TEST OF TRAIL CONE SYSTEM TO CALIBRATE STATIC ...The cone stabilized static source is an outgrowth...

11epe4rt No. 11-64-156

FINAL REPORTPreject No. 320.205-02X

TEST OF TRAIL CONE SYSTEMTO CALIBRATE STATIC PORTSFOR BAROMETRIC ALTIMETERS

HAT COeYyMICROFICHE $. 6. P

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DDC

DECEMBER 1964 FEB 18 1965

DDC.IRA E

FEDERAL AVIATION AGENCY REPRODUCED FROM

Systems Research & Development Service BEST AVAILABLE Copy

Atlantic City, New Jersey

______C Do______

FINAL REPORT

TEST OF TRAIL CONE SYSTEM TO CALIBRATESTATIC PORTS FOR BAROMETRIC ALTIMETERS

PROJECT NO. 320-205-02X

REPORT NO. RD-64-156

Prepared by:

Jack J. Shrager

December 1964

This report has been approved for general availability. It doesnot necessarily reflect FAA policy in all respects and it doesnot, in itself, constitute a standard, specification or regulation.

Joep Blatt

Direct , Systems ResearchDevelopment Set-rice

Federal Aviation Agency

Experimentation DivisionNational Aviation Facilities Experimental Center

Atlantic City, New Jersey

REPRODUCED FROMBEST AVAILABLE COPY

Page

ABS ACT ........................................... iv

I M ODU/CTION ...................................... I

TRAIIM(G CONE SYSTE4 DESCRIPTION .................. 2

TEST PROCEDURES ................................... 9

Phototheodolite/Atmospheric Observation ...... 9Vertical Aerial Camra/Anosup'.eric Observation 13.1Tower T2y-P ................................. 14Radar fracking/Pressure Survey ................ 18Radar Phototheodolite/Radiosonde .......... 1Pa.cer/Ch-as. ie .. ........................... . 18Temperature Profile Survey ..... ....... 3.8

w-2 Pcer Aoircaft ................. . .18

F80 Pacer Aircraft ........................... 23Ground Base Eq•Lui•xnt ................... .6........ 25

MM.OD OF ANALYSIS ................................ 29

TEST ESULTS AND ANALYSIS ......................... 39

CONCLUSIONS ....................................... 55

RBCONIMATIONS. ................................... 55

LIST OF SYMB OLS ................................... 56

R CES..... .......................................... 5T

LTST OF TTJSM'TRATTONS

Figure Page

1 British High-Speed Pitot Static Tube ............ 3

2 Aircraft Trailing High-Speed Tube ............... 4

3 Douglas Cone Assembly ........................... 5

4 NAFEC Cone Assembly ............................. 6

5 Reel Assembly .................................. 7T

6 Cone in Flight Behind Aircraft .................. 8

7 Drag Curve for Standard Cone ................... 10

8 Aircraft Flight Path - Field Layout ............. 12

9 Aircraft Flight Path Over Vertical Camera ....... 13

10 Aircraft Flight Path for Tower Fly-By .......... 15

11 Pressure Survey at Altitude .................... 16

12 Aircraft Overtake Method ....................... 18

13 Flight Pattern at Varying Airspeeds and ConstantAltitude ....................................... 19.

14 TV-2 Test Aircraft .... ......................... 20

15 Photopanel of TV-2 ...................... 22

16 F8C Test Aircraft ..... ...................... 24

17 Photopanel of F8C ..... . ..................... 26

18 Phototheodolite System ......................... 27

19 Aerial Camera ......... ........................ 28

20 X-Band Radar (TAR ...... ...................... 30

ii

LIST OF ILLUSTRATIONS (contlnued)

Figure Page

21 Radiosonde ...................................... 31

22 Barometer ........ ............................... 32

23 Psychron Unit ................................... 33

24 Position Error Versus Indicated Airspeed -TV-2 Pacer (Low Altitude) All Data ............ 42

25 Position Error Versus Indicated Airspeed -

TV-2 Pacer (Low Altitude) Selected Data ... . 43


F80 Pacc.r (Tow Alt1tude) All Data ............. 44

2T Position Error Versus Indicated Airspeed -

F8C Pacer (Low Altitude) Selected Data ........ 45


F8C Pacer (High Altitude) All Data ............ 46


F8C (High Altitude) Selected Data ............. 48


All Selected Data ............................. 49

31 A P/qC Versus Mn (TV-2 Pacer) .................. 50

32 A P/qC Versus Mn-F8C Pacer (Low Altitude) ...... 51

33 AP/q0 Versus Mn-F8C Pacer (High Altitude) ..... 52

34 A P/qc Versus Mn-All Data ...................... 53

35 A H Versus IAS-All Data ........................ 54

iii

Federal Aviation Agency, Systems Research and Development Service,National Aviation Facilities Experimental Center, ExperimentationDivision, Atlatitlc City, N. J.TES OF TRAIL CONE SYSTEM TO CALIBRATE STATIC PORTS FOR BAROMETRICALTIMETERS by Jack J. Shrager, Final Report, December 1964, 59 PP.incl. 35 illus.(Project No. 320-205-02X, Report No. RD-64-156)

ABSTRACT

An experimental trail cone stabilized static source system wasfabricated and tested using various towing aircraft at speeds up to1.12 Mach Number to determine its suitability as a "standard" methodfor determining altimeter static pressure error. Flight tests wereconducted at sea level, 5,000, 10,000, 15,000, 25,000, 30,000 and4 0,000 feet, to determine the static pressure defect of the trailingcone system. Test results indicated that the experimental systemhad a nominal "position error" equal to + 12 feet at sea level and+ 30 feet at 30,OO0 feet. -

IV

INTRODUCTION

The pilot of an aircraft determines his flight altitude byterring to the aircraft's barometric pressure altimeter. Sinceis instrument' a presentation is a function of the "sensed" ambientessure, it is necessary to know the difference between sensed andae ambient pressure to determine actual pressure altitude. Thisfference, generally referred to as "position error," is affectedmany factors including (a) type of the aircraft's static pressure

obe, (b) location of the static pressure probe, (c) angle of attackthe aircraft, and (d) flight speed of the aircraft as related toe velocity of sound, Mach Nwumber (M).

