Appendix 9
d. Data Reduction.
Vpect (ViB VicB Vpec ) (ViT VicT ) kts H pect (H iB H icB Vpec ) (H iT HicT ) kts ft VpecT and H pecT are calculated throughout the speed range in each configuration and plotted as shown in figures 3 and 4.
6. Tower Fly-by Method. The tower fly-by method is one of the methods which results in a direct determination of static error in indicated pressure altitude.
Since the altimeter and airspeed system use the same static source, it is possible to correlate the altimeter position error directly to the airspeed error. This correlation assumes that there is no error in the total head system.
The tower fly-by method can be modified to use radar altimeter or differential GPS to determine geometric/tapeline height above a ground based pressure measuring station.
Figure A9-6 Tower Fly-By Method
a. Procedures and Test Conditions for Tower Fly-by
(1) Air Quality. Smooth, stable air is needed for determining the error in pressure altitude.
(2) Weight and C.G. Same as for calibrations of the airspeed indicating system.
(3) Speed Range. The calibration should range from 1.3 VSO to 1.8 VS1. Higher speeds up to VMO or VNE are usually investigated so that errors can be included in the AFM for a full range of airspeeds.
A9-16
-- Appendix 9
(4) Test Procedures.
(i) The test technique is to fly the aircraft along a ground reference line, past the tower, in stabilized flight at a constant airspeed and at the approximate height of the tower. The primary piloting task is to maintain a constant indicated altitude during the run. The tower is equipped with a sensitive altimeter and a means of determining the relative angle ( ) of the aircraft. The data recorded during each run are the indicated pressure altitude of the tower, (Hitower), the angle ( ), and the aircraft's indicated pressure altitude, airspeed and temperature (HiA/C, ViA/C, and TiA/C) as it passes the tower. Note that the tower altimeter must be at the zero grid line position in the tower.
(ii) Repeat step (i) at various airspeeds in increments sufficient to cover the required range at each flap setting.
(5) Data Acquisition. Data to be recorded at each test point:
(i) Airplane Airspeed ViA/C knots.
(ii) Airplane indicated pressure attitude, HiA/C knots.
(iii) Tower observer indicated pressure altitude, Hitower.
(iv) Angle of aircraft above the tower.
(v) Wing flap position.
(vi) Landing gear position.
(vii) Fuel used in airplane.
(viii) TiA/C and Titower
(6) Data Reduction. The actual pressure altitude of the aircraft is Hcref where TsHcref = (Hitower + Hictower) + D tan Tt Where Ts is the standard day absolute temperature at the test altitude and Tt is the test day temperature in absolute units.
A9-17
Appendix 9 TThe s temperature correction is to convert the geometric height of the aircraft above the Tt reference zero grid line in the tower (D tan ) to a pressure height that can be added to the pressure altitude of the tower Hctower. The difference between the actual refere nce pressure altitude of the aircraft and the aircraft's instrument-corrected pressure altitude is the position error correction.
Hpec = Hcref - (HiA/C + HicA/C) Ts = [(Hitower + Hictower) + D tan ] - (HiA/C + HicA/C)Tt Hpec is calculated for every speed and aircraft configuration flown past the tower and the data are plotted as per Figure A9-4.
The airspeed system position error corrections can be obtained from the tower fly-by method if it is assumed that the pitot tube (total head) errors are zero.
The hydrostatic equilibrium equation states that the pressure error correction at the static source is ps = - g Hpc and from Section 3.
2 4 2 41 2 M c M c 1 2 M ic M icpd pp ps oVc (1 ....) oVic (1 ....)2 4 40 2 4 40 Since it is assumed that p 0 and for low speed aircraft, compressibility effects can be 1 2 2ignored, then p p (V V )d s o c ic 2 The above equation is used to calculate VC at every test point, then Vpec Vc Vic . The data are then plotted as per Figure A9-5.
