Calibration of Transonic and Supersonic Wind Tunnels" 19 6

, NASA CR 2920 C.1 ,.

1 ( I NASA Contractor Report 2920

7 LOAN C9-V: RETI..lF, AFWL TFr! '?'fCfi1.. LIBRARf ._ KIRTLAND AFB, N. M. - - Calibration of Transonic *.? - and Supersonic Wind Tunnels

T. D. Reed, T. C. Pope, and J. M. Cooksey

CONTRACT NAS2-8606 NOVEMBER 1977

NASA ~-

TECH LIBRARY KAFB, NM

NASA Contractor Report 2920

Calibration of Transonic andSupersonic Wind Tunnels

T. D. Reed, T. C. Pope, and J. M. Cooksey Vought Corporation Dallas, Texas

Prepared for AmesResearch Center under Contract NAS2-8606

National AeroMubics and Space Administration Scientific andTechid Information Office

1977 -" . . FOREWORD

In April, 1970 areport was issuedby an ad hoc NASA-USAF groupon tran- sonicscale effects and testingtechniques. This report assessed transonic testingtechniques and recommended, among otherthings, that a transonic wind- tunnelcalibration manualbe writtenwhich reviewed the state-of-the-art. This was viewedas a necessary step toward the development of more accurate and standardizedtunnel calibration procedures.

For this purpose,the present manual was jointly fundedby: (1) the U. S. Navy throughthe Office of Naval Research, (2) the U. S. Air Forcethrough the Air Force Flight Dynamics Laboratory and Arnold Research Organization, (3) NASA throughthe Washington Headquarters and theLewis, Langley and Ames Research Centers. The contract was administered by NASA Ames. Mr. F. W. Steinleserved as technicalmonitor.

A rough draft of this manual was reviewed by personnel of NASA Ames ResearchCenter and ArnoldResearch Organization. The comments ofthe various reviewerswere compiled by Mr. F. W. Steinleat Ames and Mr. F. M. Jackson at ARO. Themanual was improvedconsiderably by the constructive comments that were received, and we wishto thank all those involved for their time and efforts.

Our thanks go to Mr. C. J. Stalmach ofthe Vought Corporationfor the discussionof hot wires and filmswhich is given in Appendix 1. Finally, we wishto acknowledgethe superior typing and secretarialassistance provided by Ms. F. H. Deason.

..Ill .

TABLE OF CONTENTS

Sect ion

1 . INTRODUCTION...... 1 A . Gackground ...... 1 B . HistoricalSketch ...... 2 c . Calibration Procedures ...... 4 References It. TUNNEL VARIABLES ...... 7 A . Types of Tunnels ...... 7 B . OperationalParameters ...... 8 1 . PressureControl ...... 8 2 . Calibration Accuracy. Flow Uniformity and Relationship to Model Testing ...... 12 References C . Flow Parameters and UncertaintyRelationships . 22 1 . Pressures ...... 22 2 . Temperature ...... 34 3 . MachNumber ...... 36 4 . Flow Angularity and Curvature ...... 40 5 . Reynolds Number ...... 42 6 . Unsteadiness.Turbulence. and Noise .... 47 7 . Humidity ...... 52 8 . Test Mediums ...... 56 References

111. CALIBRATION PROCEDURES AND INSTRUMENTATION...... 59 A . Settling Chamber Pressure ...... 59 References B . Total Temperature ...... 63 References C . PitotPressures ...... 68 References D . TestSection Static Pressures ...... 78 1 . TransonicSurvey Pipes ...... 79 2 . Transonic Static Pressure Probes ...... 86 3 . SupersonicStatic Pressure Probes ..... 105 4 . Orifice-Induced Static Pressure Errors ... 110 5 . GeneralPurpose StaticPressure Probe ... 116 References E . Measurement of Flow Angularity ...... 124 1 . DifferentialPressure Yaweters: 2-D ... 124 2 . DifferentialPressure Yawmeters:3-0 ... 128 3 . Hot Wire/Film Yawmeters ...... 134 4 . ForceBalance Yawmeters ...... 137 References

V Sect ion Page

." ~ .;r .... " F . Measurement of UnsteadyFlow Disturbances .... 144 1 . DynamIc PressureMeasurements ...... 147 References G . TransonicTunnel Boundary Conditions and WallInterference ...... 162 1 . ConventionalVentilated Walls ...... 162 2 . AdaptiveWall Studies ...... 165 3 . BoundaryLayers and Wall Generated Noise . . 169 H . StandardModels ...... 174 1 . AGARD ForceModels ...... 174 2 . TransonicPressure Models: 2-0 ...... 175 3 . TransonicPressure Models: 3-0 ...... 175 References 1 . Optical Methods ...... 182 1 . SupersonicTunnels ...... 182 2 . TransonicTunnels ...... 182 3 . Newer Methods ...... 183 References J . Humidity Measurements ...... 185 References

IV. ERROR AND UNCERTAINTY IN CALIBRATION MEASUREMENTS .... 189 A . Random Error ...... 189 B . FixedError ...... 190 C Uncertainty 191 ...... D . ErrorPropagat ion ...... 192 References

.... v . .. CONCLUSIONS AND RECOMMENDATIONS ...... 195 A . Summary ofState-of-the-Art of Transonic and Supersonic Wind Tunnel Calibration ...... 195 B . TransonicTunnel s ...... 198 C . SupersonicTunnels ...... 202

APPENDICES

.. 1 . AND HOT WIRES HOT FILMS ...... 203 Nomenclature ...... 217 References I1 . LASERDOPPLER VELOCIMETERS ...... 222 Nomenclature ...... 249 References Ill. EFFECTS OF VIBRATION OF A CIRCULARCYLINDER ON STATIC PRESSURE MEASUREMENTS ...... 255 .. IV. FACILITIES WHICH RESPONDED TO QUESTIONNAIRE A . Table 1: Facilities ...... 260 B . Table II: TestSection CharacteristIcs ..... 267

vi " LIST OF ILLUSTRATIONS -. Fiqure . Title Page

1.c.1 Dataand ErrorFlow Diagram,Ref. 1 ...... 5 2. B. 1 Jackson's Flow Quality Criteria for Transonic Tunnels,Ref. 1 ...... 14 2. B. 2 Allowable Linear MachNumber Gradient Over Model Lengthfor BouyancyDrag Coefficient Contributionof 0.0001, ...... 17 2. B. 3 Effectsof Reynolds Number on Calibration of the PWT-16T Tunnel at Mm = 0.6 and 0.8 for ew = 0 and ~=6...... 18

2. B. 4 MachNumber Gradient OverModel Lengthas Percent of Average MachNumber for Bouyancy Drag.CoefffcIent of 0.0001 ...... 19 2. c. 1 Afterbody DragData at anAverage MachNumber of 0.95...... 23 2.c.2 AfterbodyDrag Data With Tunnel MachNumber Given to ThreeDecimals ...... 24 2.c.3 . The Uncertainty of Pitot-to-Static Pressure as a Functionof NumberMach ...... 27. 2.C.4 The Sensitivity of Dynamic Pressureto Stagnation PressureError, Transonic Operation ...... 29 2.c.5 The Sensitivity of Dynamic Pressure to Static Pressure Error,Transonic Operation...... 31 2.C.6 The Sensitivity of Dynamic Pressure to MachNumber Error,Supersonic Operation ...... 32

2.C.7 The Relation of Stagnation to Static Temperature as a Function of MachNumber ...... 35 2.C.8 The Sensitivity of MachNumber to Static Pressure and StagnationPressure Error ...... 38 2.C.9 The Sensitivity of MachNumber to Static Pressure and StagnationPressures...... 39 2.c. 10 Change in FlowDirection With Increment of Mach Number, Ref. 3 ...... 41

vii ' Flgure Title Page

2.c. 11 The'Sensitlvity of Unit .Reynolds Number to Statfc Pressure Error...... 43 2.c.12 The Sensitivity of Unit Reyno1,ds Number to Stagnation Pressure Error...... 44 2.C. I3 The Sensitivity of Unit Reynolds Number to Stagnation Temperature Error . , ...... 45 2.C.14 The Sensltlvity of Unlt Reynolds Number to Mach Number Error...... 46

2.C.15 Flow Disturbances in Transonic Tunnels, Ref. 5 . . . . 49 2.C. 16 Flow Disturbances in Supersonic and Hypersonic Tunnels, Ref. 5 ...... 50 2.C. 17 The Ratio of Relative Humidity in the Stream to Reservoir as a Function of Mach Number...... 53 2.C. I8 Reservoir TemperatureRequired to Avoid Condensation, Ref. 10 ...... 54 3.8.1 Total Temperature Probes...... 64 1 Isentropic Stagnation Pressure Probe, Ref. 8. 3.C...... 70 3.C.2 AEDC Supersonic Mach Number Probes...... 72 3.c.3 Mach Number Probe forSmall Pilot LEHRT Facilities, Ref.9...... 73 3.0.1 R.A.E. Subsonic Static-Pressure Probe ...... 83 3. D. 2 Typical Pressure Distributions Along Probeat Two Locations on Tune1 Center1 ine,M = 0.74 (choked), R/L = 19.7 x 10 z per meter...... 84 3.0.3 Variation of Static-Pressure ReadingWith Position of Static Holes and Nose Shape at I4 = 1.6, Ref. 8. . . 87 3.D.4 Transonic Pressure Distributions ona 20 deg Cone- Cylinder Wtth 0.008% Blockage, Ref. 12 ...... 90 3. D. 5 Transonic Pressure Distributionson a 20 deg Cone- Cy1 inder , Ref. 20 ...... 93 3.D.6 Dimensions of the R.A.E. Static Pressure Probes . . . . 96 3-0.7 Transonic Characteristics of the TwoR.A.E. Probes. . . 97 3.D.8 Effect of Orifice Location Utilizinga Double Wedge Support Strut, Ref. 32...... 103

viii F I qure Title Paqa

3.D.9 General Criteria for ProbeSurvey Rakes, Ref. 33. ... 104 3.D.10 Orifice-InducedStatic Pressure Errors, Ref. 50 ...'. 113 3.0.11 Transonic/SupersonicStatic Pressure Probe...... 117 3.E. 1 Two DimensionalYameters ...... 126 3. E. 2 PyramidYameter...... 129 3.E-3 Sensitivity of 60 degConical Yawmeter...... 131 3.E.4 Split Hot Film, 20° Wedge Probe CallbrationBr.idge VoltageDlfference vs Flow Angle,Ref. 23 ...... 136 3.E.5 Geometry of AEDC ForceBalance Yawmeter ...... 138 3.E.6 Sensitivity of the AEDC ForceBalance Yameter. .... 139 3. F. 1 FrequencySpectra of Noise from a Turbulent Boundary Layeron a SolidWall, Ref. 3 ...... 145 3. F. 2 Noise FrequencySpectra for Some ExistingContinuous Windtunnels at M- = 0.80, Ref. 3...... 145 Small Piezoelectric Dynamic Pressure Probe,Ref. 14 . . 154