There are several methods currently employed to determine thesition error. The magnitude of the position error at flighttitudes below 1000 feet up to the limit of the operationalaracteristics of the aircraft can readily be determined. Theree several theoretical proced,&res for extrapolating low-altitudesults to cover the entire operational envelope of the aircraftt even these procedures include certain assumptions not readilyrified.

The purpose of this Federal Aviation Agency's (FAA) Researchd Development Project was to develop, if possible, or define anceptable standard for determining the position error. This methods to be usable for general aviation, comnercial carrier aircraft,d military aircraft at all applicable altitudes and throughout1 flight speeds including subsonic, transonic, and supersoniclocities.

A secondary, though not necessarily restrictive, considerations to evolve an economically feasible system which could be readilyplemented by the aviation industry. This secondary considerations not to be restrictive on the desired attainable accuracy whichs to be that within the capability of the best known selectedight instruments with special precision calibration.

A review of all known published information, Reference A towas made to determine the theoretical requirements and the

ssible design specifications of a system which would meet thesired requirements. This report covers the test procedures andsults for such a proposed system; namely, a trailing cone stabilizedatic pressure source. This system is based on a refinement oftechnique originally suggested by the Royal Aeronautical Society,ference AH, modified by the Douglas Aircraft Ccmpany, Referenceand further modified by the FAA to meet the requirements noted

ove.

The results contained in this report include those obtained atthe FAA's National Aviation Facility Experimental Center (NAYF=),and the Naval Air Test Center (NATC) in cooperation with theBureau of Weapons, U. S. Navy.

TPAMll OONE SYSTD DESCIPTION

The cone stabilized static source is an outgrowth of a RoyalAircraft Establishment method using a high-speed Pitot static tube,Fig. 1, which was attached to a flexible, hollow molybdenum steeltube that trailed behind and below the aircraft under test, asshown in Fig. 2.

The Douglas Aircraft Company's version of this system utilizeda fiberglas cone to stabilize the tube assembly with the staticholes drilled in the tubing itself. The final Douglas version usednylon tubing in place of the flexible steel tubing, Fig. 3.

The FAA's experimental system built by Douglas Aircraft Companyto FAA specifications utilized a stainless steel insert containingthe static ports approximately 10 feet forward of the fiberglas cone.A stainless steel wire was inserted throughout the length of theassembly to carry the drag force, thus preventing separation of thenylon tubing from the metal insert. The system was later modifiedto permit free rotation of the cone without introducing a torsionalmoment to the wire insert. The system in its finas configuration,as modified by NAFEC, is shown in Fig. 4.

All of the above noted systems utilized a reel and drive motorassembly which was installed in the test aircraft to extend andretract the drag body and hollow tube behind the aircraft. Atypical reel assembly is presented in Fig. 5.

All systems in which the fiberglas cone was employed resultedin the tube assembly trailing in line with the attachment point ofthe aircraft, as shown in Fig. 6, instead of below the aircraft,Fig. 2, with the high-speed Pitot tube.

The total flow area of the static ports in the steel tubeinsert and in the drive reel mechanism was equal to or greaterthan the flow area of the plastic tube with the support wireinserted. The holes were located in a manner which wouldminimize any angle-of-attack effect of the steel tube insert.

2

.STATI cTUB F

PLAN VIEW

TOWING PIPE

FAIRING FOR PIPE FREE TO C.G. OF HlEAD VERY NEARROTATE ABOUT TOWING AXIS• TO TOWING AXIS

WING CHORD APPROXIMATELY3u NOSE DOWN RELATIVE TO

XIS OF HEA.D\ TAIL BOOM INCIDENCEADJUSTABLE ABOUTT2-HIS AXIS

SIDE ELEVATION

PIG. 1 BRITISH 1IIGII-SPEED PITOT STATIC TUBE

3

TOWING CABLE

TRAILING HEAD

FIG. 2 AIRCRAFT TRAILING :IuGH-SPEED TUBE

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a. DOUGLAS CON

4. ~a DOG a NILN 1ohi

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FIG. 3 DOUGLAS CONE ASSEMBLY

a. NAFEC CONE

24 at- tol"~ &M~IN

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11411C KRAM .14 RIA

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c OVERALL ASSEM4BLY

F'IG. 4 NAFEC CONE ASSDIBLY

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The drag forces of the cone assembly, Fig. 7, were so greatwitn respect to the mass of thne complete tuDing ana cone assemoiythat the change in the angle of attack of the tube could be con-sidered to be ccnstant and approaching zero. This can be seenin the sample problems below at a nominal sea-level velocity of16C knots, 0.24 Mn (q. = 0.940 inch Hg.) and 450 knots,0.75 Mn (qc= 8.00 inches Hg.)

4D

W=cone assembly weight=0.3 pounds W=cone assembly weight=0.3 pounds

D-=Drag Force at 160 knots=17 pounds D=Dreg Force at 450 knots=148 pounds

tan C = 0.3 = 0.01765 tan C = 0.3 = 0.002031T

S= 1.('° C =0.120

The system may be considered as the rigid nose boom of thefiberglas cone with an angle of attack equal to the angle of attackof the symmetrical conical body. The static ports, located forwardof the cone by approximately 10 times the effective body size,should result in a pressure ratio error approaching zero, up through1.02 Mn as noted in References L and AA.

TEST PROCEDJRES

The test procedures employed to determine the position errorof the trailing cone system were, as listed below, based on thephysical resources available or their limitation under the conditionsas employed. For report reference purposes only, they are furtheridentified by sequential Roman numerals.

I. Phototheodolite/Atmospheric Observation

Prior to takeoff, all instruments both airborne and ground-located were read and recorded. When available, a radiosondicsurvey of the atmosphere up to the maximum test altitude was alsomade.

