TRUE PHOENIX: STATIC SOURCE POSITION ERROR
CORRECTIONS FOR THE PITOT-STATIC
SYSTEMS ON AN F-16 FALCON
by
JASON THOMAS DIGIACOMO, B.S., M.B.A.
A THESIS
IN
SOFTWARE ENGINEERING
Submitted to the Graduate Faculty of Texas Tech University in Partial Fulfillment of the Requirements for the Degree of
MASTER OF SCIENCE
IN
SOFTWARE ENGINEERING
Approved
Hector J. Hernandez Chairperson of the Committee
Susan A. Mengel
Accepted
John Borrelli Dean of the Graduate School
December, 2005
ACKNOWLEDGEMENTS
I would like to thank my advisor, Dr. Susan Mengel, for supporting me through this project and guiding me as I traveled around the world in support of Operation Enduring
Freedom and through my change of station to Edwards AFB, CA.
I would also like to thank the U.S. Air Force in providing me the opportunity to attend Test Pilot School. The U.S. Air Force Test Pilot School is a highly competitive and respective program that allowed me to combine my two passions, aviation and engineering, into a single focus and execute amazing programs such as this one.
I also owed special acknowledgement to Lt Col Michael Taschner for his pursuit of the perfect three-leaf cloverleaf and his mathematical genius in improving our brute force spreadsheet into an elegant and efficient data reduction tool. His efforts provided me with an independent verification and validation of the cloverleaf data reduction tools.
Finally, I would like to express my gratitude to my parents and wife for supporting me in the pursuit of my dreams and always keeping me firmly planted on earth until they knew I was ready for the stars.
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TABLE OF CONTENTS
ACKNOWLEDGEMENTS...... ii
TABLE OF CONTENTS...... iii
ABSTRACT...... v
LIST OF TABLES...... vii
LIST OF FIGURES AND ILLUSTRATIONS...... viii
LIST OF ABBREVIATIONS...... xi
CHAPTER
1. INTRODUCTION ...... 1
1.1 MOTIVATION OF THESIS ...... 2
1.2 PROGRAM CHRONOLOGY...... 3
1.3 TEST ITEM DESCRIPTION ...... 4
1.4 GOAL OF THESIS...... 7
1.5 THESIS ORGANIZATION...... 8
2. LITERATURE REVIEW ...... 9
2.1 TOWER FLY-BY FLIGHT TEST TECHNIQUE ...... 9
2.2 CROSS-PACE FLIGHT TEST TECHNIQUE...... 13
2.3 ANGLE OF ATTACK FLIGHT TEST TECHNIQUE...... 16
2.4 CLOVERLEAF FLIGHT TEST TECHNIQUE...... 18
2.5 LEVEL ACCELERATION / DECELERATION FLIGHT TEST TECHNIQUE....21
2.6 SUMMARY...... 23
3. RESEARCH METHODOLOGY...... 24
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4. RESEARCH RESULTS ...... 28
4.1 POSITION ERROR CORRECTIONS ...... 29
4.2 TOTAL AIR TEMPERATURE PROBE RECOVERY FACTOR ...... 38
4.3 CLOVERLEAF FLIGHT TEST TECHNIQUE ANALYSIS ...... 40
5. CONCLUSIONS AND FUTURE WORK ...... 44
5.1 CONCLUSIONS...... 44
5.2 RECOMMENDATIONS FOR FUTURE WORK ...... 46
BIBLIOGRAPHY...... 48
APPENDICES
A: GRAPHICAL ILLUSTRATIONS...... 49
B: LESSONS LEARNED...... 114
C: INSTRUMENTATION LIST ...... 118
D: END-TO-END GROUND TEST RESULTS ...... 120
E: RAW DATA...... 122
F: DATA ANALYSIS PLAN...... 131
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ABSTRACT
This report presents the objectives, procedures, and results of the TRUE PHOENIX test project. TRUE PHOENIX calibrated a “Coral Phoenix” F-16B, S/N 92-0457, to serve as a high-speed pacer aircraft for the calibration of other aircraft air data systems and to replace an older F-16B (S/N 80-0633). Ground testing of the Pitot-static system was conducted from19 February to 5 April 2004, and flight testing was conducted from 7
April to 30 April 2004 with a total of eight successful missions.
The pacer instrumentation on F-16B S/N 92-0457 utilized the production F-16B noseboom-mounted air data probes to collect data for both total and static pressure systems. A total air temperature probe was mounted on the underside of the left forebody strake and provided the pacer air data computer with an air temperature measurement.
The pacer instrumentation did not affect the performance or flying qualities of the aircraft.
The overall test objective was to determine the air data system position error corrections of F-16B S/N 92-0457. This objective was met with marginal results; however, some of the sub-objectives were successfully met. The position error corrections were determined using the tower fly-by and F-15B pace flight test techniques.
Independent validation of these position error corrections were attempted using the level acceleration/deceleration and cloverleaf flight test techniques. In addition, a limited investigation into angle of attack effects was accomplished using the constant airspeed turn flight test technique.
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Overall, while position error corrections were determined for calibrating F-16B
S/N 92-0457, the reliability of the position error corrections at some altitudes were questionable and considered MARGINAL for calibrating F-16B S/N 92-0457 as a high- speed pacer aircraft. Through tower fly-by data analysis, the ground-level position error correction curves were accurately determined. The curves determined by the pace mission had some inaccuracies, particularly at 10,000 feet and 20,000 feet pressure altitude. The data from the level acceleration/deceleration and cloverleaf flight test techniques were not reliable enough to alter these pace curves. It was also determined that the aircraft Pitot-static system was sensitive to small changes in angle of attack.
Overall, the difference in results from the various flight test techniques requires further investigation to explain the data scatter. The total air temperature probe testing resulted in a recovery factor of 0.98, slightly below that expected from a flight test probe, and should be investigated for defects. The Pitot-static pacer calibration system software installed in this aircraft was a stable system, satisfactory for use in calibrating future aircraft air data systems.
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LIST OF TABLES
1. Test Mission Chronology...... 3
2. Total Air Temperature Probe Calibration Summary ...... 39
3. Cloverleaf Wind Analysis...... 41
4. Cloverleaf Inter-leg Wind Analysis...... 42
5. End-to-End Ground Test Results...... 121
6. Tower Fly-by Aircraft Raw Data...... 123
7. Fly-By Tower Raw Data (7 April 2004)...... 125
8. Fly-By Tower Raw Data (12 April 2004)...... 126
9. Fly-By Tower Raw Data (13 April 2004)...... 127
10. Fly-By Tower Raw Data (30 April 2004)...... 128
11. Pace Raw Data ...... 129
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LIST OF FIGURES AND ILLUSTRATIONS
1. F-16B Production ADS with Pacer Connections...... 6
2. Cloverleaf (Overhead View)...... 26
3. Altitude Position Error Correction (Tower Fly-By) ...... 50
4. Radar Altimeter vs. Tower Flyby Grid Reading...... 51
5. Altitude Position Error Correction (10,000 ft PA Pace)...... 52
6. Altitude Position Error Correction (20,000 ft PA Pace)...... 53
7. Altitude Position Error Correction (30,000 ft PA Pace)...... 54
8. Altitude Position Error Correction (35,000 ft PA Pace)...... 55
9. Altitude Position Error Correction (40,000 ft PA Pace)...... 56
10. Altitude Position Error Correction (Tower Fly-By & Pace)...... 57
11. Altitude Position Error Correction (Tower Fly-By and Extrapolatiion)...... 58
12. Altitude Position Error Correction (10,000 ft PA Cloverleaf)...... 59
13. Altitude Position Error Correction (20,000 ft PA Cloverleaf)...... 60
14. Altitude Position Error Correction (30,000 ft PA Cloverleaf)...... 61
15. Altitude Position Error Correction (35,000 ft PA Cloverleaf)...... 62
16. Altitude Position Error Correction (10,000 ft PA Level Accel/Decel)...... 63
17. Altitude Position Error Correction (20,000 ft PA Level Accel/Decel)...... 64
18. Altitude Position Error Correction (30,000 ft PA Level Accel/Decel)...... 65
19. Altitude Position Error Correction (35,000 ft PA Level Accel/Decel)...... 66
20. Altitude Position Error Correction (40,000 ft PA Level Accel/Decel)...... 67
21. Engine Airflow Effects (10,000 ft PA Level Accel/Decel) ...... 68
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22. Altitude Position Error Correction (10,000 ft PA Constant Airspeed Turn) ...... 69
23. Zero Total Pressure Validation ...... 70
24. Airspeed Position Error Correction (Tower Fly-By)...... 71
25. Airspeed Position Error Correction (10,000 ft PA Pace)...... 72
26. Airspeed Position Error Correction (20,000 ft PA Pace)...... 73
27. Airspeed Position Error Correction (30,000 ft PA Pace)...... 74
28. Airspeed Position Error Correction (35,000 ft PA Pace)...... 75
29. Airspeed Position Error Correction (40,000 ft PA Pace)...... 76
30. Airspeed Position Error Correction (Tower Fly-By & Pace) ...... 77
31. Airspeed Position Error Correction (Tower Fly-By and Extrapolated)...... 78
32. Airspeed Position Error Correction (10,000 ft PA Cloverleaf) ...... 79
33. Airspeed Position Error Correction (20,000 ft PA Cloverleaf) ...... 80
34. Airspeed Position Error Correction (30,000 ft PA Cloverleaf) ...... 81
35. Airspeed Position Error Correction (35,000 ft PA Cloverleaf) ...... 82
36. Airspeed Position Error Correction (10,000 ft PA Level Accel/Decel) ...... 83
37. Airspeed Position Error Correction (20,000 ft PA Level Accel/Decel) ...... 84
38. Airspeed Position Error Correction (30,000 ft PA Level Accel/Decel) ...... 85
39. Airspeed Position Error Correction (35,000 ft PA Level Accel/Decel) ...... 86
40. Airspeed Position Error Correction (40,000 ft PA Level Accel/Decel) ...... 87
41. Mach Number Position Error Correction (Tower Fly-By) ...... 88
42. Mach Number Position Error Correction (10,000 ft PA Pace)...... 89
43. Mach Number Position Error Correction (20,000 ft PA Pace)...... 90
44. Mach Number Position Error Correction (30,000 ft PA Pace)...... 91
45. Mach Number Position Error Correction (35,000 ft PA Pace)...... 92
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46. Mach Number Position Error Correction (40,000 ft PA Pace)...... 93
47. Mach Number Position Error Correction (Tower Fly-By & Pace)...... 94
48. Mach Number Position Error Correction (10,000 ft PA Cloverleaf)...... 95
49. Mach Number Position Error Correction (20,000 ft PA Cloverleaf)...... 96
50. Mach Number Position Error Correction (30,000 ft PA Cloverleaf)...... 97
51. Mach Number Position Error Correction (35,000 ft PA Cloverleaf)...... 98
52. Mach Number Position Error Correction (10,000 ft PA Level Accel/Decel)...... 99
53. Mach Number Position Error Correction (20,000 ft PA Level Accel/Decel)...... 100
54. Mach Number Position Error Correction (30,000 ft PA Level Accel/Decel)...... 101
55. Mach Number Position Error Correction (35,000 ft PA Level Accel/Decel)...... 102
56. Mach Number Position Error Correction (40,000 ft PA Level Accel/Decel)...... 103
57. Mach Number Position Error Correction (10,000 ft Constant Airspeed Turn) ...... 104
58. Position Error Pressure Coefficient Angle of Attack Effects ...... 105
59. Altitude Position Error Correction Angle of Attack Effects...... 106
60. Mach Number Position Error Correction Angle of Attack Effects...... 107
61. Total Air Temperature Probe Recovery Factor (Tower Fly-By & Pace) ...... 108
62. Total Air Temperature Probe Recovery Factor (10,000 ft PA Level Accel/Decel) .109
63. Total Air Temperature Probe Recovery Factor (20,000 ft PA Level Accel/Decel) .110
64. Total Air Temperature Probe Recovery Factor (30,000 ft PA Level Accel/Decel) .111
65. Total Air Temperature Probe Recovery Factor (35,000 ft PA Level Accel/Decel) .112
66. Total Air Temperature Probe Recovery Factor (40,000 ft PA Level Accel/Decel) .113
67. Rear Cockpit Pacer Controls And Time Display...... 115
68. Rear Cockpit Data Display ...... 116
69. Tower Fly-By Geometry...... 133
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LIST OF ABBREVIATIONS
AATIS Advanced Airborne Test Instrumentation System
AFFTC Air Force Flight Test Center
AFMC Air Force Material Command
AGL Above Ground Level
AoA Angle of Attack
ARDS Advanced Range Data System
FTT Flight Test Technique
GPS Global Positioning System
HUD Heads Up Display
JON Job Order Number
KCAS Knots Calibrated Airspeed
KIAS Knots Indicated Airspeed
LCD Liquid Crystal Display
M Mach Number
MSL Mean Sea Level
MOP Measure of Performance
OPSEC Operations Security
PA Pressure Altitude
PC Personal Computer
PCM Pulse Code Modulation
PCMCIA Personal Computer Memory Card International
xi
RCP Rear Cockpit
S/N Serial Number
TFB Tower Fly-by
TMP Test Management Project
TPS Test Pilot School
USAF United States Air Force
xii
CHAPTER 1
INTRODUCTION
The TRUE PHOENIX project calibrated an F-16B, S/N 92-0457, to serve as a pacer aircraft for the calibration of other aircraft air data systems. The aircraft used the production Pitot-static probe to feed sensitive pressure transducers and the production air data computer. Temperature data were collected and recorded to determine the total air temperature probe recovery factor. The corrected airspeed and altitude results were displayed in the cockpit and recorded by a digital data acquisition system for additional analysis.
This project established the static source position error corrections for the Pitot- static systems on F-16B S/N 92-0457 and calibrated the total air temperature probe.
Several flight test techniques (FTTs) were used to support data collection to include: tower fly-by, pace with the Air Force Flight Test Center (AFFTC) F-15B pacer, and GPS techniques including the cloverleaf, level acceleration/deceleration and constant airspeed turn FTTs.
Flight testing was performed at the AFFTC in April 2004 by the TRUE
PHOENIX test team. Eight F-16B test missions were successfully accomplished to support the test program, one of which was a two-ship mission with an F-15B (S/N
76-0132) for the pace flight. An Advanced Range Data System (ARDS) pod was carried on missions requiring precise GPS position and velocity.
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1.1 Motivation of Thesis
Two AFFTC pacer aircraft, F-15B S/N 76-0132 and F-16B S/N 80-0633, had previously been used by the flight test center with re-calibrations once per year. This year, F-16B S/N 80-0633 was replaced with a Coral Phoenix F-16B, S/N 92-0457, as the new AFFTC F-16B pacer aircraft.
Historically, pacer aircraft were used in the speed range from 200 KCAS to 0.93
Mach number and from the surface to 50,000 feet pressure altitude. The 0.93 Mach number “limit” was based on the shape of the Pitot-static position error correction curves.
The F-15B pacer traditionally used the same “limit” solely because a higher Mach number requirement had not been identified by a customer. The TRUE PHOENIX test team conducted testing from approach speed (11 degrees angle of attack) to 1.4 Mach number.
This thesis presents the evaluation procedures, concept, and rationale for the
TRUE PHOENIX test project. This project calibrated an F-16B, tail #92-0457, to serve as a high-speed pace aircraft for the calibration of other aircraft air data systems. This aircraft has a highly accurate calibrated air data system and a reliable data recording system to serve this function.
The overall test objective was to determine the air data system position errors using the tower fly-by and cross pace FTTs. Additional FTTs were conducted to compare the effectiveness and accuracy of other calibration techniques.
Descriptions of pace air data system calibration, test support equipment, instrumentation, test methods, and test procedures are provided within this document.
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Data analysis, data products, and reporting documentation required to support ground and flight tests are also discussed. Flight tests were conducted using F-16B and F-15B aircraft operated at the USAF Flight Test Center, Edwards Air Force Base, California.
Several flight test techniques were performed to include tower fly-by’s, cross pace with the F-15B, level accelerations/decelerations, constant airspeed turns, and cloverleaf
FTTs to collect data to be used in the calibration of the pace aircraft.
The final calibrations allow the pacer aircraft to conduct air data calibration of other aircraft at the Air Force Flight Test Center.
1.2 Program Chronology
The flight test missions were accomplished on the following dates: Table 1. Test Mission Chronology Date Flight Test Technique(s) Mission Duration Notes (hours) 7 April 04 Tower Fly-By 1.5 12 April 04 Tower Fly-By 1.5 13 April 04 Tower Fly-By 1.5 14 April 04 Pace 1.7 16 April 04 Cloverleaf 1.9 20 April 04 Cloverleaf, Level 1.4 Acceleration/Deceleration 23 April 04 Level 1.5 Acceleration/Deceleration, Constant Airspeed Turn 27 April 04 None 0.3 In-flight emergency 30 April 04 Tower Fly-By, Level 1.3 HUD failure Acceleration/Deceleration, Cloverleaf
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1.3 Test Item Description
F-16B Pacer Air Data System
A schematic of the production F-16B air data system is illustrated in Figure 1.
This figure has been modified to depict where the pacer sensitive pressure transducers
were connected. The production air data system included a Pitot-static probe mounted on
the nose that provided a dual source of static and total pressures. A second air data probe
was mounted on the forward right side of the fuselage, which provided another source of static and total pressures for a production Pitot-static system. A total air temperature probe was mounted on the underside of the left forebody strake and provided the pacer air data computer with an air temperature measurement.
The pacer instrumentation on F-16B S/N 92-0457 utilized the production F-16B noseboom-mounted air data probes to collect data for both total and static pressure systems. The air data probe incorporated two separate static and Pitot ports comprising two Pitot-static systems. Each of the Pitot-static systems was connected to calibrated sensitive pressure transducers. The sensitive transducers provided input signals to the
Advanced Airborne Test Instrumentation System (AATIS) which outputted engineering unit data to the pacer cockpit displays, a PC-104 flashcard memory, and a Mars II digital recorder.
The pacer cockpit displayed calibrated data (data corrected for both instrument and position errors) in a digital format. Both the Pitot-static system source presented on the display screens and the pacer system data recording rate were selectable from the rear
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cockpit. The PC-104 was the primary pacer data recording system and recorded
calibrated data from both Pitot-static systems for post-flight analysis. The Mars II tape
recorder was used to record the AATIS pulse code modulation (PCM) data, voice, time
code, and 1553 avionics multiplexer bus data for post flight analysis. The AATIS PCM
stream included instrument-corrected static and total pressure, as well as total air
temperature from both Pitot-static systems. Major AATIS components included the
following:
1) MARS II digital recorder: recorded all AATIS instrumentation parameters, to
include pacer system control unit (SCU)-3 outputs and 1553 avionics MUX bus
data.
2) PS-7000 Sonix pressure transducers: converted total and static pneumatic
pressure to digital format for AATIS.
3) PC-104 computer: configured and programmed for serial input, digital input,
digital output, and PCMCIA flashcard recording capability. This computer was a
commercial-off-the-shelf IBM computer for industrial embedded applications.
4) GPS time code generator: provided automatic synchronization with GPS
satellites to generate the IRIG-B Time Code.
Further information on the pacer air data system was presented in the F-16B S/N 92-0457
Modification Flight Manual, Reference 1.
5
Pacer ADS Connections #1 #2
Figure 1. F-16B Production ADS with Pacer Connections
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Test Aircraft
The F-16B was a two-seat fighter aircraft which also served in a training role.
The fuselage was characterized by a large bubble canopy, forebody strakes, and an under
fuselage engine air inlet. The aircraft was powered by a single F100-PW-220
afterburning turbofan engine with a maximum thrust of approximately 25,000 pounds.
For a complete description of the F-16B, refer to the F-16B Flight Manual and the F-16B
Supplemental Flight Manual, References 2 and 3.
The test aircraft was flown with two external tanks on the wings and an Advanced
Range Data System (ARDS) pod on a wingtip station (when required). A single ARDS
pod on the wingtip was assumed to have negligible affects on the aircraft’s Pitot-static system.
1.4 Goal of Thesis
The overall goal of the thesis was to determine the air data system position error
corrections of F-16B, S/N 92-0457. In order to accomplish this goal, a series of specific
objectives were identified:
• Determine the air data system calibration using the tower fly-by technique
• Determine the air data system calibration using the pace technique
• Determine the effect of angle of attack on the position error corrections
• Compare the air data system calibration using the cloverleaf flight test technique
to the air data system calibration using the tower fly-by and pace techniques
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• Compare the air data system calibration using the level acceleration and
deceleration flight test techniques to the air data system calibration using the
tower fly-by and pace techniques
1.5 Thesis Organization
The rest of this thesis is organized as follows: Chapter 2, presents an overview of literature and research pertinent to the air data system calibration process. Chapter 3 describes the design and implementation of a series of flight test techniques (FTT) to calibrate an aircraft to serve as a pacer aircraft for the calibration of other aircraft air data systems. Chapter 4 summarizes the position error corrections and total air temperature recovery factor and analyzes the cloverleaf flight test technique. Conclusions and potential future research directions are presented in Chapter 5.
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CHAPTER 2
LITERATURE REVIEW
This chapter presents an understanding of the air data system calibration process
and how some of these methods have been used in the past to calibrate an aircraft’s air
data system. A list of evaluation and success criteria is given for each technique along
with how the data is gathered and processed. The TRUE PHOENIX project executed
several of these techniques to calibrate an F-16B.
The overall goal of the thesis was to determine the air data system position error
corrections of F-16B, S/N 92-0457. In order to accomplish this goal, a series of
objectives were identified that have been historically achieved by performing specific flight test methods. The methods are listed below and detailed in this chapter:
• Tower fly-by flight test technique
• Cross pace flight test technique
• Angle of attack flight test technique
• Cloverleaf flight test technique
• Level acceleration and deceleration flight test technique
2.1 Tower fly-by flight test technique
The tower fly-bys are an independent evaluation of the aircraft pitot-static system.
It provides a direct comparison of the actual pressure altitude and the aircraft indicated
pressure altitude. The air data system calibration will be determined by verifying several
measures of performance (MOP). The MOP are static position error pressure coefficient,
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position error corrections, total pressure error, total air temperature probe recovery factor,
and total air temperature probe bias.
2.1.1 Measure of performance 1
Position Error Corrections is the standard measure of performance for this flight
test technique.
2.1.1.1 Success criteria
A minimum of three data points across the speed range is required.
2.1.1.2 Evaluation criteria
Any data collected will be satisfactory and meet the evaluation criteria. Sufficient data will be collected to produce the required plots.
2.1.1.3 Final data product
Plots of altimeter position correction error (Hpc) versus Mach number (Mic), velocity position correction error (Vpc) versus Mach number (Mic), and Mach position correction error (Mpc) versus Mach Number (Mic). All raw data gathered will be compared to the recorded data plots.
2.1.1.4 Data requirements
The aircraft data will come from the onboard PCMCIA flash card and/or from data processed from the onboard magnetic tape. The engineers in the tower will record
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the following data: 1) Pressure altitude measured at the height of the zero grid line 2)
Ambient air temperature measured at the height of the zero grid line 3) Grid reading as
the aircraft passes the tower 4) Time as the aircraft passes the tower. The engineers will
record the pressure altitude and ambient air temperature approximately every five
minutes from 30 minutes prior to the first pass until 30 minutes after the last pass.