7. Ground Run Airspeed System Calibration. The airspeed system is calibrated to show compliance with commuter category requirements of § 23.1323(c) during the accelerate-takeoff ground run, and is used to determine IAS values for various V1 and VR speeds. The airspeed system error during the accelerate-takeoff ground run may be determined using a trapped static source reference or a distance measuring unit that provides readouts of ground speed that can be converted into CAS.
a. Definitions.
(1) Ground Run System Error. System error during the accelerate-takeoff ground run is the combination of position error, instrument error, and the dynamic effects, such as lag, which may be caused by acceleration on the runway.
(2) Trapped Static Source. An airtight bottle with sufficient internal volume so as to be infinite when compared to an airspeed indicator's internal changes in volume while sensing various airspeeds. The bottle should be insulated to mi nimize internal bottle temperature changes as testing is in progress. For short periods of time, it can be assumed that the bottle will reflect true static ambient pressure to the test indicator.
A9-18
Appendix 9
(3) Production Airspeed Indicator. An airspeed indicator that conforms to the type certification design standards. The indicator should be installed in the approved instrument panel location since these tests involve the dynamic effects of the indicator that may result from acceleration.
(4) Test Airspeed Indicator. A mechanical airspeed indicator with known dynamic characteristics during acceleration or an electronic transducer that can provide airspeed information.
(5) Test Reference Altimeter. An altimeter that indicates the altitude of the air trapped in the bottle or local ambient static air if the valve is opened.
(6) Ground Run Position Error. Ground run position error is the static-pressure error of the production static source during ground runs with any ground effects included. Any contributions to error due to the total-pressure (pitot) are ignored.
(7) Instrument Error. See paragraph 302a(3)(b).
(8) Dynamic Effects on Airspeed Indicator. The dynamic effects on airspeed indicators occur as a result of acceleration and rapid change in airspeed during takeoff. This causes many airspeed indicators to indicate an airspeed lower than the actual airspeed.
Note: It is possible for electronic airspeed indicators driven by an air data computer to also have errors due to dynamic acceleration effects because of characteristics inherent in the basic design.
(9) Distance Measuring Unit. An instrumentation system normally used to record takeoff distance measurements. One output of these systems provides the ground speed versus time as the airplane accelerates during the accelerate-takeoff ground run. Ground speed may be converted into a corresponding CAS value by applying wind and air density corrections at intervals during acceleration where the ship's airspeed indications have been recorded.
b. Trapped Static Source Method. The trapped static source method consists of comparing instantaneous readings of airspeed, as indicated on a test airspeed indicator, with readings on a production airspeed indicator while accelerating on each other. Readings may be recorded by film or video cameras for mechanical airspeed indicators or by electronic means if a transducer type device is being utilized.
See Figure A9-7 for system schematic.
(1) Test Conditions.
(a) Air Quality. The surface winds should be light with a minimum of gusting.
(b) Weight and C.G. Ground run calibrations are not sens itive to c.g. The dynamic effects of acceleration may be affected by weight. Test weight variations should be sufficient to account for any measurable effects due to weight.
(c) Speed Range. The speeds should range from 0.8 of the minimum V1 to 1.2 times the maximum V1, unless higher values up to VR are required for expansion of takeoff data.
A9-19
Appendix 9
(d) Configuration. The airspeed system should be calibrated during the accelerate-takeoff ground run for each approved takeoff flap setting.
(2) Test Procedures.
(a) Align the airplane with the runway.
(b) With idle engine power, and with the cabin door open, open the valve to expose the bottle to static ambient conditions, then close the valve. Record the test altimeter reading.
(c) Close the cabin door.
(d) Conduct a takeoff acceleration using normal takeoff procedures. The camera should be recording speeds from the two airspeed indicators in increments sufficient to cover the required airspeed range.
A9-20
Appendix 9 Figure A9-7 - Trapped Static Source Schematic
TO SHIP'S TO SHIP'S
PITOT PITOT
SOURCE SOURCE
PRODUCTION TEST AIRSPEED
SHIP'S INDICATOR
AIRSPEED (OR
INDICATOR ELECTRONIC DEVICE, I.E.