9

110

100

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0 70

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60

50

t 0 40 ;_ __ _

30 o

20

0 2 3 4 6

DYNAMIC PRESSURE (qc) AT CENTER OF WIND TUNNELC

IN INCHES OF MERCURY

FIG. 7 DRAG CURVE FOR STANDARD CONE

3 10

After takeoff, the aircraft was flown in non acceleratinglevel flight at both 0.55 Mn and the minimum indicated airspeedapplicable for maximum aross weight in the airnraft's nruiiseconfiguration. Each test point, which was repeated twice, wasmade at an altitude of less than 500 feet above the nominalrunway elevation.

At the completion of these reference speed points, the aircraftwas flown in non accelerating level flight at various airspeedsfrjm minimum controllable to maximum continuous or maximum allowable,whichever was greater, at low altitude (less than 500 feet abovemean aiort elevation), 5,000- and 10,000- foot altitudes in closeproximilit: vu NAFEC. Both airborne and ground recordings were takenat each te: t point from runway threshold to the end of the runway.The aircraft's fLight path, Fig. 8, was designed to permit com-putation of aiL peed from ground tracking information and to allowfor stabilization of the aircraft prior to crossing the runwaythreshold.

The last data points for any flight were a repetition of thereference speed points, 0.55 Mn and minimum indicated airspeedcruise configuration. The aircraft would then land, taxi to thepoint where the original ground readings were made, and allinstruments would be read and recorded. In addition, a radiosondicsurvey of the atmosphere would be repeated.

II. Vertical Aerial Camera/Atmospheric Observation

The test procedure employed was similar to that described inTest Procedure I with the following noted exceptions:

The aircraft's flight path, Fig. 9, was directed over ahigh-precision aerial camera. At the instant the aircraft wasdirectly overhead, a photograph of the aircraft was taken. Thisprocedure, which is a modification of the tower fly-by method, isoutlined in detail in Reference AC.

This method as well as the tower fly-by method describedin Test Procedure III is used only for tests at low altitude (lessthan 19000 feet above camera elevation).

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III. Tower Fly-By

The test procedure was similar t) those described in TestProcedures I and II with the following exceptions:

J2he aixrcrafts fiignc patn, i'tg. iu, was over aprescribed ground track at an altitude which by visual referenceappeared to the pilot to be that equal tn the tower helight. Anobserver in the tower would note and record the aneular displace-ment of the aircraft with respect to a scribed scale or referencemark at the instant it passed the tower.

TV. Radar Tracking/Pressure Survey

The test procedures prior to takeoff, at low-level referencespeed following takeoff, just prior to landing, and after landingwere the same as those described in Test Procedures I and II.

Following the initial low-level flights, the aircraftascended directly to the scheduled test altitude. The aircraftwas flown at a constant speed in straight and level flight, anda pressure survey of the test area, as shown in Fig. 11, was madebased on flying a constant indicated altitude while being trackedby a precision radar system.

The aircraft was then flown at the same indicated testaltitude through its entire speed range using the flight patternshown in Fig. 11 to facilitate the computation of airspeed.

Precision radar was used to determine the absolutegeometric height of the aircraft during the pressure survey andall altitude test points.

V. Radar Phototheodolite/Radiosonde

This method is similar to that described in Test Procedure IVwith the exception that the ambient pressure at altitude was derivedfrom atmospheric information telemetrLcally obtained from a precisionradar-tracked, specially calibrated radiosonde.

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VI. Pacer/Chase

The test aircraft, after the ground and low-altitude refererin-'nhase described in Test Procedures I. II or III. ascended to thegiven test altitude. It then flew by a slower moving aircraft atincreasing speed increments generally following the pattern shownIn Fig. 12. These overtakes were based on visual reference of thepilot in the overtaking aircraft. Attempts were made to maintaina minimum horizontal separation between the two aircraft duringthe overtake so that an observer in one of the aircraft involvedcould determine the relative vertical separation at the instant ofovertake.

Recordings of all pertinent instruments in both aircraft weremade at the instant of passing.

VII. Temperature Profile Survey

The test procedure prior to takeoff, at low-level referencespeed following takeoff, just prior to landing, and after landingwas the same as that described in Test Procedures I and II.

Following the initial low-level, flights, the aircraftascended in incremental altitudes up to the scheduled test altitude asshown in Fig. 13. At each of these incremental altitudes the aircraftwas flown at low-level reference speed velocities and conditions.

The aircraft was then flown at the same indicated testaltitude th•'oughout the entire speed range using the flight patternshown in Fig. 13.

After the last speed point, the aircraft descended inaltitude increments in the same manner as that preceding the altitudetests.

VIII. Density Profile Survey

The test procedure was similar to that noted in Test ProcedureVII.

TEST EQUIPMENT

TV-2 Pacer Aircraft

The test aircraft, Fig. 14, was a modified jet trainerjpecially equipped for this program. The special modificationincluded: (a) nose boom, (b) trailing cone assembly with reel,

1T

IT

LOW SPEED PACERAIRCRAFT FLIGHT PATH

HIGH SPEED PACERAIRCRAFT FLIGHT PATH

FIG. 12 AIRCRAFT OVERTAKE METHIOD

a

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CON(CCO5- %TAIRSPEE

FOR FAGUl Rt;N F-,C

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(c) special flight sensors (described below), (d) instrumented photo-

panel observer seat (e) special radar transponder equipment, and(f) telemetric digital clock.

The related test parameters were measured by the sensor and

readout equipment noted below:

1. Ship's Indicated Altitude

The ship's normal flush static ports were connected to aselected low-hysteresis altimeter with vibrator. The repeatabilityof the selected altimeter was 1 10 feet based on special calibrations.This instrument and other photopanel instruments are shown in Fig. 15.

All pressure instruments were precision-calibrated by NAECagainst standards whose accuracy of ± 0.0011 inch Hg., is traceable tothe National Bureau of Standards' (14S) pressure references.

2. ihip's Indicated Airspeed

The indicated airspeed of the ship's normal pitot/flushstatic system was connected to a precision-calibrated, low-hysteresisairspeed indicator.

