at the conditions specified and
use a sound analyser to record and analyse the sound
pressure levels.
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
FLIGHT CONDITION
See Test 5a (or 5b for propellor driven aeroplanes)
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
RECORDED
PARAMETERS
EVALUATION NOTES
The object of this test is to ascertain that the sound
pressure levels and frequency distribution for the
specified flight conditions remains as implemented at
the original qualification and does not significantly
degrade over time. The test must be performed for the
Master QTG prior to the initial evaluation and the
results retained as a ‘baseline’ against which future
recurrent tests can be compared, when the tolerance
below will be applied.
TOLERANCES
Cannot exceed ±5 dB on three consecutive bands at
recurrent evaluations, compared to initial evaluation
The average of the absolute differences between
initial and recurrent evaluations cannot exceed 2 dB
))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))))
MANUAL TESTING
5E-2
The simulator should be set up in each specified
configuration, with the sound system and all aircraft
systems running, and allowed to stabilise before
taking the SPL readings.
EXAMPLE
The example shown in figure 5e-1 is taken from the ICAO Manual.
Figure 5e-1
Example of Recurrent Frequency Response Test Tolerance
5E-3
5E-4
APPENDIX
A
FLIGHT TEST DATA CONSIDERATIONS
A-1
APPENDIX A
FLIGHT TEST DATA CONSIDERATIONS
A1.0
DATA QUALITY
The requirements for aerodynamic data accuracy and fidelity for the
evaluation of advanced simulators are greater than almost any other
application in the aerospace industry, including aircraft certification. Final
design data are based on predicted/wind tunnel aerodynamic data. Autopilot
and stability augmentation systems can be designed and built to less accurate
aerodynamic data because prototypes of the "electronic boxes" are usually
manufactured to be finely tuned to the exact aerodynamic characteristics of
the particular aircraft in flight test. Aircraft certification flight test data often
demonstrates that the aircraft meets some particular characteristic rather than
defines that characteristic. For example, an aircraft must only demonstrate,
for certification, that it possesses a minimum positive longitudinal static
stability about a trim point, rather than define the value of that stability. A
qualified simulator, on the other hand, must duplicate the same level of static
longitudinal stability as the aircraft. Therefore, all aircraft certification flight test
data are not necessarily acceptable for simulator design and evaluation. Only
that flight test data obtained using normal flight test standards and procedures
where a sufficient number of parameters including all pertinent aircraft
configuration, trim, flight and atmospheric conditions are recorded and
properly documented will suffice for these purposes. A simulator must
replicate every performance and stability and control characteristic of the
aircraft. Validation of this fact requires that these particular characteristics of
the aircraft be determined accurately and that the simulator be evaluated by
comparison with these aircraft data.
Recognising the cost and difficulty of acquiring flight test data that meets
these requirements for simulator validation, the regulatory authorities position
is that the airframe manufacturer should be the primary source of these data.
This position is based on the concept that the manufacturer has the greatest
familiarity with its aircraft and can best identify representative data which
accurately defines it. The manufacturer usually correlates the flight test data
with his predicted or wind tunnel design data in order to prove or improve
analytical skills. He may also employ every possible independent method
within reason to verify the validity of the final flight test data realising the
liability incurred upon the publication of the data. Further, design loads data
are based upon the manufacturer's final predicted or wind tunnel data and any
significant differences between these data and the final flight test data must be
addressed in terms of airframe structural load limits, limit speeds, centre of
gravity limits, etc. Clearly the manufacturer has a vested interest in the
A-2
quality, fidelity and validity of the flight test data.
Data for simulators should be of the highest quality. It should be repeatable
and collected in well-defined conditions with calibrated instrumentation using
industry accepted procedures and highly qualified personnel. Each test run
must be started from a fully trimmed, steady state condition with all
parameters which can affect the test being known and recorded. Often, in
practice, however, this has not been achieved; even when data has been
collected with the participation of the simulator manufacturer. Some initial
evaluations of simulators have provided a number of examples of flight test
data of questionable quality. The most common are:
a)
Lack of essential parameters such as angle-of-attack, sideslip, roll or
even pitch. Comments on simulator validation packages and
discrepancies identified during the evaluation have included Vmca tests
without heading information or sufficient yaw rate data to accurately
evaluate the test, no runway deviation data for Vmcg, and no yaw or
sideslip data for rudder response.
b)
Poor test procedures and improperly trimmed conditions. There have
been numerous cases where small offsets or accelerations must be
included in the simulator initial conditions or a small control input is
required upon release of the simulator but prior to the test input in order
to match the response of the aeroplane. Often the only data available
begins immediately before a test input and the aircraft's actual initial
conditions are unknown, or at the least, uncertain. In some cases,
incorrect or unusual pilot inputs are made in standard tests, or the
aircraft is allowed to drift excessively from the required attitude, such as
roll during a phugoid test.
c)
Testing in unsatisfactory atmospheric conditions. Too often anomalies
in the aircraft data or discrepancies between the simulator data and the
aeroplane data are explained by "atmospheric effects" when little or no
wind data is available. This has been most evident in otherwise
unexplained changes in airspeed, angle-of-attack, altitude or pitch. The
lack of adequate atmospheric data is a problem in many data packages.
Data pack information and explanations on initial conditions, test methods and
data acquisition systems, often fall short of that needed to fully support
simulator model development and accurate validation of simulator responses.
Too often assumptions have to be made to explain what happened during the
flight test. Additional questions then have to be submitted to the manufacturer
for clarification.
A-3
Manufacturers also must provide design data where no other data exists,
usually when a simulator is needed for initial instructor and crew training for a
new aircraft for which flight data is not yet available. The problem with design
data is that it may not represent all of the changes which have been made to
the aircraft since the design data was released and it may represent certain
simplifications which do not impact on its use as a design tool but may not be
accurate enough for simulator use. Unfortunately, this is unknown until actual
test data becomes available. However, this risk is minimised by close
coordination and cooperation with the manufacturer and in review of
preliminary flight test data prior to final approval of initial simulator training.
A2.0 REQUIREMENTS FOR FLIGHT TEST DATA
To achieve the level of quality needed in the flight test data for simulator
support, the following requirements should be followed:
a)
The aircraft should be maintained at the trim condition prior to the start
of the test for sufficient time to ensure that it has reached stabilised
flight.
b)
The initial conditions for the trimmed aircraft should be completely
specified for each test (i.e. gross weight, centre of gravity, static air
temperature, indicated airspeed, flap, gear and stabiliser or trimming
surface positions, wind conditions, power settings, etc).
c)
Data recording devices should begin recording several seconds prior to
the start of the test.
d)
The tests must be conducted in the most stable atmospheric conditions
obtainable and with all atmospheric conditions properly noted.
e)
All pertinent parameters should be measured, especially
angle-of-attack, sideslip, roll and pitch, as well as the stability of axes
angular rates, accelerations and control surface positions.
f)
Flight tests must be conducted using calibrated flight test
instrumentation and data acquisition systems.
g)
Qualified personnel familiar with the flight test procedures and
objectives should conduct the tests.