APPENDIX I

A.l.l Correlationof Convective Heat Transferfrom Transverse Cy1 inders, Ref. 3...... 206 A. 1.2 Fluctuation Diagram for 1 Percent Mass Flow Fluctuations with Varyinq Degrees of Correlation, Ref.7...... 206

A. 1.3 Fluctuation Diagrams for 1 PercentTurbulent VelocltyFluctuations (Vorticity Made),Ref. 7. 209

A. 1.4 FluctuationDiagram for 1 PercentTemperature Spottiness(Entropy Mode), Ref. 7 ...... 209 A. 1.5 Fluctuation Diagram for SoundWaves that are Almost MachWaves Having I% PressureFluctuations, Ref. 7. .. 21 0 A.1.6 Fluctuation Diagram forUncorrelated Modes at M11.75; TemperatureSpottiness of 0.1 PerCent: TurbulentVelocity Fluctuations of 0.2 Percent; SoundWaves (Detectable) 0.1 Percentof Mass Flow Fluctuations.(Dotted Lines Show SeparateContri- butions.)Ref. 7...... 21 0

fx Fiqutc Title Page 1 A.1.7 Comparison of Pitot Probe and Hot-wire .Measurements of Free-Stream Pressure Fluctuations in a Conventional, Mach 5 Nozzle, Ref. 14 ... 1 ...... 214

APPENDIX II A.II.l Dual Beam Laser Doppler Velocimeter, with Optional Forward and Backscatter Modes...... 224 A. 11.2 Generation of Interference Fringesin Measur- ing Volume of Dual Beam Laser Doppler Velocimeter . . 225 A. 11.3 Light Scattered by a Small Particle ...... 226 A. 11.4 Laser Anemometer Signal From Photodetector...... 226 A. 11.5 Effect of Particle Diameteron Frequency Response ... 238 A. 11.6 Time Constant Asa Function of Particle Diameter For Various Mach Numbers, Particle Density= 1 gm/cc ...... 240 A.11.7 Maximum Frequency For No More Than 5% Attenuation of Sinusoidal Velocity Variations, Particle Density- 1 gm/cc...... 240 A. 11.8 Effect of Velocity Biasing on Mean Velocity Measurements in Turbulent Flow ...... 243 A. 11.9 Sensitivity CoefficlentrFor Determination of Mach Number From Velocityand Stagnation Temperature Measurements...... 248

APPENDIX Ill A.lll.1 Pressure Distribution ona Circular Cylinder in Crossf low, Ref. 1 ...... 2 58

X .... . NOMENCLATURE*

A amplitudeof sinusoidal oscillation, or probeinterference functionintroduced in Eq. (3.0.1)

0 functionintroduced in Eq. (3.D.l) as a measure of probe- crossflowinterference

fixed error 1 imit for Mach number, Eq. (4.0.2) *M

Chapman-Rubesin viscosity parameter

AC dragcoefficient increment produced by a linearpressure. DG gradientin the test section

RMS value of fluctuating static pressure coefficient

RMS fluctuatingstatic pressure coefficient per unit band widthat frequency n.

D diameterof a transverse,cylindrical, probe support

DS distance between centersof slots in tunnel wall d diameterof static pressure probe dl diameterof Pitot probe orifice diameter

F (n) nondimensionalspectral function which is ameasure ofthe intensity of staticpressure fluctuations per unit band width atthe frequency n, ACp = 1; F(n)dn f oscillation frequencyof

frequency of static pressure fluctuations fP f r finenessratioprobe of nose(2Ln/d) H total head or stagnationpressure testsectionin * Separate lists of symbolsappear in Appendices I and II.

xi total head in settling chamber

Pitot pressure at a = 0 (eithersubsonic or supersonic)

time-averaged, totalpressure behind a normalshock

RMS offluctuating total pressure behind a normalshock

slot parameter, Eq. (3.6.7)

model length

nose 1 eng t h

distance from cone-cylinder juncture to nearest static pressure orifice a distancefrom a static pressure orifice to beginning of a probeenlargement, e.g., flare or support

Mach number based on static pressure in plenum chamber

Machnumber intest section

mass flow per unit area through ventilated wall

mass flowper unit area in freestream of test section n reducedfrequency of static pressure fluctuations, fpwT/u,

"d generaldesignation for direction normal to a wall

P staticpressure in freestream of test section

RMS of fluctuating static pressure

measured, unsteady staticpressure

staticpressure in settling chamber

true,unsteady static pressure of undisturbed freestream

time-averaged, truestatic pressure

staticPressure measuredon a probe ortunnel sidewall

indicated dynamicpressure, HI-P

incompressible definition of dynamicpressure, H-P

dynamic pressurein settling chamber

dynamic pressureof freestream in test section

porosityparameter, Eq. (3.G.5)

xi i viscous parameter for flow through slots RS Re Reynolds numberReynolds Re

Rex Reynolds number based on wetted length S wingreference area, or width of a strutsupport for a probe

compressible yameter sensitivity. SY Eq. (3.E.2) sy* incompressible yameter sensiflvity, (3.E.l)Eq. T period of sinusoidaloscillation

t time

u(t1 total,unsteady velocity along a probeaxis

ut4 toea1uncertainty Interval for Machnumber, Eq. (4.0.3)

"m velocity of freestream in testsection

us veloclty of sound source 112 U turbulentfriction velocity, T (T~IP) v mode 1 vo 1 ume

Vn(t) total,unsteady velocity normal to a probeaxis

vn average velocity normal to a ventilated wall

WS width of slots

WT square root ofcross-sectional area of testsection

X Cartesiancoordinate measured alongthetunnel axis

Y Cartesiancoordinate measured normal tothe tunnel sidewalls

2 Carteslancoordinaee measured normal to the top and bottom walls of tunnel

Greek Letters

a angle of attack

8 (I" 2 1 112

Y ratio of specificheats 6 anglebetween orifice planesof a yameter

ll azmuthalangle polar or coordinate angle

xiii e semi-vertex angle of a cone

9w tunnel wall angle (positive for divergent walls)

'e 'e viscosity coefficient at edgeof boundary layer viscosity coefficient at wall temperature 'W 'W 0 density of gas am standard deviation in Mach number along tunnel centerline

T wall porosity

TW shear stress ata solid wall

$ perturbation velocity potential, Eq. (3.6.1) rlr yaw angle

0 angular velocity, rad/sec

xiv I-

1. INTRODUCTION

1 .A. Background

The use of a wind tunnel for aerodynamic measurements requiresa knowledge of the test environment. Furthermore, a definite relationship obviously exists between the accuracy with which the test conditions are known and the uncertainty in the final results. The demand for increased wind tunnel data accuracy follows naturally from the demand for improved full scale vehicle performance and accuracy of performance prediction. A sustained effort has been directed toward improving the accuracyof test data from existing wind tunnel facilitfcs. In addition, requirements have been established for new wind tunnel facilities with more complete simulation capabilities.

The results of one of the first comprehensivetest programs to study the correlation of wind tunnel data from several transonic facilitieswere reported by Treon et al. in Ref. (I). Since the same model, instrumentation and support sting were used in each of the three tunnels, this unique series of tests allowed a comparative evaluation of the effects of facility flow environmentand calibration upon data agreement. The results of this series of tests, using state-of- the-art techniques and instrumentation, were considered goodbut deficient relative to current goals.

The purpose of this report is to review the current state-of-the-artof . wind tunnel calibration techniques and instrumentation, evaluate the expected results and, where possible, recommend improvements. This program was carried out by (1) acquiring information from eighty-eight wind tunnel facilitiesby means of a c.omprehensive questionnaire,(2) a detailed literature search, (3) personal visits and telephone conversations,and (4) independent analyses.

This report documents the results of these investigations. In addition to the above background information, SectionI also presents (I) a brief historical sketch of attemptsto improve wind tunnel flow quality and calibration procedures and (2) a sumnary of tunnel cal ibrat ion tasks.Section I I discusses tunnel variables and how uncertaintyin the measurementsof various flow quantities affect test results. The details of measuring staticand total pressures, temperature, flow angularity, flow unsteadiness, and humidity are all discussed in Section 111. This section also includes a review of the

I transonic-wall-interferenceproblem, the use of standard models, and the role whichoptical methodscan have duringtunnel calibrations. Section IV discusses thevarious types of errors in calibration measurements and their effects on finalresults. In additionto presenting conclusions and recmendatlons, a summary of thequestionnaire results is given in Section V. Themanual concludes with four appendfces.Appendices I and ti review,respectively, the use of hot-, wirts/films and laserDoppler velocimeters. Appendix 111 discussesthe effects of vibrationon a cylindrical,static pressure probe. Finally, Appendix IV sumnarizes the characteristics of tunnels for which questionnaires were received.

1 .B. HistoricalSketch

The need for good flowquality in wind tunnel was recognizedby the earliestinvestigators. As reportedby Pritchard in Ref.(21, the Counci 1 of theRoyal Aeronautical Society agreed in 1870 toprovide funds for the construction of 'la suitable and well-finishedinstrument having the means of instantly setting various plane surfaces at anydesired angle and capable of registering bothhorizontal and verticalforces simultaneously for all degrees ofinclina- tion. The resultsto bepublished for the benefit of the Society.'' As a result of this action,the first wind tunnel was constructed by Wenharn and Browning, anda series of tests on flat plates wereundertaken.

A laterreport onthe results noted that "these experiments would have been more satisfactory hada steady and continuouscurrent been obtained,but the fluctuationscarried by each arm ofthe fan, as it revolved,exerted an appreciable influenceon the result," Theneed for improved flowquality was alsorecognized by laterexperimentors, e.g.,see Ref. (3).

Dr. Ludwig Mach (son ofErnst Mach) constructed a tunnelat Vienna in 1893 with a test section of I8 x 25-cm which was used for flow observation and photography.This apparatus used a wirescreen over the inlet to straighten theflaw. In 1896, Sir Hiram Maxim constructed a 91 x 91-cm tunnel and used a form of honeycomb to remove fan-inducedswirl and straightenthe flow upstream ofthe test section. The Wrightbrothers' tunnel, constructed in 1901, included bothscreens anda honeycomb. A tunnelconstructed by Dr. A. F. Zahm at Washington in 1901 includedscreens of cheese cloth and wire to smooth theinlet flow.