The following parameters will be recorded on the DAS and those with an asterisk
(*) will be backed up on hand-held data cards:
Fuel weight* Instrument corrected altitude (Hic)* Aircraft zero fuel weight* Instrument corrected airspeed (Vic)* Instrument corrected total air temperature (Tic) Angle of Attack
2.1.1.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
2.1.1.6 Test methodology
The tower fly-by method will be used to determine the
F-16B static position error pressure coefficient, ∆Pp/qcic. The aircraft will fly past the
tower on the fly-by line at a target height of 100 ± 50 feet above ground level (AGL) with a tolerance of ± 10 feet. The pilot will establish the target airspeed, (± 5 KCAS), and maintain it within ± 2 KCAS as the aircraft approaches the tower.
2.1.1.7 Expected results
Results should be consistent with previous F-16 pacer calibrations.
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2.1.2 Measure of performance 2
Total Air Temperature Probe Recovery Factor is the standard measure of
performance for this flight test technique.
2.1.2.1 Success criteria
Six points across the applicable speed range will be collected for the tower fly-by
FTT.
2.1.2.2 Evaluation criteria
A total air temperature recovery factor which is positive and less than one will be satisfactory.
2.1.2.3 Final data products
A series of graphs will be made showing the instrument corrected total air
2 temperature parameter, (Tic/Ta) –1, as a function of Mach parameter, M /5. All raw data will be given to the customer.
2.1.2.4 Data requirements
Same as MOP 1 (2.1.1.4).
2.1.2.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
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2.1.2.6 Test methodology
Same as MOP 1 (2.1.1.6).
2.1.2.7 Expected test results
The total air temperature recovery factor will be positive and less than one.
2.2 Cross-pace flight test technique
The cross pace flight calibration technique assumes the other aircraft (F-15B) has been properly calibrated. The air data system calibration will be determined by verifying several measures of performance.
2.2.1 Measure of performance 1
Position Error Corrections is the standard measure of performance for this flight test technique.
2.2.1.1 Success criteria
At least three data points to fair a position correction error curve for the airspeed range listed from 200 KIAS to .93M at 10000, 20000, 30000, 35000 and 40000 feet.
2.2.1.2 Evaluation criteria
Any data collected will be satisfactory and meet the evaluation criteria. Sufficient data will be collected to produce the required plots.
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2.2.1.3 Final data product
Plot altitude position error correction, ∆Hpc, as a function of instrument corrected
Mach, Mic. Plot airspeed position error correction, ∆Vpc, as a function of instrument corrected Mach, Mic. Plot Mach position error correction, ∆Mpc, as a function of instrument corrected Mach, Mic.
2.2.1.4 Data requirements
The following parameters will be recorded:
Parameters collected on F-16: Fuel weight Instrument corrected altitude (Hic) Aircraft zero fuel weight Instrument corrected airspeed (Vic) Aircraft center of gravity Instrument corrected total air temperature (Tic) Angle of Attack
Parameters collected on F-15: Calibrated altitude (Hc) Ambient temperature (Ta) Calibrated airspeed (Vc)
2.2.1.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
2.2.1.6 Test methodology
The cross pace points will be flown with the F-15B in the lead and the F-16B as the wingman. The F-15B will setup near the test point, ± 500 feet and ± 5 KCAS and will call “stable” when stabilized within ±10 feet and ± 2 KCAS relative to the stable point. The wingman will call “ready, ready, read” when the F-16B is stable relative to
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the lead and approximately 2 F-15 wingspans (80-100 feet) line abreast. The F-15B will
then call “cleared to xxx KCAS” prior to accelerating/decelerating to the next calibration point. The calibration points should be planned from fast (0.95M) to slow (200 KCAS) within an altitude.
2.2.1.7 Expected test results
Results should be consistent with previous F-16 pacer calibrations. When comparing final pacer calibrations, the outputs of the two systems should match within ±
20 feet, ± 1.0 KCAS, and ± 0.5 degree K ambient air temperature.
2.2.2 Measure of performance 2
Total Air Temperature Probe Recovery Factor
2.2.2.1 Success criteria
Three points across the applicable speed range will be collected from each cross
pace altitude.
2.2.2.2 Evaluation criteria
A total air temperature recovery factor which is positive and less than one will be
satisfactory.
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2.2.2.3 Final data products
Plot the instrument corrected total air temperature parameter, (Tic/Ta) –1, as a function of Mach parameter, M2/5. Annotate the total air temperature bias on the plot.
All raw data will be given to the customer.
2.2.2.4 Data requirements
Same as MOP 1 (2.2.1.4)
2.2.2.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
2.2.2.6 Test method
Same as MOP 1 (2.2.1.6)
2.2.2.7 Expected test results
It is expected the total air temperature recovery factor will be positive and less than one.
2.3 Angle of attack flight test technique
At different altitudes and aircraft weights, the air data system position errors will vary at the same Mach number due to angle of attack effects. By determining the error associated with a specific angle of attack, a more precise air data system calibration curve can be faired.
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2.3.1 Measure of performance 1
Position Error Corrections is the standard measure of performance for this flight
test technique.
2.3.1.1 Success criteria
Each data points with valid angle of attack data will be evaluated.
2.3.1.2 Evaluation criteria
Any data collected will be satisfactory and meet the evaluation criteria. Sufficient
data will be collected to produce the required plots.
2.3.1.3 Final data product
A graph showing the airspeed position error correction, ∆Vpc, as a function of angle of attack. All raw data will be compared to the recorded data from the jet.
2.3.1.4 Data requirements
GPS data from the ARDS pod is required to determine the aircraft ground speed and direction. Outside air temperature and angle of attack are required from the test aircraft, which is used in conjunction with the GPS data in order to determine the airspeed position correction error and graph it as a function of angle of attack.
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2.3.1.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
2.3.1.6 Test methodology
First, perform a wind calibration by maintaining a constant heading in level flight with no sideslip and no significant vertical velocity for at least 10 seconds. Maintain altitude and airspeed within ±50 feet and ±2 KCAS. Roll into a level turn, stabilize at a constant airspeed, Mach number, load factor, and remain within the data band of +/- 1000 feet and a tolerance of +/- 100 feet and ±5 KCAS during the turn. Continue the turn for a full 360 degrees once stable.
2.3.1.7 Expected test results
As angle of attack is increased, the air data system position correction errors will increase.
2.4 Cloverleaf flight test technique
The significance of the cloverleaf technique is its independence from other pacer aircraft and, therefore, does not propagate a calibration error from one pacer aircraft to another. However, the cloverleaf flight calibration technique requires a very accurate source of ambient air temperature because it is an airspeed comparison technique. Also, the length of each cloverleaf is believed to determine data quality. Leg length will be
18
varied on some data points but a two minute leg length will be accomplished at all data points.
2.4.1 Measure of Performance 1
Cloverleaf Flight Test Technique is the standard measure of performance for this flight test technique.
2.4.1.1 Success Criteria
At least three data points to fair a position correction error curve for the airspeed range listed from 200 KIAS to .93M at 10000, 20000, 30000, and 35000 feet.
2.4.1.2 Evaluation Criteria
Any data collected will be satisfactory and meet the evaluation criteria. Sufficient data will be collected to produce the required plots.
2.4.1.3 Final Data Products
Plots of altimeter position correction error (Hpc) versus Mach number (Mic), velocity position correction error (Vpc) versus Mach number (Mic), and Mach position correction error (Mpc) versus Mach Number (Mic) at 10000, 20000, 30000, and 35000 feet. All raw data gathered will also be given to the customer.
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2.4.1.4 Data Requirements
A source of ambient air temperature or total air temperature must exist to permit analysis of the cloverleaf data. The temperature could come from the calibration aircraft or from a rawinsonde. Stable ground track and airspeed are directly related to the quality of the data.
2.4.1.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
2.4.1.6 Test Methodology
The pitot-static cloverleaf maneuver is an airspeed technique used to obtain pitot- static position error corrections. Ideally, the pilot makes three passes through a point in the sky at the same pressure altitude and calibrated airspeed at three different ground track headings approximately 120 degrees apart. The actual headings are not critical in this technique so the pilot may choose to fly based on aircraft heading vice ground track.
A data band of ±1000 feet and ±5 KCAS and a data tolerance of ±100 feet and ±2 KCAS will be used with an emphasis on stable airspeed and altitude for each pass. The quality of the collected data is directly related to a stable ground speed and ground track. An
ARDS pod will be carried to determine accurate ground speed and ground track.
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2.4.1.7 Expected test results
The position error corrections from the cloverleaf technique will agree with the position error corrections from tower flyby and cross pace techniques.
2.5 Level acceleration and deceleration flight test technique
As airspeeds increase towards the transonic region, previous tests have shown that the air data system begins to demonstrate transonic effects. The level acceleration and decelerations will cover a range of airspeed points through the transonic region, enabling a more accurate calibration in the near-transonic region.
2.5.1 Measure of Performance 1
Position Error Corrections is the standard measure of performance for this flight test technique.
2.5.1.1 Success criteria
Each test run at a valid altitude and collection of airspeeds will be evaluated.
2.5.1.2 Evaluation criteria
Any data collected will be satisfactory and meet the evaluation criteria. Sufficient data will be collected to produce the required plots.
21
2.5.1.3 Final data products
Plots of altimeter position correction error (Hpc) versus Mach number (Mic),
velocity position correction error (Vpc) versus Mach number (Mic), and Mach position
correction error (Mpc) versus Mach Number (Mic). These plots will be compared with
those gathered using other calibration techniques. All raw data gathered will also be
compared to the data obtained from the jet.
2.5.1.4 Data requirements
The aircraft data will come from the onboard PCMCIA flash card and/or from
data processed from the onboard magnetic tape. GPS coordinates and tapeline altitude
will be collected from the ARDS pod.
The following parameters will be recorded on the DAS and those with an asterisk (*) will
be backed up on hand-held data cards:
Fuel weight* Instrument corrected altitude (Hic)* Aircraft zero fuel weight* Instrument corrected airspeed (Vic)* Aircraft center of gravity Geographic position (by GPS) Instrument corrected total air temperature (Tic) Angle of Attack
2.5.1.5 Algorithm/Processes
The equations and algorithms used are included in Appendix F.
22
2.5.1.6 Test methodology
The level acceleration and deceleration methods will be used to determine the F-
16B static position error pressure coefficient, ∆Pp/qcic. The aircraft will fly between two way points at 300 KIAS, holding a constant heading and constant altitude. The aircraft
will then fly the same course at the same altitude, slowly accelerating from 200 KIAS
through 1.1 Mach. The tapeline altitudes obtained from the GPS ARDS pod will be used
to calibrate the air data system. This method assumes that the aircraft has already been
successfully calibrated at 300 KIAS using the tower flyby and cross-pace methods.
2.5.1.7 Expected test results
The position error corrections from the acceleration/deceleration technique will
agree with the position error corrections from tower flyby and cross pace techniques.
2.6 Summary
This chapter discusses the different techniques used to calibrate an air data system
and describes the success and evaluation criteria needed to satisfy the data requirements.
It also details a list of data products that is needed in order to evaluate each technique.
This chapter explains where the data will be obtained and how it will be processed. A
brief description of the technique itself and the expected results were also discussed so
that a better understanding of the technique can be gained. All of these methods have
been used in the past to calibrate an aircraft’s air data system and the TRUE PHOENIX
project executed several of these techniques to calibrate an F-16B.
23
CHAPTER 3
RESEARCH METHODOLOGY
This chapter presents the methodology used for the TRUE PHOENIX test project.
The methodology encompasses a series of flight test techniques (FTT) to calibrate an F-
16B for the TRUE PHOENIX project.
3.1 Test Methodology
Prior to flight test, a ground checkout occurred from 19 February to 5 April 2004.
During this checkout, the pressure transducers for static and total pressure were calibrated
in the calibration lab using standard laboratory procedures. Next, installed end-to-end
system checks were performed, using a TTU-205 test set to generate pneumatic pressures
(Appendix D). The purpose of this calibration was to exercise the pacer system to verify
it was operating correctly and did not have any leaks in the system. The leak check was
performed at 400 KCAS and 20,000 feet pressure altitude, and had a leak rate tolerance
of less than 1 KCAS per minute in airspeed and 100 feet per minute in pressure altitude.
Near the end of flight testing (28 April 2004), another end-to-end system check and leak
check were accomplished to verify no changes in the Pitot-static system had occurred
during flight testing.
The following procedures were used in accomplishing each flight test technique (FTT):
1) Tower fly-by – this FTT was flown in accordance with AFFTCI 11-1, Reference 4.
The aircraft flew past the fly-by tower on the fly-by line at a target height of 100 ± 50
feet above ground level (AGL) with a tolerance of less than 100 feet per minute vertical
24
velocity. The target airspeed was established within 5 KCAS and maintained within ±2
KCAS as the aircraft approached the tower. A build-up approach from mid-speed to
high-speed and mid-speed to low-speed data points was followed. If all mid-speed data
points had been accomplished, a minimum of one mid-speed point was practiced prior to
any high- or low-speed data points.
2) Pace – this FTT was flown with the F-15B in the lead and the F-16B as the wingman.
The F-15B set up near the test point ±500 feet and ±0.01 Mach number and called
“stable” when stabilized within ±10 feet and ±2 KCAS relative to the target stable point.
The wingman called for both aircraft to simultaneously record their data values when the
F-16B was stable relative to the lead and approximately two F-15 wingspans (80-100
feet) line abreast. The F-15B then cleared the flight to the next calibration point after the
data were collected. The calibration points were planned from 200 KCAS to 0.95 Mach
number within each altitude.
3) Constant airspeed turn – this FTT was flown by first performing a wind calibration, maintaining a constant heading in level flight with no sideslip and no significant vertical velocity for ten seconds at the target altitude (±1000 feet) and at the target Mach number
(±0.02M). Altitude and airspeed tolerances were ±100 feet and ±5 KCAS. Next, the aircraft was rolled into a level turn and the load factor was increased by 0.5g. Once
stabilized at a constant airspeed/Mach number, the turn was continued for a full 360
degrees. These level turns were repeated with increasing load factor until the maximum
constant airspeed load factor was reached using military power. A wind calibration was
repeated at least every ten minutes while performing turns at a similar Mach number, and
25
a final wind calibration was accomplished after the turn sequence was complete for a
target Mach number. The target altitudes were planned from lowest to highest, and the
target airspeeds were planned from highest to lowest.
4) Cloverleaf – this FTT was flown by accomplishing three passes through a point in the
sky at the same pressure altitude and calibrated airspeed on three different courses
approximately 120° apart (Figure 2). The actual courses were not critical in this
technique, and the team chose to fly aircraft heading vice ground track. A data band of
±1000 feet and ±0.02 Mach number with a data tolerance of ±100 feet and ±2 KCAS
were used emphasizing the need to be stable on airspeed and altitude for each pass. An
Advanced Range Data System (ARDS) pod was carried to determine accurate ground
speed and ground track. Crew judgment determined when a stable ground speed and
ground track were achieved. Further information about this FTT is presented in the
Cloverleaf Flight Test Technique Analysis section.
N
W E
S
Figure 2. Cloverleaf (Overhead View)
5) Level acceleration/deceleration – this FTT was flown within the Black Mountain
Supersonic Corridor, with the first pass consisting of a slow acceleration from approach speed or a performance-defined minimum airspeed through 1.4 Mach number or 580
26
KCAS (whichever was lower) at a rate of 5 KCAS per second. This was immediately followed by a deceleration from 1.4 Mach number or 580 KCAS (whichever was lower) to approach speed at a rate of 5 KCAS per second. Next, the aircraft flew the same course at the same altitude in the opposite direction at a constant airspeed (300 KIAS or
250 KIAS). The tapeline altitudes obtained from the GPS ARDS pod were used to calibrate the air data system. This method assumed that the aircraft was successfully calibrated at the chosen steady airspeed using the tower flyby and pace methods.
27
CHAPTER 4
RESEARCH RESULTS
This chapter presents detailed analysis of the test results for the TRUE PHOENIX
test project. TRUE PHOENIX calibrated an F-16B, S/N 92-0457, to serve as a pacer
aircraft for the calibration of other aircraft air data systems.
The following paragraphs detail the test results of the TRUE PHOENIX project.
Overall, the position error correction curves resulted in a MARGINAL rating for use in a high-speed pacer aircraft. However, the marginal data was primarily at the 10,000 feet and 20,000 feet pace altitudes, with satisfactory data collected in the tower fly-by and pace missions at 30,000 feet, 35,000 feet, and 40,000 feet. Update the F-16B, S/N 92-
0457 pacer system calibration curves using the tower fly-by and pace curves at 30,000 feet, 35,000 feet, and 40,000 feet. (R1)1
The specific test objectives were rated as follows:
SATISFACTORY – Determine the air data system calibration using the tower
fly-by technique.
MARGINAL – Determine the air data system calibration using the pace
technique.
MARGINAL – Determine the effect of angle of attack on the position error
corrections.
1 Numerals preceded by an R within parentheses at the end of a paragraph correspond to the recommendation numbers tabulated in the Conclusions and Recommendations section (Chapter 5) of this report.
28
SATISFACTORY – Compare the air data system calibration using the cloverleaf
flight test technique to the air data system calibration using the tower fly-by and pace
techniques.
SATISFACTORY – Compare the air data system calibration using the level
acceleration and deceleration flight test techniques to the air data system calibration using
the tower fly-by and pace techniques.
Throughout this section, all figures that are referenced are located in Appendix A.
All altitudes referenced throughout this section are pressure altitude unless otherwise
noted.
4.1 Position Error Corrections
The primary objective of this project was to determine the position error
corrections of F-16B, S/N 92-0457, throughout the range of altitudes from tower fly-by
altitude (approximately 2300 feet pressure altitude) to 40,000 feet pressure altitude and
from approach speed (11 degrees angle of attack) to 1.4 Mach number. The tower fly-by
and pace FTTs were used to determine the curves between 200 KCAS and 0.93 Mach
number at these altitudes. The cloverleaf, level acceleration/deceleration, and constant airspeed turn FTTs were used to independently validate these position error corrections.
In addition, the constant airspeed turn FTT was used to determine the effect of angle of attack on the position error corrections.
29
Altitude Position Error Correction
Tower Fly-By
The results from the four tower fly-by flights formed a tightly-fit calibration curve
with a maximum deviation of ±6 feet from the best-fit curve (Figure 3). Figure 4
displays the readings made by the tower observer versus the aircraft’s radar altimeter
output, with the actual grid reading versus altitude above ground level plotted as a
theoretical curve. This figure clearly demonstrates that the tower observer correctly
obtained the aircraft altitude within a reasonable amount of error as it passed the tower.
Most of the radar altimeter data were less than expected by approximately 3 to 4 feet, the equivalent of approximately 0.1 grid reading. The data scatter was approximately ±5 to 6 feet relative to a line offset 3 to 4 feet less than the theoretical relationship.
Pace
The altitude position error correction, ∆Hpc, formed a family of curves at different
altitudes; ranging from 10,000 feet to 40,000 feet. The data collected from each pace
mission were used to develop a model at each altitude that was compared to the
extrapolated tower fly-by curves (Figure 5 through Figure 9). These curves were also
compared to the cloverleaf, constant airspeed turn, and level acceleration/deceleration
FTT results. Since the tower fly-by data was the best data collected over the entire speed
range, the Mach number position error correction curve data were converted to altitude
position error corrections and extrapolated up to each altitude (Figure 11). The curves
were created using Herrington’s Flight Test Engineering Handbook (Reference 7).
30
Overall, this extrapolation matched the pace model position and shape at low altitudes;
however, as altitude increased the extrapolation fell further from the pace model. Since
this extrapolation was a data standardization technique, it was only used as a verification
for each pace model at each altitude.
The data collected at 10,000 feet (Figure 5) fit the model curve well, with the
exception of the 0.49 Mach number data point. This data point was the first data point
collected during the mission, and the F-15 aircrew noted that the test aircraft was at a
higher altitude than the pacer jet. Disregarding this point, the maximum deviation from
the pace model was 15 feet at 0.54 Mach number. However, the data at the high-speed
end of this curve appeared to be erroneous for a number of reasons. First, Figure 10
demonstrates that the curve crossed the 20,000 feet and 30,000 feet position error lines,
resulting in a final transonic swing that does not coincide with the family of curves.
Second, Figure 12, Figure 16, and Figure 22 demonstrate that the cloverleaf, level
acceleration/deceleration, and constant airspeed turn FTTs produced a curve with a larger
correction than the pace model. The tower fly-by extrapolation also supported this
hypothesis since the high-speed portion of the curve had a larger correction than the pace model. This data indicated that the curve should remain above the 20,000 feet pace model and fit the family of curves.
The data at 20,000 feet (Figure 6) did not fit a curve well, with maximum deviations, from the pace model, of 34 feet at 0.45 Mach number and 0.69 Mach number.
Reasons for the deviations at these data points are unknown. However, the level acceleration/deceleration data (Figure 17) and tower fly-by extrapolation support the
31
general shape and position of the curve, although the extrapolation was 10 to 15 feet above the curve. The data at 30,000 feet, 35,000 feet, and 40,000 feet (Figure 7, Figure 8, and Figure 9) all fit a curve well. The shape of each curve also matched that expected from an altitude position error correction curve based on the tower fly-by results. These best-fit curve models were plotted together in Figure 10. Investigate why the pace data at
10,000 feet and 20,000 feet did not match the data from the other flight test techniques and the tower fly-by extrapolation. (R2)
Cloverleaf
The data from the cloverleaf FTT exhibited larger than expected scatter when compared to the tower fly-by, pace, and acceleration/deceleration data. Figure 12 through
Figure 15 compared the collected data with the best-fit model created from the pace data at each test altitude. Significant scatter around this best-fit model was most likely attributed to an invalid constant wind assumption and angle of attack instability. With the exception of the data points at 10,000 feet, the altitude calculated by applying the cloverleaf altitude position error correction was within 0.5 percent of the altitude predicted by the pace model, which in most applications would be acceptable. The data at 10,000 feet was not included due to the unreliability of the pace model at speeds above
0.6 Mach number at this altitude. In this case, the cloverlead data were not reliable enough to validate the data gathered from the pace FTT. Reference the Cloverleaf FTT
Analysis in this section for conclusions and recommendations.
32
Level Acceleration/Deceleration
The level acceleration/deceleration flight test technique provided the best
resolution for determining whether the pace data were accurate due to the large number
of data points gathered. Unfortunately, for this FTT to be accurate, the altitude position
correction error was required at the airspeed used during the constant airspeed run.
Multiple data points were taken at or near 300 KCAS at each altitude during the pace
mission; however, during the 10,000 feet and 20,000 feet these points were separated by
23 and 36 feet, respectively. For all altitudes, the best fit line through the pace data was
used as the truth correction to apply to the constant airspeed pass. The resultant level acceleration/deceleration curves are displayed in Figure 16 through Figure 20. In all of the graphs, at some point (depending on altitude), the curves created with the level acceleration and the level deceleration split. This split, expanded in Figure 21, was most likely due to engine effects since the five knots per second acceleration was accomplished through a slow throttle movement, and the five knots per second deceleration was accomplished with the throttle at idle power. This difference in engine setting likely changed the pressure field near the nose of the aircraft which resulted in the shift in altitude position error correction. Investigate the effects of engine airflow on
Pitot-static position error corrections. (R3) Due to the fact that this aircraft will be used as a pacer with the throttle set at a position to maintain a constant airspeed, the acceleration data were primarily used when comparing to the pace results since it more accurately represented the engine airflow required during a typical pace mission.
33
In each case, the level acceleration/deceleration data were plotted against the pace model, with results matching the model except in the 10,000 feet and 40,000 feet case.
The reasoning behind the 10,000 foot data not matching the pace data was discussed above. The acceleration/deceleration data collected at 40,000 feet were marginal at best due to the lack of specific excess power at this altitude and the resultant throttle position.