3. Nose Boom Indicated Altitude

The static pressure ports of the Kollsman pitot/static headwere connected to an altimeter of the type noted in item 1, and ahigh resolution, servo-driven, absolute pressure transducer. Theservo-driven transducer, whose matched tape scale of 165 inches inlength, had an absolute accuracy of ± 0.004 inch Hg,., sad arepeatability and resolution of 0.002 inch Hg.

4. Nose Boom Indicated Airspeed

The dynamic pressure sensed by the special nose boom wasconnected to a precision-calibrated airspeed indicator similar tothat used for item 2, and to a differential pressure transducersimilar to the servo-driven absolute described in item 3.

5. Trailing Cone System Indicated Altitude

The sensed static pressure of the trailing cone system was

instrumented in the manner prescribed for item 3.

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6. Air Temperature

The stagnation air temperature was sensed by anaerodynamically-compensated temperature probe with a recovery factorequal approximately to unity and readout in degrees centigrade on. adial indicator. The ability of the systems to accurately indicate,temperature was t O.1 0 C with a guaranteed recovery factor of 0.995of calibration throughout the entire flight envelope.

7. Fuel Flow

The total fuel flow was sensed by the normal aircraft fuelmeter and presented digitally on the photopanel.

8. Angle of Attack

A null-seeking pressure transduction-type angle-of-atacksensor and associated dial presentation was calibrated to 0.2angular displacement.

F8C Pacer Aircraft

The Navy-furnished FBC, Fig. 16, was a specially instrumentedsupersonic fighter-type aircraft. Included in the specialinstrumentation were: (a) instrumented nose boom, (b) tail-mountedtrailing cone system with internally stored power-driven reel,(c) special radar transponding equipment, and (d) internally mountedphotopanel. The parameters read out on the photopanel, Fig. 17,were similar to that of the TV-2 pacer aircraft.

1. Ship's Indicated Altitude

The photopanel presentation was a sensitive precision-calibrated, aneriod barometric altimeter, the calibration of whichwas traceable to NBS pressure reference standards. This unit wasin parallel to the ship's normal altimeter located in the pilot'spanel.

2. Ship's Indicaied Airspeed

A calibrated aircraft airspeed indicator was hooked inparallel to the aircraft's normal dynamic pressure system.

3. Nose Boom Indicated Altitude

The readout of the nose boom sensed static pressure wassimilar to that employed for the ship's system.

23

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4. Nose Boom Indicated Airspeed

The sensed dynamic pressure of the nose boom was read out

on a standard precision-calibrated airspeed indicator.

5. Trailing Cone System Indicated Altitude

The sensed static pressure was read out on a selectedprecision-calibrated aneriod barometric altimeter similar to theship's static system. D'i addition, a second altimeter was locatedabove the pilot's normal flight panel for pilot usage.

6. Nose Boom/Trailing Cone Indicated Airspeed

The sensed total pressure of the nose boom and sensedstatic pressure of the trailing cone system were coupled to aprecision airspeed indicator.

7. Angle of Attack

A calibrated vane-type angle of attack sensor was employedto determine aircraft angle of attack. Other existinginstrumentation, shown in Fig. 17, was not employed for this project.

Ground Based Equipment

The following equipment was employed where applicable todetermine geometric altitude of the test aircraft or to obtainatmospheric information:

1. NAMFE Phototheodolite System

The phototheodolite system, Fig. 18, was employed todetermine the geometric height and aircraft ground speed for alltests up to 10,000 feet which occurred at NAFEC. The reportedaccuracy of this system as employed in this program was ± 2.0feet in elevation and + 3.0 feet per second (fps) in computedground speed.

2. Vertical Aerial Camera

The aerial camera, Fig. 19, was used to obtain low-levelaircraft altitudes where the phototheodolite system was inoperativeor not available. A detailed description of this method iscontained in Reference AC. The accuracy of this method, as verifiedby the NAFEC phototheodolite system, was within 0.5 feet of thetheodolite data.

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3. NAFEC Tracking Radar System

The X-band radar system, Fig. 20, was employed to obtaingeometric elevation and ground speed of the test aircraft duringthe nigh-altitude tests at NAFEC. The reported accuracy of thissystem is 25 feet in determining height.

4. NATC Radar/Phototheodolite System

The NATC system employs a radar-coupled acquisitionphototheodolite system for determining the geometric height of theaircraft. The accuracy of this system is not known, but would bedepeLdent cn the optical limitati on due to the equipment and agiven atmospheric condition. Data reduction using this system wasaccomplished by INATC and is contained in Reference R.

,. Radiosonde

The atmosphere was surveyed by a calibrated radiosondesimilar to that shown in Fig. 21. The accuracy of the aneroidsystem was 0.01 inch Hg., and that of the temperature sensingsystems O.5'C.

6. Barometer

The barometer shown in Fig. 22 was the type used at eachof the ground-measuring stations noted in Fig. 8. These deviceswere calibrated against a known standard prior to and after theirapplication to this project.

T. Air Temperature/Humidi-y

The wet and dry-bulb temperatures at each of the ground-measuring stations were obtained by use of the self-operatingpsychron unit shown in Fig. 23. This unit was calibrated and hasan accuracy of + 0.10F.

METHOD OF AflALYSIS

Positiori error, as defined herein, is the difference betweenthe local pressure at the static ports and the free stream ambientpressure. Therefore, to derive this difference,AP,one must computethe value of the local presssure and that of the free stream ambient.

The value of the local pressure was obtained by threeindependent methods with respect to the readout instruments:

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The first method was to correct the absolute valueobtained for the indicated local static pressure, PIT (trailingcone system altitude). for known scale error. The scale errorcorrections included both the precision calibration term and anybils which was reflected during the preflight calibration. Thisresultant value is defined as the absolute calibrated localpressure at the static pressure ports, pIC"

Pic = IT - (PG + PB) :)

where P = Calibrated local pressure,Ic

P IT = Indicated local pressure,

P PG - nCalibration correction term of scale for agiven value of IJ

PB = Preflight calibration bias.