The IATA Flight Simulator Design and Performance Data Requirements
(Reference 12) document provides guidance on flight test data packages, and
several other documents describe many of the standard flight test procedures
A-4
which can be used to collect data for simulator validation, taking into
consideration the items discussed above.
A3.0 PRESENTATION OF VALIDATION DATA
A continuing problem in simulator evaluation, especially for some older
aeroplanes, is the format of the aircraft data. The evaluation of the simulator
requires comparing its response to that of the aircraft. Many times the
resolution and general quality of aircraft flight test plots are such that the data
cannot be accurately read within the tolerance of the specified standard.
Another aspect of this problem is that some time histories are so noisy they
are difficult to read correctly.
Time history plots should be presented on scales which allow ease of
evaluation. Noisy data should be carefully filtered to reduce the noise but
preserve the fundamental characteristic of the parameter. It is recommended
that validation flight test data be plotted to standard engineering scales (1, 2 or
4 units or multiples of 10 thereof per inch or two centimetres) such that all
specified tolerances required for a QTG are easily readable. All pertinent
parameters for each specific test should be plotted, not just the parameters for
which a tolerance is specified. Any differences between the test aircraft and
the production aircraft should also be provided.
A-5
A-6
APPENDIX
B
DYNAMIC DATA ANALYSIS
B-1
APPENDIX B
DYNAMIC DATA ANALYSIS
B1.0
INTRODUCTION
Whilst the subject of the analysis of oscillatory motions may be familiar to
most engineers, this may not necessarily be the case for a simulator
evaluator. It is therefore the intent of this appendix to permit the evaluator who
finds unfamiliar such terms as period, damping and time to half amplitude to
gain at least some understanding of them and their importance when used to
quantify certain aeroplane characteristics. Their application to aeroplane flight
simulator performance and handling qualities should thus follow naturally,
since the aerodynamic and other mathematical models used to generate the
flight simulator software programs must of necessity replicate such
characteristics very accurately. The information contained herein has been
gleaned from a variety of sources and due acknowledgement is therefore
made to those engineers and mathematicians who have over the course of
time generated the equations and techniques employed in such analyses.
B2.0
AEROPLANE GROUPS OF MOTION
The analysis of aeroplane stability characteristics is greatly simplified by the
separation of the longitudinal and lateral groups of motion. In general, it is
found that the lateral (or "asymmetric") motions are unaltered by any changes
in any of the longitudinal motions, though the reverse is true only for small
motions. For example, if an aeroplane is flying straight and level in
unaccelerated flight it may experience a change in forward velocity (perhaps
through some elevator control adjustment or possibly because of some
atmospheric disturbance such as a wind gust). This change in forward velocity
will more than likely be accompanied by small downward and pitching
velocities, but these changes will be the same for both wings and thus no
rolling moment is produced because of the symmetric change in speed. For
the same reason no yawing moment is produced because the change in drag
force is also symmetrical. Hence no anti-symmetric motions, namely yaw, roll
and sideslip, can develop from a change in longitudinal ("symmetric") motion.
Naturally this depends on the aeroplane being symmetrical about its centreline and whilst this is not exactly the case due to manufacturing tolerances in
practice this may be assumed to be true.
The reverse case, that of lateral changes causing variations in longitudinal
motion, has to be approached slightly differently. Consider, once again, an
B-2
aeroplane flying straight and level in unaccelerated flight when it is disturbed
by, for example, a crosswind gust. The aeroplane will sideslip to, say, the right
with some velocity, resulting in a tendency to roll and yaw and also, perhaps
surprisingly, to pitch and acquire changes in forward and downward velocities.
However it can be shown that the changes in longitudinal motion are in fact
dependent on the square of the sideslip velocity and it follows from this that,
for small deviations in lateral or directional velocities, the square of these
terms can to all intents and purposes be neglected, leading to the conclusion
that no coupling essentially exists between the longitudinal and lateral groups
of motion.
B3.0
AEROPLANE STABILITY
Aeroplane stability is subdivided into two main types. The first describes the
change of forces and moments on an aeroplane due to a slight displacement
and is termed static stability. The second involves the subsequent history of
the motion and the changing values of forces and moments and is termed
dynamic stability. Both terms apply to deviations in any motion, longitudinal
and lateral, and refer to the behaviour of the aeroplane in a disturbance
without the interference of pilot-operated control surfaces. Consequently, for
an aeroplane to be defined as stable, it must possess inherent stability without
the aid of an external impetus in the form of an independent control. When
applied to computer controlled aeroplanes, it follows that the basic airframe
need not necessarily possess inherent stability provided the automatic
manipulation of the control surfaces produces an effective stability. In practice,
most if not all transport category aircraft which are computer controlled (to
whatever extent) have airframes which to a greater or lesser degree are
inherently stable as well. Typically the use of various control computers on
these aeroplane types is for reasons other than to compensate for an unstable
(though probably more manoeuvrable) airframe design.
An aeroplane is defined as statically stable when a small change in motion
produces a force and/or moment system which tends to return the aircraft to
its undisturbed state. If the tendency is in the opposite sense and the
force/moment system helps the disturbance, the aeroplane is statically
unstable. A third condition, known as neutral stability, occurs when a small
disturbance produces no force or moment system - either stabilizing or
destabilizing.
An aeroplane is defined as being dynamically stable when, after a disturbance
from a steady flight state, the subsequent motion causes the aeroplane to
regain its initial steady state. Naturally it follows that if the aeroplane does not
regain its former steady state it is classed as dynamically unstable. This ability
to regain the former state may take many minutes or it may happen in only a
B-3
few seconds - the time taken is immaterial to the definition. There are several
possible motions, and these are perhaps best visualised with the aid of time
histories, albeit stylised and simplified.
The plot (which could for example be bank angle during a spiral stability
manoeuvre) in Figure B-1 indicates that the aeroplane is both statically and
dynamically stable in this mode. This is because of the tendency to return to
the original state after the disturbance has been removed (e.g. a pulsed input
of, say, control wheel). The damping can be described as being at its critical
value, allowing no overshoot.
6
4
2
0
-2
-4
-6
0
10
20
30
40
50
60
Percentage Timescale
70
80
90
100
Figure B-1
Example of Critical Damping (Statically and Dynamically Stable)
In Figure B-2 below, the aeroplane is also both statically and dynamically
stable, even though the parameter plotted (e.g. pitch rate or sideslip angle)
shows some overshoot. The damping is still positive however, and the
tendency is still for the aeroplane to return to its original state. The
characteristics required to be measured in this case would be period and time
to half amplitude (or damping coefficient).
