Appendix 9
a. Test Conditions.
(1) Air Quality. Smooth, stable air is needed for calibrating the airspeed indicating system using a trailing bomb or trailing cone.
(2) Weight and C.G. Same as speed course method. See paragraph 2a(2) of this appendix.
(3) Speed Range. The calibration should range from 1.2 V STALL to VMO/VNE or maximum level flight speed, whichever is greater. If the trailing bomb becomes unstable at high airspeed, the higher airspeed range may be calibrated using another accepted method; that is, trailing cone or pace method.
(4) Use of Bomb. Care should be exercised in deploying the bomb and flying the test to ensure that no structural damage or control interference is caused by the bomb or the cable.
At higher speeds, the bomb may become unstable and porpoise or oscillate. A means for a quick release of the trailing bomb should be provided, in the event an emergency arises.
Flight tests using a bomb should be conducted over open (unpopulated) areas.
(5) Free Stream Air. The bomb hose should be of adequa te length to assure bomb operations in free stream air. This should include consideration of all airplane test configurations that could possibly impart body interference upon the bomb. It will usually require that the bomb be at least one-half wing span away from the airplane.
(6) Qualifications For Use. Under stabilized flight conditions at constant airspeed and altitude, trailing cones and airspeed bombs are considered excellent airspeed reference systems. See paragraph 17b of this AC for additional discussion.
b. Test Procedures.
(1) Stabilize airplane in level flight approximately 30 seconds just above stall with flaps and gear retracted. Record data.
(2) Repeat step (1) at sufficient increments to provide an adequate calibration curve for each of the configurations.
c. Data Acquisition. Data to be recorded at each test point:
Altimeter Method Pressure Differential Method
1. Airplane Airspeed (V i) 1. Airplane Airspeed (V i)
2. Airplane indicated altitude (H iA/C) 2. Airplane indicated altitude (H iA/C)
3. Trailing Cone/Bomb altitude (H iref) 3. Pressure Differential P P PS Sref SA C /
4. Flap position 4. Flap position
5. Landing gear position 5. Landing gear position
6. Fuel used 6. Fuel used A9-8
Appendix 9
d. Data Analysis. The data are analyzed according to the methods and equations presented above. The data could be presented in the form as shown in Figures A9-3 and A9-4. Data that fall outside the FAR/JAR limits fail the airworthiness codes.
2. Speed Course Method. The speed course method consists of using a ground reference to determine variations between indicated airspeed and ground speed of the airplane. An accurately measured ground course is required. The course distance should be se lected to be compatible with the airspeeds being flown. Excessively long times to traverse the course will degrade the test results.
Generally, airspeeds above 250 knots should be flown over a 5-mile course. Below 100 knots, limit the course to 1 mile. Perpendicular "end lines" (roads, powerlines, etcetera) should be long enough to allow for drift and accurate sighting of end line passage. A one-second error at 200k is 6k on a 2-mile course.
a. Test Conditions.
(1) Air Quality. The air should be as smooth as possible with a minimum of turbulence and wind. The wind velocity, while conducting the test, should not exceed approximately 10 knots.
(2) Weight and C.G. Airspeed calibrations are usually not c.g. sensitive but may be weight sensitive, especially at low airspeeds (higher angles-of-attack). Initial airspeed calibration tests should be conducted with the airplane loaded at or near maximum takeoff gross weight.
Additional tests should be conducted at near minimum weight and at low airspeeds to spot check the maximum weight airspeed calibration results. If differences exist, an airspeed system calibration should be accomplished at minimum weight.
(3) Altitude. When using a visual reference on the airplane for timing, the altitude throughout the test run should be as low as practical but should be maintained at least one and one-half wing spans above the highest ground elevation so that the airplane remains out of ground effect. When conditions permit using the airplane shadow for timing, speed course altitudes of 500-2000 feet AGL can be used. All run pairs should be conducted at the same altitude.
(4) Speed Range. The speed should range from 1.3 V S1 to the maximum level flight speed, to extrapolate to VD.
(5) Run Direction. Reciprocal runs should be made at each speed to eliminate wind effects, and the ground speed obtained in each direction should be averaged to eliminate wind effects.
Do not average the time flown in each direction.
(6) Heading. The heading should be maintained constant and parallel to the speed course throughout the run, allowing the airplane to "drift," if necessary, so that the effect of crosswinds can be eliminated.
(7) Configuration. The airspeed system should be calibrated in each landing gear and wing flap configuration required in §§ 23.45 through 23.77. This normally consists of gear up/flaps up, gear up/flaps takeoff and gear down/flaps down.
A9-9
Appendix 9
b. Test Procedures.
(1) Stabilize airplane in level flight at test speed, with gear and flaps in the desired configuration, prior to entering the speed course.
(2) Maintain constant speed, altitude, and heading through speed course. Record data.
(3) Repeat steps (1) and (2) of this paragraph on the reciprocal speed run.