2 Dr. Zahm also was concerned with flow uniformityand the accuracy of calibration of thetunnel velocity. He developed an extremely sensitive manometer for measuring the pressures generatedbv a.Pitot-statlc tube which was used for velocity measurements. In describing this instrument, he used the term "wind tunnel" for the first time in the literature. Zahm also used a toy balloon moving with the flowto obtain a time-of-flight measurement of the velocity. Another calibration procedureused by Zahm involved measurement of the force on a llpressure plate" or drag plateat the same time the flow velocity was measured. This method allowed determination of the flow velocityduring .later testsby observing the forceon the pressure plate,Ref. 4.

Add i t iona 1 discussion of early wind tunnels and measurement techniquesis also given in an article by Goin -(Ref. 5). From the beginning, the development of wind tunnel facilities has usually been a precursor of improved flight vehicles as outlined by Goethert in Ref. 6. The development of new and improved wind tunnels has, in turn, required new calibration procedures, techniques and instrumentation in the struggle to provide experimental data with the accuracy requiredby vehicle designers.

3 I.C. CalibrationProcedures

Both the quality of, the wind tunnel flow environment and the accuracy with whichthis environment is known contributeto the accuracy of aerodynamic measurements. The totaluncertainty in aerodynamicdata isthe result of a large number of error sources, as is discussed inSections 11.8.2 and IV.

Figure I.C. 1, from Ref. (71, illustrates the many sources of error and the manner in whicherrors propagate to a typical test result suchas drag coefficient.Considering the total number oferror sources,the necessity to minimizethose due totunnel calibration is obvious. This flow diagram is helpfulsince it isolatesthe facility flow environment and calibration elements whichare discussed herein.

Both the quality of the flow and theaccuracy with which the flow conditionsare known areconsidered as part of the calibration contribution. It is suggestedthat the calibration effort include the following elements:

1. Initial evaluation of performancecharacteristics and flow quality, and determination as to need for corrective action.

2. Determinationof optimum tunneloperational parameters such as wallangle and porosity,control systemperformance, etc.

3. Diagnostic measurements to investigate a specificflow problem or deficiency.

4. Measurement of mean, unsteady and spatialdistributions of test sectionflow conditions for the selected tunnel configuration and variousoperating conditions.

5. Standard model testsfor inter-facility comparisons.

6. Periodicre-evaluation of basic tunnel calibration for control or monitoringpurposes. This may be accomplished in partby tests on a standard faci 1 i ty model.

Consideringthe above taskdescriptions, it can be observedthat flow quality improvements, verificationtests and basictunnel calibrations are intimatelyrelated. The accuracyrequirements may vary,depending upon the type of calibration task and theprimary purpose of the facility, but are most stringentfor items 4 and 6 sinceerrors in the measurements can contribute directly to the random or fixed error in the final data.

4 0FACILITY

1 FORCE PRESSURE & I'TEMPERATITRF: TRANSIENTS I

Figure l.C.1. DATAAND ERROR FLOW DIAGRAM, Ref. 7 R. J.: "FurtherCorrelation of Data From Investigationsof a High Subsonic-Speed Transport Aircraft Model in ThreeMajor Wind Tunne 1 AlAA Paper 71-291, Albuquerque, New Mexico, 1971-

2. Pritchard, J. L.:"The Dawn of Aerodynamics,''Journal of theRoyal AeronauticalSociety, V. 61, March 1957.

3. Randers-Pehrson, H. H. : "Pioneerblind Tunnels," V. 93, No. 4, Smithsonian MiscellaneousCollect ions.

4. Bird, K. D.: "Old Tyme Wind Tunnels,"Perspective Sept.-OCt., 1957, Cornel1Aeronautical Laboratories, Inc.

5. Goin, K. L.: "The History,Evolution, and Use of Wind TunneIs," AlAA StudentJour., Feb. 1971.

6. Goethert, B. H.: Transonic Wind TunnelTesting, pp 2-31, Perrnagon Press, New York, 1961.

7. Picklesimer, J. R., Lowe, W. H., and Cumrning, D. P. : ''A Study of Expected Data Precision in The Proposed AEDC HlRT Facility," AEDC-TR- 75-61, August, 1975.

6 I I. TUNNEL VARIABLES

II.A.l Types of Tunnels

The materialpresented herein is directed toward wind tunnels operating in the Machnumber range from 0.4 to 3.5. Themodes of operation of the variousfacilities surveyed include: (1) continuousflow, (2) blowdown, and (3) intermittent.

In thecase of intermittenttunnels, e.g., a Ludwiegtube, the very shortrun times require special provisions for measurement and recording systems. Pressure measurementscan be accomplished using eitherhigh-response pressuretransducers or a capturesystem which permits measurements of pressureafter the run. However, the same basicprocedures must be followed in orderto calibrate the facility as for a long-run-timefacility. Thus, the specialproblems associated with the short run times of intermittent tunnels arenot discussed, but the general discussions of calibration procedures are applicable.

Althoughtransonic tunnels with high-aspect-ratio (2-D) testsections aregenerally operated at higher Reynolds numbers, thistype of tunnel is not discussedseparately becausethey share the same calibration problemsas sym- metricaltunnels.

Discussionsof the various topics are of a generalnature where possible. Subdivisionsinto transonic and supersonicareas are made where dictatedby thepeculiarities of these regions. Further subdivisions are made, as appropriate,in discussions of details.

7 1I.B. OPERATIONAL PARAMETERS

II.B.l Pressure Control

Pressure controls are incorporated in some form in all wind tunnels (with the possible exception of supersonic,indraft tunnels). The methods of control are obviously different for transonic and supersonic wind tunnels and for intermittent, blowdown and continuous wind tunnels. This section is limited to discussions of pressure control systems as they influence tunnel calibration programs and the effects of variations introducedby these systems on tunnel flow quality and measurement accuracy.

Continuous Wind Tunnels Continuous wind tunnels may be either pressure tunnels or atmospheric- vented tunnels. For the pressure or variable density wind tunnel, the stagnation pressure is determined by the static or wind-off pressure and the pressure added by the fan or compressor drive system. The drive system pressure ratio may be controlled by varying compressor speed, blade angle,or auide-vane ang 1 e.

Vented tunnels usually operate at atmospheric stagnation pressure,but some facilities of this type operate at atmospheric test section static pressure or theatmospheric vent may be located at some other part of the tunnel circuit so that neither the stagnation nor the static pressureis atmospheric. The drive system is controlled by the same techniquesas for the pressure tunnel to achieve the desired pressure ratio across the nozzleand test section.

For supersonic tunnels the Mach number(and all Mach dependent test section conditions) are determinedby the nozzle geometry and stagnation conditions. The supersonic nozzle is not normally considered a static pressure control (although it does perform that function). Several tunnels include automatic control of nozzle geometry. The pressure control is simplest for the atmospheric-stagnation-pressure, supersonictunnel since the prime function of the drive system is to create the pressure ratio necessary to startand maintain nozzle flow. For pressure tunnels, both the tunnel pressurization and main drive system control the stagnation pressure.

8 Transonictunnel operation requires additional control of the test section staticpressure. In addition to control of compressorpressure ratio,the static pressureis controlled by some typeof plenum evacuation system. At supersonic Mach numbers aboveabout 1.4 a variable geometry,convergent-divergent nozzle is usually used. Also, a specified Yach number can be attainedover a range of tunnelpressure ratios byother control variables. Tunnel .pressure ratio may therefore beone ofthe variables investigated for tunnel flow optimization, in terms ofboth flow uniformity and minimum power consumption. Plenum evacuation canbe accompli shedby ejector flaps which usethe main stream flow to pump the plenum, or auxi 1 iary pumping systemscan beused.

Almost all ofthe continuous tunnels responding to the questionna ire use ma nua 1 con t ro1 of total and staticpressure, although several haveava i lable automatic systems toindicate the measured testconditions to the operators, and a fewinclude closed-loop, automatic control. The response of a continuous tunnel,particularly a large one, tocontrol inputs is influenced by thetime constantsinvolved. These timeconstants are a functionof the circulating air mass, therotational inertia of themain drive, etc., and aregenerally large. A beneficialeffect of thelarge time constants is that short-term disturbances tend to be heavilyattenuated andsmoothed. Precise, smooth controlis possible by manual control,but changes inlevel require a longerperiod than for small timeconstant systems. Fluctuationsin the controlled pressure tend to occur atthe system naturalfrequency, which is theinverse of thetime constant. The period of thesefluctuations can be verylong - up to 10-15 seconds. In order to obtain a measurement of the mean valueof tunnel flow conditions, measure- mentsover atleast one periodare required.

Blowdown Wind Tunnels The control systems for blowdown windtunnels can have a significanteffect ontunnel flow quality. In addition to the automatic stagnation pressure control system, automaticcontrol systems are used in a majorityof the transonic, blowdown windtunnels for Mach number controlalso.

The stagnationpressure control for ablowdown windtunnel uses a control valve between thestorage reservoir and the stilling chamber tocontrol pressure inthe chamber. Constantstagnation pressure is the normal mode ofoperation,

9 butthe system can also be computer or program controlled to maintain constant Reynolds number asthe stagnation temperature drops during the run, or the pressure may be increasedlinearly with time to investigate Reynolds number effects,explore flutter boundaries, etc. In general, the functional capability ofthe stagnation pressure system has become more sophisticated with the intro - duction of digital computer control.

From theflow quality standpoint, the mostimportant performance parameter of the system isthe accuracy of pressure control or, in more specificterms, thevariance of the stagnation pressure about the mean level. The periodof this variation is typically about 1 secondand thecurrent state-of-the-art appears to be about a 0.1 percentstandard deviation; much largerperturbations, up to 1/2%, can easilyresult due toelectrical noise, mechanical friction or misadjustment ofthe control computer. The stagnationpressure control system

must operate continuously to overcome thedisturbance created by the decreasir ‘9 reservoirpressure. Thusa controllerof higher order than that used for a continuoustunnel pressure control is usually required to achievethe desired accuracy. A simpleregulator is normally inadequate.

The shocksystem generated downstream ofthe blowdown-wind-tunnel Control valve may introduceexcessive flow unsteadiness. Test-section flow angularity may alsovary with valve position (and therefore,time). Thus, theentire flow channel,from the storage reservoir to the stilling chamber, mustbe considered when designingthe stagnation pressure control system. Considerablework has been accomplished inrecent years to identify and correct flow problemscaused bythe stagnation pressure control system. Corrective measureshave included choked-flowdevices in series downstream of thevalve, specialized valves, acousticsilencers and honeycombs in the stilling chamber.

A second pressurecontrol system used intransonic tunnels, the Mach number control,functions to maintain a desired Mach number by controlling staticpressure as a functionof stagnation pressure. Almost all blowdown, transonictunnels use a choked throat downstream of the test section to control subsonic Mach numbers.The primaryadvantage ofthis control mode isthat the Machnumber is determinedby the test section geometry (at a fixed model angle ofattack) and istherefore independent of fluctuationsin stagnation pressure.