The second method was to correct the indicated groundcheck reading for any difference between indicated ground stationpressure and known ground station pressure, and then subtract fromthis pressure value the indicated pressure at the test condition.The resultant value is defined as change in calibrated localpressure of the trailing cone system's static ports, APIC.

A -PQGI - - (2)

where A PIC = Change in calibrated local pressure,

PGI = Indicated ground station pressure,

P = Known ground station pressure.

The third method was a modification of the secondprocedure where the indicated local pressure, PIR, at a givenreference speed at the test altitude das used in place of theindicated ground pressure reading.

APIC=Pn-PIT (P3)where Pn = Indicated local pressure at a given reference

speed.

The value of the free stream ambient pressure at the testcondition altitude was determined by five different methodsdependent upon attainable accuracy of observed information.

34

The first method consisted of computing the correctedstation pressure reading for each of the observed ground readings.The observed reading was corrected for scale pressure and temperatureerror, then extrapolated to the station pressure ground elevation,PGC, using the model atmospheric lapse rate. The extrapolated valuesof PGC were then averaged. The ambient pressure at test altitude,PC, was derived by subtracting from the averaged value for stationpressure, IrGC, .the computed change in pressure, A PC, as a functionof density.

PC- 7GC - APc (4)

ZN[PGI~ (PGP + P(;+PK Ah)] (5)

N

APc = - pg AH (6)

IC= to STs xI T

where PC = Computed pressure at test altitude,

7GC = Mean computed ground station pressure,

,PC = Change in computed station pressure due to A H,

PGT = Calibration correction factor due to ambienttemprature of the barometer,

PK = Rate of change in pressure at station elevationbased on a model atmosphere,

Ah = Difference in elevation between observed ground

reading and station elevation,

N = Number of readings

iC = Computed local density of air mass,

g = Local gravity,

35

A -- Difference in geometric height of station elevation

and test altitude,

S3 = Model atmosphere density for a given elevation,

Ps - Standard day pressure at station elevation,

TS =Stndd day air temperature at station elevation,

C- Mean ground station temperature.

The secomd method of computing ambient pressure at thetest altitude was to determine the mean pressure lapse rate, 7KC,based on observed corrected barometric pressure, Pj, measured atvarious elevations. This computed lapse rate was multiplied by thechange in elevation, AH, from the closest accurately determinedbarometric pressure and the resultant value subtracted from thebarometric pressure.

Ape Pc-7KC AR (8)

N

1 : 1 (9)N

where Pp, Ob-:Tved corrected barometric pressure closest4, the. test altitude,

.mean computed pressure lapse rate,

Observed corrected barometric pressure of thebase of a given air column,

PB,: Observed corrected barometric pressure at someknown point in a given air column.

A third method of obtaining a value for ambient pressureat altitude was to assume a mean reference pressure, •R, based onthe horizontal profiling of the atmosphere covered in Test ProcedureIV. The value was then corrected by the model atmospheric lapse

rate for any difference between the test altitude and the altitudeat which the mean reference pressure was obtained. The lapse rateused was that applicaýle for the reference pressure altitude.

36

PC -" R - - A H (10)

N N' (1)

2

where = Mean ambient reference pressure,

PH CorTected reference pressure based on horizontalpressure survey preceding test,

PR'. Corrected reference pressure based on horizontalpressure survey following test,

N - Number of readings preceding test,

N'= lamber of readings following test.

An additiLonal method of obtaining ambient pressure ataltitude was by computing the pressure by Boyle's Law for an idealgas. The temperature employed was that computed based on the verticalprofile of the air column temperature.

PC PTO (12)I

where R = Gas constant for dry air,

TC = Computed temperature at test altitude bbued onvertical profile of air column temperature.

The final method employed to service the ambient pressureat the test altitude was to compute the density lapse rate as afunction of geometr. c altitude, then use the computed temperature tocalculate the ambient pressure.

PC = eRTC (Assuming air as a perfect gas) (13)

weC CopGC den s ast T b oi)

where OIGC Comput•.d density at the base of the air column.

37

Tre were two mehotds employed to obtain calibrated airspeedof the test aircraft depending upon available test observations and

pocedures.

The first method was to compute the calibrated airspeedbased o the average of tracked ground speed based on flight pathswhich were 1800 vith respect to each other at the same indicatedvelocity. The equivalent dynamic pressure was computed by

wrz oulli's relation of pressure to velocity and density.

VGl + VG (15)2

where V - Calibreted velocity of aircraft,

V = Tracked ground speed in direction A,01

VG2 , Tracked ground speed in direction A + 180°,

and = tV V2 (for incompressible flow) (16)

where qC = Dynamic calibrated pressure.

The second method of deriving calibrated airspeed was bycorrecting the indicated airspeed for instrument and position error.The position error of the ship's system or nose boom system wasassumed to be the difference between sensed static of the particularsystem and the trailing cone static pressure.

vc C v= -' V I + (VCS - Vcc•] UT')

where VC = Calibrated airspeed,

V, = Indicated airspeed,

VIE= Scale Error of indicator,

VCS= Calibrated airspeed from aircraft system indication,

VCC- Calibrated airspeed from trailing cone indication.

38

The equivalent value of qc for this was obtained from ReferenceG.

The Mach Number was also obtained from Reference G based oncomputed equivalent values of altitude and airspeed.

The position error of the trailing cone system was obtained bydetermining the difference between the reference value used for PCor A PC and the related value for Pic or A Pic.

TEST RESULTS AMu ANALYSIS

The results shown in Fig. 24 are a composite of all validateddata obtained using the TV-2 pacer aircraft at altitudes ofapproximately 200, 5,000, and 10,000 feet.

The position error, A P, for test conditions at 200 feet wascomputed by solving for PIC by equation (1), and PC by equation (4),and Test Procedure I.

A p = P IC " PC

AP [I - (PGP + RB)] - ["C- A PC] (18)

All data shown in Fig. 24 were verified by independently solvingfor A P using equations (2) and (6), or through equations (2) and (8).

AP= APic- APc

AP =[PGI - (PGI - PG) -PIT] + Og A H (19)

Ap =[PGI - (PGI - PG) -PIT] - PBCN - FKC AH (20)

The position error at 5,000 and 10,000 feet was determined andverified by combining Test Procedures I and VIII.