In order to accelerate the aircraft, the throttle was immediately set to maximum power, instead of using a slow throttle movement. In the deceleration, the throttle was set to idle power as discussed above.
Constant Airspeed Turns
The constant airspeed turn flight test technique provided a third comparison to validate the pace data. Since this technique was the test team’s lowest priority, data were only collected at 10,000 feet. Figure 22 displays the results of this FTT compared against the pace model. As previously discussed, this constant airspeed turn data matched the level acceleration/deceleration and cloverleaf data, supporting the conclusion that the pace data at 10,000 feet above 0.6 Mach number were erroneous. Overall, the constant airspeed turn data were considered marginal due to gusty winds during the data collection. A more in-depth analysis is made in the Angle of Attack Effects section below.
34
Zero Total Pressure Error Analysis
When calculating the airspeed and Mach number position error corrections from the altitude position error corrections, the total pressure error present in the system was assumed to be zero. To determine whether this assumption was valid, the difference between the airspeed position error correction as calculated directly from the pacer and the airspeed position error when calculated from the altitude position error correction was determined. Figure 23 illustrates this difference (or “calibrated airspeed error”) plotted versus instrument corrected Mach number. Typically, a zero total pressure error can be assumed when the data are centered on zero with scatter less than ±1 KCAS.
The majority of the data points in Figure 23 were centered on zero and were within ±3.3 knots of zero error. There were three points well outside of the data band. A data point at 0.57 Mach number and 30,000 feet as well as a data point at 0.68 Mach number and 40,000 feet were attributed to the two aircraft being at different airspeeds, thereby inducing error into the calculation and resulting in a calibrated airspeed error of greater than ±10 knots. A point at 0.94 Mach number and 10,000 feet also had a calibrated airspeed error of greater than 10 knots, but the cause of this large error could not be determined. This point was removed from the data analysis using engineering judgment. Only four other points had errors greater than ±3 KCAS. There were eight test points with errors between ±2 and 3 KCAS, and twenty test points with errors between ±1 and 2 KCAS. Only 29 of the 66 total test points had errors within ±1 KCAS.
The cause of the larger-than-normal data scatter was not determined.
35
Airspeed Position Error Correction
The airspeed position error correction, ∆Vpc, was determined from the tower fly-
by altitude position error correction data (Figure 24) and from the pace altitude position
error correction data (Figure 25 through Figure 29), assuming zero total pressure error.
The results of those two flight test techniques are summarized in Figure 30. The tower
fly-by results were extrapolated to higher altitudes and are presented in Figure 31.
The results from the cloverleaf flight test technique, an airspeed measurement
technique, are presented in Figure 32 through Figure 35. The results from the level
acceleration/deceleration flight test technique are presented in Figure 36 through Figure
40. The analysis, conclusions, and recommendations made in the altitude position error
correction section directly apply to the airspeed position error correction data since the
same data were used to derive these results.
Mach Number Position Error Correction
The Mach number position error correction, ∆Mpc, was determined from the tower fly-by altitude position error correction data (Figure 41) and from the pace altitude position error correction data (Figure 42 through Figure 46). The Mach number position error correction fairings are summarized in Figure 47. Mach number position error corrections were also determined from the cloverleaf FTT data (Figure 48 through Figure
51), from the level acceleration/deceleration data (Figure 52 through Figure 56), and from the constant airspeed turn data at 10,000 feet (Figure 57). Ideally, this data should fit a single curve, assuming that angle of attack effects are negligible. Determine why Mach
36
number position error corrections at all altitudes did not fit a single curve. (R4) As in the
airspeed position error corrections, the analysis, conclusions, and recommendations made
in the altitude position error correction section directly apply to the Mach number
position error correction data since the same data were used to derive these results.
Angle of Attack Effects
In order to analyze angle of attack effects on the position error corrections, a
series of constant airspeed turns were conducted with varying load factor. Unfortunately,
one of the primary assumptions was that the winds at the target altitude were constant in
magnitude and heading for the duration of the turn. During the mission when this
maneuver was accomplished, the winds were extremely gusty, making it difficult for the
aircraft to maintain a constant airspeed over an entire 360 degree turn. Thus, most of the
data were removed from the analysis due to the resultant airspeeds being well outside of
tolerances. The data that are presented had a sinusoidal variation of airspeed over the
duration of the turn but were bounded near the tolerance of ±5 KCAS. Data collection
was limited to 10,000 feet due to fuel restrictions and aircraft system failures.
Figure 58 presents the position error pressure coefficient, ∆Pp/qcic, versus angle of
attack for each data point along with a 10,000 foot pace data best-fit curve. All of the
data presented from this flight test technique fell around a best-fit line indicating the
position error pressure coefficient was dependent on the angle of attack changes due to
changes in Mach number. When comparing points at a constant Mach number the data indicated an angle of attack effect existed. The 0.9M points had the strongest correlation
37
between changes in angle of attack and changes in the position error pressure coefficient,
while the 0.7M points demonstrated a minimal change position error pressure coefficient
as the angle of attack changed. In general, the constant Mach number position error pressure coefficients did not generally follow the pace best-fit line (changes due to Mach number) as the angle of attack varied. Therefore, an angle of attack effect existed in the position error pressure coefficient which was independent of the angle of attack effect caused by changing Mach number. The exact relationship between angle of attack and position error pressure coefficient was unknown due to the limited data and the turbulent conditions in which the data were collected. Figure 59 and Figure 60 show similar relationships between altitude position error corrections and Mach number position error corrections. Each of these graphs indicated a relationship between Mach number and changes in angle of attack that affected the position error corrections; the strongest correlations was at 0.9M and the weakest correlation at 0.7M.
4.2 Total Air Temperature Probe Recovery Factor
The total air temperature probe recovery factor was determined using both the
tower fly-by and pace FTTs, and validated using the level acceleration/deceleration FTT.
The recovery factor was successfully determined, as all data fell within ±0.05 units of the
best fit line of the instrument corrected total air temperature parameter, 5(Tic/Ta -1). For the tower fly-by data, the instrument corrected temperature on the aircraft was compared to the ambient air temperature measured in the fly-by tower. During the pace, the temperature was compared to the ambient air temperature determined by the F-15B pacer
38
jet. For the level accelerations and decelerations, the temperature was compared to the ambient air temperature from the rawinsonde balloons. The resultant recovery factor
(from the tower fly-by data) was 0.98, which was lower than historical results (Figure
61). While enough data were collected to provide confidence to this result, the difference with historical results called the condition of the total air temperature probe into question.
Additional recovery factors were calculated using the acceleration/deceleration data and weather balloon data. The temperature recovery factors were calculated using the subsonic deceleration portion of the data due to the lower rate of change in airspeed.
These results are presented in Figure 62 through Figure 66 and Table 2 with most altitudes agreeing with the tower fly-by data except for the 40,000 foot data. All of the acceleration/deceleration data were questionable due to the time difference between the test runs and when the weather balloon was at the test altitude. The 40,000 feet deceleration difference was most likely due to the significant time difference
(approximately 1hr and 20 min) between the balloon data and the aircraft data.
Determine the condition of the total air temperature probe prior to future aircraft calibrations. (R5)
Table 2. Total Air Temperature Probe Calibration Summary Flight Test Technique Pressure Altitude Calculated Calculated Bias (1000 feet) Recovery Factor (N/D) (N/D) Tower Fly-By 2.3 0.9818 -0.0416 Acceleration/Deceleration 10 0.9811 -0.0402 Acceleration/Deceleration 20 0.9457 -0.0238 Acceleration/Deceleration 30 0.9998 -0.0284 Acceleration/Deceleration 35 1.0644 -0.0638 Acceleration/Deceleration 40 0.8602 0.0780
39
4.3 Cloverleaf Flight Test Technique Analysis
Flying proper cloverleaf patterns took practice to accomplish with any degree of accuracy. The higher altitude patterns were the most challenging due to higher winds and poor engine performance. To minimize possible wind changes and to accomplish as many data points as possible during the flight, a minimum time to fly a maneuver was computed and used. To standardize the maneuver and ensure minimum maneuver time, a turn radius chart consisting of planned load factor (2g or 3g), airspeed, and altitude was used to compute a no-wind fly-out distance from the cloverleaf center point. This provided a starting point to adjust fly-out distance real-time based on the test day winds.
The legs with strong headwinds on the outbound portion were the most challenging because, if the fly-out distance was misjudged, little time was available to stabilize on an airspeed before the center point was crossed. Because of the small airspeed tolerances, calibrated airspeeds were computed for each test point to provide a target airspeed.
Using these procedures, data were gathered within the established data bands and tolerances, but the graphs yielded significant scatter with no definable fairing.
First, the quality of the cloverleaf data was examined to yield any possible causes for the scatter. The 0.85M test points at all altitudes were consistently above a normal fairing of the data. The 0.85M test points were re-flown after the second ground test of the Pitot-static system and, while the results were closer to expected values, they still were significantly above the expected fairing of the data. Since the cloverleaf data accuracy was better when using constant airspeeds between legs, data points were
40
selected with a focus on matching calibrated airspeeds. Each leg’s data points were
within 1 knot of each other and within fourteen seconds of the cloverleaf center point to
attempt to satisfy the constant wind assumption.
Additionally, the constant wind assumption was examined to yield any possible
causes for the scatter. The cloverleaf FTT data reduction assumed steady state winds
throughout each approximately four minute maneuver. This assumption was evaluated
by comparing the wind calculations from the cloverleaf data reduction for the different
test points at the same altitude. If the calculated winds were consistent across five test
points spanning approximately 20 minutes, the constant wind assumption was considered
to be reasonable. The results of this analysis are shown in Table 3 with the worst wind
deviation in magnitude and heading displayed in the rightmost column. The 10,000 foot
data points were flown on two separate days, 0.5M through 0.8M on 23 April 2004 and
0.85 through 0.9M on 16 April 2004.
Table 3. Cloverleaf Wind Analysis Alt 0.50M 0.60M 0.70M 0.80M 0.85M 0.90M Deviation hdg/kts hdg/kts hdg/kts hdg/kts hdg/kts hdg/kts hdg/kts
10K 230/14.3 199/17.7 215/20.2 230/25.8 37/25.4 34/30.6 31/8.1 20K 29/60.7 30/60.9 29/60.2 30/61.3 31/61.3 29/61.3 1/1.1 30K 42/85.5 45/84.6 44/82.3 45/86.3 44/86.1 3/4.0 35K 47/88.1 46/89.3 47/90.4 46/87.6 46/88.7 1/2.8
Based on Table 3, the winds changed from 1.1 to 8.1 knots. Wind changes can
have a significant effect on the ground speed used in the data reduction. For example, a change in wind velocity of 4 knots translated to an associated Mach number change of
0.011M at 30,000 ft. The winds further changed between legs on each cloverleaf. Table
41
4 shows the wind speed and heading along with the greatest deviation between a set of
three cloverleaf legs. The winds were quite variable with the greatest heading change of
18 degrees and the greatest speed change of 8 knots. A change of 8 knots at 35,000 ft equates to a Mach number change of 0.025M. These wind changes violated the constant wind assumption used in the cloverleaf FTT and caused scatter in the data at a 0.01M resolution. Overall, the constant wind assumption was not valid.
Table 4. Cloverleaf Inter-leg Wind Analysis Leg 1 Leg 2 Leg 3 Deviation hdg/kts hdg/kts hdg/kts hdg/kts 10K, 0.5M 252.1/19.9 204.5/22.5 233.9/15.7 18.2/4.2 20K, 0.8M 57.3/12.2 64.7/16.9 59.3/13.9 7.4/4.7 30K, 0.8M 31.1/84.4 31.8/84.3 25.9/82.4 5.2/2.0 35K, 0.8M 33.9/88.5 34.5/92.5 32.5/84.2 2.0/8.3
Finally, another possible reason for the scatter was airspeed stability during the
three legs of the cloverleaf maneuvers. While the target airspeed was attained within data
tolerances during each leg, the length of time at the target airspeed varied between legs.
Based on the data, the first leg of every maneuver was the most stable with airspeed
stability lasting more than five seconds. For some legs, the target airspeed was only
attained for a second of time while the aircraft decelerated or accelerated through the
airspeed. This was common on legs where the aircraft exhibited poor engine
performance (higher altitude points) and when the inbound distance was shortened by
tailwinds. This airspeed instability translated into angle of attack instability and, as
discovered during this project, angle of attack appeared to introduce some error into the pressure measurements which adversely affected the data. Increasing leg lengths would
42
provide more stabilization time and remove possible angle of attack effects. Increase cloverleaf leg lengths to remove possible angle of attack effects from the data. (R6)
Based on the ∆Mpc vs. Mach number cloverleaf graphs (Figure 48 through Figure
51), the Mach number calculated by applying the cloverleaf Mach number position error correction was within 0.7 percent of the Mach number predicted by the pace model. A cloverleaf flight test technique test on the previous AFFTC pacer F-16B also yielded results within 1 percent of the calibration curves calculated via other FTTs (Reference 6).
Therefore, based on this error spread, the most realistic position correction resolution was
0.005M for ∆Mpc. For some aircraft Pitot-static systems where this degree of error is acceptable, the cloverleaf FTT is a recommended technique. If a higher degree of accuracy is necessary, the cloverleaf FTT is not recommended as an air data system calibration FTT. Utilize the cloverleaf Pitot-static calibration method if an error of ±1 percent is acceptable. (R7)
43
CHAPTER 5
CONCLUSIONS AND FUTURE WORK
This chapter presents the overall conclusions of the TRUE PHOENIX test project and lists recommendations in order to improve the project in the future.
5.1 Conclusions
The overall test objective was to determine the air data system position errors of
F-16B, S/N 92-0457. Position errors from approximately 2300 feet to 40,000 feet at approach speed to 1.4 Mach number were determined. The final calibration curves determined the system position errors between 200 KCAS and 0.93 Mach number.
However, the reliability of the position error correction curves at 10,000 and 20,000 feet were questionable.
The pace position error correction model at 10,000 feet did not fit the family of curves formed by the data at 30,000 feet, 35,000 feet, and 40,000 feet. The curve at
10,000 feet (Figure 10) crossed the 20,000 feet and 30,000 feet position error correction curves, resulting in a final transonic swing that did not coincide with the family of curves.
Second, Figure 11, Figure 16, and Figure 22 demonstrated that the cloverleaf, level acceleration/deceleration, and constant airspeed turn flight test techniques produced a curve above the 10,000 feet pace model. The data at 20,000 feet (Figure 6) did not fit a curve well, with maximum deviations of 34 feet at 0.45 Mach number and 0.69 Mach number. Reasons for the deviations at these data points are unknown.
44
The level acceleration/deceleration data showed an engine airflow effect on Pitot-
static position error corrections. At some point (depending on altitude), the curves
created from the level acceleration and the level deceleration data split. This split,
expanded in Figure 21, was most likely due to engine effects since the five knots per
second acceleration was accomplished through a slow throttle movement forward, and
the five knots per second deceleration was accomplished with the throttle at idle power.
This might adversely effect position error corrections and introduce error in future
calibrations.
The Mach number position error correction, ∆M pc , formed a family of curves at
different altitudes. In a similar fashion to the airspeed position error correction, the Mach
number position error correction was also derived from the altitude position error
correction, assuming zero total pressure error in the Pitot-static system. Ideally, this data
should fit a single curve, assuming that angle of attack effects were negligible.
The total air temperature probe recovery factor was evaluated using both the
tower fly-by and pace FTTs. The resultant recovery factor was 0.98, which was lower
than other historical results. The difference with historical results calls the condition of the total air temperature probe into question.
The cloverleaf data showed a data scatter around the position error correction family of curves. While the target airspeed was attained within data tolerances during each cloverleaf leg, the length of time at the target airspeed varied between legs. For some legs, the target airspeed was only attained for a second of time while the aircraft decelerated or accelerated through the airspeed. This airspeed instability translated into
45
angle of attack instability and, as discovered during this project (Figure 60), angle of attack appeared to introduce some error into the pressure measurements.
Based on the ∆Mpc vs. Mach number cloverleaf graphs (Figure 48 through Figure
51), the cloverleaf data were within 1.2 percent of the position correction error from the data gathered from the other flight test techniques. A cloverleaf flight test technique test on the previous AFFTC pacer F-16B also yielded results within 1 percent of the calibration curves calculated via other flight test techniques (Reference 6).
5.2 Recommendations for Future Work
Since the reliability of the position error correction curves at 10,000 and 20,000 feet were questionable, recommend updating the F-16B, S/N 92-0457 pacer system calibration curves using the tower fly-by and pace curves at 30,000 feet, 35,000 feet, and
40,000 feet. (R1) This would improve the position errors of the aircraft with data that is more acceptable.
Recommend investigating why the pace data at 10,000 feet and 20,000 feet did not match the data from the other flight test techniques and the tower fly-by extrapolation. (R2) The pace position error correction model at 10,000 and 20,000 feet did not fit the family of curves formed by the data at 30,000 feet, 35,000 feet, and 40,000 feet and the deviations at these data points are unknown.
Since the level acceleration/deceleration data showed an engine airflow effect on
Pitot-static position error corrections and curves split, recommend investigating the effects of engine airflow on Pitot-static position error corrections. (R3)
46
Recommend determining why the Mach number position error corrections at all
altitudes did not fit a single curve. (R4) The Mach number position error correction,
∆M pc , should fit a single curve.
Recommend determining the condition of the total air temperature probe prior to
future aircraft calibrations since the resultant recovery factor was 0.98, which was lower
than other historical results. (R5)
Recommend increasing the cloverleaf leg lengths to remove possible angle of
attack effects since the data showed scatter around the position error correction family of
curves. (R6) Increasing leg lengths would provide more stabilization time and remove
possible angle of attack effects.
The cloverleaf flight test technique yielded results within 1 percent of the calibration curves. For some aircraft Pitot-static systems where this degree of error is acceptable, the cloverleaf FTT is a recommended technique. (R7)
47
BIBLIOGRAPHY
“F-16B Modification Flight Manual,” USAF Series, F-16 A/B Aircraft, F-16B Serial Number 92-0457, Air Force Flight Test Center, Edwards AFB, CA, 1 October 2003.
Flight Manual, USAF/EPAF Series Aircraft, F-16 A/B Block 15, Technical order 1F-16A-1, Change 14, 15 August 2003.
Supplemental Flight Manual, USAF/EPAF Series Aircraft, F-16 A/B Block 15, Technical Order 1F-16A-1-1, Change 10, 15 February 2003.
AFFTC 11-1, Aircrew Operations, 14 January 2004.
“Pitot Statics Calibration and the Standard Atmosphere”, USAF Test Pilot School, Edwards AFB, CA, January 2004.
Olson, W.M., “Pitot-Static Calibrations Using a GPS Multi-Track Method,” Presented at SFTE Symposium, Reno, NV, 1998, http://www.camasrelay.com/aircraftperformance.htm.
Herrington, R.M. et. al., “Flight Test Engineering Handbook,” AF Technical Report No. 6273, May 1951.