TRANSDUCER) TEST
REFERENCE
ALTIMETER
PRODUCTION OPENABLE OPEN TO
RATE-OF- VALVE AMBIENT STATIC
CLIMB AIR CONDITIONS
TO SHIP'S
STATIC
SOURCE
(e) The takeoff run should be continued, if possible, until beyond the maximum required speed then aborted. When at rest with engi nes idling, open valve again and observe the test altimeter. Any significant jumps or changes in indicated altitude may indicate a system leak, too much runway gradient or other factors WHICH WILL INVALIDATE THE RESULTS OF THE RUN.
(f) Repeat steps (a) through (e) of this paragraph until there are sufficient runs to provide adequate calibration curves for the required configurations.
A9-21
Tfi)IE TRAP PED
(1) TS CORREC TED SHIP 1SIAS
(l)SHIP 1S
CORRECTED
(2) AIRS PEED STATIC (TS) IAS AIRSP EED TSIA S (KN OTS)
AIRSPEED SHIP 'SIAS ERRO R (KNOTS) INS
TRUMENT INST RUM ENT
CORRECTION
CORRECTIO N 7:41 :45
50.7 0
50.7 49 0 49
1.7 :46
56.1 I
56.1 54 I 54 +2.1 :47
61.4 I
61.4 61 I 61 +o.4 :48
66.9 I
66.9 66 I 66 +-0.9 :49
71.9 I 7l.9 72 I 72 -0.1 50
76.7 I
76.7 77 I 77 -0.3 :51
82.1 I
82.1 83 I 83 -0.9 :52
86.8 I
86.8 88 I 88 -l.2 :53
91.5 I
91.5 91 I 91 +o.5 :54
96.5 I
96.5 99 I 99 -2.5 :55
100.9 I I00.9 102 I !02 -I.I :56 !05.2 I !05.2 !07 I 107 -1.8 :57
110.1 I IIO.l 113 I IIJ -2.9 :58
114.4 I
114.4 119 I ll9 -4.6 :59
118.2 I
118.2 123 I 123 -4.8 7:42 :00
122.9 V
122.9 128 V 128 -5.1 11/16/2011 AC 23-8C Appendix 9
(3) Data Acquisition and Reduction. Read the recorded data (film or video) at increments of airspeed arbitrarily selected within the required range. See Table A9-2 for a sample data reduction. Record and perform the following:
Table A9-2 - Trapped Static (TS) Data Reduction NOTES: 1. Obtain from instrument calibration.
2.
Corrected trapped static IAS minus corrected ship's IAS.
3.
Corrections are added.
A9-22
Appendix 9
(a) Production indicated airspeed, test indicated airspeed, and configuration.
(b) Correct the test indicated airspeed for instrument error and, in the case of electronic devices, any known dynamic effects. Static pressure in the bottle is assumed to result in no position error. These corrected airspeed values may be assumed to be CAS.
(c) Calculate the amount of system error (difference between corrected trapped static indicated airspeed and production indicated airspeed).
(d) Plot IAS versus CAS within the required range of speeds. See Figure A9-8 for a sample plot.
A9-23
lJ' V
V ,,, ,,, / I/ V V
V , ,,, / I/ V V , , ,, / 1,,1 V / , ,,, ,,, / I/ V / I' ,,, ,,, 11/16/2011 AC 23-8C Appendix 9 Figure A9-8 - Ground Airspeed Calibration 110
GROUND AIRSPEED
CALIBRATION 100 90 CAS KNOTS 80 70 60 50 40 40 50 60 70 80 90 100 110 120
IAS - KNOTS A9-24
Appendix 9
c. Distance Measuring Unit Method. The distance measuring unit method consists of utilizing the readouts of ground speed to obtain CAS values within the required range of speeds. These values are compared with readings at the same instant on a production airspeed indicator. Airspeed indicator readings may be recorded by film or video cameras for mechanical airspeed indicators or by electronic means if a transducer type device is being utilized. There should be a method of correlating recorded airspeeds with the CAS values obtained from the distance measuring unit system.