The average change in A P as a function of airspeed, initiallyderived by equation (3), was determined for the low-altitude testswhich preceded and followed the altitude tests.

N NZ I - PIT -: IT-(PI)

N'

'•iC =2

where ST=IC Mean change in ca3ibrated local pressure betweenreference test speeds,

N Number of data points at a given reference speedpreceding altitude tests,

N' Number of data points at a given reference speedafter the altitude tests.

The number of useful data points for a given test conditionwas nominally 10. Any given graphic validated presentation for aparticular velocity is the mean of these points.

The same procedure for obtaining EP -T at the test altitude wasemployed since the first two test points at any altitude were thereference speeds. The equation, however, takes the form shown belowin equation (22).

NS PIR "PIT

EPIC . (22)

N

The indicated value of local pressure R at a given testcondition, P:T, ,-as corrected for any difference in altitude betweenthe test elevation and the altitude reference speed test elevation,

A H. The pressure constant, PKC, was that derived based on theradiosondic survey.

AP= PIc - PKC AH (23)

This resultant A P was referenced to the low-level results bythe difference between &Pj' , as derived in equations (21) and (22).

The value obtained above was validated by subtraction from thecalibrated local pressure, P1 C' the computed absolute value of PC asreflected in equation (13). The values of pressure and temperatureto compute density were based on the atmospheric survey as definedin Pest Procedure VIII.

40

A' P IC -~ PC T , R T~l(4" " ý [PIT - (PGP + PB)] "[ PGC T 4S --S T],24

Thus, Fig. 24. represents all ciatai which could be validated bysolving for free stream pressure at a given test condition by botha computed pressure lapse rate and a compv+.ed density lapse rate.The computed first-degree fit for all the .ta is shown by thestraight line drawn through the presented data. The trailing conein the extended position was 38 feet behind the exhaust pipe of theTV-2 for all tests, both low-level and altitude.

The data shown in Fig. 25 include all those data points whoseabsolute value of A P computes to 2.0 pounds per square foot (prf)or less, and the first-degree fit represented by the straight linedrawn through the data.

Figure 26 reflects the low-altitude data obtained for thetrailing cone system using the Navy F8C pacer. Test Procedures 1,II, and II, or a combination of any two of these were used toobtain this information. The method of deriving A P was similarto that outlined above for the TV-2 at the 200 foot altitude. Thetrailing coL.- in its extended position was approximately 63 feetbehind tkhe vertical stabilizer of the F8C for all tests.

Figure 27 is similar in form for the F6C as that shown in Fig. 25for the TV-2. The straight line through the data is representative ofthe first-degree fit of the data points by linear regression analysis.

The results shown in Fig. 28 are a composite of all validateddata obtained at high altitude using the F8C pacer. The TestProcedures used to obtain this data included I, II, or III for thelow-level reference velocity points, and IV, V, VI, VII, and VIIIor any combination of these procedures for altitude tests.

The method employed to calculate A P at altitude was thatnoted in equations (23) and (24). In addition, due to thelimitation in available radiosonic data, the computation shownin equation (10) was employed and TR was assumed to be the accuraterepresentation of PC at reference speed.

P R- [PIT - PK AH] (25)

41

SF I i

-• ; . . ; i , ** ,

1 4i

L I J-----i

S .4

-I. o

L 4

- |,,I .

-1.6 1

-2.4

4 -1

uT

-3. 2

1-20 14U 160 18U ot -'U U L) 22 4u 6 1) 2. H U 00t o-t 34L, 360 3e') 400 4-1ý 440 460L 480

INDICATF.D AIR, SPEED (KNOTS)

FIG. 24• POSITION ERROR VERSUS INDICATrED AIRSPEED-TV-2 PACER (LOW ALTITUDE) ALL DATA

,42

i I ----,--------.--,- -

1 4

--- .(---- -i- -6I

0 S 0

, S

2S

0 -. -,-! I t - -"

----- - - ----- :_--:-- •-. 4-

-1 0

-1.4 1 -L- :

-14 b

I o0 120 140 160 1P 8 •00 Zo O 240 260 280 300 320 340 360 380 400

INDICATED AIR SPEED (KNOTS)

FIG. 25 POSITION ERROR VERSUS INDICATED AIRSPEED -

TV-2 PACER (LOW ALTITUDE) SELECTED DATA

43

- 4.7

t 1 F-1-f- 1t

4- -4-

4--

.4 U-i++ +--

44 ++44

1 4

.4

-.4 -

00

-i,-A

-1,4

-1. - - - - - -

-61 --- -T -r I --

I 3U 1l4) kl )1. 11u I li({ ) 2 su 27t) 190 7 -q 310 1311 30 3SU 70 390 410 410 4S•U 4"70 490

INDMGATED AIR .Sr"D 1KNOTS)

FIG. 27 POSITION ERROR VERSUS INDICATED AIRSPEED -F8C PACER (LOW ALTITUDE) SELECTED DATA

45

.6

4. ! . ,-

F8C~~~~~ PAE LHAN ALITD)SLDT

46 ..

0 4'

"-1 0-

Ito 140 160 180 400 L20 414C 2b0 ZO ý00 3zU 340 360J 380 400 4-Z0 440 460 4NO rOU q0 SD 40

INDICATED AIR •PEF•FD (KNOTS)

FIG. 28 POSITION ERROR VERSUS INDICATED AIRSPEED-FSC PACER (HIGH ALTITUDE) ALL DATA

46

An additional method of obtaining AP is to aszume thedifference between the computed velocity obtained in equations (15)and (17) where available is due to the static pressure error of the

VG1 + VG2 ) (6

AP= " [VIE + (VCs - vCc)] - (26)2

where A P can be derived from differences in velocity whenexpressed in terms of dynamic pressure. The valuesof dynamic piessure can be obtained in terms of A Pfrom appropriate charts in Reference G.

Figure 29 is a presentation of those data points whose absolutevalue of AP computes to 2.0 pounds per square foot or less.