10 Automaticcontrol of the downstream throatarea is used in anumber of facilities.Automatic control is highly desirable in order to maintain constant Mach number during model attitude variations and simultaneouslymaintain optimum plenumevacuation. More sophisticated operational modes areavailable under computer control, suchas Machnumber sweeps, etc. The performance of this system also directly influences the variation of the test section Machnumber. Currentbest performance appears to be 'about 0.001. butlarger variations are possibleat subsonic speeds, particularly at high model-pitch rates.

In the absence ofperturbations introduced by the Mach orstatic pressure controlloop, small variations in stagnation pressure cannot necessarily be takeninto account by simultaneous measurement of the twopressures because of phase lag and attenuationerrors.

An importantprocedural difference in making calibration measurements in a blowdown tunnel is thatruns should be made at each calibrated Machnumber where all tunnelvariables are held constant during an entire blowdown, in orderto detect time or valve position dependent effects. If a traversing angularityprobe is moved alongthe tunnel centerline during the run, for example, time-dependent effects will be obscuredby the spatial variations and viceversa.

Intermittent(Impulse) Wind Tunnels Intermittentwind tunnels are considered to bethose that operate in a basic blowdown mode, butwith a runtime of about 4 to 5 seconds or less. The Ludwieg tunnelis a typicalfacility of this class. The Ludwieg principle canbe applied to either a supersonicor a transonicwind tunnel. Pressure controlis limited to the initial charge tube pressure and istherefore rela- tivelystraightforward. An advantage of theLudwieg tunnel isthat the stagnationpressure downstream ofthe initial expansion tube is constant (neglecting viscouseffects). The primarycalibration measurement problemsassociated with the Ludwiegtunnel obviously arise from theshort test duration.

11 ll.B.2 CalibrationAccuracy, Flow Uniformity and Relationshipto Model Testing

The calibration of a transonicwind tunnel is significantly more difficult than calibration of a supersonictunnel due primarily to the ventilated test sectionwalls. The ventilatedwalls and thebasic nature of transonic flow preventthe determination of test section conditions from tunnelor nozzle geom- etryalone, as is thecase with a calibratedsupersonic tunnel nozzle. A measurement oftest section static pressure, in addition to stagnationpressure, isrequired during calibration and routinetest operations. Further, for fixed testsection geometry,the model orother apparatus in the test section can influence the Mach number. These factorsrequire that the tunnel calibration provide a relation between thestatic pressure distribution in the test section anda referencepressure measured in the plenum chamber or onthe ventilated wall.

Transonictunnel calibration is further complicated by theadditional degrees of freedom providedby a ventilatedwall, i.e., at each Mach number, the optimum wallangle, wall porosity (for adjustable porosity walls), plenum evacuationflow rate,tunnel pressure ratio, and choke controlposition must all be determined. Criteria for optimum adjustmentinclude uniformity of Mach number and, at supersonic speeds, shock and expansion-wave cancellation characteristics which are usuallyevaluated based on testsof cone-cylinder models. At subsonic speeds, in additionto minimizing variations in Mach number distribution, other criteria for optimizationare tunnel noise level and forces ona standardmodel. A recent report byJackson (Ref. I) provides a comprehensivediscussion of the procedures employed inselectina transonic tunnel parameters to minimize Machnumber variations.

Many ofthe transonic tunnels surveyed determine the wall angle basedon shock and expansion wave cancellationat supersonic speeds, and this angle is often maintained constant at all Mach numbers, whileothers adjust the wall angleaccording to a Machnumber schedule. Ingeneral, adjustment of wall angle with Machnumber will provide amore uniformflow.

A typicaloptimization problem at subsonic Mach numbers isbalancing of plenum evacuation and chokearea for a choke-controlled blowdown tunnel. The averagetest section Mach number can be attained with an infinite number of combinationsof plenum pumpingand chokearea. For example, the criterion usually chosen is to minimize downstream Mach number increasesor decreases from

12 theupstream value. Downstream disturbancesin Machnumber areundesirable becausethey can create bouyancy effects further upstream.Since the disturb- ancemagnitude isextremely sensitive to changes in plenumpumping at subsonic Mach numbers belowabout 0.85, the optimumpumping is determinedduring calibration and maintainedconstant for routine testing. The testsection Machnumber is controlled by varying the chokearea which does not alter the downstream disturbance.

A similarupstream disturbance occurs at Mach numbers near 1.0. Therefore, one of thepurposes of a calibration program is to determinethe region of flow along the test section within which the Machnumber deviation does not exceed variouslimits such as +O.OOl, 20.002, etc.Jackson (Ref. I) hassuggested the following criteria beadopted as anindustry standard for "good flow quality" in transonictunnels. For subsonic flows, 2a deviationsin centerline Machnumber shouldbe less than 0.005 and lessthan 0.01 inthe case of supersonicflows. Of course,the minimum Machnumber deviationis indicative of the best distribution and thereforeflow quality for a giventest section length and setof tunnel conditions.Jackson's flow quality criteria are shown inFig. 2.8.1 as a functionof Machnumber. Recent calibrationdata from the AEDC-PWT 16T TransonicTunnel isalso included for comparison.

Morris and Winter(Ref. 2) havesuggested even more stringentrequirements for supersonictunnels. These investigators havesuggested the maximum allowed variations in (I) flow angularity be+0.1 - degand (2) Machnumber be 20.003 at M = 1.4, +0.005 at M = 2, -+0.01 at !I = 3.

It should be noticedthat criteria based onthe standard deviation do not distinguish between random or periodic variations and mean flowgradients. Thus, in addition to standarddeviation criteria, consideration must be givento empty- tunnelstatic pressure gradients. The staticpressure distribution along the testsection mustbe eitherconstant (within acceptable limits) or any gradient mustbe known andrepeatable to a sufficiently high degree of accuracy so that bouyancy correctionscan be made to attain the required accuracy in measurements of model drag. It istherefore of interest to investigate, in a systematic manner, the effects of testsection pressure gradient on drag measurement accuracy and how thisrelates to flow quality requirements. 0.024

0.020 l-

0.01 6 - z - 0 AEDC-PWT 16T DATA 0- I- ,/ -a ""JACKSON S CR ITER I A FOR "GOOD" > FLON QUAL 1 TY / :0.0 12 3 I 0.008 u t.' " a I: -"

0.004

0 0 0.2 0.4 0.6 0.8 1 .o 1.2 1.4 1.6 1.8 TESTSECTION MACH NUMBER, Mm

Figure 2.8.1 JACKSON'S FLOW QUALITYCRITERIA FORTRANSONIC TUNNELS, Ref. 1. The bouyancy drag coefficient resulting from a linearstatic pressure gradient (Ref. 3) can be stated as

AC = - ” dP/dx , (2.8.1) DG qw where V isthe model volume, S is wingreference area, qm isthe average test section dynamicpressure, dP/dx isthe pressure gradient and AC isthe DG dragcoefficient increment produced by thepressure gradient. Utilizingthe above equation,lsaacs (Ref. 4) investigatedthe effects of bouyancyon thedrag of typical, transport aircraft models in a 2.44-111 (8-ft) windtunnel. Based onmodel valuesof the parameter V/S ranging from 0.069 to 0.208 meter (0.23 to 0.68 ft), lsaacs determined that --dP should be known to q,dx an accuracyof 0.00047 to 0.0014 permeter (0.00014 to 0.00043 per ft) inorder to know AC to an accuracyof 0.0001, i .e. , one dragcount. DG In a studyof bouyancy effects on drag measurement accuracy in supersonic windtunnels, Morris andWinter (Ref. 2) determinedthe allowable pressure gradientfor a bouyancydrag of 1% ofthe model drag. Based on an assumed, rectangular-wing,aircraft modeland Eq. 2.8.1, theallowable pressure gradientin terms of AP/H overthe model length was foundto range from 0.002 at M - 1.4 to 0.0005 at M = 3.0.The corresponding Machnumber gradientover the model length was approximately 0.4% ofthe average Mach number. The estimated dragcoefficient of theconfiguration considered indicated 1% of ACD was 0.00023 at M = 1.4 and0.00013 at ?l= 3.0. On a per-drag-countbasis, the allowable Mach number gradient,in percent of average Mach number, was then 0.17% at M = 1.4 and 0.31% at M -- 3.0.

Bouyancy effects may be evaluatedin a generalized way by taking into accountboth model configurationvariables and Mach number effects. Assuminga specificheat ratio of 1.4, therelations

P 2 -3.5 -x (1 0.2 ) (2.B.2) H + M

qW 2 -3.5 -D 0.7 M2(1 + 0.2 M H (2.B.3)

15 may beused to write Eq. (2.6.1)as #- - 2 (2. e.4) M (1+0.2 Hz)

Where H and qm areconsidered constant at their average values. If the Mach num- dM bergradient is assumed to be 1 inear, =may be written as AM/Ax with Ax takenas the model length, L,. AM is thenthe Machnumber variationper model length. Eq. (2.6.4) becomes

AC = "[ 2 ] AM . (2.6.5) DG SLm M(I+O.Z M')

The parameter V/SLm is a nondimensionalconfiguration parameter and is there- foreindependent of model scale.Figure 2.8.2 shows theallowable Machnumber gradient,over the model length,for a bouyancy-induced,drag coefficient error of 0.0001 as a functionof the confiauration parameter V/SL Owing tothis m' extreme sensitivity of drag measurements accurate to within one count,there are anumber of problems inachieving this goal. For example, if the random, shortwavelength variations in Machnumber aretoo large, the mean gradient may be obscured and difficult to def.ine. One approach is to useempty-tunnel pressuredistributions, measured duringcalibrations, to integrate Over the model length. However, this procedurecan be inerror because oflack of exact repeatabilityof tunnel flow conditions. lrt thecase oftransonic tunnels, the model may inducedepartures from empty-tunnel cal ibrations, e.g., Parker(Ref. 5) Inaddition, Jackson(Ref. 1 ) has foundthat a chanqe in unit Reynolds number 6 6 from 4.1 x 10 to 15.8 x 10 (permeter) can cause an increase of 0.003 intunnel Mach number, see Fig. 2.6.3. This is an effectthat is frequently ignored duringtransonic tunnel calibrations.

The dataof Fig. 2.6.2 arealso shown inFig. 2.B.4 withthe Machnumber gradientexpressed in percent of the averaae Machnumber. Pointsderived from the criteria suggestedby Morris and Winter'(Ref. 2) forsupersonic flow are shown onFig. 2.8.4 for comparison.This comparison indicates the model configuration used by Morris and Winter to establish flow uniformity criteria hada value of approximately 0.05 for V/SLm.