(4) Repeat steps (1) through (3) of this paragraph at sufficient increments (minimum of five) to provide an adequate calibration curve for each of the configurations.
c. Data Acquisition and Reduction. Data to be recorded during each run:
(1) Time to make run;
(2) Pressure altitude;
(3) Total air temperature (airplane indicator) corrected to static air temperature (SAT);
(4) Indicated airspeed;
(5) Wing flap position;
(6) Landing gear position; and
(7) Direction of run.
d. Sample Speed Course Data Reduction (Refer to Table A9-1).
DistanceSpeed = Time 1 Knot 6076.1 feet/nautical mile 1.6878 feet/second = = 3600 second/hour Ground Speed 10560 .5925 (10560) 132.8 knots = = =
(1.6878)(47.1) 47.1 G S AVG (TAS) = 132.8 + 125.6 = 129.2 knots 2 Sample Speed Course Data and Data Reduction.
a . W e i g h t C.G.
b. Course Distance 10,560 Ft.
c. Pressure Altitude 1,600 Ft. (Altimeter set to 1013 m.b.)
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J-- I I I I 11/16/2011 AC 23-8C Appendix 9 Flap Position Degrees Gear Position (Up or Down) Time Seconds I.A.S.
(Knots) Pressure Altitude Feet SAT.
Degrees Fahrenheit Ground Speed* Knots Average Ground Speed Knots Factor** Calibrated Air Speed Knots Average I.A.S.
Knots Airspeed System Instrument Position 0 Fixed 47.1 128.0 1610 55 132.8 129.2 0.975 126.0 128.5 2.5 1 1.5
49.8 129.0 1600 55 125.6
44.5 135.0 1600 55 140.5 136.7 0.975 133.3 136 2.7 0 2.7
47.1 137.0 1600 55 132.8
40.5 148.0 1600 55 154.2 149.3 0.975 145.6 148 2.4 -1 3.4
43.3 148.0 1600 55 144.3 C Course Distance (Ft)* Ground Speed = Time (Seconds) C = 0.5925 for course speed in kts or use: C = 0.6818 for M.P.H.
** Factor = chart) F from read (oreTemperaturObserved 7 . 559Hg.) (In. PressureObserved16 . 4 / O o Table A9-1 - Sample Speed Course Data and Data Reduction
(1) Density Altitude. TAS is greater than CAS if density altitude is above sea level. For density altitudes below 5000 feet and calibrated airspeeds below 200 knots, it is considered acceptable to use the term CAS = EAS = TAS o ./ In this case, density altitude is obtained from figure A7-4 in Appendix 7. At 1600 ft. pressure altitude and SAT 55 °F, we read a density altitude of about 1700 feet. This density altitude intercepts o at a value/ of .975. CAS = 129.2 (0.975) = 126.0 knots.
AVERAGE
GS ERROR (TAS) CAS IAS System = Instrument + Position
129.2 126.0 128.5 +2.5 +1 +1.5 (CAS-IAS) V inst Vpos
(2) Required Accuracy. Instrument error is determined by applying standard pitot and static pressures to the airspeed instrument and developing a calibration curve. IAS corrected for +1 instrument error = 127.5 knots. The position of the static source is causing +1.5 error.
A9-11
Appendix 9 Section 23.1323(b) requires the system error, including position error, but excluding instrument error, not to exceed 3 percent of CAS or 5 knots, whichever is greater, in the designated speed range.
(3) Compressibility. For many years CAS was used for design airspeeds. However, as speeds and altitudes increased, a compressibility correction became necessary because airflow produces a total pressure on the pitot head that is greater than if the flow were incompressible. We now use EAS as a basis for design airspeeds (§ 23.235). Values of CAS versus EAS may be calculated or you may use the chart in Appendix 7, figure A7-5, to convert knots CAS to EAS.
3. GPS Method. The Global Positioning System (GPS) method consists of using a GPS unit to determine ground speed, which is then used to calculate true airspeed. Any commercial GPS unit can be used that produces consistent results. Once true airspeed is calculated, the data reduction is nearly identical to the speed course method described previously. One difference is that the scale altitude correction factor (ΔVc the difference between CAS and EAS as shown in figure A7-5) may be significant at higher altitudes and speeds that may be flown with this method. Specifically, you will solve the following equation for (ΔVpec):
Vi + ΔVic + ΔVpec + ΔVc = Vtrue* And then, assuming that all of the pitot-static error is in the static port, you may calculate the altitude position, error, ΔHpec, as described in the Trailing Bomb/Cone Method.
a. Test Conditions.
(1) Air Quality. The air should be as smooth as possible with a minimum of turbulence. The wind velocity and direction must be constant for this method to give correct results.
(2) Weight and C.G. Same as the speed course method.
(3) Altitude. The altitude is not critical, but it should be chosen where the air is smooth and the winds are constant.
(4) Speed Range. Any speed at which the aircraft can be stabilized in level flight (see Figure A9
2).
(5) Runs . Three runs per airspeed are required to calculate one true airspeed. The three runs must be done at the same indicated speed and altitude on different headings. The headings should be 60 to 120 degrees apart. _____
(6) Configuration. Same as the speed course method.
A9-12
Appendix 9
b. Test Procedures.
(1) Stabilize the airplane in steady level flight at the desired test speed configuration. Record the indicated airspeed, pressure altitude (29.92 set), outside air temperature and configuration of the aircraft.