B-4
10
8
6
4
2
0
-2
-4
-6
-8
-10
0
10
20
30
40
50
60
70
80
90
100
Figure B-2
Example of Positive Damping (Statically and Dynamically Stable)
In Figure B-3, the aeroplane is statically stable, because there is still a
tendency for it to return to its original state once the disturbance has been
removed. However, it is dynamically unstable because of the increasing
50
45
40
35
30
25
20
15
10
5
0
-5
-10
-15
-20
-25
-30
-35
-40
-45
-50
0
10
20
30
40
50
60
70
80
90
100
Figure B-3
Example of Negative Damping (Statically Stable but Dynamically Unstable)
B-5
amplitude of the oscillations as time progresses. The damping in this case is
negative and the required measurement for the simulator tests would be time
to double amplitude. This characteristic is rare in large modern transport
aeroplanes, but some older models sometimes exhibited this tendency at, for
example, high mach number conditions.
50
45
40
35
30
25
20
15
10
5
0
-5
-10
-15
-20
-25
-30
-35
-40
-45
-50
0
10
20
30
40
50
60
70
80
90
100
Figure B-4
Example of Simple Divergence (Statically and Dynamically Unstable)
In Figure B-4, the aeroplane is both statically and dynamically unstable,
because there is no tendency to return to the original state. Again, this
characteristic is rare in large modern transport aeroplanes and in any case
would generally speaking not be acceptable during the aeroplane certification
process.
These plots can be assumed to be 'stylised' versions of such simulator tests
as Phugoid Dynamics (Test 2c(9)), Short Period Dynamics (Test 2c(10)),
Spiral Stability (Test 2d(4)) and Dutch Roll (Test 2d(7)), though the precise
shape of any of the plots shown may not be characteristic of any particular
aeroplane or class of aeroplanes.
B4.0
SIMULATOR RESULTS ANALYSIS
It is obviously not the purpose during a simulator evaluation to assess whether
or not the aeroplane itself is stable or unstable in certain modes and under
certain conditions. Thus the evaluation of dynamic test results is confined - as
in all aeroplane related tests - to ascertaining that the simulator conforms to
B-6
the aeroplane performance and handling within the tolerances prescribed in
the International Standards. For most tests this is fairly easy and can be
determined by merely scrutinising the results output by the simulator test
system. However, this is not the case with oscillatory plots such as Phugoid
(Test 2c(9)) and Dutch Roll (Test 2d(7)), for which a more mathematical
approach must be employed.
Second Order Damped Oscillations
(e.g. Phugoid)
Amplitude (e.g. pitch rate - deg/sec)
16
12
8
4
0
-4
-8
-12
-16
0
50
100
150
200
250
300
350
400
450
500
Time (seconds)
Figure B-5
Method of Determination of Time to Half Amplitude of a Second Order Oscillation
Take Figure B-5 to be a typical plot output of a parameter resulting from the
Phugoid test. Most of the information shown here will probably be seen on the
simulator output, but the two amplitudes, A1 and A2, and the times at which
they were recorded by the host computer will probably not be shown. For
manual verification of the computed results for period, time to half (or double)
amplitude and/or damping coefficient, the method described below should be
used.
The periodic time is found by merely measuring the time difference between
two successive peaks (or troughs) in the plotted value. A more generalised
method though, involves determining the time difference between, say, five
successive peaks or troughs and then dividing this value by five to obtain the
average period, which may differ very slightly between any two individual
maxima or minima. It is important to ignore the first peak or trough, which may
B-7
not be part of the aeroplane free response since the influence of the initial
control input or other disturbance may still be present. Therefore,
measurement should not begin close to the time origin of the plot, but for a
phugoid test may not be possible for more than a minute into the test.
In the equations below, the times at which amplitudes A1 and A2 occur are
denoted T1 and T2 respectively. Tp is the periodic time (in seconds) of the
oscillations, measured from the time history. Other definitions are given in the
equations.
For oscillatory motion:
The natural radial velocity of the oscillations,
Tn = 2B/Tp (B = 3.14159)
Î
The damping coefficient,
. = loge(A1/A2)
Tn (T2- T1)
Ï
The time to half amplitude,
T1/2 = 0.693/ . Tn
Ð
Hence the only parameters which need to be measured are the two
amplitudes, A1 and A2, which should be spaced as far apart as is reasonable,
along with the times at which those amplitudes occur and also the periodic
time, Tp. The tolerances for the period, time to half amplitude and damping
coefficient are found in the ICAO Manual.
B-8
APPENDIX C
EXAMPLE COMPLIANCY STATEMENTS
C-1
APPENDIX C
EXAMPLE COMPLIANCY STATEMENTS
C1.0 INTRODUCTION
In Appendix 1 of the ICAO Manual there are many items for which a Statement
of Compliancy (SOC) is required. There is no definition of precisely what a SOC
should contain, and there have probably been many different versions across the
spectrum of simulators and simulator manufacturers. The purpose of this
appendix is to provide guidance as to what is required in a SOC by way of four
examples. For the other items which require an SOC the reader is referred to the
ICAO Manual itself.
C2.0 EXAMPLE STATEMENTS
C2.1 COCKPIT SOUNDS
OBJECTIVE
Demonstrate simulator compliance with para 2.l. of reference [1],
Appendix 1: sound of precipitation, windshield wipers, and other significant
aeroplane noises perceptible to the pilot during normal operations and the
sound of a crash when the simulator is landed in excess of landing gear
limitations.
TEST PROCEDURES
This test is included as part of the Functional and Subjective Tests
checklist of reference [1], Appendix 3, para 3.j. The checklist is
reproduced in this document.
METHOD OF COMPLIANCE
Simulator compliance with this section of reference [1] is demonstrated by
performing the subjective tests included in this document.
REFERENCE and DATA SOURCE
1. [1] Appendix 1, page 4.
C-2
C2.2 GROUND EFFECT/GROUND REACTION/GROUND HANDLING
OBJECTIVE
Demonstrate simulator compliance with para 2.n. of reference [1],
Appendix 1: ground handling and aerodynamic programming to include:
(1)
Ground effect--for example: roundout, flare, and touchdown. This
requires data on lift, drag, pitching moment, trim, and power in
ground effect.
(2)
Ground reaction--reaction of the aeroplane upon contact with the
runway during landing to include strut deflections, tyre friction, side
forces, and other appropriate data, such as weight and speed,
necessary to identify the flight condition and configuration.