16 V "- sLm L/ 2

3 TESTSECTION MACH NUHEER

~l~ure2.8.2 ALLOWABLELINEAR MACH NUMBER GRADIENT OVER MODELLENGTH FOR BOUYANCYDRAG COEFFICIENTCQNTRIBUTION OF 0.0001

17 0.020

0.016

0.012

Hm - Mc

0.008

0.004 -

0 0 1 .o 2.0 3.0 4.0 5.0 6.0 7.0 Re x 10-6/ft

I~ 1 I 1 1 1 0 4 8 12 16 20 Re x 10'6/m

Flgure 2.8.3 EFFECTS OF REYNOLDSNUMBER ON CALIBRATION OF THE PWf-16T TUNNELAT M_ = 0.6 AND 0.8 FOR Ow = 0 AND T = 6%

18 I

0 .5 1.0 1.5 2.0 2.5 3.0 3.5 TESTSECTION MACH NUMBER, M

Figure 2.8.4. MACHNUHHER GRADIENT OVER NOOELLENGTH AS PERCENT OF AVERAGEMACH NUMBER FOR RQUYANCY DRAG COEFFICIENTOF 0.0001 The valueof the parameter V/SLm forseveral aircraft typical OF fighter, 2. attack and transportconfigurations are listed below.”

Aircraft W/SL -m

F-1 5 0.054 F-16 0.048 YF-17 0.043 A-7 0.071 oc-8 0.061 DC-9 0. 088 DC-IO 0.083 8-741 0.065 8-727-100 0.076 8-727-200 0.056 C-141A 0.055 C -5A 0.078

The above datademonstrates the variation in V/SLm with aircraft type is notlarge, at least Forconventional configurations, and thatthe model configurationselected by Horris and Winter (Ref.2) is representativeof supersonicfighter aircraft. It is anticipatedthat V/STOL configurations would have a largervalue of V/SL, thanthe aircraft listed above and would therefore be more sensitive to Hach number qradienteffects.

* Due tothe approximate values used for some ofthe aircraft volumes, the valuesof W/SL, shouldbe regarded as approximate.

20 I I .D. P.eferences

1. Jackson, F. M.: "Calibrationof the AEDC-PWT 16-Ft TransonicTunnel at TestSection \,la11 Porositiesof Two. Four, and SixPercent." AEDC-TR-76-13, Jan. 1976.

2. Morris, 0. E. and &linter, K. G.: "Requirements forUniformity of Flow in Supersonic \,!ind Tunnels," RAE Tech Note AERO 2340 (l?54).

3. Glauert, H.: "Wind Tunnel Interference onVinqs, Yodies and Ailerons," A.R.C. R&M 1566 (1933).

4. Isaacs, 0.: "Calibrationof the R.A.E. Bedford 8-ft. x 3-ft. Wind Tunnel at Subsonic Speeds, Includino a Discussion of theCorrections Applied to the Measured Pressure Distribution to Allow for the Direct and Crlockaae Effects due tothe Calibration Probe Shape,': A.R.C. R&M 3583 (1569).

5. Parker, P.. L.: "Flow GenerationProperties of FiveTransonic Wind Tunnel TestSection Wall Confiaurations," AEOC-TR-75-73, Aug. 1975. 1I.C. FLOW PARAMETERS AND UtKERTAlNTY RELATIONSHIPS

The proper measurement of stream properties to allow the accurate determination of the various flow parameters is necessary for the meaningful

interpretation of wind tunnel test results. For example, the desirability of a Mach number accuracy of 0.001 has been suggested (i.e., Ref. I). The necessity of such a requirement may be illustratedby the afterbody dataof Fig. 2.C.I.

This data appearsto have substantial scatter but may be correlated using Mach number measurements with a precision of 0.001 as shown in Fig. 2.c.2.* It also may be noted that for a typical fighter aircraft configuration the transonic drag riseis such that a Mach numt.er uncertaintyof 0.001 is "equivalent" to 0,0002 (2 counts) in drao coefficient. Similarly. other parameters must be computed to high degrees of accuracy. The sensitivities of the several flow parameters to the various measurements are presented in this sectionto illustrate the consequencesof measurement uncertainty on accuracy.

II.C.1.Pressures

The pressure of a fluid is one of its most significant properties, The knowledge of static and stagnation pressures in a wind tunnel is necessary to define characteristic flow conditions such as Mach number and Reynolds number and to properly normalize the various data coefficients.The following discussion concerns the measurementof these two pressures.

Static Pressure: During transonic operation static pressure is obtained from a reference pressure (wall or plenum) and a predetermined relation (calibration) of this pressure to the test section static pressure. During supersonic operation static pressure is usually obtained from stagnation pressure and the Mach number previously obtained during calibrationof the facility with the particular -nozzle setting. Figures 2.C.l and 2 were obtained through private communication with Mr.Jack Runkel. NASA Langley Research Center. This requirement for a Mach number accuracy Of at least 0.001 is also substantiated by the recent nozzle-afterbody tests reported by Spratley and Thompson (Ref. 17). 22 NASALANGLEY TAILINTERFERENCE MODEL M = .95

.28

cD .24

1 2 3 4 5 6 "jdP

Figure 2.C.l AFTERBODY DRAG DATA AT ANAVERAGE MACHNUMBER OF 0.95 h) W NASA LANGLEY TAILINTERFERENCE MODEL

.2a

.24 %

.20

.16

Figure 2.C.2 AFTERBODYDRAG DATA WITHTUNNEL MACHNUMBER GIVEN TOTHREE DECIMALS Inthe transonic region a staticpressure probe,or array of probes

may beused torelate the reference and testsection static pressures. With

regard to thehigher Mach numbers it has been illustrated (Ref.2), that the

uncertaintyin Mach number may be related to uncertainties in static P-,

isentropictotal, Hs, and Pitot, H2, pressuresby the following relations

(assuming theratio of specific heats is 1.4):

"-aHS a H2 35 (M2 - 112 = aM[ 3 (2.c.1) HS H2 (M2 + 5) (7M2 - 1)

If it is assumed thatthe total pressure is measured inthe stilling

(2.c.2) a nd

Solving for inthe first equation and substitutinginto the latter, -H thefollowing expression is obtained.

"aH2 5 (M2-1)2 'Pm - =o (2.C.4) H2 M2 ( 7M2- 1 ) Po0 which yields

aH2 H2 2. 2 5 -1) -I- (M (2.C.5) pm a M2(7M2-l) '

2 3.5 2.5 Since -"2 = [ F] [4"] ,then aH2- (2.C.6) pm 7H -1 a pm

canbe simplified to:

Hence the ratio of uncertaintyof Pitot-to-static pressure becomes a simple

functionof Mach number and is shown inFig. 2.C.3. It may benoted that the

ratio becomes 1 near M 1.6. Thus, for a specifiederror in Mach number at an

M 1.6, theerror in static pressure may be greaterthan ,.the error in * Pitotpressure. For Mach numbers greaterthan 1.6 thereverse is true. This occursbecause the staticpressure becomes verysmall at high Mach numbers, andsmall absolute errors in the .measurement of Pm produce relatively large errorsin calculated Mach number. For example, Fig. 2.C.3 shows thatat

Mach 3 theabsolute error in Pitot pressure canbe approximately seven times

thestatic pressure error for the same errorin calculated Mach number. Thus

the use of staticpressure for the determination of Mach number is generally restricted to Mach numbers lessthan 1.6; whilePitot pressure is employed

(withstagnation pressure) at higher Mach numbers.

* Also see Fig. 2.C.9,p.39.

26 5

2 StagnationPressure: The pressure of thetest medium is measured withthe fluld at rest either in the settling chamber or by means of a total headtube.

The settling chamber (isentropicstagnation) pressure, Hs, isgenerally used for bothtransonic and supersonicoperat ion. Because of the afore- mentioned sensitivity of Mach number to static pressure, after-shock total

(pitot)pressure, H2, is employedabove a nominal 1.6 Mach number.

Dynamic Pressure: Dynamic pressure, q, is perhapsthe most frequently employed flow parameterused tonormalize wind tunnel data. Thus theaccuracy of q is directlyreflected in the accuracy of coefficient data. In mostinstances, after static pressure hasbeen obtained bymeasurement (transonic)or by

inference(supersonic) it is used with Mach number to compute q from

q = M2Pm . (2.c.8)

Inthe transonic range,both Pm and !i are measured. Errorsin either affect S Mach number. Fig. 2.C.4shows thesensitivity of q to HS whichresults solely from Mach number erroras determined from the following.

2 Since q+pm ,

= YMP- -aM , (2.C.9) aHS a HS and ( aq/q )/(aHS/Hs) - 5 = - (aM/M)/(aHS/Hs)= 2 . q aHs

It will subsequently be illustrated (seeSection Il.C.3) that 5 2 aM/(aHS/Hs) 31 7~ (1 + .2M ,

(2.C. IO)

28 I

C -5 1.0 1.5 2 .o Mach Number Figure 2.C.4 THE SENSITIVITY OF DYNAbEC PRESSURE To STAGNATION PRESSURE ERROR, TRANSONIC OPERATION Equation 2.C.10may besubstituted into the preceding equation to obtain

(2.c.11)

Errors in Pm affect q by means of the Pm term and theerroneous Mach number as illustrated in Fig. 2.C.5. From q = 7Y2 M Pm :

aM 2 = yMPm - + -$ M (2.c.12) a a

1 (2.C.13)

I-(aM/(aPm/Pm)) I (2.C. 14) M + .

It will beshown inSection Il.C.3 that

2 aM/ (aPJP,) = (1 +.2M ) , (2.C.15) - -7M which upon substitution yields

2 (aq/q)/(aPm/Pm) 1 - 7~z(M2+5) . (2.C.16)

DuringSupersonic operation, calibrated Mach numbers are known forthe facility geometry setting and are employed with H forq determination. . S However, an error in defining the calibrated Mach number will affect q as shown inFig. 2.C.6. The functionillustrated in this figure was obtained as follows:

30 1

I I 0 I I 05 1.0 Mach er

-2

-3

-4

Figure 2.C.5 THEI Sli3KLTIVITY OF DYNAMIC PRESSURE 'K) STATIC PRESSURE ERROR, TRANSONIC OPERATION

31 2

\ Mach Number

-1

-2

-3

-4

Figure 2.c.6 THE SENSITIVITY OF DYNAMIC PRESSURE To MACH NUMBER ERROR, SUPERSONIC OPERATION (2.C.'17)

(2.C. 18)

= 2 + (aP /P )/,(~M/M) (2.C. 19) WOD

As shown inSection ll.C.3,

5 2 aM/(aPw/Pw) = - -7M (1 + .2M ) , (2.C.20)

hence 2 (aM/M)/(aP /P 1 = - - (I + .2M )- (2.C.21) ww 7M2

Then

(2.C.22)

(2.C.23)

In a similar manner, errors in HS canbe shown to havea one-to-onerelationship with errors in q.