48
APPENDIX A
GRAPHICAL ILLUSTRATIONS
49
95 . t) 0 0f (230 l e Mod y b y Fl
0.85 r e w o T ired a F B
F T il r 0.75 p t: Hand i A 30 e F er Fly-By) v r w u C B ic F T 5 , M l r 6 . ri e 0 p b ection (To A r 13 ach Num
B F T l i r p 0.55 Indicated M 12 A Position Error Cor e B F T il 5 r 4 p . 0 A 7 Figure 3. Altitud s 04 k n r a p se i B ng T i st u 30 A 6 0.35 e 1 - F L W
04 - : n: Cr r A t 457 f ght T i 0 G a - r 2 c : Fl r 9 i 7 Ap e : guratio i s A o 370 t #: w s il ourc a T Te T Conf Date S 0.25 0 0 0 0 0 0 0 80 60 40 20
20 18 16 14 12 10
pc pc ∆ ) eet (f H , n o ecti r r Co ror r E tion Posi tude ti l A
50
4.0 + 31 .5 3 .48*x es u l a = 31 V e y d v e r u rv se al C c i Ob et 3.0 heor T g .5 2 Readin . Tower Flyby Grid Reading id r G 2.0 Figure 4. Radar Altimeter vs s 04 ank .5 1 e t B ng T i s uis 30 Apr 6 e 1 - Cr F 04 - : AL W ft 0457 ra ation: : Flight T rc i ur 7 Apr ig ce s: A o 370G f te n w st il #: 92- o a e a T T T C D Sour 0 . 1
80 60 40
160 140 120 100
) t e e (f t u p t u O r te e tim l A r a d a R
51
0
. 1 n o ati ol ap r Ext B 9 F . 0 T K 10 l 8 . de 0 ce Mo ired Fa K Pa nd 10 a 7 . 00 ft PA Pace) it: H ic F e , M r e b Curve K Pac 10 60 . ch Num 0 a M
d te a dic In 5 . Position Error Correction (10,0 e 40 . 0 2 s 3 k Figure 5. Altitud n -01 a T
76 g se i B n i N: st 6 e 3
. S/ W Cru T / t 04 L F-1 : n: h 7 r B A t f g 5 p 5 io i l t 0G A -04 : F rcra i ura F-1 37 14 ig A o te: nf urce cer: w st N: 92 e T S/ Co T Da So Pa 20 . 0 0 0 0 0 0 50
25 20 15 10
pc ∆ et) (fe H , tion ec rr Co r o r r E tion i s o P tude Alti
52
1.0 n o ati apol tr Ex B .9 F 0 d e r 20K T Fai
d n a H : t 0.8 Fi e Model rv u C 20K Pace .7 00 ft PA Pace) 0 ic e er, M b m 20K Pac Nu 0.6 ach M
ed cat i d In .5 0 Position Error Correction (20,0 e 0.4 s Figure 6. Altitud ank T 76-0132 : t B ng i s uise 6 r .3 e 1 0 F-
04 AL W ft: on: C ght T i l 15B / S/N ra - F rc i 14 Apr r: igurati ce: F A o 370G e t w s e T T S/N: 92-0457 Conf Date: Sour Pac 0.2 0 50
250 200 150 100
pc pc ∆ ) eet f ( H on, i t c e r Cor r o r r E tion i Pos Altitude
53
0
. 1 ion apolat tr x B E .9 F 0 T ed K 30 Fair l e 0.8 Fit: Hand e v Mod e Cur 30K Pac 7 . 00 ft PA Pace) 0 ic e er, M b Pac K 30 ch Num 0.6 a
cated M i Ind .5 0 Position Error Correction (30,0 e 4 0. s 132 nk 0 Figure 7. Altitud a e 76- : B ng T i st uis 6 r e 1 .3 - 0 T t 04 h 7 r AL W t: F f g p ion: C i t 15B / S/N A 0G a 045 Fl - : rcra F- 2 : r igur 9 o 37 e : te: 14 nf urce c w st Ai N / o a e T S Co T Da S P 0.2 0 0 0 0 0 50
25 20 15 10
pc pc ∆ ) eet f ( H , on i rrect o C ror r E on i t si o P ude t i t l A
54
0 1. ion polat tra B Ex .9 F 0 K T 5 3 0.8 odel M e ed r K Pac Fai 5 d 3 .7 00 ft PA Pace) 0 Han : t ic , M e Fi e r e rv b u m C u N 35K Pac 0.6 ach M
d e t dica In .5 0 Position Error Correction (35,0 e 0.4 2 s 3 k Figure 8. Altitud n -01 a 6 T e t ng i s uis 3 r e
. 0 F-16B 04 ht T 7 r B / S/N: 7 AL W t: p 5 ion: C A raf 0G at 045 c - : Flig r i F-1 ur : e r ig c : 14 A 92 o 37 e : te w st a onf e T S/N C T D Sour Pac 0.2 0 50
250 200 150 100
pc ∆ ) eet f ( H n, ectio r r Co r o r r E sition o P Altitude
55
0
. 1 n io polat a tr B Ex .9 F 0 K T 0 4 l 8 e 0. Mod e K Pac d 40 ire .7 a 00 ft PA Pace) 0 F
ic M , e r e b it: Hand m K Pac u e F 40 N v 6 . ch 0 a Cur M
d e t ca di n I 5 . 0 Position Error Correction (40,0 e 4 . 0 2 s 3 k 1 Figure 9. Altitud n 0 - a 6 T
e 7 B ng N: i st uis e
.3 S/ W
T 4
0 Cr
F-16 L : n: ht r 0 B / A t f g p io i G l 457 15 at - A 0 F : rcra i ur F 14 : r ig ce : 92- o 370 r e te c u w st A N: onf o e T S/ C T Da S Pa 2 0. 0 0 0 0 0 50
25 20 15 10
pc ∆ ) eet f ( H ion, t c e r r o C ror r E tion Posi titude Al
56
0 1. del del .9 0 Mo Mo e e c c a a K P K P 20 40 0.8 .7 0 ic del del , M r Mo Mo e e e b c c a a m u K P K P N 10 35 0.6 rrection (Tower Fly-By & Pace) ach o M
d e t rror C dica t) In 0f .5 ition E 0 230 Pos e odel ( del M
y b Mo e c Fly a r 0.4 e w K P o 2
T 30 s 013 r 04 ank Figure 10. Altitud 76- Ap : se ng T i st e 3 -30 16B . - W 0 / S/N T
L n: Crui ht B r 04 A t: F f o g i i t l a 15 r - 0457 : F ura F 370G : g r i s: 7 Ap ce Airc 92- o f r e t : n te c u w s N o a / e T S Co T Da S P 0.2 0 50
250 200 150 100
pc ∆ ) eet f ( H n, io t c e r Cor ror Er n tio i Pos de Altitu
57
0 1. ion ion polat polat a a .9 tr tr 0 B Ex B Ex F F K T K T 20 40 atiion) l 0.8 ion ion .7 0 polat polat ic a a tr tr , M r e B Ex B Ex b F F m u K T K T N 10 35 0.6 on (Tower Fly-By and Extrapo ach M
d e t rrecti o ) dica t f In rror C .5 n 0 2300 olatio del ( ition E ap tr Mo y b Pos y l e B Ex F F r e T 0.4 w K o T 30 4 s 0 r ank p A se ng T i st e 3 - 30 16B . - W T 0 Figure 11. Altitud L n: Crui ht r 04 A t: F f o g p i i 0457 t l a r : F ura 92- 370G g i s: 7 A ce Airc o f r t n te u w s il #: o e a T T Co T Da S 0.2 0 50
250 200 150 100
pc ∆ ) eet f ( H , tion ec Corr r o r r E tion i Pos de Altitu
58
1.0 n o ati apol tr Ex B .9 F 0 10K T 0.8 10K Cloverleaf .7 0 ic 0,000 ft PA Cloverleaf) er, M Model b
e m Nu 0.6 ach rrection (1 10K Pac M o
C ed cat i rror d In .5 0 ition E Pos e 0.4 s 04 ank T Figure 12. Altitud 30 Apr t B ng - i s uise 6 r e .3 1 0 04 F- Pod
AL W ft: on: C ght T i Apr 0457 l ra rc i igurati ce: F A o 370G t #: 92- l ne ARDS w s i e a T O T T Conf Dates: 16 Sour 0.2 0 50
300 250 200 150 100
pc ∆ ) (feet H , ion t c e r Cor r o r r E on ti i s o P Altitude
59
0 1. olation ap tr B Ex F .9 0 20K T 0.8 20K f lea over l C .7 0 ic , M r 0,000 ft PA Cloverleaf) odel e b M e m u N K Pac 0.6 0 ach rrection (2 2 o M
d C e t rror dica In .5 0 ition E Pos e 0.4 s 04 k r n p a A T e Figure 13. Altitud t 30 ng - i s uis 3 r e . 0 7 04 F-16B 5 Pod ht T AL W t: 4 ion: C 0 - raf 0G at c : Flig r 92 i ur 16 Apr e : ARDS ig c s A o 37 #: l te w i st onf a a e T T C T One D Sour 0.2 0 50
250 200 150 100
pc ∆ ) eet f ( H ection, Corr r o Err on i Posit de Altitu
60
1.0 n o ati ol trap x E B .9 F 0 T K 30 K 8 0 0. 3 f a e erl ov Cl .7 0 ic , M del 0,000 ft PA Cloverleaf) ce Mo Number Pa 6 . K ch 0 a rrection (3 30 o
d M C e t ca rror Indi .5 0 ition E Pos e 0.4 s r 04 nk p a A T Figure 14. Altitud
0 se B ng i st - 3 6 .3 e d 4 0 T W Crui 7 : r 0 Po n ht AL o S ig Ap -045 0G ati r 16 : Fl rcraft: F-1 92 i u 37 ARD g i A o e t #: n tes: nf urce w s il e a T T Co T O Da So 2 0. 0 0 0 0 0 50
25 20 15 10
pc ∆ ) eet f ( H on, ti c e r r o C r o r r E tion si o P e titud l A
61
0 1. ion polat tra B Ex .9 F 0 K T 5 3 K 0.8 35 af le er v o l C .7 0 ic l , M r 5,000 ft PA Cloverleaf) e b Mode
m e u N 0.6 ach rrection (3 35K Pac o M
d C e t rror dica In .5 0 ition E Pos e 0.4 s 04 ank Figure 15. Altitud se 30 Apr ng T i st 3 . e 16B - 0 W 04 - T
L Pod n: Crui ht A t: F f S o g i i 0457 t l a Apr D r 6 R : F ura 1 92- A 370G : g i ce Airc o f r t ne n tes u w s il #: a o e a T T Co T O D S 0.2 0 50
250 200 150 100
pc ∆ et) (fe H ion, ect Corr ror Er on i Posit de Altitu
62
Model e 1.4 10K Pac l e Dec l/ .2 e 1 c Ac 1.0 ic M , er b m 0,000 ft PA Level Accel/Decel) u ach N ction (1 .8
0 ed M rre o dicat In rror C ition E 0.6 Pos e s ank T e .4 0 t B ng i s uis 6 e W 04 : F-1 AL ft ion: Cr 0457 t light T DS Pod ra Figure 16. Altitud a r : F rc 92- i 370G 23 Apr e igu c : A o r t ne AR te w s il #: onf a e a O T T T C D Sou 0.2 0 0 0 0 0 0 0 00 00 00 00 80 60 40 20
-2 -4 -6 -8
120 100
pc ∆ ) eet f ( H , n o i t c e rr o C r ro Er n o i t i s Po e d u t i t l A
63
l ode M
e 1.4 20K Pac l e Dec / l .2 e 1 c Ac 1.0 ic M er, mb u 0,000 ft PA Level Accel/Decel) N h c a M .8 ction (2 0 ed rre o cat di In rror C ition E 0.6 Pos e s k n a T e .4 s t i 0 ng i s u r e 16B Figure 17. Altitud T F- t 04 Pod n: C AL W t: 457 f o i a -0 DS r 0G at r Fligh : 92 e 20 Apr igu Airc e AR o 37 t #: n te: urc w s il a onf e a T T C T O D So 0.2 0 0 0 0 0 200 000 80 60 40 20
-200 -400 -600
1 1
pc ∆ ) ) eet f ( H , ion rrect o C Error on i sit Po e d u tit l A
64
Model e 1.4 K Pac 30 l e l Dec .2 e 1 c Ac 30k Level Accel/Decel) 1.0 ic , M r e 0,000 ft PA Numb ch a M .8 ction (3 0 rre o cated i Ind rror C ition E 0.6 Pos e s ank T e .4 g s t i 0 n i s u e Figure 18. Altitud T F-16B 04 Pod
: ht t -0457 Apr : Flig rcraf 92 i uration: Cr 370GAL W e ig o w il #: st A a e One ARDS T T T Conf Date: 20 Sourc 0.2 0 800 600 400 200
-200 -400 -600 -800
1200 1000
pc ∆ et) et) (fe H n, io Correct ror Er on i Posit de u t i Alt
65
Model
e 1.4 35K Pac l ce De
l .2 1 ce c A k 5 3 1.0 ic M mber, 5,000 ft PA Level Accel/Decel) ach Nu .8 ction (3 0 rre o cated M i d In rror C ition E 0.6 Pos e s ank e .4 s t i 0 ng T i s u r e
d W Figure 19. Altitud C 7 Po : F-16B n: AL t f io Apr 04 ra at 70G : Flight T rc ur e ig o 3 f n w st Ai il #: 92-045 o e a T One ARDS T T C Date: 20 Sourc 0.2 0 0 0 0 0 0 0 0 0 0 0 0 0 00 00 80 60 40 20
-2 -4 -6 -8
12 10
pc ∆ t) t) (fee H n, io Correct ror Er n o i Posit e d u t i t Al
66
Model
e 1.4 40K Pac l ce De
l .2 1 ce c A k 0 4 1.0 ic M er, 0,000 ft PA Level Accel/Decel) Numb ach M .8 ction (4 0 rre o cated i Ind rror C ition E 0.6 Pos e s ank e .4 t 0 B ng T i s e Figure 20. Altitud T 04 57 : F-16 n: Cruis AL W t : Flight rcraf 92-04 i uratio e ig o 370G f te: 30 Apr w st A il #: a e a T One ARDS Pod T T Con D Sourc 0.2 0 0 0 0 0 0 0 00 00 00 00 80 60 40 20
-2 -4 -6 -8
120 100
pc ∆ t) t) (fee H n, io ect Corr Error on i t si Po de u t i Alt
67
85 . 0 l de o ce M a K P 80 . 10 0 l ation ce e D eler / l c 5 c A cce 0.7 A 0 0.7 ic Level Accel/Decel) M , r A on Numbe ti h a r c 0.65 e a s (10,000 ft P d M Decel te ca ndi I 60 . 0 55 . 0 s nk Figure 21. Engine Airflow Effect a T e s t ng i s ui e 50 16B W -
Cr 0. L Pod 04
ht T A t: F 457 S g i G l D -0 raf 0 ation: c 7 : F r ur 92 e AR : g i c Ai o 3 e t # n nf w s il e a T T Co T O Date: 23 Apr Sour 5 0.4 90 80 70 60 50
150 140 130 120 110 100
pc ∆ ) eet (f H on, i t c e r Cor ror r E on ti i s o P tude ti Al
68
1.0 .5g 2 10K/ rns u T .9 0 tant G s ) Con 0.8 10K/2g 10K/3.7g rns rns u u T T .7 0 ic tant G tant G ns ns er, M b Co Co m 0,000 ft PA Constant Airspeed Turn 0.6 ach Nu M on (1 ed 10K/3.5g 10K/1.5g rrecti cat s s i o d rn rn u u In .5 0 rror C G T G T tant tant ition E Cons Cons Pos e 0.4 s nk 10K/3g a s n r e u s t B ng T odel i s ui 6 e .3 1 M 0 G T e Figure 22. Altitud t T t F- 04 Pod : h AL W ft tan g 0457 RDS : Fli rcra i uration: Cr e 10K Pac Cons g i c A o 370G #: 92- l w i st e a T T Conf T One A Date: 23 Apr Sour 0.2 0 50
250 200 150 100
pc ∆ eet) (f H n, o i t c e r r o C r o r r E on i t i s o P de tu i t Al
69
1.0 .9 0 0.8 .7 0 ic mber, M Nu
h essure Validation c r 0.6 a e 40K Pac dicated M In e .5 c 0 Pa K 35 Figure 23. Zero Total P e 0.4 K Pac 30 s e 132 c 0 - ank T 76 e s t i ng i s u 20K Pa .3 e 0 T F-16B L W
04 : e r A t ight G l 15B / S/N: - -0457 : F rcraf i e r: F iguration: Cr A o 370 e t 10K Pac w s e T T S/N: 92 Conf Date: 14 Ap Sourc Pac 0.2
4.0 3.0 2.0 1.0 0.0
-1.0 -2.0 -3.0 -4.0
r o r _er c (knots) V Error, Airspeed ed brat i Cal
70
0.95 Model y b y l F r e w o T 0.85 B F T ril 30 Ap 0.75 B F l T i ic 13 Apr d e er, M ir 0.65 b a m F B d F n a l T i r Correction (Tower Fly-By) H ach Nu : t M i rro F ed 12 Apr e v r cat i 0.55 d Cu sition E In B F T 7 April 0.45 Figure 24. Airspeed Po s nk r 04 a e s t B ng T i s 30 Ap ui 6 0.35 e 1 T t F- 04 - : h AL W ft g 0457 : Fli rcra i uration: Cr 7 Apr e : g i c A o 370G #: 92- l w i st e a T T Conf T Dates Sour 0.25
3.0 2.8 2.6 2.4 2.2 2.0 1.8 1.6 1.4 1.2 1.0
pc ∆ ) s t o (kn V , ion rrect Co r rro E n o i sit Po y cit Velo
71
0 1. olation ap tr Ex B 9 . F 0 K T 0 1 l 8 e ed r 0. Mod e Pace) A Hand Fai K Pac : 0 t 1 7 . 0 rve Fi ic u C , M e r e b m 10K Pac Nu 6 h c 0. a r Correction (10,000 ft P M
rro ed cat i d n I sition E .5 0 4 . 0 s 32 Figure 25. Airspeed Po k n 01 a e 76- t B ng T N: i s uis 6 r e
.3 W T 0
C
: 04 7 ht r B / S/ AL t: F-1 5 ion lig 45 1 raf 0G - at Ap -0 4 : F rc F 2 : e r igur c : 1 Ai 9 o 37 e t : nf te w s e T T S/N Co Da Sour Pac 2 0. 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ nots) (k V , n io rrect o C r ro Er n o i t i s Po y t ci o Vel
72
0
. 1 n io t ola trap ed Ex r B F .9 Fai 0 d n K T 0 2 Ha
: t e Fi rv l u 8 . de C 0 Mo e c Pace) A K Pa 20 .7 0 ic , M e r 20K Pac Numbe 6 h c 0. a r Correction (20,000 ft P
rro ted M ndica I sition E 5 . 0 4 0. 2 s k Figure 26. Airspeed Po n 013 - a 6 T 7 e
g : B n i st uis 6 e 3
1 . Cr 0 / S/N T F- L W 04 ht r B A t: g ion: i G 457 15 Ap raf at -
0 Fl - rc ur r: F ig ce: : 14 Ai 92 o 370 f r e t : e n t c u w s N o a o e T S/ C T D S Pa 2 0. 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ nots) (k V , n io rrect Co r Erro on i t si Po y t ci lo Ve
73
0
. 1 n o i at l o trap x E d B .9 F ire 0 a T K 30 Hand F
t: i l 8 . 0 Curve F Mode e c a Pace) A K P 0 3 .7 0 ic , M r e Pac K Numbe 30 6 h c 0. a r Correction (30,000 ft P
rro ted M ndica I sition E 5 . 0 4 0. 2 s k Figure 27. Airspeed Po n 013 - a 6 T 7 e
g : B n i st uis 6 e 3
1 . Cr 0 / S/N T F- L W 04 ht r B A t: g ion: i G 457 15 Ap raf at -
0 Fl - rc ur r: F ig ce: : 14 Ai 92 o 370 f r e t : e n t c u w s N o a o e T S/ C T D S Pa 2 0. 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc pc ∆ ) nots (k V , n o i t rrec Co r Erro n o i t si Po y t ci o Vel
74
0
. 1 n io t ola trap Ex ed B r F .9 0 K T 5 3 Hand Fai : t l 8 . de rve Fi 0 u C Mo e c Pace) A K Pa 35 .7 0 ic , M e r 35K Pac Numbe 6 h c 0. a r Correction (35,000 ft P
rro ted M ndica I sition E 5 . 0 4 0. 2 s 3 Figure 28. Airspeed Po nk -01 a se t B ng T i s 6 3 e .
1 W 0 T
F-
: n: Crui ht 7 r 04 B / S/N: 76 AL t f o 5 i t lig 1 ra 0G - Ap 045 F - : rc i ura F 2 14 : e g r i : c A 9 o 37 f e : w st e T T S/N Con Date Sour Pac 2 0. 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ nots) (k V , n io rrect o C r Erro n o i t i Pos y t ci o Vel
75
0
. 1 on i lat ired a apo tr Ex
nd F B .9 F Ha 0 t: i K T 40 Curve F 8 . del 0 Mo e Pace) A 40K Pac .7 0 ic M e er, mb K Pac u 0 N
4 6 ch 0. a r Correction (40,000 ft P M
d rro e t ca i d n I sition E 5 . 0 0.4 2 s Figure 29. Airspeed Po 3 nk -01 a se t B ng T i s 6 .3 e
1 W 0 T
F-
: n: Crui ht 7 r 04 B / S/N: 76 AL t f o 5 i t lig 1 ra 0G - Ap 045 F - : rc i ura F 2 14 : e g r i : c A 9 o 37 f e : w st e T T S/N Con Date Sour Pac 2 . 0 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ nots) (k V , n o i t rrec Co r Erro n o i t si Po y t i c Velo
76
0
. 1 .9 0 0.8 l l e e Mod Mod e e c c a a .7 0 20K P 40K P ic M er, b m Nu 6 . ch 0 a Correction (Tower Fly-By & Pace) M
r l l ed rro cat i Mode Mode
d e e In .5 sition E 0 K Pac K Pac 10 35 t) 0f 0.4 230 2
l ( s 3 e 1 04
nk 0 Figure 30. Airspeed Po - a del Mod T Apr e 76 y : Mo t B b
ng i s y uis 30 6 e l e .3 - W F 0 Cr F-1 :
04 : 7 ht T AL t f wer lig o 15B / S/N Apr ra ation - 045 T 30K Pac 70G : F F rc i ur
7 e : : 3 r ig c A 92- o f e t : c ur w s N a / o e T T S Con Dates S P 0.2
3.5 3.0 2.5 2.0 1.5 1.0 0.5 0.0
pc ∆ ) s (knot V , n io rrect o C r Erro n o i t si Po y t i c o Vel
77
1.0 .9 0 8 . 0 apolation apolation tr tr x x B E B E F F .7 0 ic 20K T 40K T , M r mbe u ch N ion (Tower Fly-By and Extrapolated) 0.6 a M
d te polation polation a a a c r Correct tr tr Indi rro B Ex B Ex .5 F F 0 sition E 10K T 35K T t) 0f 0.4 230 s lation 04 r ank apo p tr Model ( e y Figure 31. Airspeed Po t B b ng T Ex i s
uis y 6 l .3 e 30 A B W 0 T F
F L : F-1 n: Cr A t 04 - f o wer K T G light o 457 0 ra ati 0 T 3 Apr
: F rc i 7 e igur c o 370 f ur w st A N: 92- / o e T T S Con Date: S 0.2
3.5 3.0 2.5 2.0 1.5 1.0 0.5 0.0
pc ∆ ) s t o n (k V n, o i rrect Co r Erro on i t Posi y cit Velo
78
0
. 1 n io t ola trap Ex B 9 . F 0 K T 0 1 l 8 . de 0 Mo e c K Pa Cloverleaf) A 10 .7 0 ic M , es av e mber erl u v o 6 ch N Cl 0. a K
Correction (10,000 ft P 10 r d M te rro a Indic .5 sition E 0 4 0. 4 s k r 0 Figure 32. Airspeed Po n p a A
T 0 e B ng i st uis - 3 6 .3 e d 1 0 o 04 W Cr T 7
t L r F- 5 P : h p t f g ion: i l t A -04 ra 0GA a F RDS : 16 rc 92 ur
37 A
: ig ce s o e f r #: l n te u w st Ai i o o e a T T C T On Da S 2 . 0
6 5 4 3 2 1 0
pc ∆ ) nots (k V , n o i t rrec Co r Erro n o i t si Po y t ci o Vel
79
0
. 1 on i lat apo r t Ex B .9 F 0 K T 20 l 8 e . 0 Mod e Cloverleaf) A 20K Pac .7 0 ic M , r e aves mb e l Nu
6 over l ch 0. a
M Correction (20,000 ft P d 20K C r e t rro ca i d n I 5 . 0 sition E 0.4 s 04 k r n Figure 33. Airspeed Po a Ap
T e g 30 B n i st uis - 6 .3 e d 4 1 0 0 Cr T
7 F- L W Po
ht A t: S g ion: i 045 G - raf at 2 Fl rc 9 ur 16 Apr
: ig ce: s Ai e ARD o 370 f r t e n t n u w s il #: a o o e a T T C T O D S 2 . 0 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ ) nots (k V , n o i ct rre Co r Erro n o i t i s Po y t ci o Vel
80
0
. 1 on i lat apo r t Ex B .9 F 0 K T 30 l 8 . e 0 Mod e Cloverleaf) A 30K Pac .7 0 ic , M r aves le Numbe h over l c 0.6 a
Correction (30,000 ft P 30K C r ted M rro ndica I .5 0 sition E 4 0. s 04 nk a Figure 34. Airspeed Po Apr 0 e 3 s t i
B ng T i s u 6 3 r e . d W 0 T o
C
04 - : r 57 P ht AL t: F-1 ion Ap 04 lig
DS raf 0G at R : F rc 16 92- : e A igur s c Ai o 37 t te nf w s il #: e a T One T T Co Da Sour 2 0. 5 0 5 0 5 0 5 0
3. 3. 2. 2. 1. 1. 0. 0.