(3)
Ground handling characteristics--steering inputs to include
crosswind, braking, thrust reversing, deceleration, and turning
radius.
(1)
The aeroplane manufacturer aerodynamic model for the X740-200
includes separate data for lift, drag, pitching moment and downwash
in ground effect. The aerodynamic programming included with the
simulator is a full implementation of this data as described in
reference [2].
(2)
The aeroplane manufacturer shock strut model, combined braking
and cornering model, and low speed tyre model include, among
other parameters, strut deflections, tyre friction, side forces, weight
on gear, and tyre velocities to simulate ground handling. These
models are implemented in full as described in reference [3].
(3)
The aeroplane manufacturer aerodynamic and ground handling
models (references [2] and [3]) combined reproduce ground
handling characteristics including weather-cocking, braking, reverse
thrust effects, deceleration and turning radius of the aircraft.
Simulator compliance with this section of reference [1] can be
demonstrated by performing the following objective tests in this document:
1A1 Minimum Radius Turn
1A2 Rate of Turn vs Nosewheel Steering Angle
C-3
1B2
1B6
1E1
1E2
2E1
2E3
2E4
2E8
2F1
Minimum Control Speed, Ground (Vmcg)
Crosswind Takeoff
Stopping Time and Distance, Wheel Brakes - Dry Runway
Stopping Time and Distance, Reverse Thrust - Dry Runway
Normal Landing
Crosswind Landing
Directional Control with Reverse Thrust
Ground Effect Trim
1.
[1] Appendix 1, page 4.
2.
[2] D999N101-1, Rev F
3.
[3] D999N104, Rev E
C2.3 RUNWAY CONDITIONS
OBJECTIVE
Demonstrate simulator compliance with section 2.q. of reference [1],
Appendix 1: representative stopping and directional control forces for at
least the following runway conditions based on aeroplane related data:
(1)
(2)
(3)
(4)
(5)
(6)
Dry
Wet
Icy
Patchy Wet
Patchy Icy
Wet on Rubber Residue in Touchdown Zone
INITIAL CONDITIONS
1.
2.
3.
4.
5.
6.
7.
8.
C-4
Gross Weight (lbs) . . . . . . . . . . . 198000
Centre of Gravity (%MAC) . . . . . . . . . 21
Flap Position (deg) . . . . . . . . . . . . . . . . 0
Gear Position . . . . . . . . . . . . . . . DOWN
Calibrated Airspeed (kts) . . . . . . . . . . 120
Total Net Thrust (lbs) . . . . . . . . . . . 2000
Ground Spoilers (deg) . . . . RETRACTED
Wheel Brakes . . . . . . . . . . . . . . . . . OFF
TEST PROCEDURES
MANUAL:
(1)
(2)
(3)
(4)
(5)
(6)
Use the QTG utility to run patchy wet case manually. Autotest will
set the runway condition for each case automatically.
Position flap and gear levers as required above.
When terminal indicates automatic setup is complete press the
autopilot disconnect (APD) switch to disengage the backdrives.
<CR> to start the test.
After 1 second apply and hold full brakes until a full stop has been
reached.
Repeat steps 1 to 2 for :
- patchy ice
- dry
- rubber residue on wet runway
- wet
- icy
AUTOMATED :
1.
Use the QTG autotest utility to run each case automated.
RECORDING DETAILS
1.
Autotest will record the following time histories :
a.
b.
c.
d.
e.
f.
Ground speed (kts)
Longitudinal acceleration (ft/sec**2)
Ground distance (ft)
Brake pressure (psi)
Pitch Angle (deg)
Heading angle (deg)
DATA EVALUATION
For DRY, WET and ICY results see tests 1E1, 1E3, and 1E4
RUNWAY COND: Patchy WET Patchy ICE
Dist / Time Dist / Time
WET RUBBER
Dist / Time
SIMULATOR ....
FAA ...................
C-5
TOLERANCE
Subjective
The X740-200 simulator includes a full implementation of all runway
friction coefficient data in reference [3] including the effects on braking,
cornering and rolling friction. The available data includes dry, wet and icy
runways. Runways contaminated with wet rubber data is taken from data
of reference [8].
Simulator compliance with requirements (4), (5) and (6) of this section of
reference [1] is demonstrated by performing the above tests in a manual
or automated mode. In addition, the following objective tests should be
performed:
1.E.1
1.E.3
1.E.4
Stopping Time and Distance, Wheel Brakes, Dry Runway for
requirement (1)
Stopping Time and Distance, Wheel Brakes, Wet Runway for
requirement (2)
Stopping Time and Distance, Wheel Brakes, Icy Runway for
requirement (3)
1.
2.
[3] Pages 163 to 179, 216 to 227, 285 to 321
C2.4 BRAKE AND TYRE FAILURE DYNAMICS
OBJECTIVE
Demonstrate simulator compliance with section 2.r. of reference [1],
Appendix 1: representative brake and tyre failure dynamics (including
antiskid) and decreased brake efficiency due to brake temperatures based
on aeroplane related data.
INITIAL CONDITIONS
1.
2.
C-6
Gross Weight (lbs) . . . . . . . . . . . 190000
Centre of Gravity (%MAC) . . . . . . . . . 20
3.
4.
5.
6.
7.
8.
9.
Flap Position (deg) . . . . . . . . . . . . . . . 30
Gear Position . . . . . . . . . . . . . . . DOWN
Calibrated Airspeed (kts) . . . . . . . . . . 120
Total Net Thrust (lbs) . . . . . . . . . . . 2000
Ground Spoilers (deg) . . . . . EXTENDED
Wheel Brakes . . . . . . . . . . . . . . . . . OFF
Runway . . . . . . . . . . . . . . . . . . . . . . DRY
TEST PROCEDURES
Brakes & Antiskid
1.
2.
3.
4.
5.
6.
Set up simulator as per initial conditions and select FLIGHT
FREEZE.
Select one of the following from the I/F pages.
A) Failure of inboard ANTISKID power unit resulting in loss of
antiskid protection on the inboard tyres.
B) Brake actuator failure resulting in left inboard and outboard
brakes inoperative.
Reset brake temperature and release all freezes
Apply brakes and bring A/C to a stop.
Evaluate the required pilot recognition and control task.
Repeat step 1 to 4 for the other malfunctions as desired.
Tyre
1.
2.
3.
Select the following tyre failure.
a) Failure of a single main gear tyre.
Perform a take-off roll and abort take-off when tyre failure occurs.
Evaluate:
a) bank angle experienced with progressive tyre failures.
b) simulator yawing moments with and without brake application.
c) tyre vibration characteristics.
d) sound effect.
Brake efficiency due to brake temperature
MANUAL :
1.
2.
3.