At lowsubsonic Mach numbers, thepressure ratio Pw/HS approaches unity, so that determination of the Mach number anddynamic pressurefrom measurements ofthe individual pressures becomes increasinglyInaccurate. At these

low Mach numbers (belowabout 0.4)a preferredprocedure is to measure the differential (Hs - P,) directly with a low rangetransducer and to compute the dynamicpressure from:

(2.C.24)

33 At low Mach numbers, only the first term of the'seriesis usually required.

For example, the error is only 0.14 percent at M = 0.5 using only the first term. At M = 1.0,the first three terms yieldresults accurate to 0.1 percent.

I I .C.2 Temperature

As a fundamental stateproperty, stream (static) temperature is of substantialimportance in establishing the character of thefluid flow.

Thus anaccurate value of temperature is required in wind tunnel testing to determineseveral correlation parameters which define the nature of the flow.

The determination of static temperature in a gasstream conventionally involvesan indirect measurement. Stagnationtemperature is a convenient measurement to make since it is relatively easy to obtain, and thereare establishedprocedures for computing static temperature from thestagnation value and flow Mach number. Figure 2.C.7 illustratesthe relation of stagnation-to-statictemperature for a perfect gas (y = 1.4) in an adiabatic process. This relation'(To/T = 1 + 2My-1 2) is used in wind tunnels which operateat moderate pressures and temperatures and where real gas effects arenegligible. It can beseen that an errorin the measurement of stagnation temperature, To, isdirectly reflected in the static temperature.

34 Mach Number I I. C. 3 t,iach Number

As previouslydiscussed Mach number is computedusi.ng settling chamber

pressure and either static pressure or after-shock (Pitot) stagnationpres-

sure. Inthe transonic region, Mach number is computedfrom

(2.C.25)

The sens it ivity of Mach number to sett ling chamber pressure measurementcan

be derived byobtaining the partial de rivative of the above expression with

respectto HS, i.e.,

(2. C .27)

Thisexpression may be non-dimensionalitedto obtain

2 aM/(aHS/Hs) = (-4"5 -7M 7 (2.C.28) pW

Similarly,the non-dimensional sensitivity of Mach number to Poo is found to

2 aH/(aPw/Pw)= (1 + .2M ) , - -7M (2.C.30)

which i 1 lustrates that ":$,;. . . .. aM/(aPw/Pw) = - aM/(aHS/Hs) . (2.C.31) The relationof stagnation pressure behind a normalshock, H2, to HS

as aMach number function is:

(2.C.32)

Thisrelation will notyield an explicitexpression for Mach number, therefore, the sensitivity of Machnumber to H2 was evaluatedusing a numer ical,finite interval approach. As previously shown

(2.C.33)

These sensitivitiesare illustrated in Fig. 2.C.8.

This figure consistently showsa larger Machnumber errorper percent

errorin Hs and H2 thanper percent error in H and Pa However, when S . nominalvalues of HS, PW and H aresubstituted appropriately, the relative 2 magnitude of Machnumber errorper N/m2 errorin the measurement illustrates

thesuperiority of H2 over PIP at supersonic speeds (seeFig. 2.C.9).

37 2.4

2 .o

1.6

1.2

0.8

0.4

for M = f (HS, P,)

0 1 2 3 4 5 Mach Number

Figure 2.C .El THE SENSITIVIm OF h4ZH NUMBER TO STATIC PRESSURE ANI) STAGNATION PRESSURE ERROR Hs = 2.75 x 105 N/m2

-1600

-1300

"200

-1OOO x lo6 N/m2

-800

-600

-400 L -2oc

C 1 2 3 4 5 Mach Number

Figure 2 .c. 9 THE SENSITMTY OF MACH NUMBER "0 STATIC PRESSW AND STAGNATION PRESSURES I1.C.4 FlowAngularity and Curvature

F low angu larity and curvature can result from nozzlecontour errors,

irregu larities ordiscontinuities in the internal surfaces of a tunnel,

in-flowor out-flow due to leakage, and swirl or curvaturepropagated from

upstream of thenozzle or contraction. The resultingnon-uniformity pro-

duces localperturbations in the flow which result in gradients or varia-

tionsin flow properties including static pressure and therefore, Mach

number (seeFig. 2.C. 10). Thus stepsare takento dissipate these disturb-

ancesby means appropriate to the particu lartunnel configuration. These

correctiveactions include nozzle contour corrections,installation of

honeycombs inregions of low Machnumber flow, and more recently,perforated

platesin regions wherean uncontrolled,high-pressure-ratio shockdown would

generateadditional undesirable perturbations (Ref. 4).

Because ofthe acute sensitivity of certain model configurationsto non-uniformityof flow suchas localflow direction and Mach number, it

is necessary to define via calibration any flowanomalies that may exist

inthe test section. Probes for measuringflow angularity are discussed in

Section I I1.E.

40 0.6

0.5

I cu

0, 0.2

Ed2 Fu

0.1

c & 2 3 4 5 6 7 Mach Number

Figure 2.C .10 CHANGE IN FLOW DIIECTION WIT)i INCRplENT OF MACH “BER, Ref. 3

41 I I .C.5 Reynolds Number

The ratioof inertial to viscous forces in the test medium is

obtained from wind-tunnel measurements asa dimensional unit Reynolds

number given by PU R/E = - (2.C.34) lJ

Thiscan be expressed in units of m-l in terms of To, M and Pm as fol lows:

6 P,,H R/I1 = 2.29 x 10 - (1 + .2M ) To2 ’ 2, 2, ( 1 :.,,I (2.C.35)

6 HSM R/E = 2.29 x 10 TZ (1 + .2H 2 ) 1.5

Since P,/Q is alinear function of Pa and HS, thesensitivity to these parameters is one-to-one;that is, agiven error in either of ,these will

be reflectedin the same percenterror in R/R. However, inthe transonic

range poo and H are used toobtain Mach number which is alsoa variable in S theabove expressions. Thus errorsin P,, and Hs can be reflectedin R/%

througherrors in M. Figures 2.C.11 and 2.C.12 illustrate these sensitivities 6 forselected unit Reynolds numbers (5, 25,50 C 100 x 10 /meter) at anominal stagnationtemperature of 311 OK (100 OF). The sensitivity Df .?eynolds number to measurements of stagnationtemperature is shown in Fig. 2.C.13.

Intunnels where calibrated Mach numbers areobtained andconsidered constantfor subsequentoperation with the same facility configurations, any errors in Machnumber due to calibration or adissimilar configuration will con-

tributeto errors in R/Q . Thiseffect is shown inFig. 2.C.14.

42 Figure 2. C .11 TI& SENSITiVI!iT OF UNIT REYNOLDS NUMBER TO STATIC PRESSURE EHROR

43 100

x 60

I I I 1 2 3 4 5 hch €,Jwnber

Figure 2 .C .12 THE SENSITIVITY OF UNIT REYPIOLDS FWER TO STAGNATiON PRESSURE ERRCR

44 To = 3UoK

-160 2- \ \ ”

-40

I I I I I 1 2 3 4 5 Mach Number

45 1 2 3 4 5 Mach Number

Figure 2 .C .14 THE SENSITIVITY OF UNIT REYNOLDS NUMBER To MACH NUMBER ERROR

46 ll.C.6. Unsteadiness,Turbulence and Noise

Large,continuous flow tunnels often have small-amplitude, low-frequency oscillationsin the mean flowconditions. For example, the11-Ft Transonic

Tunnel at NASA Ames has a characteristicperiod of approximately 10 seconds.

Of course, thistype of variation should be calibrated and used to establish

routinetesting procedures.

Accordingto ‘Jestley (Ref. 5), measurements in AEDC, Langley, and Modane windtunnels indicate the maximum axial and transverseturbulence intensities areapproximately 1.0% and 0.4% for Mach numbers near one. InPart 11.2 of

Ref. 6, it is notedthat wind tunnel turbulence hasbeen used at MLR (Netherlands) and ONERA (France) toexcite model flutter modes.” However, Time (Ref. 7) cautions,that turbulence not only can mask the initiation of flutter, but may also excite response modes whichare not true flutter modes. Also,Time pointsout that Mabey (RAE) foundthe transonic buffet boundary to be very sensitiveto flow unsteadiness, In addition, freestream turbulence introduces errorsin static pressure measurements (seeSection 1II.D) and affects bounda ry

layertransition, separation phenomena atleading and trailing edges, and shock- boundarylayer interactions.

The followingare known to be sources ofnoise in transon icwind tunnels:

1. porous wa 11s which can generatedistinct frequenc ies known as

edgetones and/ororgan tones,

2. slottedwalls which generate broad-band disturbances due to

shearingin the slots between themovino air in the test

section and the air in the surrounding plenum chamber, * A preferredprocedure would appear to be a controlled excitation of the model via either a mechanicalexcitor, pressure pulse generator, or loudspeakers.

47 3. reverberation of tunnelwalls,

4. plenum chamber surges, 5. turbulent boundary layersalong the tunnel walls, 6. diffuserflow instability,

7. compressors incontinuous wind tunnels,

8. controlvalves in blowdown windtunnels,

9. vibrationof tunnel sidewalls,

10. workingsection cutouts, and

11. model supports and struts.

,The noisesources, which usually dominate at various Mach numbers, areindicated in Figs. 2.C.15 and 16.

It is noted inthe review paperby b!estiey (Ref. 5) that CL buffet max ’ onset, transonicdrag rise, boundarylayer transition and separati.on, sk n.

friction drag, shock shapes and locat.ions, etc.’,, may’ali. .I be.affected by I. tunnel -generatednoise. Hence, windtunnel data will not be representat ve of free-flightconditions in caseswhere this, istrue. Our presentstate of knowledge does notallow a quantitative definition of the complex interactions between turbulence,noise, and aerodynamic testingin wind tunnels. The funda- mentalobjective of currentresearch in this ‘area is to obtain a betterunder- .. standing of this phenomena via a systematictesting programwhich uses standardized

instrumentation. P. list of 25 recommendationsconcludes the paDer by Westley

(Ref. 5). Theserecommendations mainlyconsist of:

(1)decisions which need to be made tostandard izeinstrumentat ion

and testprocedures, and

(2) new experimentalprograms. TRANSONIC

I I Y M,< 0.3 TUR.BULENCE u C Dominant Souta (valves. compressor 1

H I DIFFUSER p 4- "

0.3 < M,< I WALL HOLE I RESONANCE

" I+ II n 1 JET No'SE IC

Figure 2.C.15 FIxlW DISTURBANCES IN TRANSONIC TUNNELS, Ref. 5 SUPERSONIC - HYPERSONIC - I I M,. I -3 Usually - Dominant

M, = 3 - IO RADIATED NOISE .( c Usuolly 20 (cold flow) ""-Dominant M, > IO ENTROPY Arc tunnels Shock 9 MHD .