pc ∆ ) nots (k V , n io rrect Co r Erro n o i t i Pos y t ci Velo
81
1.0 n o polati a tr B Ex .9 F 0 35K T 0.8 Model e Cloverleaf) A 35K Pac .7 0 ic s er, M b eave l er 6 ov 0. ach Num
Correction (35,000 ft P 35K Cl r ed M rro dicat In .5 0 sition E 0.4 s ank Figure 35. Airspeed Po T
e g s t 30 Apr 04 B n i s ui - 6 e .3 4 1 0 0 Pod AL W t: F- f on: Cr ght T i 0457 l a ati : F 16 Apr 92- e : gur i c o 370G r #: l ne ARDS w i st Aircr a e T O T T Conf Dates Sou 0.2
3.5 3.0 2.5 2.0 1.5 1.0 0.5 0.0
pc ∆ ) s t o (kn V , ion rrect Co Error on i Posit y Velocit
82
Model
e 4 1. 10K Pac l ce e D / l .2 1 cce A Level Accel/Decel) 0 A 1. ic , M r e ach Numb M
ection (10,000 ft P .8 d 0 r Corr Indicate rro sition E 0.6 s k n a T e .4 0 B ng i Figure 36. Airspeed Po st uis 6 r e
T
7 t Pod F-1
: r 04 AL W 45 S p igh 0 - A 0G ation: C
2 : Fl rcraft i 37 9 e igur : 23 A e ARD o n nf w il #: st o e a O T T T C Date Sourc 0.2 8 6 4 2 0
-2 -4 -6
12 10
pc ∆ ) nots (k V on, ti c e r r o C r o r r E tion i s o P ty ci o l e V
83
del Mo
e 4 1. 20K Pac l ce e D / l .2 1 cce A Level Accel/Decel) A 1.0 ic Number, M ach
ection (20,000 ft P .8 0 ted M ca i r Corr Ind rro sition E 0.6 s ank e .4 0 B ng T i Figure 37. Airspeed Po st e W
T F-16 t L
04 57 : n: Cruis h A t f g io i G t : Fl rcra i 20 Apr gura i ce o 370 f r te: u w il #: 92-04 st A a o e a One ARDS Pod T T T Con D S 2 0. 8 6 4 2 0
-2 -4 -6
12 10
pc ∆ ) s (knot V ection, Corr Error n o i sit Po y t ci o Vel
84
del Mo e 4 1. K Pac 30 l ce De l .2 1 cce A k 0 3 Level Accel/Decel) A 1.0 ic Number, M ach
ection (30,000 ft P .8 0 ted M ca i r Corr Ind rro sition E 0.6 s ank e .4 0 B ng T i Figure 38. Airspeed Po st e W
T F-16 t L
04 57 : n: Cruis h A t f g io i G t : Fl rcra i 20 Apr gura i ce o 370 f r te: u w il #: 92-04 st A a o e a One ARDS Pod T T T Con D S 2 0. 8 6 4 2 0
-2 -4 -6
12 10
pc ∆ ) nots (k V n, io Correct ror Er on i sit Po y cit Velo
85
12 Test Aircraft: F-16B 35k Accel Decel 35K Pace Model Tail #: 92-0457 10 Configuration: Cruise Two 370GAL Wing Tanks
) One ARDS Pod Date: 20 Apr 04
nots 8 Source: Flight Test (k pc V
∆ 6
, n o i t c
e 4 rr o
2 86
Error C n o i
sit 0 Po y
cit -2 Velo
-4
-6 0.20.40.60.81.01.21.4
Indicated Mach Number, Mic
Figure 39. Airspeed Position Error Correction (35,000 ft PA Level Accel/Decel)
Model e 1.4 Pac 40K l ce e D
l 2 . 1 ce c A k 0 4 Level Accel/Decel) A 1.0 ic , M mber u ach N 8
ection (40,000 ft P . 0 cated M i r Corr Ind rro sition E 0.6 s ank e .4 0 t B ng T i Figure 40. Airspeed Po s uis 6 e d o T W L 04 : F-1 A t 457 f on: Cr 0 ati : Flight rcra 92- i ur e ig c o 370G f w st A il #: our e a T One ARDS P T T Con Date: 30 Apr S 2 0. 8 6 4 2 0
-2 -4 -6
12 10
pc ∆ ts) o (kn V n, ectio r r Co r o Err n sitio Po city o Vel
87
1.0 l e d o y M yb l F r e .9 w 0 o T B F l T i 0.8 30 Apr B F ed 7 . ir 0 l T i ic Fa d er, M 13 Apr b Han : Num e Fit B ch 0.6 rv F a u l T M C i Error Correction (Tower Fly-By) ed cat i 12 Apr d In .5 0 B F l T i 7 Apr 0.4 s k n r 04 a p T Figure 41. Mach Number Position A e t ng i s - 30 e
16B .3 - T 4 0 ht r 0 AL W t: F f g p 0457 a - r ation: Cruis 2 70G : Fli rc i e gur i c A o 3 #: 9 l tes: 7 A ur w st i e a T T Conf T Da So 0.2 0 5 0 5 0 5 0 5 0
0.006 0.005 0.005 0.004 0.004 0.003 0.003 0.002 0.002
pc ∆ M n, io Correct Error n tio i Pos er b Num Mach
88
1.0 odel M y b Fly r e .9 w 0 o T odel 0.8 M e 10K Pac d e .7 ir 0 ic Fa e nd er, M b m t: Ha i 10K Pac Nu e F 0.6 rv ach u M C rror Correction (10,000 ft PA Pace)
cated di In .5 0 0.4 s 0132 ank T e s t ng i s ui Figure 42. Mach Number Position E e
16B .3 0 ght T i l 15B / S/N: 76- raft: F- - ation: Cr 0457 70GAL W : F rc i e gur r: F i c A o 3 e t w s e T S/N: 92- Conf T Date: 14 Apr 04 Sour Pac 0.2 8 7 6 5 4 3 2 1 0
0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00
pc ∆ M , ion rrect Co Error n sitio Po er mb Nu Mach
89
1.0 l e d o y M yb l F .9 r 0 e w d o e T ir Fa l 0.8 it: Hand F Mode e rve u C K Pac 20 .7 0 ic e er, M c mb Pa K Nu 20 ch 0.6 a M rror Correction (20,000 ft PA Pace) ed
cat i d In .5 0 0.4 s 132 -0 ank T e g s t B n i s ui 6 Figure 43. Mach Number Position E e 1 .3 - 0 F
04 : n: Cr r B / S/N: 76 AL W ft o p ight T l 457 15 ati - -0 : F rcra i ur F e : g r i c A o 370G e te: 14 A w st a onf e T T S/N: 92 C D Sour Pac 0.2 8 7 6 5 4 3 2 1 0
0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00
pc ∆ M , n io rrect Co r Erro n o i t i Pos er b m Nu ch Ma
90
0 1. l e d o y M yb l f r e .9 0 w o T ed r l e 8 . 0 Mod
Hand Fai e : t c e Fi K Pa rv 30 u C 7 . 0 ic M , e r K Pac 30 6 ch Numbe 0. a rror Correction (30,000 ft PA Pace)
ated M Indic .5 0 4 . 0 2 s 13 nk 0 - a T 76 e g t B n N: i s uis 6 / r Figure 44. Mach Number Position E e
1 3 S . C T 0 / 04 t F-
: n: h r B AL W t 57 g io i raf 0G at Ap Fl 4 rc i ur 2-04 e: 37 ig : 1 o f er: F-15 n te c urc w st A N: 9 o a / o e T T S C Da S P 2 0. 8 7 6 5 4 3 2 1 0 00 00 00 00 00 00 00 00 00
0. 0. 0. 0. 0. 0. 0. 0. 0.
pc ∆ M , ction e r Cor r o r r E sition o P mber Nu h Mac
91
0 1. l e d o M y yb l F r 9 e . 0 w o T ed r del 8 0. Mo e Pac Hand Fai K : t 35 7 . rve Fi 0 u ic C , M e r 35K Pac Numbe 6 h c 0. a rror Correction (35,000 ft PA Pace)
ted M ca i d n I .5 0 4 0. s 32 nk -01 a T e t B ng N: 76 i s uis 6 / Figure 45. Mach Number Position E e 1 .3 S W Cr 0 T F- / t L 04 r B t: 5 ion: GA ligh 457 1 Ap raf 0 at -
0 F - 7 : rc i ur e 14 ig : c 92 o 3 f : e n t cer: F ur w st A N o a a e T S/ C T D So P 2 0. 8 7 6 5 4 3 2 1 0 00 00 00 00 00 00 00 00 00
0. 0. 0. 0. 0. 0. 0. 0. 0.
pc ∆ M , n io rrect o C r ro r E n o i t i s o P er b m Nu Mach
92
0 1. l de Mo y b y r Fl 9 e . w 0 o T d l e r 8 0. Fai
Mode e Hand : K Pac t 0 4 7 . rve Fi 0 u ic C , M e r c K Pa 40 Numbe 6 h c 0. a rror Correction (40,000 ft PA Pace)
ted M ca i d n I .5 0 4 0. s 32 nk -01 a T e t B ng N: 76 i s uis 6 / Figure 46. Mach Number Position E e 1 .3 S W Cr 0 T F- / t L 04 r B t: 5 ion: GA ligh 457 1 Ap raf 0 at -
0 F - 7 : rc i ur e 14 ig : c 92 o 3 f : e n t cer: F ur w st A N o a a e T S/ C T D So P 2 0. 8 7 6 5 4 3 2 1 0 00 00 00 00 00 00 00 00 00
0. 0. 0. 0. 0. 0. 0. 0. 0.
pc ∆ M , n io rrect o C r Erro n o i sit Po er mb u N h Mac
93
1.0 .9 0 0.8 Model Model e e .7 0 ic er, M 20K Pac 40K Pac b Num 0.6 ach M r Correction (Tower Fly-By & Pace) ed
del del cat i Mo Mo e e Ind 5 . 0 10K Pac 35K Pac 0.4
s nk 0132 r 04 Model a
y e b Model s e t B ng T i s 30 Ap ui 6 Fly e 1 r .3 e 0 Figure 47. Mach Number Position Erro F- 04 - w : ht T 7 AL W o ft on: Cr g i T 30K Pac 15B / S/N: 76- -045 : Fli rcra i F- urat 7 Apr e : : g r i c A o 370G e w st e T S/N: 92 Conf T Dates Sour Pac 0.2
0.008 0.007 0.006 0.005 0.004 0.003 0.002 0.001 0.000
pc ∆ M n, Correctio r Erro on i Posit er b Num Mach
94
1.0 odel M y b Fly r e .9 w 0 o T 0.8 Model
e 10K Pac .7 0 ic er, M b overleaf Num 0.6 ach 10K Cl M r Correction (10,000 ft PA Cloverleaf) ed
cat i Ind 5 . 0 0.4 4 s nk a e s t 30 Apr 0 B ng T i s ui 6 e 1 .3 04 - 0 F- Figure 48. Mach Number Position Erro : ht T AL W ft on: Cr g i 0457 : Fli rcra i urat 16 Apr e : g i c A o 370G #: 92- l ne ARDS Pod w i st a e T T Conf T O Dates Sour 0.2
0.012 0.010 0.008 0.006 0.004 0.002 0.000
pc ∆ M , ion rrect Co Error n sitio Po er b Num h Mac
95
1.0 odel M y b Fly r e .9 w 0 o T odel 0.8 M e 20K Pac .7 0 ic er, M b Num 0.6 ach 20K Cloverleaf M r Correction (20,000 ft PA Cloverleaf) ed
cat i Ind 5 . 0 0.4 4 s nk a e s t 30 Apr 0 B ng T i s ui 6 e 1 .3 04 - 0 F- Figure 49. Mach Number Position Erro : ht T AL W ft on: Cr g i 0457 : Fli rcra i urat 16 Apr e : g i c A o 370G #: 92- l ne ARDS Pod w i st a e T T Conf T O Dates Sour 0.2
0.008 0.007 0.006 0.005 0.004 0.003 0.002 0.001 0.000
pc ∆ M n, Correctio r Erro on i Posit er mb Nu Mach
96
0 1. l de Mo y b y 9 er Fl . w 0 o T l de 8 0. Mo e c K Pa 30 7 . 0 ic f , M r a e erl ov Numbe Cl 6 h K c 0. a 30 r Correction (30,000 ft PA Cloverleaf)
ted M ca i d n I .5 0 4 0. 4 s k n pr 0 a A
T 0 e g B n i st uis - 3 6 e d 1 04 .3 Cr T 7 F- L W 0 Po r
Figure 50. Mach Number Position Erro ht p A t: S g ion: i 045 G A - raf at 6 2 Fl rc 9 i ur
ig ce: A o 370 e ARD f r t n n tes: 1 u w il #: s o o a e T T C T O Da S 2 0. 8 7 6 5 4 3 2 1 0 00 00 00 00 00 00 00 00 00
0. 0. 0. 0. 0. 0. 0. 0. 0.
pc ∆ M , n io rrect o C r ro r E n o i t i s o P er b m Nu Mach
97
1.0 Model y b y l F r e .9 w 0 o T 0.8 Model e 35K Pac .7 0 ic er, M eaf l b over Num 0.6 ach 35K Cl M r Correction (35,000 ft PA Cloverleaf) ed
cat i Ind 5 . 0 0.4 4 s nk a e s t 30 Apr 0 B ng T i s ui 6 e 1 .3 04 - 0 F- Figure 51. Mach Number Position Erro : ht T AL W ft on: Cr g i 0457 : Fli rcra i urat 16 Apr e : g i c A o 370G #: 92- l ne ARDS Pod w i st a e T T Conf T O Dates Sour 0.2
0.008 0.007 0.006 0.005 0.004 0.003 0.002 0.001 0.000
pc ∆ M n, Correctio r Erro on i Posit er b Num Mach
98
l e d o M 4 y 1. yb l F r e w o T l .2 1 de Mo
e c a K P 10 Level Accel/Decel) 1.0 ic l M , ce De / l cce A ch Number a .8 rrection (10,000 ft PA 0
ted M a Indic 0.6 s k n a T
e s .4 t B ng 0 i s e 04 T
F-16 Pod r n: Crui ht AL W t: 457 o lig Ap DS Figure 52. Mach Number Position Error Co raf 0G c F R : r 92-0 urati e 37 ig c Ai e A o t #: l n tes: 23 nf ur w i s a a e T T Co T O D So 0.2 0 5 0 5 0 5 0 5 2 1 1 0 0 0 1 1 0 0 0 0 0 0 0 0 . . .
0. 0. 0. 0. 0.
-0 -0 -0
pc ∆ M , n io rrect Co r rro E n o i sit o P er b m Nu ach M
99
l e d o M 4 y . 1 yb l F r e w o T .2 1 Model e 20K Pac 1.0 ic ft PA Level Accel/Decel) l ce De / l cce A ach Number, M .8 rrection (20,000 0
ted M a Indic 0.6 er Position Error Co s ank T e g .4 n 0 i st e d W o
T 04 7 t L P : F-16B n: Cruis h A t 45 g i Figure 53. Mach Numb 0 G - Fl RDS : 20 Apr rcraf 92 i uratio
: : ig ce A e A o 370 r n u w st il # o e a T T Conf T O Dates S 0.2 0 5 0 5 0 05 10 15 02 01 01 00 00 .0 .0 .0
0. 0. 0. 0. 0.
-0 -0 -0
pc ∆ M , ion ect Corr Error on i sit Po er mb Nu ch a M
100
l e d o 4 y M . 1 yb l F r e w o T 2 l . e Mod e c Pa 30K Level Accel/Decel) 01 . 1 ic l , M er ce e b D m
l Nu cce h A c k a 0 8 3 . rrection (30,000 ft PA
ted M Indica 60 . 0 s k n a T e 4 . t B ng i s uis
6 e 1 - W Cr 7 : L r 04 F 5 : n ht T p A 4 ft g io 0 a A - r 0G at 0 2 Figure 54. Mach Number Position Error Co c Fli : r i 2 ur 9 : e 37 : ig s A o e ARDS Pod t # l n nf urc w i s T O Te Ta Co Date So 20 . 0 0 5 0 5 0 5 0 5 2 1 1 0 0 0 1 1 0 0 0 0 0 0 0 0 . . .
0. 0. 0. 0. 0.
-0 -0 -0
pc ∆ M , n io rrect Co r rro E n o i sit Po er mb Nu Mach
101
Model
4 y . b 1 y l r F e w o T .2 1 Model
e c 35K Pa 1.0 ic l ce De
mber, M l Nu cce h A c k a 5 M 3 .8 0 rrection (35,000 ft PA Level Accel/Decel)
dicated In 0.6 s ank e .4 t ng T 0 i s e W T 04 Cruis 7 : r Pod : F-16B AL t ion t raf Figure 55. Mach Number Position Error Co 70G : Flight rc 20 Ap : e igura o 3 w st Ai il #: 92-045 e a T One ARDS T T Conf Dates Sourc 0.2 0 5 0 5 0 2 1 1 0 0 05 10 15
0.0 0.0 0.0 0.0 0.0
-0.0 -0.0 -0.0
pc ∆ M on, i Correct Error n o i t si Po er b Num Mach
102
l e d o y M 1.4 yb l F r e w o T l .2 e 1 Mod e c a 40K P 1.0 ic l er, M Dece mb Nu
ccel A k ach 0 4 M .8 rrection (40,000 ft PA Level Accel/Decel) 0
ed dicat In 0.6 s k n a T e .4 t B 0 ng i s uis
e 04 7
r Pod : F-16 ht T p AL W t f g i Figure 56. Mach Number Position Error Co 045 l ra ation: Cr : F rc 92- 30 A e : igur c o 370G f ne ARDS w st Ai il #: e a T O T T Con Dates Sour 0.2 .005 .010 .015
0.020 0.015 0.010 0.005 0.000 0 0 0
- - -
pc ∆ M , n io rrect Co r rro E n o i sit Po er b m Nu ch Ma
103
0 1. 5g 3. 2g / / K K 0 0 1 1 9 . 0 urns urns T T G G t t n n a a t t s s 8 Con Con 0. nt Airspeed Turn) 5g 7 . 1. 3g / / 0 K K ic 0 0 1 1 l
, M r s s de urn urn Mo T T y b y Numbe 6 h nt G nt G c 0. a er Fl nsta nsta w o ection (10,000 ft Consta Co Co T
ted M ca i d n I .5 0 5g 7g . . 2 3 / / K K 0 0 1 1 s s l e 4 urn urn 0. T T
s Mod G G e nk nt nt a T nsta nsta se K Pac t B ng 0 i s
6 1 Co Co e 4 1 W - T Crui
7 .3 F : Pod
0 : n pr 0 AL t 45 f o i light A DS ra 3 2-0 F 70G rc i 2 urat Figure 57. Mach Number Position Error Corr 9 : e: AR 3 g i s c A e o f r #: l n n te w st i e a T O T T Co Da Sou 2 0. 8 7 6 5 4 3 2 1 0 00 00 00 00 00 00 00 00 00
0. 0. 0. 0. 0. 0. 0. 0. 0.
pc ∆ M , n o i t rrec Co r Erro n o i sit Po er b m u N h Mac
104
12 s ank l e d c i t ng T o i st a uise st e W M T P 04 F-16B e - : n: Cr t t n f e 0457 - ra ambi rc 92 i 10 K Pac iguratio ce: Flight A o 370GAL 0 =P p ur 1 w st il #: e a P T T T Conf Dates: 23 Apr So ∆ ch a 9 M 8 0. ch a M (deg) 0.8 ck, tta 6 A h of efficient Angle of Attack Effects c a
ngle M A re Co 0.7 4 h ror Pressu c r a 0.6 M ch a 2 M Figure 58. Position E 0.5 0 0 5 0 5 0 5 0 0 0 1 1 2 2 3 0 0 0 0 0 0 ......
0.0
-0 -0 -0 -0 -0 -0
c ci p ∆ /q P , t en ci i f f e o C re u s es Pr r ro Er n o i t si Po
105
0 1 s ank e ch ng T i st uis e 16B Ma - 04 ht T 0.9 t: F 457 g i Apr -0 raf 0GAL W ation: Cr c : Fl r 23 e : igur o 37 #: 92 l ne ARDS Pod urc w i st Ai 89 h o a e T T Conf T O Dates S 0.8 Mac h c 67 0.7 Ma ) gle of Attack Effects eg (d ck a t t ch A
rrection An o of e 0.6 Ma ngl rror C A 45 ition E h Pos e 0.5 Mac 23 Figure 59. Altitud 01 5 0 5 0 5 0
75 50
22 20 17 15 12 10
∆ ) (feet Hpc , ion rrect o C r Erro on i t i s o P e d u t i t l A
106
0 1 s k n a e ch ng T i st uis e 16B - 04 F r ht T p AL W t: 0.9 Ma 457 g i l -0 raf 0G ation: Cr 2 : F rc i 23 A e : igur o 37 #: 9 l ne ARDS Pod urc w i st A h o a e T T Conf T O Dates S fects f 0.8 Mac 789 h 6 ) 0.7 Mac eg (d ck a t t h A
Correction Angle of Attack E of e 0.6 Mac ngl A h c 0.5 Ma Figure 60. Mach Number Position Error 012345 0 5 0 5 0 5 0 5 0
0.008 0.007 0.007 0.006 0.006 0.005 0.005 0.004 0.004
∆ Mpc , n io Correct Error n o i sit Po r e b m Nu Mach
107
0.95 ne i st Fit L e 0.85 B e c 0.75 Pa B F 0.9818x - 0.0416 r T 65 . p = 0 2 A y c ower Fly-By & Pace) 30 M , red a B u F q T r 0.55 p er S b m 13 A ecovery Factor (T ch Nu B a TF obe R 0.45
r r d M p e A rat 12 Calib B F 0.35 T r p 7 A 0.25 s nk 0132 r 04 a e s t B ng T i s 30 Ap ui 6 e Figure 61. Total Air Temperature P 1 / S/N: 76- T 15 t . F- 04 - : h 0 AL W ft g 15B 0457 : Fli rcra i F- uration: Cr 7 Apr e : g r: i c A o 370G e w st e T S/N: 92- Conf T Dates Sour Pac 0.05
1.0 0.9 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0.0
a c ti ) -1 /T T 5(
108
5 0 . 1 95 0. ) 85 . Decel 0 c i 0402 n o s b u (S 5 7 ear . n 0 i L 9811x - 0. 2 c M = 0. cel y d, 65 re De 0. c i n qua o s S b r Su 5 5 . ry Factor (10,000 ft PA Level Accel/Decel) 0 Numbe h
c ecel a D M ccel/ ed 5 t 4 obe Recove K A 0. r bra 0 li 1 a C 35 0. 5 2 s . 0 nk a se t B ng T i s
6 e 1 W T
04 F- r
Pod 15 : n: Crui ht 7 AL t f o S Figure 62. Total Air Temperature P 0. i Ap t lig
ra 0G 045 F - : rc i ura 23 2 : e g i s c A 9 o 37 f : e ne ARD t w st e T O T S/N Con Da Sour 05 0. 2 0 8 6 4 2 0
1. 1. 0. 0. 0. 0. 0.
a tic ) -1 /T 5(T
109
5 5 . 1 5 3 1. ) Decel c i n o s b u 5 (S 1 1. ear n i L 2 c M cel d, re De c i 95 n qua o 0. s S b r Su ry Factor (20,000 ft PA Level Accel/Decel) Numbe h
c ecel a D M 0238 75 ccel/ 0. ed t obe Recove K A r bra 0 li 2 a C 9457x - 0. 55 = 0. 0. y s nk a se t 35 B ng T i s
6 0. e 1 W T
04 F- r
Pod : n: Crui ht 7 AL t f o S Figure 63. Total Air Temperature P i Ap t lig
ra 0G 045 F - : rc i ura 20 2 : e g i s c A 9 o 37 f : e ne ARD t w st e T O T S/N Con Da Sour 15 0. 4 2 0 8 6 4 2 0
1. 1. 1. 0. 0. 0. 0. 0.
a tic -1) /T T 5(
110
5 1.8 1.65 ) Decel c i n o s 45 . b 1 u ear (S n i L 2 c l M 1.25 d, e Dece r
c i qua on s S b r Su mbe 05 u . ry Factor (30,000 ft PA Level Accel/Decel) 1 N
cel ch a De l/ M
e c d c e t a obe Recove K A r br 0 li 3 a 0.85 C 0.65 s ank e = 0.9998x - 0.0284 t B y ng T i s uis 5 6 e d 1 o T 04 0.4 F- L W r : n: Cr p A t f Figure 64. Total Air Temperature P 57 G light ra atio 04 : F rc 20 A : e igur c o 370 f ne ARDS P ur w st Ai N: 92- / o e T O T S Con Dates S 0.25
1.6 1.4 1.2 1.0 0.8 0.6 0.4 0.2 0.0
a tic ) -1 /T T 5(
111
5 0 . 2 5 8 1. ) Decel c i n o s 65 b 1. u (S ear n i L 2 c M 45 cel 1. d, re De c i n qua o s S b r Su 5 2 . ry Factor (35,000 ft PA Level Accel/Decel) 1 Numbe h
c ecel a D M ccel/ ed t obe Recove K A r 5 bra 5 0 li 3 . a 1 C 85 0. 0638 s nk a se t B ng T i s
6 e 65 1 W T
0644x - 0. 0.