Figure 30 is a composite of all selected data obtained for thetrailing cone system at all test altitudes. Approximately 75 percentof the data shown exhibits a A P of less than-± 1.0 psf. This isequal to approxizmately + 12 feet or less at sea level, and ± 30 feetor less at 30,000 feet.- These absolute values of A P are within theoverall accuracy of each of the several systems involved in themeasurement of the various parameters.

Figures 31 through 34 , inclusive, reflect the position erroras a furction of dynamic pressure and Mach Number. This presentationis the more comon technical representation of the static pressuredefect.

Figure 35 depicts the statistically-derived position error interms of feet. Shown are the curves for both sea level and 30,000feet based on a model atmosphere. Also shown are the respectivevariations of 75 percent of the observed data which was appi ximately+ 12 feet at sea level, and + 30 feet at 30,000 feet.

The test results shown in Figs. 24 through 35 are based onaircraft performance with the aircraft in steady-state unacceleratedlevel flight with no control surface movement or power changes. Itwas noted during the data gathering and testing that aircraftacceleration, or surface control movement at a time when data wasbeing taken, had adverse effects on the data.

)ý7

- 0 ]

0

0/

0 .-. - 'S o - -- - ---

* 0 I 1 1 . -

0 0

I- - - - - - - - - - - - - - - - -

0,

- i . .

- - - -- - - - - - - - - - - - - - - - - -

-o 102,I

1-.4

InI

Oc4-

-o -0~-

- Kfit

'T--------

( u -M r"

--4- 1• 4 IT

F I

-- - & -

zo

T 0

F

01 1

F.I G. 3 1 &Pq-VRU n T-2 AE

- --. 00o

:1 -K ---

_ 4 4

:i _iT: .

0G. 31 ,P/ VERSUS Mn TV, PACEIR

•, \C II•1.•I ~

FIG.31 P/q VESUSo V- AE

[E± lILA 1~-B

0 .006

005

i- t i -4

S003

' 001

ool

-, 003

-. 004

-. u I- -

M(ill

i:,.G. 32 AP/q c VERSUS Mn F8C LOW ALTITUDE

.-.-.

,iI -Jjllii iij24-.

o1,

001?

* 008- - t- 2 ILw 07 ~ . ---- ,--i--=.006

z ,

I 0054

%- .003•

0

.ooi- - ----- *1-

w wo

0, -. 002

4 -. 00oz

u. --- 4~.

-. 004 -O

-I--I-.- t----i---'---------. 007 - 1--7 1:-

N - - u:..0 .1 ,• ,3 .4 .5 .b .7 .8 ,9 1.0 1. 1 1.z 3

MACHI NUMBEPS

FIG. 33 &P/q VERSUS Mn FSC HIGH ALTITUDE

C

52

+-7

U .004 ---

.003 --- E.- v.

0 -ON~-

p- 00

001

0053

2 iCHANCE E IN HP[LIGI I IN F LE F

pu Q I I t

CA

00

0 %

I II

> >

l r, ri ViH>

.2.. ....

,3NCLU,2 ON:..

tlasel on trie tes t . r suits shown, tt.e rollowing concluslonsmay be dravn:

1. The trailing cone system does not exhibit any significanterror as a Vunction of the towing aircraft when the aircraft isflown in unaccelerated, level flight (Figs. 25 and 27).

2. The system does noL (xijibit anr' noticuable Mach ef'fectup to 1.12 Mn (Fig. 34).

3. Due to the basic design, the system does not have ameasurable angle-of-attack sensitivity.

Based on observed flight performance and pilot comments, thefollowing is concluded:

1. The system does not interfere with nor affect thecontrollability of the towing aircraft.

2. The use of the system as a component of the airspeedsystem gives a satisfactory instrument presentation without thenon readable oscillatory behavior normll'y encountered whenH lerating through M..h 1.0.

RECOMMENDATIONS

It is recommended that the trailing cone system be employedas a stanidard to determine the position error of the static pressureportion of' an aircraIt'. airspeed system. Position errordetermination using this methcd must be accomplished when theaircraft is in unaccelerated level flight.

559

•Ir OF 32.MBOLOI

D -Drag .g Locel gravity.

,all Difference in geometric height of station elevation andtest altitude.

18h Al • i eUTevEL±Uon between observeq ground readingand station elevation.

N Number of' observations relating to a given defined grouping.N'= Number of observations reJating to a given defineu grouping

other than that used for N.&P = Position error.PB = Preflight calibration bias.

P3C = Observed corrected barometric pressure of the base of agiven air coluxm.

NC'= Observed corrected barometric pressure at some known pointin a given air column.

PBCN = Observed corrected barometric pressure closest to the testaltitude.

PC = Computed pressure at test altitude.APO = Change in computed station pressure due to A H.

PG = Known ground station pressure.76c = Mean computed ground station pressure.P0 I = Indicated ground station pressure.PGp = Calibration correction term of scale for a given value

of PIT'PUT = Calibration correction factor due to ambient temperature

of the barometer.S= Calibrated local pressure.

APIC = Change in calibrated local pressure.-c Mean change in calibrated local pressure between reference

test speeds.PI = Indicated local pressure at a given reference speed.

PIT = Indicated local pressure.PI = Rate of change in pressure at station elevation based on

a model atmosphere.= Mean computed pressure lapse rate.= Corrected reference pressure based on horizontal pressure

survey preceding test.7R = Mean ambient reference pressure.PRI= Corrected reference pressure based on horizontal pressure

survey following test.PS = Standard day pressure at station elevation.c, = Dynamic calibrated pressure.R = Gas constant for dry air.

TO = Computed temperature at test altitude based on verticalprofile of air column temperattu ..

7GC = Mean ground station temperature.

LIST OF SDWBOLS (continued)

T= Atandard dry air temperature at station eleva+ion.7 Cal1ib:at~ ve-÷.iL~i~ vf alrcra...

V= Calibrated airspeed.VnC= Calibrated airsp"ed f-rom trpal!ing cone indicaticn.