(1) Test Conditions.
(a) Air Quality. The surface wind velocity should be steady, as low as possible, and not exceed 10 knots. The wind direction should be as near as possible to the runway heading.
(b) Weight and C.G. Same as for the trapped static source method.
(c) Speed Range. Same as for the trapped static source method.
(2) Test Procedures.
(a) Align the airplane with the runway.
(b) Conduct a takeoff acceleration using normal takeoff procedures. The distance measuring unit should be recording/determining the ground speeds. The camera should be recording speeds from the production airspeed indicator and the time or counting device utilized to correlate speeds.
(c) The takeoff may continue or be aborted when beyond the maximum required speed.
(d) Record surface wind velocity and direction; surface air temperature and runway pressure altitude for each run.
(e) Repeat steps (a) through (d) of this paragraph until there are sufficient runs to provide adequate calibration curves for the required configurations.
(3) Data Acquisition and Reduction. Read the recorded data (film or video) at increments of airspeed arbitrarily selected within the required range. For these same increments, determine the ground speeds from the distance measuring unit system. See Table A9-3 for a sample data reduction. Record and perform the following:
A9-25
<--, TIME DMl WIND TAS (I) SHIP'S (2) SHIP'S CORRECTED (3) GROUND SPEED COMPONENT (KNOTS) CAS !AS AIRSPEED SHW'SIAS GROUND (KNOTS) DOWN THE (Ki'lOTS) (KNOTS) INSTRUMENT AIRSPEED
RUNWAY CORRECTION ERROR 07:00 :09 48.0 3 51.0 50.1 49 0 49 +LI :IO 52.8 I 55.8 54.8 54 I 54 +0.8 :II 56.8 I 59.8 58.7 59 I 59 -0.3 :12 61.0 I 64.0 62.8 63 I 63 -0.2 :13 64.2 I 67.2 66.0 68 I 68 -2.0 :14 67.3 I 70.3 69.0 71 I 71 -2.0 :15 70.9 I 73.9 72.5 75 I 75 -2.5 :16 74.0 I 77.0 75.6 78 I 78 -2.4 :17 77.2 I 80.2 78.7 82 I 82 .3.3 :18 80.7 I 83.7 82.2 83 I 83 -0.8 :19 83.9 I 86.9 85.3 87 I 87 .1.7 :20 87.0 I 90.0 88.3 89 I 89 -0.7 :21 90.6 I 93.6 91.9 92 I 92 -0.1 :22 93.8 I 96.8 95.1 95 I 95 +0.1 :23 96.9 I 99.9 98.1 101 I IOI -2.9 :24 100.3 I 103.3 101.4 103 I 103 -1.6 <--, :25 103.6 I 106.6 104.7 106 I 106 .IJ :26 106.6 V 109.6 107.6 110 V I IO -2.4 11/16/2011 AC 23-8CAppendix 9 Table A9-3 - Sample Ground Airspeed Calibration Using a Distance Measuring Unit Test Conditions:
Notes:
Pressure Altitude - 1240 ft. 1. CAS = TAS Temperature - 52 °F 2. Obtain from instrument calibration - 0.982 3. CAS minimum corrected Ship's IAS Runway - 1 4. Corrections must be added Wind - 350/3 A9-26
Appendix 9
(a) Production indicated airspeed, ground speed, surface air temperature, runway pressure altitude, wind velocity and wind direction with respect to runway heading.
(b) Compute a CAS value for each data point. This is accomplished by identifying the wind component parallel to the runway; computing the corresponding true airspeed; computing the air density ratio; then computing the calibrated airspeed.
(c) Calculate the amount of system error (difference between CAS and production indicated airspeed).
(d) Plot IAS versus CAS within the required range of speeds. See Figure A9-8 for a sample plot.
A9-27