Use the TMS QTG utility to run case manually.
Position flap and gear levers as required above.
autopilot disconnect (APD)
switch to disengage the backdrives.
C-7
4.
5.
6.
Apply full brakes until A/C comes to a complete stop.
Autotest will then reset the aircraft back to the initial conditions
without resetting the brake temperature.
NOTE: Brake Temp can be monitored on MAINT INDEX - BRAKES
7.
8.
9.
10.
autopilot disconnect (APD) switch to disengage the backdrives.
Repeat steps 5 to 8 until aircraft has come to a stop for the third time.
Test ends when aircraft has come to a stop for the third time.
AUTOMATED :
1.
Use the Autotest utility to run each case automated.
RECORDING DETAILS
1.
Autotest will record the following as a snapshot :
1.
2.
3.
4.
5.
6.
Equivalent airspeed (kts)
Left brake force (lbs)
Right brake force (lbs)
A/C stationary flag
Brake temperature (deg C)
Ground Distance (ft)
DATA EVALUATION
Verify that the time (sec) and distance (ft) to stop the A/C increases with
increasing brake temperature.
#1
SIMULATOR ....
FAA ....................
TOLERANCE
Subjective
C-8
#2
#3
The selection of representative brake and anti-skid failures was based on
a review of the aeroplane manufacturer published brake and anti-skid
failure analysis documents (reference [3]). Failures were chosen to
provide a good selection of symmetric and asymmetric control tasks to the
pilot, and are considered to be good training conditions. Pneumatic
braking will be available to the pilot in the hydraulic pressure failure cases.
It should be noted that many other combinations of component failures
could be selected leading to these same failure conditions but nothing
would be added in terms of the required pilot recognition and control task.
Proper dynamic response to these failures is insured by selection of the
friction coefficient limits corresponding to the above failures. When
hydraulic pressure is lost to a brake, that tyre will be limited to rolling
friction unless pneumatic braking is invoked by the pilot. In the event of
loss of antiskid protection the pilot may skid the tyres by commanding a
braking force beyond the friction limit. In this case, the limiting tyre friction
will be reduced to the skidding tyre level, rather than the normal max
braking coefficient. (skidding friction is determined from reference [3]).
Proper total aircraft response will be generated through the equations of
motion, since individual brake contributions are added into all applicable
aircraft forces and moments, and are therefore integrated into the
equations for translational and rotational velocities of the aircraft.
Representation of proper tyre forces in the event of tyre failure is assured
through the following :
1)
Tyre normal force characteristics are computed for individual tyres.
This allows simulation of the change in vertical spring constant of the
gear assembly when tyres fail. As a result, bank angle will change
with progressive tyre failures.
2)
Tyre drag forces and side forces for failed tyres are based on
aeroplane manufacturer ref. [3]. The friction coefficient for a blown
tyre with a free rolling wheel is set at 0.03. The friction coefficient for
all tyres failed is 0.05.
3)
Brake torque limits for each landing gear assembly are adjusted for
tyre failures. When a single tyre fails, the torque limit for that gear
assembly is reduced by 50%.
The effects of brake temperature are implemented as per aeroplane
manufacturer data on Brake Fade and Torque Peaking as in reference [3]
pg. 251.
C-9
1.
C-10
APPENDIX
D
MOTION SYSTEM ENVELOPE
D-1
APPENDIX D
MOTION SYSTEM ENVELOPE
D1.0
INTRODUCTION
No specific requirements have been set down in the ICAO Manual for
the minimum specification of a simulator motion system. However,
during the International Standards Working Group deliberations on
motion system requirements it was deemed appropriate to further
discussions on this subject by setting up a motion performance subgroup. This sub-group met on two occasions during 1992 and
formulated the information given in section 3.5 below. The content of
section 3.5 has not been included as such in the ICAO Manual and
thus the information it contains cannot be said to have the force
implied by that document. Nevertheless, it is considered that the
figures given form the basis for realistic guidelines on motion system
performance capability.
D1.1
CONCLUSION OF THE MOTION PERFORMANCE SUB-GROUP
a) Introduction
A statement of compliance of motion performance with respect to all
the requirements below shall be included in the QTG. A
demonstration of motion performance will not normally be required,
though if requested the following procedures should be used.
b) Procedures
i) Compliance of the motion system should be demonstrated by
driving the system with single axis commands whilst keeping
the remaining axes at the neutral (zero) position. These
command signals should replace the normally calculated
motion platform demands.
ii) Only the outputs of the six actuator position transducers and a
test accelerometer(s) should be used to verify compliance.
iii) The resolution, signal to noise and frequency response of the
test accelerometer(s) must be adequate for the test being
conducted.
c) Motion Envelope
D-2
Pitch
Maximum excursion
Maximum Velocity
Maximum Acceleration
±23 degs
±20 degs/sec
±100 degs/sec2
Roll
Maximum excursion
Maximum Velocity
±25 degs
±20 degs/sec
±100 degs/sec2
Yaw
Maximum excursion
Maximum Velocity
±25 degs
±20 degs/sec
±100 degs/sec2
Vertical
Maximum excursion
Maximum Velocity
±30 ins
±24 ins/sec
±0.8 g
Lateral
Maximum excursion
Maximum Velocity
±35 ins
±28 ins/sec
±0.6 g
Longitudinal
Maximum excursion
Maximum Velocity
±35 ins
±28 ins/sec
±0.6 g
Onset acceleration in each of the rotational axes ±300
degs/sec2/sec.
Onset accelerations in the linear axes:
Vertical
Lateral
Longitudinal
±6 g/sec
±3 g/sec
±3 g/sec
d) Frequency Response (Heave)
FREQUENCY
0.1 TO 0.5 Hz
0.51 to 1.0 Hz
1.1 to 2.0 Hz
MAX. PHASE SHIFT
±30 degs
±60 degs
±110 degs
AMPLITUDE RATIO
±2 dB
±4 dB
±8 dB
e) Leg Balance
When the motion platform is driven sinusoidally in heave through a
displacement of 600mm peak to peak at a frequency of 0.2 Hz, the
phase shift between any two actuators should not exceed TBD
D-3
degrees.
f) Turn Around
When the motion platform is driven sinusoidally in heave through a
displacement of 150mm peak to peak at a frequency of 0.5Hz, the
maximum deviation from the desired sinusoidal heave acceleration
should not exceed 0.05g.
D-4
APPENDIX
E
DISCUSSION OF MATH PILOTS
E-1
APPENDIX E
DISCUSSION OF MATH PILOTS
E1.0 INTRODUCTION
This appendix examines the use of math pilots or closed-loop controllers in the
development and validation of simulator models and discusses the reasons why
a math pilot can be a useful and appropriate tool. Comparisons of simulation
data with actual flight test results are shown for the purpose of validating the use
of closed-loop controllers. Also provided is a hypothetical example of the misuse
of closed-loop controllers which might result in a close match to the tolerance
parameters, but mask a deficiency in the simulation model.