Pigure 2.c.16 i!TQW DISTURBANCES IN SUPERSOKIC AND HYPERSO1:IC TUNNELS, Ref. 5 One ofthe primary recamnondations is thatstandard Instrtnnentation beadopted to measure free-streamdisturbances. This probim of noise measurements in transonictunnels is discussedIn Section 1II.F.

51 I I.C,7 Humidity

The acceleration of air from rest involves the reduction of static pressure and temperature. Such expanslon to evenmoderate speeds results inthe rapid approach to water-vaporsaturation. Figure 2.C.17 illustrates this condltion in terms of the ratio of the relative humidity of the stream to that of air at rest as in a reservoir. The extentof the effect of conden- sationon aerodynamic testdata, and thusthe amount ofcondensation which can be tolerated, has not been firmlyestabl ished (Ref. 8). Forexample, the investigationreported by Norton,et al. (Ref. 9) indicatesvery little differencein data obtained on the same model in moist air ascompared with thatobtained in dry air.

In the absence of a watersurface or a precipitant(such as a droplet or foreignnuclei), humid air can be cooledwell beyond thetheoretical satura- tionpoint before condensation occurs. This is because theprocess istime dependent and therate of expansion (which defines the temperature history of theflow and isusually related to the tunnel size) defines the amount of 0 supercoolingthat can be attained.Supercooling of as much as 100 C hasbeen 0 experimentally measured usingsubstantial temperature gradients (100 C/cm), e.g., Ref. 10;and theoretical work hasbeen accomplishedwhich indicates that thesaturation vaporpressure may beexceeded by a factorof 4, Ref. 11. It hasbeen demonstratedthat supercooling of 30 OC can be accomplished with negligiblelikelihood of condensation, Ref. 12. However, even withthis tolerance, it may be seen inFigure 2.C.18 that for an arbitrary dew point of 2 OC extremereservoir temperatures would be required to avoid condensat ion at low supersonlc Mach numbers. Therefore, it isgenerally not practical (because ofairstream stagnation temperature 1 imits, suchas those of Ref. 8) to employ reservoirheating as a means foravoiding condensation. In practice, air dryers are usually used to reduce dew points to as lowas practical ; althoughthis may be above thestream temperature, the total water content is small, and condensationeffects are negligible.

As noted by Pope and Goin(Ref. 12), the effect which humidity has on tunnel Machnumber depends onwhether the flow is subsonic or supersonic. In the case of subsonicflow, water vapor tends toincrease the Mach number and

52 100

Mach Number Figure 2 .C .17 THE RATIO OF RELATIVE HUMIDITY IN THE STREAM TO RESERVDIR AS A FUNCTION OF MACH NUMBER

53 Assumptions: Dew Point Temperature = 2 c Allowable Supercooling = 30 C Allowable Stream Temperature = -29 c

1 2 3 4 Mach Number Figure 2 .c .l8 RESERVOIR TEMPERATLTRE: REQUIRED TO AVOID CONDEXSATION, Ref. IO

54 reduce staticpressure; whereas, theopposite occurs in supersonic flow. This effect has also been substantiatedby analyses at AEDC.* These resultsindicate a negative Machnumber gradientoccurs when moisture condenses in supersonic flow.

The absence of condensationduring tunnel calibration (i.e., empty tunnel) -doesnot preclude the possibility of local condensation in proximity of a model duringproduction testing. It has been observed inthe AEDC Aerodynamic Wind Tunnel 4T (Transonic 4T) thattransonic.force data is unaffected bymoisture contentuntil condensation can be seen (anominal water-vapor content of 0.002 gm/gm of air). However, testsinvolving surface pressure measurements are more sensitive, and experienceat AEDC indicatesthis type of transonic ** testing shouldbe conducted withhumidity -< 0.0015 gm H20/gm air. An additional procedure forreducing the effect of humidity in transonic tunnels is to adjust wallangle according to the test medium dew point, e.g.,Ref. 8. Inthe super- sonicregime, experience at NASA Ames has shown that 0.0004 gm H20/gm of air is a good rule-of-thumbfor model testswith M 3.5.' Forexample, mass flow through an inlet model is found tovary about 1% at M = 3.0 when themoisture contentvaries from 0.0002 to 0,001. Because ofthe facility variables which affectthe allowable moisture content, it isdesirable to establish the level whichcan be tolerated in a particular facllity by conductingtests on arepre- sentativeconfiguration and varyingonly humidity. This type of test was included inthe workreported by Corson, et al. (Ref. 8).

* Privatecommunication, Mr. J. D. Gray, AEDC. ff Privatecommunication, Mr. J. Gunn, AEDC. ' Privatecommunication, Mr. F. W. Steinle, NASA Ames. 55 I I .C.8 Test Mediums

Air is almost,universally usedas the test medium in transonic and

supersonicwind tunnels. Although these facilities have different operating

characteristics with the air being subjected to different pressure and tempera-

turelevels during the various cycles, it isgenerally allowable to consider

the gas to beideal. Real gas effects may become relevantat extremeconditions

such as in a Ludwiegtube facility (Ref. 6). Departures from theideal gas

relations may occur when othertest mediums are employed.However, it has been

found thatthe ideal relations are suitable for the very low temperature

nitrogen used inthe Langley IO-Meter Transonic Cryogentc Tunnel (Ref. 14).

Recenttransonic wind tunnel tests of airfoils have indicatedan effect of

varying y , Refs. 15 and 16. Althoughno effect was detected for subcritical

flows, a systematicreduction in local peak Mach numbers was observedfor

supercriticalflows. Tutla, et al. (Ref. 16) suggest thistrend is assoclated

withthe effects of y on transonic-shock/boundary-layer interactions.This is

relevantto the calibration of empty, transonictunnels if a conventional,

static-pressureprobe is used to measure freestream Mach number in different

test gases, As discussed inSection lll.D.2, a transonic shock alwaysforms

ona conventionalstatic-pressure probe. If variationsin y can affect super-

..cr,itical pressure distributions, this may also change the location on a probe

atwhich freestream pressure exists. Research on this phenomena iscontinuing

at NASA Ames.

56 1 I .C. References

1. Hi11, Jacques A. F. et al, "MachNumber Measurements in High SpeedWind Tunnels," MIT, NavalSupersonic Laboratory Technical Report 145, Februaty 1956.

2. Thompson, J.. S. and Holder, D. Y., "Noteson Wind TunnelPressure Measurements from the Operator's Point of View," RAE TN Aero.2547, February 1958.

3. Raney, D. J., "Flow Direction Measurements in Supersonic Wind Tunnels," Her Majesty'sStationery Office, London 1956.

4. Cooksey, J. M. and Arnold,J. W., "Flow quality Improvements in a Blowdown Wind TunnelUsing a Multiple Shock EntranceDiffuser,'' AlAA Journal, Vol. 10, No. 9, September 1973.

6. MiniLaWsWorking Group, "A Further Review ofCurrent Research Aimed at the Design and Functionof Large Windtunnels," AGARD-AR-83, Sept. 1975.

7. Time, A., "Effectsof Turbulence and Noise onWind-Tunnel Measurements

At Transonic Speeds ,I' AGARD-R-602, Apr i 1 1973.

8. Corson, Blake, W., etal. "Calibrations of the Langley 16-foot Transonic Tunnel withTest Section Air Removal," NASA TR R-423, August 1974.

9. Norton,Harry T. Jr., Runckel,Jack F., and Pendergraft,Odis C. Jr., "Tran- sonicPerformance of Two Convergent-DivergentEjector Nozzles Designed forCorrected SecondaryFlows of 3 and 9.4 Percent," NASA TM X-909, 1964.

IO. Lundquist, G. A., "Recent Experimental Work at NOL on Condensation in Compressible Flows,"Geophysical Research Paper No. 37, ARDL, July 1955.

57 14. Adcock, Jerry B., Kilgore,Robert A. an'dRay, Edward J., "CryogenicNitrogen asa Transonic Wind TunnelTest Gas," AlAA Paper 75-143, January 1975.

IS. Gross, A. R. and Steinle, F. W.: "PressureData from a 64010 Airfoilat

Transonic Speeds in Heavy Gas Media of Ratio of Specific Heatsfrom 1.67 to

1.12 ,'I NASA TH X-62468, Aug. 1975.

16. Tuzla, K.; Wai,J. C.; and Russell, 3. A.: "y-Effectson 2-Dimensional Transonic

Aerodynamics,"Proc. AIM 9th Aerodynamic TestingConference, June 1976.

17. Spratley, A. B., Thompson, E. R., and Kennedy, T. L.: l'Reynolds Number and

and Nozzle Afterbody Configuration Effects on Model Forebody and Afterbody Drag,"

AIAA Paper77-103, January 1977.

58 111. CALIBRATION PROCEDURES AND lNSTRUMENTATlON

A. Settling Chamber Pressure

As discussed in Section 11.A., there'servoi r total pressure is a fundamental variable which is usually measured directly in the settlin$chambers of both transonic and supersonictunnels. The Machnumber and dynamic * pressure inthe settling chamber aredetermined by thecontraction ratio, Ao/A , where A. is thecross sectional area at the settling chamber and A* is the choked throatarea correspondingto the test section Mach number. The maximum stilling chamber Mach number normallyoccurs at Mach 1.0 in a transonic-supersonictunnel. At a contraction ratio of IO,.for example, the corresponding stilling chamber Mach number is 0.058. The flow may be consideredincompressible and theratio of dynamic tostagnation pressure determined frbm

-qS = "5 - ps = 1 - (-+,pS HS HS HS S

where (PS/Hs) isdefined by thesettling chamber Mach number. At a contraction ratio of 10, the stilling chamber dynamic pressureis 0.235 percentof the stagnationpressure. Thus, theerror in measured total head, inducedby using a static orifice in place of a Pitot probe,would be 0.235 percent.This would contribute a Mach number errorof 0.002. Therefore, if a Mach number accuracy of 0.001 is to be achieved and static orifices are used to measure settling chamber pressure,the error must be eliminated via calibration with Pitot probes.