04 F- r
Pod : n: Crui ht 7 AL t f o S Figure 65. Total Air Temperature P i Ap t lig
ra 0G 045 = 1. F - : rc i ura 20 y 2 : e g i s c A 9 o 37 f : e ne ARD t w st e T O T S/N Con Da Sour 45 0. 8 6 4 2 0 8 6 4
1. 1. 1. 1. 1. 0. 0. 0.
a tic ) -1 /T 5(T
112
5 8 . 1 5 6 1. ) Decel c i n o s b u 5 (S 4 1. ear n i L 2 c M cel d, re De c i 25 n qua o 1. s S b r Su ry Factor (40,000 ft PA Level Accel/Decel) Numbe h
c ecel a D M 05 ccel/ 1. ed t obe Recove K A r bra 0 li 4 a C 85 0. 078 s nk a se t 65 B ng T i s
6 0. e 8602x + 0. 1 W T
04 F- r
Pod : n: Crui ht 7 AL t f o S Figure 66. Total Air Temperature P i Ap t lig
= 0. ra 0G 045 F y - : rc i ura 30 2 : e g i s c A 9 o 37 f : e ne ARD t w st e T O T S/N Con Da Sour 45 0. 6 4 2 0 8 6 4 2
1. 1. 1. 1. 0. 0. 0. 0.
a tic ) -1 /T T 5(
113
APPENDIX B
LESSONS LEARNED
114
Human Factors
F-16B S/N 92 0457 had several special controls and displays installed in the rear cockpit (RCP) for interfacing with the pacer instrumentation and data acquisition system.
The controls included an event mark button, a display configuration selector, and DAS recorder controls on the left console, aft of the time display, Figure 67. Both the MARS
II tape and the PCMCIA DAS recorders had power and record/stop toggle switches.
Additionally, the PCMCIA recorder had high and low data rate recording options. All controls were clearly labeled. The event button was in an awkward location that induced operator arm fatigue and interfered with throttle operation. It was easy to bump the throttle during low power settings when operating the event button or any other pacer system controls. Due to this, it was best to hold one’s finger on the event button throughout all maneuvers. This awkward position led to muscle fatigue in the left arm.
Place an event button on the RCP instrument panel in an easily accessible location. (R8)
Time Display
Figure 67. Rear Cockpit Pacer Controls And Time Display
115
The displays included GPS time on the left console, Figure 67, and a multi- window LCD display of altitude, airspeed, and event correlation number on the left side of the instrument panel, Figure 68. Recording handheld data required looking down to read the time, and then looking up to read the other displays. This required significant
FTE workload during elevated load factor test points. Additionally, at near idle power settings, the throttle blocked viewing the time display. Move the time display so that it is not blocked from view during flight and is within view of the altitude, airspeed, and correlation number displays. (R9) The multi-display was not externally labeled, but once power was applied “Hc”, “Vc”, and “CN” appeared on the LCDs. However, adding external labels would improve usability. Add labels to the altitude, airspeed, and correlation number displays. (R10)
Figure 68. Rear Cockpit Data Display
An event was actuated by holding the event button down for approximately 1 second. This was an excessive time delay between the event button actuation and the actual event being recorded in the DAS. Remove the time delay so that the computer
116
events the data immediately upon event button actuation. (R11) After event actuation, the altitude, airspeed, and correlation number displays froze for approximately 10 seconds to allow for hand recording of the data, which was a good feature of the system.
Overall, the human factors design of the pacer system was MARGINAL, and should be corrected prior to use as a developmental test pacer aircraft.
117
APPENDIX C
INSTRUMENTATION LIST
118
Flyby Tower Instrumentation Item Part Number Serial Number Setra Digital Pressure Gauge 370 2017712 Omega Digital Thermometer HH-41 305107
Aircraft Instrumentation Item Part Serial Description Number Number TTU-205F 1256 Dual Sonic Encoder PS7000 0011 Aircraft System 1 pressure transducer Dual Sonic Encoder PS7000 0014 Aircraft System 2 pressure transducer
119
APPENDIX D
END-TO-END GROUND TEST RESULTS
120
Table 5. End-to-End Ground Test Results 19 Feb 04 Calibration 28 Apr 04 Spot Check Test Set Readings Pacer Display Readings Pacer Display Readings Pressure Pressure Airspeed Altitude Calibrated Altitude Calibrated
Hp (ft) (KCAS) (hc) Airspeed (Vc) (hc) Airspeed (Vc) 5000 150 5008.3 150.34 5000 180 5028.7 181.31 5000 250 5062.6 252.13 5000 300 5087.2 302.41 5089.6 302.53 5000 350 5114.2 352.67 5000 400 5139.5 402.7 5141.3 402.73 5000 450 5158.2 452.57 5000 500 5161.8 502.28 5162.8 502.4 5000 550 5176.7 552.15 5000 650 4710.6 647.21 5000 750 4489 749.38 20000 180 20048 181.66 20000 250 20089 252.08 20093 252.22 20000 300 20119 302.23 20122 302.49 20000 350 20137 352.11 20141 352.23 20000 375 20140 376.96 20000 400 20143 401.82 20148 401.9 20000 425 20155 426.82 20159 426.78 20000 450 20505 455.3 20508 455.27 20000 550 19367 546.12 20000 650 19617 659.69 40000 180 40079 181.75 40000 200 40094 201.82 40102 202.11 40000 225 40106 226.7 40000 250 40107 251.53 40118 251.98 40000 275 40118 276.45 40000 300 40681 305.89 40702 306.23 40000 350 39335 345.87 40000 400 39591 401.12 35000 250 35111 251.83 30000 250 30105 252.03 25000 250 25094 252.05 15000 250 15076 252.23 10000 250 10067 252.33
121
APPENDIX E
RAW DATA
122
Table 6. Tower Fly-by Aircraft Raw Data
Grid Angle Reading Psic Ptic Ttic of Date Time (L) Run (N/D) (in Hg) (in Hg) (deg. K) Attack 4/7/2004 07:37:35 300 2.8 27.575163 31.996853 293.00728 4.8450 4/7/2004 07:42:10 325 3.5 27.569563 32.981541 295.86416 4.2078 4/7/2004 07:46:12 350 3.3 27.577963 33.410034 297.10646 3.7244 4/7/2004 07:50:00 375 3.1 27.596659 34.980003 301.20682 2.9443 4/7/2004 07:53:37 400 3.0 27.613056 35.584435 302.94674 2.6917 4/7/2004 07:57:12 425 2.2 27.64695 36.85321 305.55704 2.4170 4/7/2004 08:00:42 450 2.6 27.643753 38.295898 308.16787 1.9885 4/7/2004 08:04:43 500 3.0 27.629055 41.58176 315.38157 1.5491 4/7/2004 08:08:32 525 3.1 27.627754 42.096355 316.62575 1.3513 4/7/2004 08:12:28 550 2.5 27.636053 45.050892 322.59969 1.1865 4/7/2004 08:16:15 575 2.0 27.655451 45.575439 323.96917 1.0767 4/7/2004 08:20:47 275 2.5 27.581263 31.095503 293.25568 5.1746 4/7/2004 08:25:36 250 2.1 27.581861 30.520515 291.88951 6.2512 4/7/2004 08:30:12 200 2.9 27.538868 29.410025 288.90922 9.4263 4/7/2004 08:34:54 180 3.4 27.515871 29.01055 288.2884 11.6455 4/7/2004 08:39:39 190 3.5 27.524269 29.182552 288.78505 10.7007 4/7/2004 08:44:23 225 2.7 27.565865 30.160276 291.76532 6.6687 4/7/2004 08:48:36 325 2.6 27.604158 33.009323 299.09436 3.1201 4/7/2004 08:52:43 475 2.3 27.65765 39.811287 314.26192 1.4612 4/12/2004 07:34:13 300 1.4 27.679146 32.178722 293.37988 5.2515 4/12/2004 07:38:13 325 1.3 27.689445 33.021812 295.61571 3.8232 4/12/2004 07:41:59 350 1.6 27.688944 33.870399 298.10037 3.4607 4/12/2004 07:45:31 375 1.2 27.707144 34.850994 300.58547 2.9004 4/12/2004 07:48:55 400 1.9 27.702244 35.887852 303.07102 2.5488 4/12/2004 07:52:19 425 1.7 27.709341 37.185028 305.18411 2.0984 4/12/2004 07:55:32 450 2.2 27.703342 38.238396 308.66523 1.7908 4/12/2004 07:58:40 475 2.0 27.714941 39.780087 311.64981 1.5381 4/12/2004 08:01:45 500 2.2 27.712641 41.194214 314.13752 1.4392 4/12/2004 08:04:49 525 2.2 27.703043 42.821766 317.37232 1.2195 4/12/2004 08:08:02 550 2.1 27.705442 44.511684 320.60803 1.0217 4/12/2004 08:11:13 575 2.1 27.704742 46.616936 324.71623 0.9558 4/12/2004 08:15:09 275 1.4 27.668749 31.423567 292.63468 4.6912 4/12/2004 08:19:11 250 2.1 27.641453 30.76544 290.89602 5.6250 4/12/2004 08:23:19 225 1.7 27.640953 30.139891 289.40589 6.9214 4/12/2004 08:27:41 200 2.3 27.620655 29.574333 288.41256 8.3167 4/12/2004 08:32:30 190 1.9 27.623653 29.370947 287.79175 9.3494 4/12/2004 08:37:14 180 1.9 27.621155 29.197344 287.41928 9.7229 4/12/2004 08:41:25 300 1.4 27.686844 32.206799 295.61571 3.5046 4/12/2004 08:45:17 180 1.7 27.631155 29.188948 287.79175 9.5911
123
Grid Angle Reading Psic Ptic Ttic of Date Time (L) Run (N/D) (in Hg) (in Hg) (deg. K) Attack 4/12/2004 08:51:24 170 2.2 27.612757 29.049625 287.79175 10.0745 4/13/2004 07:31:12 300 2.2 27.625154 32.299232 297.10646 4.7131 4/13/2004 07:35:08 325 2.3 27.634052 33.175003 299.46712 3.8013 4/13/2004 07:38:46 350 2.2 27.653151 34.518436 303.19531 3.0872 4/13/2004 07:42:20 375 2.4 27.65255 35.093204 304.43827 2.7356 4/13/2004 07:45:42 400 2.1 27.671947 36.524193 307.29753 2.4390 4/13/2004 07:48:54 425 2 27.680647 37.217834 309.16261 2.0874 4/13/2004 07:52:21 450 2.5 27.673647 37.976879 310.9036 2.0105 4/13/2004 07:55:35 475 2.4 27.676947 39.497868 314.38632 1.6699 4/13/2004 07:58:54 500 2.6 27.675947 41.346153 317.9945 1.3403 4/13/2004 08:02:26 525 2.2 27.682447 42.93552 321.47934 1.2085 4/13/2004 08:05:56 550 2.4 27.671646 44.765488 325.71239 1.0547 4/13/2004 08:09:25 575 1.9 27.685045 46.848412 330.07169 1.0327 4/13/2004 08:13:24 275 2.1 27.621355 31.391788 296.23684 4.5813 4/13/2004 08:17:30 250 1.8 27.618256 30.689497 294.49775 5.6030 4/13/2004 08:21:41 225 2.5 27.59466 30.078535 293.00728 6.9763 4/13/2004 08:26:18 200 2.2 27.596258 29.575432 291.76532 8.8000 4/13/2004 08:31:13 180 2.2 27.585461 29.210436 290.89602 9.9536 4/13/2004 08:36:04 190 2.5 27.58466 29.368347 291.51695 9.2065 4/13/2004 08:40:17 300 2.5 27.627554 32.130756 299.09436 3.5266 4/13/2004 08:44:14 325 2.2 27.644953 33.071976 301.82819 2.9004 4/13/2004 08:47:12 275 2.1 27.625854 31.430861 297.35493 4.1748 4/13/2004 08:53:07 170 3 27.558165 29.023642 291.14439 10.7336 4/13/2004 08:56:38 450 2.4 27.683546 38.556812 314.63513 1.2195 4/30/2004 07:22:18 300 2 27.707144 32.466911 295.11883 4.4824 4/30/2004 07:25:44 500 3 27.725039 41.414524 315.25716 1.4612 4/30/2004 07:30:04 200 2 27.669149 29.748936 288.04007 9.6460
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Table 7. Fly-By Tower Raw Data (7 April 2004) Zero Grid Alt Correction (ft)1 Time Setra (in Hg) Setra (ft) (ft) Ta (°F) 06:59:00 27.5801 2237.2 -3 2234.2 46.11 07:04:00 27.5819 2235.5 -3 2232.5 46.27 07:09:00 27.5819 2235.5 -3 2232.5 46.55 07:14:00 27.5845 2232.9 -3 2229.9 46.89 07:19:00 27.5828 2234.6 -3 2231.6 47.35 07:24:00 27.586 2231.4 -3 2228.4 47.99 07:29:00 27.5872 2230.2 -3 2227.2 48.52 07:34:00 27.5878 2229.6 -3 2226.6 48.85 07:39:00 27.5872 2230.2 -3 2227.2 49.31 07:44:00 27.5878 2229.6 -3 2226.6 49.83 07:49:00 27.5887 2228.7 -3 2225.7 50.27 07:54:00 27.591 2226.5 -3 2223.5 50.71 07:59:00 27.5913 2226.2 -3 2223.2 51.28 08:04:00 27.5922 2225.3 -3 2222.3 51.88 08:09:00 27.5928 2224.7 -3 2221.7 52.32 08:14:00 27.594 2223.5 -3 2220.5 52.99 08:19:00 27.5952 2222.3 -3 2219.3 54.04 08:24:00 27.5949 2222.6 -3 2219.6 54.41 08:29:00 27.5952 2222.3 -3 2219.3 55.20 08:34:00 27.5963 2221.2 -3 2218.2 55.85 08:39:00 27.5969 2220.6 -3 2217.6 56.26 08:44:00 27.596 2221.5 -3 2218.5 56.67 08:49:00 27.596 2221.5 -3 2218.5 57.31 08:54:00 27.5946 2222.9 -3 2219.9 58.22 08:59:00 27.5931 2224.4 -3 2221.4 59.03 09:04:00 27.5922 2225.3 -3 2222.3 60.20 09:09:00 27.5919 2225.6 -3 2222.6 60.36 09:14:00 27.5934 2224.1 -3 2221.1 60.96 09:19:00 27.5931 2224.4 -3 2221.4 62.00 09:24:00 27.5925 2225.0 -3 2222.0 62.19 1 The Setra pressure transducer was recalibrated after the test flights and found to have a 3 ft error.
125
Table 8. Fly-By Tower Raw Data (12 April 2004) Correction (ft)1 Time Setra (in Hg) Setra (ft) Zero Grid Alt (ft) Ta (°F) 07:00:30 27.635 2183.0 -3 2180.0 47.27 07:05:30 27.6356 2182.5 -3 2179.5 47.00 07:10:30 27.6374 2180.7 -3 2177.7 46.77 07:15:30 27.6383 2179.8 -3 2176.8 46.95 07:20:30 27.64 2178.1 -3 2175.1 47.40 07:25:30 27.6397 2178.4 -3 2175.4 48.07 07:30:30 27.6412 2176.9 -3 2173.9 48.58 07:35:30 27.6433 2174.9 -3 2171.9 48.98 07:40:30 27.6445 2173.7 -3 2170.7 49.92 07:45:30 27.6442 2174.0 -3 2171.0 50.27 07:50:30 27.6454 2172.8 -3 2169.8 50.13 07:55:30 27.646 2172.2 -3 2169.2 50.27 08:00:30 27.6468 2171.4 -3 2168.4 50.48 08:05:30 27.6495 2168.7 -3 2165.7 51.13 08:10:30 27.6506 2167.7 -3 2164.7 51.34 08:15:30 27.6504 2167.9 -3 2164.9 51.59 08:20:30 27.6504 2167.9 -3 2164.9 52.19 08:25:30 27.6516 2166.7 -3 2163.7 52.66 08:30:30 27.6522 2166.1 -3 2163.1 53.56 08:35:30 27.6533 2165.0 -3 2162.0 54.07 08:40:30 27.6545 2163.8 -3 2160.8 54.34 08:45:30 27.6539 2164.4 -3 2161.4 54.97 08:50:30 27.6539 2164.4 -3 2161.4 55.01 08:55:30 27.6545 2163.8 -3 2160.8 55.57 09:00:30 27.6533 2165.0 -3 2162.0 56.38 09:05:30 27.6533 2165.0 -3 2162.0 57.10 09:10:30 27.6533 2165.0 -3 2162.0 57.34 09:15:30 27.6542 2164.1 -3 2161.1 59.34 09:20:30 27.656 2162.3 -3 2159.3 59.95 1 The Setra pressure transducer was recalibrated after the test flights and found to have a 3 ft error.
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Table 9. Fly-By Tower Raw Data (13 April 2004) Correction (ft)1 Time Setra (in Hg) Setra (ft) Zero Grid Alt (ft) Ta (°F) 06:55:00 27.6105 2207.2 -3 2204.2 51.57 07:00:00 27.6102 2207.5 -3 2204.5 52.56 07:05:00 27.6105 2207.2 -3 2204.2 51.90 07:10:00 27.6126 2205.1 -3 2202.1 51.62 07:15:00 27.6126 2205.1 -3 2202.1 52.48 07:20:00 27.6126 2205.1 -3 2202.1 53.37 07:25:00 27.6144 2203.4 -3 2200.4 53.07 07:30:00 27.6161 2201.7 -3 2198.7 53.08 07:35:10 27.6188 2199.0 -3 2196.0 53.78 07:40:00 27.6214 2196.5 -3 2193.5 54.46 07:45:00 27.6235 2194.4 -3 2191.4 55.23 07:50:00 27.6235 2194.4 -3 2191.4 55.94 07:55:00 27.6235 2194.4 -3 2191.4 56.68 08:00:00 27.6232 2194.7 -3 2191.7 57.41 08:05:00 27.6214 2196.5 -3 2193.5 57.88 08:10:00 27.6217 2196.2 -3 2193.2 58.50 08:15:10 27.6211 2196.8 -3 2193.8 58.53 08:20:00 27.6214 2196.5 -3 2193.5 58.73 08:25:00 27.6223 2195.6 -3 2192.6 58.97 08:30:00 27.6235 2194.4 -3 2191.4 59.41 08:35:00 27.6241 2193.8 -3 2190.8 59.06 08:40:00 27.6241 2193.8 -3 2190.8 59.52 08:45:00 27.6226 2195.3 -3 2192.3 60.80 08:50:00 27.6229 2195.0 -3 2192.0 60.69 08:55:00 27.6235 2194.4 -3 2191.4 61.71 09:00:00 27.6232 2194.7 -3 2191.7 62.22 09:05:00 27.6229 2195.0 -3 2192.0 62.21 09:10:00 27.6217 2196.2 -3 2193.2 62.75 09:15:00 27.6203 2197.6 -3 2194.6 63.25 09:20:00 27.6206 2197.3 -3 2194.3 63.83 09:25:00 27.6214 2196.5 -3 2193.5 63.56 09:30:00 27.6223 2195.6 -3 2192.6 64.40 1 The Setra pressure transducer was recalibrated after the test flights and found to have a 3 ft error.
127
Table 10. Fly-By Tower Raw Data (30 April 2004) Correction (ft)1 Time Setra (in Hg) Setra (ft) Zero Grid Alt (ft) Ta (°F) 06:30:00 27.6713 2147.3 -3 2144.3 49.00 06:35:00 27.6734 2145.2 -3 2142.2 49.92 06:40:00 27.6752 2143.4 -3 2140.4 50.59 06:46:00 27.6773 2141.4 -3 2138.4 50.70 06:50:00 27.6796 2139.1 -3 2136.1 50.76 06:55:00 27.6835 2135.2 -3 2132.2 50.79 07:00:00 27.6846 2134.2 -3 2131.2 51.20 07:05:00 27.687 2131.8 -3 2128.8 51.61 07:10:00 27.6882 2130.6 -3 2127.6 51.18 07:15:00 27.6908 2128.1 -3 2125.1 51.23 07:20:00 27.6923 2126.6 -3 2123.6 50.27 07:25:00 27.6944 2124.5 -3 2121.5 47.81 07:30:00 27.6967 2122.2 -3 2119.2 48.51 07:35:00 27.697 2122.0 -3 2119.0 49.54 07:40:00 27.697 2122.0 -3 2119.0 50.93 07:45:00 27.6994 2119.6 -3 2116.6 52.35 07:50:00 27.7029 2116.1 -3 2113.1 52.82 07:55:00 27.7044 2114.7 -3 2111.7 52.94 08:00:00 27.7041 2115.0 -3 2112.0 53.78 1 The Setra pressure transducer was recalibrated after the test flights and found to have a 3 ft error.