Also mentioned is the possible need for regulatory standards for the use of math
pilots. Additional requirements, in terms of parameters and their tolerances, are
examined and discussed. If the validation tests are intended to provide an
objective comparison between flight and simulation results, then the use of
closed-loop controllers needs to be recognised and addressed.
E2.0 BACKGROUND
Flight simulator performance is objectively evaluated by comparing flight test
data to test results generated with the simulator. In generating these
comparisons, closed-loop controllers (math pilots) are often used in the
secondary axes, such as controlling bank angle with wheel during a longitudinal
manoeuvre. In recent years, closed-loop controllers are being used more
frequently in the primary axis when the pilot was actively controlling the during
the flight test manoeuvre. While the use of math pilots to aid simulator-to-flight
matching has become a relatively common practice, there is no regulatory or
other guidance to govern their use. Consequently, the acceptability of a match
using this technique is left largely to engineering judgment. While this may not
be a problem, the potential for abuse exists. When using closed-loop controllers,
conventional tolerances on kinematic variables such as angular rate and attitude
can be meaningless. Closed-loop controllers could in theory be misused to mask
a deficiency in the simulation model, which could potentially have a negative
impact on training.
The use of math pilots or closed-loop controllers has become a common practice
to primarily address three issues:
!
!
E-2
To cope with flight test data uncertainties
To adequately follow lengthy time history matches
!
To prevent the need for database tailoring
An attempt to discuss each of these issues is presented below and also to
examine whether additional requirements on math pilot usage are needed and
what form those requirements might take.
E3.0 FLIGHT TEST DATA UNCERTAINTIES
Although improvements are continually being made to the quality of flight test
data used for model development and validation, uncertainties still exist. Math
pilots can address these uncertainties such as small measurement errors or
small unmeasurable atmospheric disturbances.
Figure E-1 provides a time history match
of a flap change. The match was driven
with the elevator position, stabiliser
deflection and flap handle. For this test
the tolerances specified by the ICAO
Manual are airspeed (± 3 knots), altitude
(± 100 feet) and pitch attitude (± 1.5o).
The figure presents the tolerance
parameters and bank angle. During the
configuration change the flight test is
stimulated by an asymmetry that results
in the rolling off. The change in bank
angle is left unchecked by the pilot. The
simulator response is not excited by the
same asymmetry and does not roll off.
The result is that prior to the required 15
second interval after completion of the
configuration change the simulator
match exceeds the tolerances on
airspeed, altitude and pitch attitude. If
only the longitudinal parameters for this
match were examined, one might
conclude that a shortcoming existed in
the model. The match demonstrates
how a disturbance or slight asymmetry in
the secondary (lateral) axis can impact
the primary axis and the tolerance
parameters. To compensate for the
missing disturbance in the simulator
Figure E-1
Flap Change Match - Open Loop Roll Axis response, a math pilot can be used
during the match.
E-3
Figure E-2 provides the same
match except that in this
example, a wheel controller is
used to track bank angle, roll
rate and roll acceleration during
the flap change. The result is an
excellent match of the response
in both pitch and roll. The
magnitude of the wheel
controller input and resulting
surface deflections are checked
to ensure that the longitudinal
characteristics are not
significantly influenced by any
spoiler or aileron input. If the
use of a math pilot for this case
is to be acceptable, then the
lateral control inputs, particularly
the spoilers, need to be minimal
to reduce any coupling effects.
This case represents a common
application of closed-loop
controllers on a secondary axis.
Although the source of the
upset may not be typical, it
demonstrates how a closed-loop
controller can account for an
unmeasurable atmospheric
Figure E-2
Flap Change Match - Closed Loop Roll Axis
E-4
E4.0 LENGTHY TIME HISTORY MATCHES
In the past, validation packages produced by some aeroplane manufacturers
typically included snapshot results for a number of tests including climbs,
longitudinal trims, wind-up turns, speed stability, Vmca, engine out trim, steady
sideslip, and ground effect. In response to regulatory authority input there has
been a trend to provide more time history matches as illustrated in Table 1
below.
For matches of short duration this has not been an issue. However for time
histories of considerable length, small errors are liable to integrate into large
errors. These errors may result in exceeding the tolerances on a parameter for
a given test. A review of the most recent data packages produced show that
many of the conditions in the table above are also some of the longest time
history matches.
Table E-1 Snapshot Match Trends in Validation Documents
Climbs
Longitudinal trims
Wind-up turns
Speed Stability
Vmca
Engine out Trim
Steady Sideslip
Ground Effect
Aeroplane A
(1990)
a
a
a
a
a
a
a
a
Aeroplane B
(1996)
a
a
Aeroplane C
(2002)
a
a
a
a
a
The use of a closed-loop controller or math pilot can prevent atmospheric
disturbances or small errors from integrating into large errors. This is especially
important when an open-loop matching technique is inherently unstable, i.e.
where deviation from the flight test profile results in a divergent error.
An example of this divergence can be seen in ground effect. As an aeroplane
nears the ground it encounters a nose down pitching moment resulting primarily
from the reduction in downwash. The magnitude of the nose down pitching
moment increases the closer the aeroplane gets to the ground. If an open loop
match is performed of a landing and the simulator altitude response begins to
deviate lower than flight test, the simulator will encounter a larger nose down
pitching moment. Left unchecked, this nose down moment will result in an even
further deviation from the flight test altitude profile.
E-5
Figure E-3 presents an open loop
match of a landing flight condition
provided in a recent validation data
package. This match was generated
by driving with flight test thrust,
stabiliser position and elevator
deflection. The resulting simulator
altitude match begins to deviate from
flight as ground effect is encountered.
As the simulated response
approaches the ground, the match
continues to degrade and exceeds the
tolerances for airspeed (± 3 knots),
altitude (± 10 feet), angle of attack (±
1.5o) and pitch attitude (± 1.5o).
Figure E-3
Open Loop Landing Match
Figure E-4 presents the same flight
condition matched closed-loop. In
this example an elevator math pilot
is used to close the loop on a limited
number of pitch axis parameters
such as pitch attitude, pitch rate and
airspeed.
E-6
Figure E-4
Closed Loop Landing Match
The result is an excellent match of the tolerance parameters. The error, whether
it is in the flight test data or in the mathematical models, is accounted for in the
difference between the flight test and simulator elevator position. But is this an
acceptable match? Engineering judgement must be used to determine whether
the difference between the simulator and flight test elevator position is
acceptably small. As seen in Figure E-5, the elevator error remains centered
about zero, is small in magnitude and the error is not sustained for any
considerable length of time.