(a) Record both ground track and ground speed from the GPS unit once the values are stable (this can take up to 10 seconds after stabilizing).
(b) Turn 60 to 120 degrees either direction and record the new ground track and speed once restabilized at the same airspeed and altitude on the new heading. Minor variations in altitude (up to 100 feet) are much preferred to any variation in airspeed from the initial value. A one knot change in indicated airspeed will cause at least a one knot change in true airspeed, but 100 feet of altitude only causes on the order of 1/2 of 1 percent change in the density ratio, σ.
(c) Turn another 60 to 120 degrees in the same direction and record a third set of ground track and speeds.
(2) Once you have three sets of ground track and speed for a given indicated airspeed and configuration, repeat steps (1)(a) through (1)(c) above at sufficient increments, to provide an adequate calibration curve for each of the configurations.
A9-13
Appendix 9
c. Data Reduction. The best way to calculate true airspeed from the three sets of ground tracks and speeds is with a personal computer spreadsheet. The following solution was developed assuming that the three legs flown had the same true airspeed (indicated airspeed, OAT, and pressure altitude were identical) and that the wind did not change during the three legs. The table shows the spreadsheet equations for one popular spreadsheet program that will solve the problem. Note that wind speed and direction are intermediate outputs. If a series of points are done at nearly the same time, altitude and geographic location, then the consistency of the calculated wind speed and direction will be an indicator of the validity and accuracy of the calculated true airspeeds.
A B Result 1 Ground Speed 1 184 184 2 Track 1 265 265 3 Ground Speed 2 178 178 4 Track 2 178 178 5 Ground Speed 3 185 185 6 Track 3 82 82 7 X1 =B1*SIN(PI()*(360-B2)/180) 183.3 8 Y1 =B1*COS(PI()*(360-B2)/180) -16.0 9 X2 =B3*SIN(PI()*(360-B4)/180) -6.2 10 Y2 =B3*COS(PI()*360-B4)/180) -177.9 11 X3 =B5*SIN(PI()*(360-B6)/180) -183.2 12 Y3 =B5*COS(PI()*(360-B6)/180) 25.7 13 M1 =-1*(B9-B7)/(B10-B8) -1.17 14 B1 =(B8+B10)/2-B13*(B7+B9)/2 6.71 15 M2 =-1*(B11-B7)/(B12-B8) 8.77 16 B2 =(B8+B12)/2-B15*(B7+B11)/2 4.42 17 Wx =(B14-B16)/(B15-B13) 0.2 18 Wy =B13*B17+B14 6.4 19 Wind Speed =SQRT(B17^2+B18^2) 6.4 20 Wind Direction =MOD(540-(180/PI()*ATAN2(B18,B17)),360) 177.9 21 True Airspeed =SQRT((B7-B17)^2+(B8-B18)^2) 184.4
4. Pace Airplane Method. An airplane whose pitot static systems have been calibrated by an acceptable flight test method is used to calibrate the pitot static systems of a test aircraft.
a. Test Conditions. Smooth ambient flight conditions
b. Test Procedures. The pace airplane is flown in formation with the test airplane at the same altitude and speed. The aircraft must be close enough to ensure that the relative velocity is zero yet far enough away so that the pressure fields of the two airplanes do not interact. Readings are coordinated by radio.
c. Data to be recorded
(1) Test Airplane airspeed (ViT) kts
(2) Test Airplane Pressure Altitude (HiT) ft A9-14
Appendix 9
(3) Pace Airplane airspeed (Vip) kts
(4) Pace Airplane Pressure Altitude (Hip) ft
(5) Configuration for both airplanes
(6) Fuel used in both airplanes
d. Data Reduction. Correct all the instrument readings for instrument errors and the pace aircraft readings for the known position error.
VpecT (Vip Vicp Vpec ) (ViT VicT )kts H pecT (Hip H icp H pec ) (HiT HicT ) ft Calculate VpecT and H pecT for all data points in each configuration and plot in a manner similar to Figure A9-4 and Figure A9-5.
5. Pitot-Static Boom Data. If a flight test Pitot-Static boom is mounted on an airplane such that the pitot tube (total head) is not affected by flow angularity and the static source is outside the pressure field of the aircraft, then it can be assumed that the boom data is without position errors. The boom data can then be taken as the pace data. The Pitot-Static boom should be calibrated for high alpha and high speed.
a. Test Conditions. Smooth ambient flight conditions
b. Test Procedures. The nose or wing-mounted boom is flown in formation with the test airplane at the same altitude and speed. The aircraft must be close enough to ensure that the relative velocity is zero yet far enough away so that the pressure fields of the two airplanes do not interact. Readings are coordinated by radio.
c. Data to be recorded
(1) Test Airplane airspeed (ViT) kts
(2) Test Airplane Pressure Altitude (HiT) ft
(3) Nose or wing-mounted boom airspeed (Vip) kts
(4) Nose or wing-mounted boom Pressure Altitude (Hip) ft
(5) Configuration for both airplanes
(6) Fuel used in both airplanes A9-15
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.
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