VcS = Calibrated airspeed from aircraft system indication.VGI = Tracked ground speed in direction A.VG2 = Tracked ground speed in direction A + 18C0.

VI = Indicated airspeed.V Scale error of Indicator.

• = Weight in pounds.S = Angle of attack.

ec = Computed local density of air mass.aGC = Computed density at the base of the air column.

=s - Model atmospheric density for a given elevaticn.

5'

A Anonymous, Accuracies of Radiosonde Data. U.S.A.F. (MATS)Tech. Report 105-133.

B. i~oymuMnrwu on [~~t . ue, U.8.A-,F.--t(A7Tech. Report TN 105-101.

a Anonymous, Optical Tracker Evaluation and Accuracy Test System,U.S. Army (Sig Corps) ASITA 270638.

D Anonymous, Radiosonde Observation Computation Tables andDiagrams, U.S.A.F. (A.W.S.) Tech. Report 105-64.

E Anonymous, Dupirical Frequency Distribution of Pressure,STemperature and Air Densitz at Levels of Constant Altitude,U.S. Arnm (Missile Command) Tt 61-26.I

SF Anonymous, StDS Technical Facilities, RD P 6000.2.

G Bartlett, E. P., Performance Flight Test Handbook, U.S.A.F.AFFITC-TN-29-22.

H Butler, C. M. and Haller, N. M., An Investigation of the* Feas1bility of Indicating Altitude by Gravimetric Measurements,

ASTIA 27T112.

I Deleo, R. V. and Hagen, F. W., Evaluation of New Methods forFlight Calibration of Aircraft Instrumen ystems, U.S.A.F.(WADt) tr 59-295.t .....

i J Fine, R L., Flight Test Evaluation of Aircraft PressureS~Altimeter Installations, U.S.A.F. (WAX•) TN 56-436.

SK Flickinger, H. S., Test and Data Reduction Procedures forStatic Pressure Position Error Calibration by titimeter-Depression Methods, U.S. Navy (NATO) ¶M 4-53.

L Gracey, W., Measurement of Static Pressure or Aircraft,NACA Report 1364.

M Gracey, W., Measurement of Static Pressure on Aircraft,NACA RM L5tA09.

N Gracey, W., Position Errors of the Service AirspeedInstallations of 10 Airplanes, NACA TN 1692.

57

REPFMR~CES (continued)

0 Gracev. W.. Measurement of' Static Pressure on Aircraft.NACA TN 4e184.

F Graoey, W-. &hd Shipp, 3'. A., Random Dleviations from CruiseAltitudes of a Turbojet Trank rt at ITTItues letwen 20000

Q Hanks, N. J., Altitude Separation ?rogrwi, U.S.A.F. Tfl456-28.

R Hesse, W. J., Position Error Determinations by StadianietricRanging, U.S. Navy (NATC) TR 2-55.

8 Holbeche, T. A., and Spence, D. A., Temperature MeasurementsBehind an Attenuating Shock Wave, R.A.E. Aerwo 2btZ3.

T Hnrkey,, E. J. , A Method of' Airapoed Calibration at High MachNumbers Witho'it TF~fwece to The~ Ground, Worth AmericanAviation IReport =A47,-775.

U James, P., LI-ndsayr, J. L., frant, Jr. and Za~lovcik, J. A.,A. Metho~d of- JClia AirsW W antallations onAipaeat ansonicand sonic do se sor Accelerometer

* an zlue-Anl measurements., RCRepor 1.5

V Icuowler, A. E., Lock, C. N. H. and Pearcey, H. H., The Effectof2!wsbility on SaticHeads, ARC Report 238

W Martin, T. A. and White,, W. E. P Transonic Investigation ofThree Pitot-Static Probes Des igned for laressure-FkrrorCompnaation, UI.B.A.F.7=~)) b2-123.

X Piocard, J., Proedý&s of Altime Conference, ABra

of Research an eeomn, ac 1 ~o

Y Price, E. N., A grstem for Providing and Measuring Free StreamStatic Pressure, D~ouglas Aircraft Co. DEV 3519.

Z Richardson, N. N., Forecasting Density Altitude, U.S.A.F.(MA.TS) Tech. Report lb5.

AA Ri~tchie, V. S., Several Methods for Aerodynamic Reduction ofStatic Pressure Sensing Errors for Aircraft at Subsonic,. Near-Sonic, and Low Supesonic Speeds, NAZA TRR-15.

I5

REFEROECES (coritinued)

E~rror ffor CAR 3 t'-pe Aircraft, FAA M 9D Report 115-122I.

AC Shrager, J. J., Cal ibrating Static Pressure Systems at UVwAltitude, FAA( S) R 64-37.

AD Siegel, n. B., Acceleration - Deceleration Method ofCalibration of the Static Pressure Position Error in theTransonic Flight Region, U. S. Navy (NATC) TM 1-60.

AE Smetaaa, F. 0., Stuart, J. W., and Wilber, P. C.,Invest•atLon of Free Stream Pressure and StagnationPrre Measurement from Transonic and Supersonic Aircraft,U.S.A.?. (WADC) 55-236.

AY Smith, A. F., Digital Optical Tracking System, U.S.A.F. (Aa)Ti-6&-358.

AG Smith, K., Pressure Lag in Pipes, R.A.E. Report Aero 2507.

AH Smith, K. W., The Measurement of Position Error at HighSpeeds and Altitude by Means of a Trailing Static Head,A.RC. 160.

AI Werner, F. D. and Geronime, R. L., Automatic Correction ofErrors in Airplane Static Pressure Surces, U.S.A.F. (WADC)56-193.

AJ Wright, W. F., A Survey for Naval Ordnance Laboratory ofHigh Altitude Atmospheric Temperature Sensors and AssociatedProblems, ASTIA 27bI52.

AK Zalovcik, J. A., A Radar Method of Calibrating AirspeedInstallations on Airplanes in Maneuvers at High Altitudesand at Transonic and supersonic Speeds, NACA Report 9b5.

59

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TEST OF TRAIL CONE SYSTEM TO CALIBRATE STATIC ...The cone stabilized static source is an outgrowth...

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