Figure E-5
Closed Loop Elevator Difference
(Simulator - Flight Test)
E5.0 DATABASE TAILORING
If engineering judgement could not be used, the closed-loop matching technique
in Figure E-4 would be deemed unacceptable. The expectation is to match the
output parameters within tolerance while driving with the measured controllers.
If this interpretation is strictly adhered to, the result might force the model to be
“tailored” to produce an open loop match.
The aerodynamic simulator model and associated databases are developed
using numerous flight conditions. Only a small subset of these conditions will
appear in the validation documents. Many conditions are flown in a manner to
isolate the various independent variables of the data tables. These conditions
are used to cover a wide range of these “independents” and are heavily relied
upon in developing the aerodynamic model. The result is a continuous database
that makes physical sense and satisfies a wide range of conditions.
E-7
Figure E-6 provides an open
loop match of the landing
used in the previous
examples. For this case a
component of the pitching
moment ground effect
coefficient increment was
modified to achieve an open
loop match to the flight test
data. The modifications
were based on the difference
between the flight test and
simulator elevator from the
closed-loop match. The
elevator error was converted
to an equivalent pitching
moment coefficient and
added to the existing data
function. The resulting
landing match with the
modified function is within
the specified tolerances for
the landing test.
Figure E-6
Open Loop Landing Match Modified Function
E-8
Figure E-7 provides a comparison of the modified and baseline pitching
moment coefficient increment in the ground effect buildup. The comparison
is made at a constant angle of attack. Although the modified function
results in an open loop match for this particular case, it had a negative
impact on other open loop landing matches. The end result of obtaining an
open loop match for this landing is a function that no longer makes physical
sense and only matches a single condition. Clearly, this is not an
acceptable solution.
Figure E-7
Comparison of Pitching Moment Ground Effect Coefficient Increment
E6.0 ABUSE
The next example in this series illustrates the improper use of a closed-loop
controller. Again the landing condition is matched closed-loop, but this time the
elevator control effectiveness is artificially reduced by 20 percent. Figure E-8
shows excellent correlation between the simulator and flight results for the
tolerance parameters (airspeed, angle of attack and altitude). The match of the
tolerance parameters is as good as the match generated with the baseline
aerodynamic model presented previously in Figure E-4.
E-9
Figure E-8
Closed Loop Landing Match 20% Reduction in Elevator
Effectiveness
One might conclude that the modified model is adequate for training, when in fact
it contains a significant deficiency. The error in the model only becomes apparent
when a comparison of the elevator position is made (Figure E-9).
The elevator trace with the modified database shows a sustained error occurring
at the end of the match. With no tolerance on the elevator deflection, the
acceptability of the match must be determined by engineering judgement, and
for this example, such judgement should lead to the conclusion that the match
is unacceptable.
E-10
Figure E-9
Elevator Error for Closed Loop Match 20% Reduction in
Effectiveness
Another potential concern with the use of math pilots is the distribution of error
among multiple parameters. The proper use of math pilots assumes that the
parameters being closed upon are matched closely; that nearly all the error will
be accounted for in the difference between the flight test and simulation
controller position. However, the gains on the closed-loop controller could be
relaxed such that more error appears in the tolerance parameters and less in the
controller position error. In the preceding example of the landing match with the
20% reduction in elevator effectiveness, allowing more error in the airspeed,
altitude and pitch attitude traces would reduce the elevator error. In such a
case, engineering judgement is again needed to assess the acceptability of the
match – and to ensure the use of closed-loop controllers is not being abused.
There have in the past been instances noted of open-loop simulator QTG tests
which have initially been run by simulation engineers in a closed-loop manner
and then the resultant primary control fed back into the simulation as if the test
was being run open-loop. This is a clear abuse of math pilot methodology, since
it renders the test result unclear as to whether the simulation and/or test is
actually valid.
E-11
E7.0 ADDITIONAL REQUIREMENTS?
The preceding examples have demonstrated that closed-loop controllers are a
useful and appropriate tool in the development and validation of simulator
models.
Math pilots can provide legitimate means to address small
measurement errors in flight data or small unmeasurable atmospheric
disturbances. The simulator models and associated databases do not need to
be compromised to match a specific condition using open loop techniques. The
model will make physical sense and satisfy a wide range of conditions.
The use of math pilots however can be abused and mask a deficiency in the
model if sound engineering judgement is not applied. So is engineering
judgement enough? If it is not, then additional parameters and their associated
tolerances need to be defined; but for which tests, and which parameters and
what magnitude of the tolerance?
The performance and handling qualities tests used to qualify flight crew training
simulators can be divided into three categories:
!
!
!
Control sweeps / calibration checks
Free responses
Tracking tasks
For the “free response” tests the pilot is not actively in the loop. The pilot initiates
the manoeuvre and allows the to freely respond. Tests categorised as free
responses would include: small control inputs, configuration changes,
longitudinal trims, phugoid, short period, roll response, roll overshoot, spiral
stability, rudder response and dutch roll. For these tests math pilots on the
secondary axis would be allowed. However in the primary axis math pilot activity
would be inappropriate. The configuration change example previously discussed
in the paper would be an example of the proper use of closed-loop controllers for
a “free response” test.
For the tracking task tests the pilot is continuously in the loop. Tests categorised
as tracking tasks would include: takeoffs, dynamic engine failure after takeoff
(during the recovery), climbs, descents, longitudinal manoeuvring stability,
longitudinal static stability, stalls, minimum control speed - air, engine-out trims,
steady sideslip, landings, go-around, directional control with asymmetric reverse
thrust and ground effect. Like the free response tests, math pilots would be
allowed on the secondary axis; however, in contrast to the free response tests,
math pilots would also be allowed on the primary axis. Since the math pilot will
force a match of some of the tolerance parameters, their use during these
“tracking task” tests requires sound engineering judgement in assessing the
controller input or additional tolerances on the controllers.
E-12
The question of additional requirements has been partially addressed. The use
of math pilots on the secondary axes would be allowed for both “free response”
and “tracking task” tests. The magnitude of the inputs on the secondary axis
needs to be minimised to ensure that the inputs do not affect the primary axis.
The use of a math pilot on the primary axis would only be allowed during a
“tracking task” test with additional tolerances on the controllers. The additional
tolerances on the controller inputs are an attempt to replace engineering
judgement with a quantifiable standard.
But what magnitude should the tolerance be and on which controller input: pilot
control deflection, pilot control force or control surface deflection? With regard
to the magnitude of the tolerance some guidance can be found in the current
tolerances on surface deflections or controller force. The static control tests and
some handling qualities tests (longitudinal manoeuvring stability, longitudinal
static stability, longitudinal trim, ground effect, steady sideslip, engine out trim)
include tolerances on controller inputs noted in the table below.