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Table 11. Pace Raw Data F-15 F-16 Angle Alt Hc Vc Ta Mc Psic Ptic Ttic of Time (Deg (Z) (ft) Run (feet) (KCAS) K) (N/D) (in Hg) (in Hg) (deg. K) Attack 22:15:10 10K 0.35 9957.8 198.67 276.2 0.36 20.6455 22.543543 282.57794 9.536 22:14:31 10K 0.4 9975 220.59 276.67 0.40 20.6327 23.031096 284.93637 7.614 22:13:36 10K 0.45 10002 250.73 276.77 0.45 20.6217 23.638834 287.04682 6.042 21:04:37 10K 0.5 9918.2 276.34 275.61 0.50 20.7281 24.425657 288.41256 5.944 21:05:34 10K 0.55 9938.8 305.66 275.87 0.55 20.6991 25.349298 291.64113 4.427 22:16:09 10K 300 9941.2 301.28 276.43 0.54 20.6786 25.214775 291.64113 3.966 21:06:33 10K 0.6 9956.5 333.49 276.21 0.60 20.6900 26.194126 294.8704 3.834 21:07:42 10K 0.65 9940.3 361.53 276.76 0.65 20.704 27.279617 298.47311 3.076 21:08:30 10K 0.7 9925.9 389.77 276.55 0.70 20.7242 28.608173 301.95247 2.505 21:09:21 10K 0.75 9890.9 421.84 277.97 0.75 20.7383 29.942726 304.81118 2.109 21:11:43 10K 0.8 9902.4 449.29 278.67 0.80 20.7205 31.54328 309.03826 1.703 21:12:25 10K 0.85 9896.8 474.94 278.14 0.84 20.7208 32.995731 312.14731 1.505 21:13:34 10K 0.9 9798.6 505.62 277.33 0.89 20.7954 34.693611 316.50133 1.176 21:14:16 10K 0.92 9761.2 518.31 276.85 0.91 20.8239 35.639538 318.74116 1.351 22:19:00 10K 0.93 9776.1 523.12 276.5 0.92 20.8003 36.803307 319.86125 1.373 22:18:15 10K 0.94 9881.3 528.71 276.31 0.93 20.8481 36.217873 320.11017 1.198 22:18:40 10K 0.95 9843.2 533.99 276.35 0.94 20.7640 36.506687 320.3591 1.252 22:10:27 20K 0.45 20018 207.32 253.71 0.46 13.7824 15.809406 263.72133 8.405 21:17:13 20K 0.5 19998 230.56 254.05 0.51 13.7889 16.397293 266.69753 7.614 21:17:58 20K 0.55 19967 251.76 254.7 0.55 13.8045 16.920404 269.5501 6.361 21:18:48 20K 0.6 19955 276.94 254.24 0.60 13.8086 17.596163 271.78282 4.867 21:19:33 20K 0.65 19941 300.5 254.18 0.65 13.8224 18.29871 274.63609 4.384 22:08:54 20K 300 20054 300.44 253.87 0.65 13.7771 18.235432 274.2639 3.790 21:20:13 20K 0.7 19934 322.33 254.04 0.70 13.8244 19.10232 277.4898 3.680 21:22:09 20K 0.75 19990 347.77 253.72 0.75 13.8126 19.882139 280.84036 3.054 21:22:53 20K 0.8 19952 371.31 254.75 0.80 13.8306 20.846212 283.9433 2.428 21:23:33 20K 0.85 19907 398.46 254.93 0.85 13.8603 22.086189 287.91591 2.065 21:24:17 20K 0.9 19860 423.4 254.42 0.90 13.8777 23.27507 291.76532 1.780 21:24:51 20K 0.92 19806 433.12 254.55 0.92 13.9076 23.750824 293.37988 1.670 21:25:22 20K 0.93 19778 439.14 254.75 0.93 13.9350 24.133986 294.8704 1.780 21:25:47 20K 0.94 19712 443.97 254.95 0.93 13.9777 24.418356 295.73993 1.802 21:26:16 20K 0.95 19694 449.39 254.76 0.94 13.9913 24.723026 296.60952 1.703 21:28:58 30K 0.55 29871 205.12 228.96 0.55 8.97909 11.229238 243.88873 8.306 21:29:47 30K 0.6 29893 223.92 228.79 0.60 8.96878 11.468969 244.88005 7.646 21:32:24 30K 0.65 29929 243.59 228.82 0.65 8.95408 11.921652 247.60634 6.526 21:33:24 30K 0.7 29961 261.5 228.59 0.70 8.9336 12.23923 249.21743 5.482 21:34:10 30K 0.75 29957 284.44 228.85 0.75 8.93847 12.986306 253.5554 4.395 22:06:39 30K 300 29932 302.15 228.37 0.79 8.94647 13.483658 255.66265 3.483 21:34:48 30K 0.8 29950 303.56 228.68 0.80 8.94037 13.551411 256.90228 3.823 21:35:30 30K 0.85 29920 325.55 228.56 0.85 8.94287 14.340459 260.1256 3.153
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F-15 F-16 Angle Alt Hc Vc Ta Mc Psic Ptic Ttic of Time (Deg (Z) (ft) Run (feet) (KCAS) K) (N/D) (in Hg) (in Hg) (deg. K) Attack 21:36:07 30K 0.9 29905 346.57 228.57 0.90 8.94737 15.034486 263.10134 2.747 21:36:30 30K 0.92 29869 354.8 228.77 0.92 8.95858 15.428144 264.46534 2.505 21:36:51 30K 0.93 29839 359.73 228.85 0.93 8.97249 15.655462 265.95344 2.439 21:37:15 30K 0.94 29811 364.05 228.93 0.94 8.99350 15.728935 266.32548 2.560 21:37:36 30K 0.95 29771 367.7 229.03 0.95 9.01241 15.939159 267.31762 2.461 21:40:33 35K 0.6 34951 200.94 215.8 0.60 7.09331 9.04177 230.01258 9.646 21:42:15 35K 0.65 34920 218.31 216.12 0.65 7.09951 9.4379921 232.73795 7.888 21:45:36 35K 0.7 34925 234.11 215.74 0.70 7.10301 9.8533001 235.33956 6.570 21:46:39 35K 0.75 34986 252.39 215.42 0.75 7.07510 10.279103 238.31299 5.153 21:47:28 35K 0.8 35023 271 214.83 0.80 7.05810 10.695312 240.54319 4.812 21:48:17 35K 0.85 35037 291.72 214.39 0.85 7.04790 11.365742 244.26047 3.823 22:04:48 35K 300 34956 298.19 215.7 0.87 7.07560 11.533125 246.11925 3.604 21:48:58 35K 0.9 35005 311.38 214.6 0.90 7.04920 11.990204 247.60634 3.043 21:49:28 35K 0.92 34974 317.81 214.73 0.92 7.06340 12.144096 248.47384 3.065 21:49:48 35K 0.93 34933 322.65 214.92 0.93 7.0758 12.348654 249.83711 2.966 21:50:08 35K 0.94 34900 326.45 215.16 0.94 7.09221 12.507942 250.82861 2.999 21:50:26 35K 0.95 34867 330.23 215.07 0.95 7.11491 12.611769 251.07649 3.219 21:55:04 40K 0.7 39944 214.85 209.49 0.71 5.56835 7.5624061 227.28733 9.020 21:55:51 40K 0.75 39854 228.21 209.97 0.75 5.60226 8.1454973 231.87077 6.658 21:56:51 40K 0.8 39818 244.99 209.42 0.80 5.61036 8.5471163 233.97679 5.625 21:57:49 40K 0.85 39873 261.88 208.66 0.85 5.58405 8.9868078 236.82625 4.768 21:58:45 40K 0.9 39914 278.97 208.35 0.90 5.56745 9.4493837 240.29539 4.010 21:59:19 40K 0.92 39894 285.7 208.4 0.92 5.56955 9.6542406 241.65834 3.593 21:59:53 40K 0.93 39864 289.21 208.33 0.93 5.57785 9.7248907 242.27788 3.384 22:01:18 40K 0.94 40106 290.75 208.21 0.94 5.51545 9.7502728 243.14525 3.604 22:01:40 40K 0.95 40094 295.4 208.23 0.95 5.52845 9.8072329 243.39307 3.768
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APPENDIX F
DATA ANALYSIS PLAN
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1.0 OVERVIEW
This data analysis plan appendix describes the source of test data and the procedures by which the data will be processed. The role in the data analysis plan will be to perform in-depth analysis of selected test points of interest in order to reach preliminary conclusions. The following data analysis sections were extracted from “Pitot
Statics Calibration and the Standard Atmosphere”, USAF Test Pilot School, Edwards
AFB, CA.
2.0 REQUIRED DATA / DATA FORMAT
The tables under the “Test Methodology” section list the data from each test point that are required for analysis.
3.0 DATA REDUCTION
PACER data will be downloaded to a PC via a PCMCIA data card.
4.0 DATA ANALYSIS
4.1 Tower Flyby Data Reduction
4.1.1 Overview
The tower fly-by FTT is an altitude comparison technique. The truth altitude of the aircraft is determined by triangulation from the tower, and then compared to the pressure altitude read in the cockpit or on a DAS. To find the altitude position correction (∆Hpc), we will
a. Find the geometric altitude of the aircraft relative to the tower b. Find the truth pressure altitude c. Correct cockpit readings for instrument errors d. Find the altitude position correction
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The altitude position correction can be found in this fashion without making any assumptions about total pressure error. If we assume the total pressure error to be zero, we can
e. Find the static port position error ratio (∆Pp/Ps)
For purposes of this discussion, we will use the following example data:
Theodolite Pressure Altitude (Hctower) 2500 feet Grid Reading 5.2 Tower Ambient Temperature (Ttest) 5 °C
Altimeter Instrument Correction (∆Hic) +15 feet Airspeed Instrument Correction (∆Vic) -2.5 knots
Indicated Altitude (Hi) 2450 feet Indicated Airspeed (Vi) 425 knots
4.1.2 Find the Geometric Altitude of the Aircraft Relative to the Tower
The geometric altitude of the aircraft relative to the tower is found by triangulation, as shown in Figure 46.
x y
∆H
z
Fly-By Line
Figure 69. Tower Fly-By Geometry
For Figure 69,
x Eyepiece to Grid Distance y Grid Division Height*Tower Grid Reading
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z Eyepiece to Fly-By Line Distance
By similar triangles, we see that
∆H y = z x
or
z ∆H = y x
If we separate the Grid Division Height from the Tower Grid Reading, we can find a grid constant
⎛ Eyepiece to Flyby Line Dis tan ce ⎞ Grid Cons tan t = ⎜ ⎟()Grid Division Height ⎝ Eyepiece to Grid Dis tan ce ⎠
For the Edwards Fly-by Tower, the grid constant is 31.4 feet/division. Therefore, the aircraft geometric altitude above the tower is given by
∆H = (Grid Constant)(Grid Reading)
⎛ ft ⎞ ∆H = ⎜31.4 ⎟()5.2 divisions ⎝ division ⎠
∆H = 163.28 feet
4.1.3 Find the Truth Pressure Altitude
To find the truth pressure altitude, we need to convert the geometric altitude above the tower to a pressure altitude above the tower. The relationship between geometric altitude and pressure altitude is dependant on the ratio of the test day temperature ratio and the standard temperature for the pressure altitude
Tstd ∆H c = ∆H Ttest
Ttest is measured in the fly-by tower. We find the standard temperature from the pressure altitude measured in the tower.
-6 Tstd = (1 – 6.87559x10 Hctower) TSL
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-6 Tstd = (1 – 6.87559x10 /feet (2500 feet))(288.15 K)
Tstd = 283.2 K
The pressure altitude difference between the aircraft and the tower is thus given by
Tstd ∆H c = ∆H Ttest
283.2 K ∆H = ()163.28 feet c ()5°C + 273.15
∆Hc = 166 feet
The truth pressure altitude of the aircraft is then this difference added to the tower pressure altitude.
Hc = Hctower + ∆Hc
Hc = 2500 feet + 166 feet
Hc = 2666 feet
4.1.4 Correct Cockpit Readings for Instrument Errors
The instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values.
Hic = Hi + ∆Hic
Hic = 2450 feet + (+15 feet)
Hic = 2465 feet
Vic = Vi + ∆Vic
Vic = 425 knots + (-2.5 knots)
Vic = 422.5 knots
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4.1.5 Find the Altitude Position Correction The altitude position correction is the difference between the instrument corrected altitude and the truth pressure altitude.
∆Hpc = Hc - Hic
∆Hpc = 2666 feet – 2465 feet
∆Hpc = 201 feet
4.1.6 Find the Static Port Position Error Ratio (∆Pp/Ps)
If we assume that the Total Pressure Error is negligible, then all of the altitude position correction arises from the error in reading the ambient pressure (difference between static and ambient pressure), or ∆Pp, defined as
∆Pp = Ps - Pa
Therefore, we calculate the static and ambient pressures
-6 5.2559 Ps = PSL(1 – 6.87559x10 Hic) (Hic ≤ 36089.24 feet)(D8)
2 -6 5.2559 Ps = (2116.22 lb/ft )(1 – 6.87559x10 (2465 feet))
2 Ps = 1934.38 lb/ft
-6 5.2559 Pa = PSL (1 – 6.87559x10 Hc) (Hc ≤ 36089.24 feet)(D10)
2 -6 5.2559 Pa = (2116.22 lb/ft )(1 – 6.87559x10 (2666 feet))
2 Pa = 1920.13 lb/ft
Then the static port position error ratio is
∆Pp P − P = s a Ps Ps
2 2 ∆Pp 1934.38 lb / ft −1920.13 lb / ft = 2 Ps 1934.38 lb / ft
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∆Pp = 0.0073667 Ps
4.2 Pace Data Reduction 4.2.1 Overview
The pace FTT is an altitude comparison technique, or a combination altitude and airspeed comparison technique. If the calibrated (“Pace”) aircraft flies by the test aircraft at the same altitude, then it is an altitude comparison technique. If the calibrated aircraft and test aircraft fly in formation, then it is both an altitude and an airspeed comparison technique. The truth altitude and truth airspeed of the aircraft are determined from the instruments in the calibrated aircraft after correcting for instrument and position error. These values are compared to the pressure altitude and airspeed read in the test aircraft cockpit or on a DAS. To find the altitude position correction (∆Hpc) and airspeed position correction (∆Vpc), we will
a. Find the position correction of the pace aircraft b. Find the instrument and position corrected (“truth”) altitude and airspeed for the pace aircraft c. Correct cockpit readings for instrument errors d. Find the altitude and airspeed position corrections e. Find the static port position error ratio (∆Pp/Ps)
For purposes of this discussion, we will use the following example data:
Pace coefficients C0 54 C1 -324 C2 850 C3 0 C4 0
Airspeed Position Correction (∆Vpcpace) +3 knots
Altimeter Instrument Correction (∆Hicpace) +20 feet
Airspeed Instrument Correction (∆Vicpace) -2 knots
Indicated Altitude (Hipace) 20,000 feet
Indicated Airspeed (Vipace) 280 knots
Test aircraft data
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Altimeter Instrument Correction (∆Hictest) -15 feet
Airspeed Instrument Correction (∆Victest) +3 knots
Indicated Altitude (Hitest) 20,010 feet
Indicated Airspeed (Vitest) 275 knots
4.2.2 Find the Position Correction of the Pace Aircraft
Note: This step assumes that the position correction curve for the pace aircraft has been supplied to you as coefficients to a polynomial of the form
2 3 4 ∆Hpcpace = (C0 + C1Micpace + C2Micpace + C3Micpace + C4Micpace + …) θstdtest alt
138
Position error curves of this form were used in the past on AFFTC pace aircraft. Current pace aircraft, such as F-16B/633 are fitted with a computer system that automatically applies instrument and position corrections to the indicated values before displaying them to the crew. Thus these values are Hpc and Vpc. If using a system like this, skip forward to “Correct cockpit readings for instrument errors.”
The polynomial above assumes that the altitude position error is only a function of Mach number and temperature ratio. This assumes that the pace aircraft is at a sufficiently high airspeed that angle of attack effects are negligible and that the Pitot-static probe is uncompensated.
To calculate the altitude position error from the polynomial above, we first need the Mach number and standard day temperature ratio at the test altitude. We can calculate these from altitude and airspeed, but the best values we can get for altitude and airspeed are the instrument corrected values.
The instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values.
Hicpace = Hipace + ∆Hicpace
Hicpace = 20000 feet + (+20 feet)
Hicpace = 20020 feet
Vicpace = Vipace + ∆Vicpace
Vicpace = 280 knots + (-2 knots)
Vicpace = 278 knots
From we can get our best number for the standard day temperature ratio at the test altitude (if we were really picky, we could iterate after finding Hpc, but the difference would most likely be insignificant, i.e. within a foot or two).
-6 θstdtest alt = 1 - 6.87559x10 Hicpace
-6 θstdtest alt = 1 - 6.87559x10 /feet (20020 feet)
θstdtest alt = 0.8623
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To calculate instrument corrected Mach number (Micpace), we need the instrument corrected pressure ratio (δicpace) and instrument corrected airspeed (Vicpace). To calculate
δicpace
-6 5.2559 δicpace = (1 - 6.87559x10 Hicpace)
-6 5.2559 δicpace = (1 - 6.87559x10 /feet (20020 feet))
δicpace = 0.4592
So Mach number would be calculated by ⎡ 2 ⎤ ⎢⎛ ⎧ 7 ⎫ ⎞ 7 ⎥ ⎜ 2 ⎟ ⎢ ⎪⎡ ⎛ V ⎞ ⎤ 2 ⎪ ⎥ ⎜ 1 ⎪⎢ ⎜ icpace ⎟ ⎥ ⎪ ⎟ M = 5⎢⎜ ⎨ 1+ 0.2 −1⎬ +1⎟ −1⎥ ic pace ⎢ δ ⎢ ⎜ a ⎟ ⎥ ⎥ ⎜ icpace ⎪⎢ ⎝ SL ⎠ ⎥ ⎪ ⎟ ⎢⎜ ⎪⎣ ⎦ ⎪ ⎟ ⎥ ⎢⎝ ⎩ ⎭ ⎠ ⎥ ⎣ ⎦
⎡ 2 ⎤ ⎢⎛ ⎧ 7 ⎫ ⎞ 7 ⎥ ⎜ ⎪⎡ 2 ⎤ 2 ⎪ ⎟ ⎢⎜ 1 ⎪ ⎛ 278 knots ⎞ ⎪ ⎟ ⎥ M = 5⎢ ⎨⎢1+ 0.2⎜ ⎟ ⎥ −1⎬ +1 −1⎥ ic pace ⎜ 0.4592 ⎢ ⎜ 661.48 knots ⎟ ⎥ ⎟ ⎢⎜ ⎪⎣ ⎝ ⎠ ⎦ ⎪ ⎟ ⎥ ⎢⎝ ⎩⎪ ⎭⎪ ⎠ ⎥ ⎣⎢ ⎦⎥
Micpace = 0.6058
Now we can calculate the altitude position correction.
2 3 4 ∆Hpcpace = (C0 + C1Micpace + C2Micpace + C3Micpace + C4Micpace + …) θstdtest alt
2 3 4 ∆Hpcpace = (54 + (-324)(0.6058) + (850) (0.6058) + (0) (0.6058) + (0) (0.6058) + …) (0.8623)
∆Hpcpace = 146 feet (units determined by polynomial coefficients)
4.2.3 Find the Instrument and Position Corrected Altitude and Airspeed for the Pace Aircraft
So the instrument and position corrected altitude (“truth” altitude) is
140
Hpcpace = Hicpace + ∆Hpcpace
Hpcpace = 20020 feet + (+146 feet)
Hpcpace = 20166 feet
Note that if we took this altitude and recalculated θstdtest alt we would get 0.8613, a difference of 0.1 percent, and would change the altitude position correction to 145.8 feet, or an error of less than a foot for using Hic to calculate θstdtest alt instead of iterating on the answer.
To calculate instrument and position corrected airspeed, the airspeed position correction would have to be provided or calculated from ∆Hpc assuming zero total pressure error.
Vpcpace = Vicpace + ∆Vpcpace
Vpcpace = 278 knots + (+3 knots)
Vpcpace = 281 KCAS
4.2.4 Correct Cockpit Readings for Instrument Errors
The instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values.
Hictest = Hitest + ∆Hictest
Hictest = 20,010 feet + (-15 feet)
Hictest = 19,995 feet
Victest = Vitest + ∆Victest
Victest = 275 knots + (+3 knots)
Victest = 278 knots
4.2.5 Find the Test Day Altitude And Airspeed Position Corrections
141
The test day position corrections are the difference between the instrument corrected altitude and the truth pressure altitude.
∆Hpctest = Hpcpace - Hictest
∆Hpctest = 20,166 feet – 19,995 feet
∆Hpctest = 171 feet
∆Vpctest = Vpcpace - Victest
∆Vpctest = 281 knots – 278 knots
∆Vpctest = 3 knots
4.2.6 Find the Position Error Ratio
The position error ratio can be calculated from the altitude position correction or the airspeed position correction. If the total pressure error is truly negligible, the results from both methods will be the same.
If we choose to start with the altitude position correction, we calculate the static and ambient pressures
-6 5.2559 Ps = PSL(1 – 6.87559x10 Hictest) (Hictest ≤ 36089.24 feet) (D8)
2 -6 5.2559 Ps = (2116.22 lb/ft )(1 – 6.87559x10 (19,995 feet))
2 Ps = 972.70 lb/ft
-6 5.2559 Pa = PSL (1 – 6.87559x10 Hpcpace) (Hpcpace ≤ 36089.24 feet) (D10)
2 -6 5.2559 Pa = (2116.22 lb/ft )(1 – 6.87559x10 (20,166 feet))
2 Pa = 965.75 lb/ft
Then the static port position error ratio is
142
∆Pp P − P = s a Ps Ps
2 2 ∆Pp 972.70 lb / ft − 965.75 lb / ft = 2 Ps 972.70 lb / ft
∆Pp = 0.007145 Ps
If we choose to start with the airspeed position correction, we calculate the following ratios
7 ⎛ 2 ⎞ 2 q ⎜ ⎛ Vpcpace ⎞ ⎟ c = 1+ 0.2⎜ ⎟ −1 (V < a ) (C100) ⎜ ⎜ ⎟ ⎟ cpace SL PSL ⎜ a SL ⎟ ⎝ ⎝ ⎠ ⎠
7 2 2 q ⎛ ⎛ 281 knots ⎞ ⎞ c = ⎜1+ 0.2⎜ ⎟ ⎟ −1 P ⎜ ⎜ 661.48 knots ⎟ ⎟ SL ⎝ ⎝ ⎠ ⎠
q c = 0.13212 PSL
7 ⎛ 2 ⎞ 2 q ⎜ ⎛ Vic ⎞ ⎟ cic = 1+ 0.2⎜ test ⎟ −1 (V < a ) (D28) ⎜ ⎜ ⎟ ⎟ ic SL PSL ⎜ a SL ⎟ ⎝ ⎝ ⎠ ⎠
7 2 2 q ⎛ ⎛ 278 knots ⎞ ⎞ cic = ⎜1+ 0.2⎜ ⎟ ⎟ −1 P ⎜ ⎜ 661.48 knots ⎟ ⎟ SL ⎝ ⎝ ⎠ ⎠
q cic = 0.12919 PSL
Then
⎛ q q ⎞ ⎜ c cic ⎟ ∆Pp = ⎜ − ⎟PSL (D32) ⎝ PSL PSL ⎠
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2 ∆Pp = (0.13212 − 0.12919)2116.22 lb / ft
2 ∆Pp = 6.2005 lb/ft
2 ∆Pp 6.2005 lb / ft = 2 Ps 972.70 lb / ft
∆Pp = 0.0063745 Ps
While this value looks much different than the value determined from the altitude position correction, starting with the value determined from the altitude position correction and working backwards gives a Victest of 277.629 knots, which is only 0.370 knots different. Thus, it would be reasonable to say that this example does not show a significant total pressure error.
4.3 Survey Data Reduction
4.3.1 Overview
The survey FTT is an altitude comparison technique. The truth altitude of the aircraft is determined by comparing its geometric altitude to the geometric altitude of a calibrated aircraft at the same location but at a different time. To find the altitude position correction (∆Hpc), we will
a. Determine the tapeline (geometric) altitude of both aircraft over a selected location b. Find the difference in tapeline altitudes c. Correct cockpit readings for instrument errors d. Convert tapeline altitude difference to a pressure altitude difference e. Find the altitude position correction
The altitude position correction can be found in this fashion without making any assumptions about total pressure error. If we assume the total pressure error to be zero, we can
f. Find the static port position error ratio (∆Pp/Ps)
4.3.2 Determine the Tapeline (Geometric) Altitude of Both Aircraft Over A Selected Location
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Because the test aircraft accelerates through the measurement corridor, the location of interest is determined by the position of the test aircraft as it accelerated through a selected airspeed.
For purposes of this discussion, we will use the following example data:
Indicated Altitude (Hitest) 19970 feet
Indicated Airspeed (Vitest) 425 knots Time 16:04:52
Altimeter Instrument Correction (∆Hictest) -15 feet
Airspeed Instrument Correction (∆Victest) -2.5 knots
At the time 16:04:52, the location data (radar or GPS) give a location of the aircraft and a tapeline altitude of
Tapeline Altitude (Htapetest) 25132 feet
Consulting the location data for the calibrated aircraft, when it passed closest to the location of the test aircraft specified above, the following data were recorded
Tapeline Altitude (Htapecal) 25202 feet Time 15:59:32
Consulting the data recorded onboard the calibrated aircraft at 15:59:32
Indicated Altitude (Hical) 20020 feet
Indicated Airspeed (Vical) 300 knots Indicated Temperature (Ti) 25 °C
Altimeter Instrument Correction (∆Hiccal) -10 feet
Altimeter Position Correction (∆Hpccal) +40 feet
Airspeed Instrument Correction (∆Viccal) -1.5 knots
Airspeed Position Correction (∆Vpccal) +1 knots
Temperature Instrument Correction (∆Tic) +1 °C Temperature Recovery Factor (Kt) 0.98
4.3.3 Find the Difference in Tapeline Altitudes
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The tapeline difference in altitude would be
∆Htape = Htapecal - Htapetest
∆Htape = 25202 feet - 25132 feet
∆Htape = 70 feet
4.3.4 Correct Cockpit Readings for Instrument Errors
For the test aircraft, the instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values.
Hictest = Hitest + ∆Hictest
Hictest = 19970 feet + (-15 feet)
Hictest = 19955 feet
Victest = Vitest + ∆Victest
Victest = 425 knots + (-2.5 knots)
Victest = 422.5 knots
For the calibrated aircraft, the instrument and position corrected altitude and airspeed are calculated by adding the instrument and position corrections to the indicated values.