Table E-2 Existing Controller Tolerances
Static Checks
Handling Qualities Tests
Pitch
± 5 lbs column force
± 2 elevator
± 5 lbs column force
± 1 elevator
o
±2
±5
aileron (or 10%)
Roll
± 3 lbs wheel force
± 2 aileron
± 3 spoiler
o
spoiler (or 10%)
Yaw
± 5 lbs pedal force
± 2 rudder
±1
rudder
o
o
o
o
o
o
Some aeroplane manufacturers have used the surface position tolerances from
the handling qualities tests for the magnitude of the “tolerances” on the closedloop controllers. The tolerances from the static checks were found to be too
lenient for the elevator and rudder control deflections. Cockpit controllers or
forces were not used because of greater confidence in the flight test surface
position measurement and a better grasp of the relevance of the control surface
error to the aerodynamic model.
Some additional internal development guidelines suggest that the error should
be about zero and that inputs approaching the controller “tolerance” only be for
a short duration. The controller and the parameters that are being closed on
must be clearly identified in the test setup. These guidelines are not absolute
and still require engineering judgement in assessing the controller error. This
E-13
leads to the original quandary that the validation tests are intended to be a
quantitative assessment of a flight simulator; yet when using closed-loop
matching techniques for the primary axis engineering judgement is required in
lieu of tolerances to determine the acceptability of the match.
Is an absolute tolerance on the controller input the answer? By adding a
tolerance on the controller input the emphasis shifts from minimising the
incremental input (from flight test) to not exceeding the tolerance. A sustained
input is now acceptable as long as it does not exceed the tolerance. The test
results might be easier to evaluate, but will the additional tolerance result in an
improvement in or maintain the current level of model fidelity? If the tolerance
is too small, the result could be a return to “tailoring” the model to match specific
flight conditions. If the tolerance is too large, the result could be a reduction in
model fidelity. Clearly, if tolerances are defined for the controller inputs, the
magnitudes must be chosen with care.
E8.0 CONCLUSION
The examples provided above have demonstrated that closed-loop controllers
are a useful and appropriate tool in the development and validation of simulator
models. Math pilots can address small measurement errors in flight data or small
unmeasurable atmospheric disturbances such that the simulator models and
database do not need to be compromised to match one condition open loop.
The model will make physical sense and satisfy a wide range of conditions. The
use of math pilots however can be abused and mask a deficiency in the model
if sound engineering judgement is not applied.
It is arguable that the utilisation of math pilots should be recognised in the
regulatory material. If the use of closed-loop controllers becomes problematic,
then their use would need to be addressed in greater detail, and perhaps
requirements in the form of additional parameters and tolerances be defined to
ensure their proper use.
E-14
APPENDIX
F
THE ELECTRONIC QTG
F-1
F-2
APPENDIX F
THE ELECTRONIC QTG
F1.0
INTRODUCTION
The publishing world has moved into the electronic era, with thousands of
documents being electronically published every day, using well-established
and widely available technology. The acceptance of electronic media is
widespread, and its application to the supply of technical documentation is
now almost universal.
The content and production process of the QTG makes it ideally suited to
electronic publication. Moving forward to an electronic media will facilitate
considerable improvements to quality and revision control, through the
increased use of automation. The electronic document can be distributed on
CD-ROM or DVD; with copies costing a fraction of that of the paper version,
multiple copies become practical, and the onerous task of reviewing a
document can now be performed in parallel by several individuals.
F1.1
REQUIREMENTS
The principal requirement for the electronic QTG (eQTG) is to simply provide
the paper document in an electronic format, but it is important that the concept
of the eQTG is understood to embrace more than just the use of electronic
storage. The electronic version must be a worthy and practical successor to
its paper prototype, if it is to gain acceptance as a complete substitute. The
document has to be globally portable, easy to access, and comprehensive.
These three features are undeniable attributes of the paper version, but not
automatically bestowed upon an electronic counterpart. These, and a few
other attributes beyond electronic storage alone, are required to make the
electronic version as useable as the paper prototype and these must be
understood to be essential elements of the eQTG concept.
All electronic formats will require the reader to have access to a suitable
computer and some software in order to access the document. The choice of
electronic format will clearly influence the portability of the document. It should
be remembered that a simulator has an expected service life in excess of
twenty years, and the eQTG must be reviewed and recreated throughout the
devices lifetime. Paper documents carefully stored are still legible a thousand
years later. Electronic media does not stand the test of time so well, in
particular the file format itself becomes rapidly out of date. An acceptable
F-3
eQTG system should employ a file format and storage media that provides
reasonable insurance against becoming unsupportable within the life of the
device. Portable Document Format (PDF) introduced by Adobe in 1993 for
use with the Acrobat Reader has been chosen by the industry as the most
standard format for the eQTG.
The QTG for a typical simulator, including reference data and supporting
materials can constitute a document spanning 10 volumes. If the simulator
has been designed to support more than one variant of the aircraft, such as
an alternative engine fit, then the size of the QTG document increases
proportionally. When such a document is transferred into electronic format a
single linear arrangement of the 5000 or so pages will not be as manageable.
A reader would need to scroll through many thousands of pages to reach any
particular item. Quite clearly an acceptable eQTG must be given some
amount of additional structure and indexing to assist the reader. Any practical
eQTG will require a hierarchical structure of automatic links to the major
sections of the document. To be intuitive to the first time reader an acceptable
eQTG should ideally be arranged with a single ‘point of entry’ leading the
reader into a simple, well indexed method of navigation throughout the
document.
We should also consider the problems posed by the need for the eQTG to be
comprehensive, and easy to extend or adapt as the simulator device evolves
during its life cycle. Any test result, or supplementary data that is required to
be included into the QTG, must also be added to the eQTG. For example, if it
becomes necessary during the life of the simulator to include a piece of
supplementary data alongside a QTG test, then it can be added to the paper
copy after no more processing than the simple operation of a suitable hole
punch. An acceptable eQTG system should also be capable of embracing
data from all external sources, via a process that demands the very minimum
level of technical skill.
Reference 23 is in effect the definitive source of information concerning both
the concepts and practical aspects of electronic QTG’s. As with many
publications and standards prevalent within the flight simulation industry, it
has been developed with the full participation and support of many facets of
the industry. Clearly the subject requires such detail as would be beyond the
scope of this Evaluation Handbook, so no attempt has been made to repeat
the information it contains, but the reader is especially referred to Section 2 of
that publication for guidance.
F-4

Aeroplane Flight Simulator Evaluation Handbook Volume 1

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