Hpccal = Hical + ∆Hiccal + ∆Hpccal
Hpccal = 20020 feet + (-10 feet) + (+40 feet)
Hpccal = 20050 feet
Vpccal = Vical + ∆Viccal + ∆Vpccal
Vpctest = 300 knots + (-1.5 knots) + (+1 knot)
Vpctest = 299.5 knots
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To determine the temperature at the test altitude, we need the Mach number of the calibrated aircraft.
-6 5.2559 δ = (1 - 6.87559x10 Hpccal)
δ = (1 - 6.87559x10-6/feet (20050 feet))5.2559
δ = 0.45858
⎡ 2 ⎤ ⎢⎛ ⎧ 7 ⎫ ⎞ 7 ⎥ ⎢⎜ ⎪⎡ 2 ⎤ 2 ⎪ ⎟ ⎥ ⎜ 1 ⎪ ⎛ Vc ⎞ ⎪ ⎟ M = 5⎢ ⎨⎢1+ 0.2⎜ ⎟ ⎥ −1⎬ +1 −1⎥ (M < 1) (C110) ⎜ δ ⎢ ⎜ a ⎟ ⎥ ⎟ ⎢⎜ ⎪⎣ ⎝ SL ⎠ ⎦ ⎪ ⎟ ⎥ ⎢⎝ ⎩⎪ ⎭⎪ ⎠ ⎥ ⎣⎢ ⎦⎥
⎡ 2 ⎤ ⎛ ⎧ 7 ⎫ ⎞ 7 ⎢⎜ ⎡ 2 ⎤ 2 ⎟ ⎥ ⎢ 1 ⎪ ⎛ 299.5 knots ⎞ ⎪ ⎥ M = 5 ⎜ ⎨⎢1+ 0.2⎜ ⎟ ⎥ −1⎬ +1⎟ −1 ⎢⎜ 0.45858 ⎢ ⎜ 661.48 knots ⎟ ⎥ ⎟ ⎥ ⎢⎜ ⎪⎣ ⎝ ⎠ ⎦ ⎪ ⎟ ⎥ ⎝ ⎩⎪ ⎭⎪ ⎠ ⎣⎢ ⎦⎥
M = .6509
4.3.5 Convert Tapeline Altitude Difference to a Pressure Altitude Difference
To convert the tapeline altitude difference to a pressure altitude difference, we need to know the standard temperature and the ambient temperature at the test altitude.
For the standard temperature
-6 Tstd = (1 - 6.87559x10 Hpccal)TSL
-6 Tstd = (1 - 6.87559x10 /feet (20050 feet))(288.15 K)
Tstd = 248.42 K
For the test temperature
T + ∆T T = i ic a 2 1+ 0.2K t M
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25 +1+ 273.15 Ta = 1+ 0.2()0.98 (0.6509)2
Ta = 276.21 K
To convert the difference in tapeline altitude to a difference in pressure altitude
Tstd ∆H c = ∆H tape Ta
248.42 K ∆H = ()70 feet c 276.25 K
∆Hc = 62.9 feet
4.3.6 Find the Altitude Position Correction The truth pressure altitude for the test aircraft can be found by subtracting the difference in pressure altitude from the pressure altitude of the calibrated aircraft.
Hctest = Hpccal - ∆Hc
Hctest = 20050 feet – 62.9 feet
Hctest = 19987 feet
The altitude position correction is the difference between the instrument corrected altitude and the truth pressure altitude.
∆Hpctest = Hctest - Hictest
∆Hpctest = 19987 feet – 19955 feet
∆Hpctest = 32 feet
4.3.7 Find the Static Port Position Error Ratio (∆Pp/Ps)
If we assume that the Total Pressure Error is negligible, then all of the altitude position correction arises from the error in reading the ambient pressure (difference between static and ambient pressure), or ∆Pp, defined as
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∆Pp = Ps - Pa
Therefore, we calculate the static and ambient pressures
-6 5.2559 Ps = PSL(1 – 6.87559x10 Hictest) (Hic ≤ 36089.24 feet)(D8)
2 -6 5.2559 Ps = (2116.22 lb/ft )(1 – 6.87559x10 (19955 feet))
2 Ps = 974.33 lb/ft
-6 5.2559 Pa = PSL (1 – 6.87559x10 Hctest) (Hc ≤ 36089.24 feet)(D10)
2 -6 5.2559 Pa = (2116.22 lb/ft )(1 – 6.87559x10 (19987 feet))
2 Pa = 973.02 lb/ft
Then the position error ratio is
∆Pp P − P = s a Ps Ps
2 2 ∆Pp 974.33 lb / ft − 973.02 lb / ft = 2 Ps 974.33 lb / ft
∆Pp = 0.0013445 Ps
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4.4 Cloverleaf Data Reduction
4.4.1 Overview
The cloverleaf FTT is an airspeed comparison technique. The truth airspeed of the aircraft is determined from measurements of ground speed and track angle, and then compared to the airspeed read in the cockpit or on a DAS. The basic process is
a. Correct cockpit readings for instrument errors b. Find the indicated Mach number c. Find the indicated true airspeed d. Find the truth true airspeed and wind velocity e. Find the truth Mach number f. Find the Mach position correction g. Find the position error ratio (∆Pp/Ps)
The airspeed position correction can be found from the position error ratio without making any assumptions about total pressure error. The calculation of the position error ratio (∆Pp/Ps) is based on the concept that total pressure error is zero. Even if this assumption is false, the calculation of airspeed correction from the position error ratio should still be correct, since both the Mach and airspeed depend on both total and static pressure.
For purposes of this discussion, we will use the following example data:
First Leg
Indicated Airspeed (Vi) 130 knots Indicated Altitude (Hi) 6000 feet Indicated Temperature (Ti) 11 °C GPS Ground Speed (Vg) 130.3 knots GPS Ground Track (Ψ) 7°
Second Leg
Indicated Airspeed (Vi) 130 knots Indicated Altitude (Hi) 6000 feet Indicated Temperature (Ti) 11 °C GPS Ground Speed (Vg) 153.8 knots GPS Ground Track (Ψ) 114°
Third Leg
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Indicated Airspeed (Vi) 130 knots Indicated Altitude (Hi) 6000 feet Indicated Temperature (Ti) 11 °C GPS Ground Speed (Vg) 122.5 knots GPS Ground Track (Ψ) 234°
Altimeter Instrument Correction (∆Hic) -20 feet Airspeed Instrument Correction (∆Vic) +2 knots Temperature Instrument Correction (∆Tic) -1°C (-1 K) Temperature Recovery Factor 1.0
4.4.2 Correct Cockpit Readings for Instrument Errors
The instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values. If the indicated values of airspeed or altitude are not identical between runs, about the best that we can do is to average the two values.
Hic = Hi + ∆Hic
Hic = 6000 feet + (-20 feet)
Hic = 5980 feet
Vic = Vi + ∆Vic
Vic = 130 knots + (+2 knots)
Vic = 132 knots
Tic = Ti + ∆Tic
Tic = 11 °C + (-1 °C) + 273.15
Tic = 283.15 K
4.4.3 Find the Indicated Mach Number
So our next step is to determine the indicated Mach number at the test conditions. We could jump straight to the Mach meter equation with calibrated airspeed and pressure altitude, but since we will need qcic/PSL later, we’ll calculate that first. This is the value in the middle of the calibrated airspeed equation.
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3.5 2 q ⎛ ⎛ V ⎞ ⎞ cic = ⎜1+ 0.2⎜ ic ⎟ ⎟ −1 P ⎜ ⎜ a ⎟ ⎟ SL ⎝ ⎝ SL ⎠ ⎠
3.5 2 q ⎛ ⎛ 132 knots ⎞ ⎞ cic = ⎜1+ 0.2⎜ ⎟ ⎟ −1 P ⎜ ⎜ 661.48 knots ⎟ ⎟ SL ⎝ ⎝ ⎠ ⎠
q cic = 0.028153 PSL
Note that because this value was calculated using Vic, this is the pressure ratio seen at the airspeed indicator, and thus contains the position error. This is why the value is qcic/PSL and not qc/PSL.
To find the Mach number we will need qcic/Ps, and to find that we need qcic/PSL and δic. We’ll find δic (which also contains the position error) seen by the altimeter from the instrument corrected altitude, Hic.
-6 5.2559 δic = (1 - 6.87559x10 Hic)
-6 5.2559 δic = (1 - 6.87559x10 /feet (5980 feet))
δic = 0.80198
We can now find qcic/Ps
q q P q 1 cic = cic SL = cic Ps PSL Ps PSL δic
q 1 cic = ()0.028153 Ps 0.80198
q cic = 0.035104 Ps
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This is the number we find in the middle of the true airspeed and Mach equations. Now we can find the instrument corrected Mach number, Mic, which is what a perfect Mach meter in the cockpit would have read. Note that because this Mach is calculated using numbers that contain the position error, it too contains the position error.
⎡ 2 ⎤ ⎛ q ⎞ 7 M = 5⎢⎜ cic +1⎟ −1⎥ ic ⎢⎜ ⎟ ⎥ ⎝ Ps ⎠ ⎣⎢ ⎦⎥
⎡ 2 ⎤ M = 5 ()0.035104 +1 7 −1 ic ⎣⎢ ⎦⎥
Mic = 0.22256
4.4.4 Find the Indicated True Airspeed
To find the indicated true airspeed (true airspeed calculated from indicated airspeed), we need to find the ambient temperature. Since we don’t have the true Mach number, we will use the indicated Mach number as our best information.
T T = ic ai 2 1+ 0.2K t Mic
283.15 T ai = 1+ 0.2()1.0 (0.22256)2
T 280.37 K ai =
With this temperature we can calculate the indicated true airspeed
V = M a θ ti ic SL
280.37K V = ()0.22256 (661.48 knots) ti 288.15K
V = 145.22 knots ti
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4.4.5 Find the Truth True Airspeed And Wind Velocity
This step involves building the solution matrix. First we need to find the components of ground speed. For the first leg
Vgx = Vg sin θ (E10)
Vgx = 130.3 sin 7°
Vgx = 15.88 knots
Vgy = Vg cos θ (E11)
Vgy = 130.3 cos 7°
Vgy = 129.33 knots
The solution matrix is given by
2 2 ⎡2V + ∆V 2V − V 2V − V ⎤⎡∆V ⎤ ⎡V − V ⎤ ti t gx1 wx gy1 wy t ⎢ g1 ti ⎥ ⎢ ⎥⎢ ⎥ 2 2 ⎢2Vti + ∆Vt 2Vgx2 − Vwx 2Vgy2 − Vwy ⎥⎢Vwx ⎥ = ⎢Vg2 − Vti ⎥ (E20) ⎢ ⎥ ⎢2V + ∆V 2V − V 2V − V ⎥⎢V ⎥ V 2 − V 2 ⎣ ti t gx3 wx gy3 wy ⎦⎣ wy ⎦ ⎣⎢ g3 ti ⎦⎥
[A][x]= [C] (E21)
To start the calculations, we will assume
⎡∆Vt ⎤ ⎡0⎤ ⎢ ⎥ ⎢ ⎥ ⎢Vwx ⎥ = ⎢0⎥ (E23) ⎢ ⎥ ⎢ ⎥ ⎣Vwy ⎦ ⎣0⎦
Calculating the numbers gives an initial matrix
⎡290.44 31.76 258.66 ⎤⎡∆Vt ⎤ ⎡− 4110.8⎤ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎢290.44 281.01 −125.11⎥⎢Vwx ⎥ = ⎢ 2565.6 ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎣290.44 −198.21 −144.01⎦⎣Vwy ⎦ ⎣− 6082.6⎦
Using the MINVERSE function in Excel to invert this matrix and multiplying for a new solution
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[x]= [A]−1[C] (E22)
⎡∆Vt ⎤ ⎡0.001191 0.000852 0.001399 ⎤⎡− 4110.8⎤ ⎡−11.22207⎤ ⎢ ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎢Vwx ⎥ = ⎢ − 0.0001 0.002135 − 0.002035⎥⎢ 2565.6 ⎥ = ⎢ 18.26476 ⎥ ⎢ ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎣Vwy ⎦ ⎣0.002541 − 0.001219 − 0.001321⎦⎣− 6082.6⎦ ⎣− 5.534321⎦
Using this new solution vector to recalculate the A matrix we have
⎡279.2179 13.49439 264.1918 ⎤⎡∆Vt ⎤ ⎡− 4110.8⎤ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎢279.2179 262.7418 −119.5779⎥⎢Vwx ⎥ = ⎢ 2565.6 ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎣ 279.279 − 216.4739 −138.4731⎦⎣Vwy ⎦ ⎣− 6082.6⎦
Using the MINVERSE function in Excel to invert this matrix and multiplying for a new solution
⎡∆Vt ⎤ ⎡0.001182 0.00105 0.001349 ⎤⎡− 4110.8⎤ ⎡−10.36863⎤ ⎢ ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎢Vwx ⎥ = ⎢ − 0.0001 0.002135 − 0.002035⎥⎢ 2565.6 ⎥ = ⎢ 18.26476 ⎥ ⎢ ⎥ ⎢ ⎥⎢ ⎥ ⎢ ⎥ ⎣Vwy ⎦ ⎣0.002541 − 0.001219 − 0.001321⎦⎣− 6082.6⎦ ⎣− 5.534321⎦
Repeating this iteration three more times gives a final solution vector
⎡∆Vt ⎤ ⎡−10.33583⎤ ⎢ ⎥ ⎢ ⎥ ⎢Vwx ⎥ = ⎢ 18.26476 ⎥ ⎢ ⎥ ⎢ ⎥ ⎣Vwy ⎦ ⎣− 5.534321⎦
From this we see that the error in true airspeed is –10.3 knots. Thus the truth true airspeed would be
Vt = Vti + ∆Vt
Vt = 145.2 knots + (-10.3 knots)
Vt = 134.9 knots
The wind vector would be found by
2 2 2 Vw = Vwx + Vwy
2 2 2 Vw = (18.26476) + (-5.534321)
Vw = 19.08 knots
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V Ψ = tan −1 wx Vwy
18.26476 Ψ = tan −1 − 5.534321
Ψ = 106.9 deg rees
Two things to note about the calculation of the wind angle. The “y” and “x” coordinate seem to be reversed. This is because compass directions are defined differently than Cartesian coordinates. Rather than being defined counterclockwise from the horizontal (x) axis, directions are defined clockwise from the vertical (north, y) axis. Using the function ATAN2(Vwy, Vwx) in Excel will return the proper direction in the proper quadrant.
Secondly, this is the direction of the wind vector, i.e. where the wind is blowing to, not from. To match the traditional method of defining wind direction as the “from” direction, simply add 180 degrees to get 286.9 degrees.
4.4.6 Find the Truth Mach Number
We can find the truth Mach number from the truth true airspeed, but we will need to know the local speed of sound at the test conditions. To find the local speed of sound, we need to know the temperature. Now that we know the truth true airspeed, we can use that to find the correct ambient temperature. We will use the alternative method, which requires knowing the value of Cp for air. Various values for CP are
m2 knots2 mph2 ft2 Cp =1005 = 3791 = 5028 = 10820 sec2 K K K sec2 K
The ambient temperature is then given by
2 K t Vt Ta = Ti + ∆Tic − 2Cp
(1.0)(134.9 knots)2 Ta = ()11°C + (−1°C)− + 273.15 ⎛ knots 2 ⎞ 2⎜3791 ⎟ ⎜ ⎟ ⎝ K ⎠
Ta = 7.6 °C = 280.7 K
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The test temperature ratio is given by
Ta θtest = Tsl
280.7 K θ = test 288.15 K
θtest = 0.97415
Now we can find the truth Mach number
V M = t asl θtest
134.9 knots M = 661.48 knots 0.97415
M = 0.2066
4.4.7 Find the Mach Position Correction
Now that we know the truth Mach number and the indicated Mach number, the Mach position correction can be found simply by subtracting.
∆Mpc =M – Mic
∆Mpc = 0.2066 – 0.22256
∆Mpc = -0.01596
4.4.8 Find the Static Port Position Error Ratio (∆Pp/Ps)
From the truth Mach number, we calculate the ratio PT/Pa
P 7 T = (1 + 0.2M 2 ) 2 (M < 1) (D68) Pa
P 7 T = (1+ 0.2()0.2066 2 ) 2 Pa
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P T = 1.03020 Pa
Doing the same for instrument corrected Mach number
7 PT 2 2 = (1 + 0.2Mic ) (Mic < 1) (D71) Ps
P 7 T = (1+ 0.2()0.22256 2 ) 2 Ps
P T = 1.03510 Ps
The position error can then be calculated directly from these ratios by
⎛ ⎞ ⎜ ⎟ ∆Pp ⎜ 1 1 ⎟ PT = ⎜ − ⎟ (D74) Ps PT PT Ps ⎜ ⎟ ⎝ Ps Pa ⎠
∆Pp ⎛ 1 1 ⎞ = ⎜ − ⎟1.03510 Ps ⎝1.03510 1.03020 ⎠
∆Pp = −0.0047563 Ps
4.5 Constant Airspeed Turn Data Reduction
4.5.1 Overview The Constant Airspeed Turn FTT is an airspeed comparison technique to calculate the effects of AoA. The truth airspeed of the aircraft is determined from measurements of ground speed and track angle, and then compared to the airspeed read in the cockpit or on a DAS. A GPS is used as a truth source for ground track and altitude. The basic process is
a. Correct cockpit readings for instrument errors b. Find the indicated Mach number c. Find the indicated true airspeed d. Find the truth true airspeed and wind velocity
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e. Find the truth Mach number f. Find the Mach position correction g. Find the position error ratio (∆Pp/Ps)
Perform a wind cal using a wind triangle
The airspeed position correction can be found from the position error ratio without making any assumptions about total pressure error. The calculation of the position error ratio (∆Pp/Ps) is based on the concept that total pressure error is zero. Even if this assumption is false, the calculation of airspeed correction from the position error ratio should still be correct, since both the Mach and airspeed depend on both total and static pressure.
4.5.2 Wind Triangle
The Constant Airspeed Turn data reduction is based on using wind triangles to calculate the true airspeed error and the wind speed and direction. The necessary assumptions are:
1. Indicated Airspeed, and thus True Airspeed is the same throughout the turn 2. Wind Speed and Direction are constant during the turn 3. All wind is horizontal (no vertical component)
Figure E1 shows the basic wind triangle.
ϖ Vw ϖ ∆Vt ϖ ϖ V V g ti
Figure E1. Basic wind triangle where
Vti True airspeed as calculated from the instrument corrected airspeed ∆Vt Position error in true airspeed Vw Wind velocity Vg Ground velocity
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The vector equation for the basic wind triangle is
ϖ ϖ ϖ Vg = Vt + Vw (E1)
Breaking the true airspeed into the true airspeed calculated from the instrument corrected airspeed and the position error in true airspeed we get
ϖ ϖ V = V + ∆V t ti t (E2)
Substituting into Equation E1 gives
ϖ ϖ ϖ V = V + ∆V + V g ti t w (E3)
Which by rearranging gives
ϖ ϖ ϖ V + ∆V = V − V ti t g w (E4)
Within this equation are three knowns and five unknowns. The known values are the magnitude of Vti, the magnitude of Vg (ground speed) and the direction of Vg (ground track). The unknowns are the direction of Vti, the magnitude of ∆Vt, the direction of ∆Vt, the magnitude of Vw (wind speed), and the direction of Vw (wind direction).
4.5.3 Correct Cockpit Readings for Instrument Errors
The instrument corrected altitude and airspeed are calculated by adding the instrument corrections to the indicated values. If the indicated values of airspeed or altitude are not identical between runs, about the best that we can do is to average the two values.
Hic = Hi + ∆Hic
Vic = Vi + ∆Vic
Tic = Ti + ∆Tic
4.5.4 Find the Indicated Mach Number
So our next step is to determine the indicated Mach number at the test conditions. We could jump straight to the Mach meter equation with calibrated airspeed and pressure altitude, but since we will need qcic/PSL later, we’ll calculate that first. This is the value in the middle of the calibrated airspeed equation.
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3.5 2 q ⎛ ⎛ V ⎞ ⎞ cic = ⎜1+ 0.2⎜ ic ⎟ ⎟ −1 P ⎜ ⎜ a ⎟ ⎟ SL ⎝ ⎝ SL ⎠ ⎠
Note that because this value was calculated using Vic, this is the pressure ratio seen at the airspeed indicator, and thus contains the position error. This is why the value is qcic/PSL and not qc/PSL.
To find the Mach number we will need qcic/Ps, and to find that we need qcic/PSL and δic. We’ll find δic (which also contains the position error) seen by the altimeter from the instrument corrected altitude, Hic.
-6 5.2559 δic = (1 - 6.87559x10 Hic)
We can now find qcic/Ps q q P q 1 cic = cic SL = cic Ps PSL Ps PSL δic
This is the number we find in the middle of the true airspeed and Mach equations. Now we can find the instrument corrected Mach number, Mic, which is what a perfect Mach meter in the cockpit would have read. Note that because this Mach is calculated using numbers that contain the position error, it too contains the position error.
⎡ 2 ⎤ ⎛ q ⎞ 7 M = 5⎢⎜ cic +1⎟ −1⎥ ic ⎢⎜ ⎟ ⎥ ⎝ Ps ⎠ ⎣⎢ ⎦⎥
4.5.5 Find the Indicated True Airspeed
To find the indicated true airspeed (true airspeed calculated from indicated airspeed), we use the ambient temperature from our previously calibrated temperature probe. With this temperature we can calculate the indicated true airspeed
V = M a θ ti ic SL
4.5.6 Find the Truth True Airspeed and Wind Velocity
First we need to find the components of ground speed.
Vgx = Vg sin θ
Vgy = Vg cos θ
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From this we find the error in the true airspeed. Thus the truth true airspeed would be
Vt = Vti + ∆Vt
The wind vector would be found by
2 2 2 Vw = Vwx + Vwy
V Ψ = tan −1 wx Vwy
Two things to note about the calculation of the wind angle. The “y” and “x” coordinate seem to be reversed. This is because compass directions are defined differently than Cartesian coordinates. Rather than being defined counterclockwise from the horizontal (x) axis, directions are defined clockwise from the vertical (north, y) axis. Using the function ATAN2(Vwy, Vwx) in Excel will return the proper direction in the proper quadrant.
Secondly, this is the direction of the wind vector, i.e. where the wind is blowing to, not from. To match the traditional method of defining wind direction as the “from” direction, simply add 180 degrees.
4.5.7 Find the Truth Mach Number
We can find the truth Mach number from the truth true airspeed, but we will need to know the local speed of sound at the test conditions. To find the local speed of sound, we need to use our measured and corrected temperature. Now that we know the truth true airspeed, we can use that to find the correct ambient temperature. The test temperature ratio is given by Ta θtest = Tsl
Now we can find the truth Mach number V M = t asl θtest
4.5.8 Find the Mach Position Correction
Now that we know the truth Mach number and the indicated Mach number, the Mach position correction can be found simply by subtracting.
∆Mpc =M – Mic
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4.5.9 Find the static port position error ratio (∆Pp/Ps)
From the truth Mach number, we calculate the ratio PT/Pa
P 7 T = (1 + 0.2M 2 ) 2 (M < 1) Pa
Doing the same for instrument corrected Mach number
7 PT 2 2 = (1 + 0.2Mic ) (Mic < 1) Ps
The position error can then be calculated directly from these ratios by
⎛ ⎞ ⎜ ⎟ ∆Pp ⎜ 1 1 ⎟ PT = ⎜ − ⎟ Ps PT PT Ps ⎜ ⎟ ⎝ Ps Pa ⎠
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