When using a Pitot probe to calibrate total pressure in a settling chamber, theprobe must be located downstream of anyscreens, honeycombs, etc.,since theseitems can cause significantpressure losses, Ref. 1. Also,the chamber crosssectlon should be surveyed for variationsin total pressure. If a single valueof total pressure is to be used(as is comnonlydone) and itscontribution to Machnumber error is to be lessthan 0.001, then 2a of spatial variations in totalpressure must be lessthan 0.05 percent (EAM = 0.0005 at M = 0.80). Un- fortunately, this is not only near the state-of-the-art of pressure measurement accuracy, it isalso very difficult to achieve this uniformity in practice. Thus, thedecision as to what is an acceptable amount ofnonuniformity in settling chamber pressure mustbe leftto individual judgment. Thisdecision * The terms"Settling chamber"and "stilling chamber" are usedinterchangeably. 59 should be based on the particular facility characteristics* and the type of testswhich are conducted in that facili,ty.

Once the spatial variations in settling chamber pressureare judged to be acceptable, it is suggested that an appropriateaverage be defined basedon measurements inthe central portion of the flow. For example, inthe 16 ft. TransonicTunnel at NASA Langley,four Pitot probes have beenmounted in the centralportion of the flow to define anaverage (Ref. 2). Ingeneral, in-itial calibrations require more measurements in order to establish a suitable average. However, oncethe average total pressure is determined for the range of operating conditions, a simplewall mounted tube (or a static orifice) can be calibrated torelate its measurements tothe average. By followingthis procedure, routine test,ingcan be accomplishedwithout any unnecessary obstructions in the central portion of the flow.

Although a widevariety of Pitot probe nosegeometries have been used in low speed flows,simple steel tubing with an internalto external diameter ratio -B 0.5 and a square-cut nose will measure total pressure in the settling ** chamber withnegligible error. A Pitot probe withthis diameter ratio is unaffectedby flow angles of 10 degrees or less, Ref. 3. Assumina that reasonablecare is taken to align the probe with the flow, this type of probe will provideadequate accuracy even if considerableturbulence exists in the settling chamber. Thisconclusion is substantiated bythe following discussion.

The problem of Pitot probe measurements in an incompressible,turbulent flow hasbeen examined by Becker and Brown (Ref. 4). These authors have analyzeddata for four different probe geometries: (1) spherical-nosed probe(a sphere ona tubularsupport), (2) a hemispherical-nosedtube, (3) a square-nosedtube, and (4) sharp-1ipped probes made by conicallytapering the exteriorof a tube. The resultsof their semi-empirical analysis for square- nosed probesindicates the following. In an isotropic,turbulent flow with a turbulenceintensity of 5 percent, a square-nosedprobe with a diameterratio * A number ofsupersonic tunnels have fixed-contour,sliding block nozzles which areroutinely operated off design. These nozzlescan have significant total pressurelosses which can only be determined by Pitot surveys within the test section. However, theaverage test section total pressure could be related to stilling chamber pressurevia calibration tests. A* This assumes the nose isfree of burrs. Finishing of orifices is briefly discussed inSection I 11.0.4.

60 of 0.5 will capturethe total pressure with an error of 0.56 x 10-4 q. For a given amount ofturbulence, the error decreases with increasing diameter ratio. Thisaccuracy is more than ample for most tunnelssince hot-wire measurements at AEDC inthe settling chamber ofthe Aerodynamic Wind Tunnel (4T), Ref. 5, 3. indicatethe longitudinal component of theturbulence intensity is of theorder, of one percentfor 0.3 M 1.2. Assuming isotropicturbulence, this means I. thetotal turbulence intensity is approximately 1.73 percent. Thus, the suggestedsquare-nosed probe can be used in most settling chambers withconfi- ; dence.

The above describedaccuracy analysis ignores a number of other possible' sources oferror. It is assumed theprobe nose islong enough toisolate it from any effectsof downstreamgeometry. Becker and Brown (Ref. 4) suggests the nose length be greaterthan six probe diameters. Also, the effect of ;i. changes inthe internal diameter is ignored. In order to eliminate internal geometry asa variable, Becker and Brown suggestthe internal diameter be constantfor a distanceof three probe diameters. In addition, the probe should be located more thantwo diameters from the nearest wall in order to avoid a reductionin measured pressure (e.g.,Ref. 3, p. 12).

Finally,the probe should be designed and mounted tominimize vibration. Winternitz(Ref. 6) has presented a simplifiedprocedure for designing canti- levered,circular cylinders to avoid oscillations induced by vortexshedding. Ower and Pankhurst(Ref. 7, p. 54) observethat for a cylinderwith a diameter of 0.8 cm (916 in.)the vortex shedding frequency in air is 40 Hz at 1.5 m/sec and 160 Hz at 6 m/sec. Hence, theyconclude resonance between vortex frequency and thenatural frequency of theprobe is unlikely in most windtunnel applications. However, in some cases thiscould be a problem atthe low speeds characteristic of stilling chambers. Thus,probes for measurements inthe stilling chamber should be designed to avoid this phenomenon.

f: The problem of internal geometry changes causingbiasing of measured mean pressuresin fluctuating flows is briefly discussed in Ref. 3, p.105. 1II.A. References

1. Loehrke, R. 1. and Nagib, H. M.: "Experimentson Management of Free-StreamTurbulence," AGARD-R-598, Sept. 1972.

2. Corson, B. W., Jr. ; Runckel, J. F. ; and Igoe, #. B. : "Calibration of

theLangley l6-Foot Transonic Tunnel with Test Section Air Removal ,'I NASA TR-R-423,Aug. 1974.

3. Bryer, D. W. and Pankhurst, R. C.: Pressure-ProbeMethods forDetermining WindSpeed and FlowDirection, National Physical Laboratory, Her Majesty's StationeryOffice, London, 1971.

4. Becker, H. A. andBrown, A. P. G.: "Response of Pitot Probes inTurbulent Strearns,l'Jour. Fluid Mech., Vol. 62, Part 1, 8 Jan. 1974.

5. Credle, 0. P.: "An Evaluationof the Fluctuating Airborne Environment

in the AEDC-PWT 16-Ft TransonicTunnel ,I1 AEDC-TR-69-236, NOV. 1969.

6. Winternitz, F. A. L.: "Effectsof Vibration on Pitot ProbeReadings," The Engineer, 1/01. 201, 30Mar. 1956, pp. 273-275and 6 April 1956, pp. 228-290,London.

7. Ower, E. and Pankhurst, R. C.: The Measurement of Air Flow, Pergamon Press, London, 1966.

62 I I I . B. TOTAL TEMPERATURE

The totaltemperature is normally monitored.in the stilling chamber during routinetunnel operation. Since the difference between total and static temperature is smallat low velocities, a shielded,high-recovery thermocouple probe isnot usually necessary. In fact, data obtained by Stickney(Ref. 1) forthese twotypes of probes show thatthe recovery factors are nearly identical ("0.999) for temperaturesnear ambient and M < 0.2. Thus, Pope and Goin(Ref. 2) note that in many cases thetotal temperature can bemeasured in the stilling chamber,* withsatisfactory accuracy, by using a simplebare-wire thermocouple junction. A schematic ofthis type of temperature probe is shown in the upper part of Figure 3.8.1. Measurementsby Stickney(Ref. 1) indicate that such unshielded temperatureprobes have amuch shorterresponse time compared to more elaborate, shieldedprobes. In the case of blowdown tunnelswhere total temperature can varyrapidly, this is an essentialadvantage. For example, if testsare con- ductedat constant Reynolds numbers, thetotal temperature mustbe monitored continuously so thattotal pressure canbe controlledautomatically. Also, smallwire thermocouples with time constants of the order of 0.1 sec. are typicallyrequired. For example, a 0.13 mm (0.005 in.)diameter wire has a * >*: timeconstant of 0.1 sec. in air at ambienttemperature and pressure and a velocityof 19.8 m/sec (65 ft/sec). Whereas, forthe same conditions, a 0.53 mm (0.021")diameter wire has a 1.0sec time constant, e.g., Ref. 4.

In response tothe questionnaire , themajori ty of tunneloperators ind i ca ted they do infact use thebare-wire thermocouple for total temperature measurements. Estimatedaccuracies varied from +0.56"C to 21.1"C (21°F to22°F). Based on therelations presented in Section ll.C.2, an uncertaintyin total temperature of 1°C will cause, at M = 1, a maximum uncertaintyof 0.5 percentat a Reynolds num- berper meter of 33 million.For most testing purposes this is acceptable. However, fewtunnels (transonic or supersonic) appear to havebeen cal ibrated for tempera- turegradients which may existacross and alongthe flow. * A comprehensivediscussion of thermocouple principles, circuits, electromotive forcetables, stability and compatibilitydata, installation techniques, etc. may be found in Ref. 3.

IA r. I. The timeconstant is heredefined as thetime required to reach 63.2% of an instantaneoustemperature change. 63 Typical Bare-Wire Probe

/Two-hole ceramicholder .229 O.D. x .033 Wall t

All Dimensions In Centimeters

AEDC-IWT 16~Probe (Ref. 5 7.635 O.D. x .089 Wall 1.12 R -.+” 1.27 v“ h v 30.48

.478 O.D. x .081 Wall L 4 Vent holes(0.17) equally spaced

Figure 3 .B. 1 TOTAL !E!&PEMTUFU3 PROBES One of the most complete and extensivecalibrati.on of temperature gradients in a transonictunnel has beendone inthe AEDC Propulsion Wind Tunnel (16T), Ref. 5. The temperaturecalibration was done toinvestigate the effects ofa special-purpose,cryogenic cooling system which consists of a liquid nitrogen system to chill the coolant in thetunnel cooler and a liquid air system for direct injectioninto the tunnel airstream. A rectangulararray of shielded temperatureprobes was locatedin the nozzle contraction region and inthe test section. A schematic of a typicalprobe is shown inFigure 3.8.1.*Since therecovery factor of all thermocouple probes need to be calibratedfor Mach and Reynolds number effects (Ref. l), the raw temperaturedata were first correctedfor these effects by Robson (Ref. 5). Subsequently,the temperature ofthe flow through the central portion of the nozzle entrance section was defined byan average ofthirteen temperatures measured over a 2 x 3.5 m (6 x 11 ft) rectangularregion. The temperatureof the test section flow was defined by an average of 17 temperaturesobtained over a2 x 2 m (6 x 6 ft) portion of thecore. The difference betweenthese two temperatures was used todefine a temperaturecalibration parameter which relates temperature at the nozzle entranceto test section temperature. In this case, thetest section flow was found to be approximately 1.1"C (2°F)lower than the nozzle flow. Deviations of 28°F wereobtained across both the nozzle and thetest section over a Mach number range0.2 to 0.8 and -22°C < To < 21°C. These detailedtemperature measurements were made because ofthe anticipated nonuniformities produced by -1. .L I\ ,. thespecial cooling system. Although smaller temperature gradients usually existin tunnels without special cooling or heatingsystems, this example illustratesthe procedure required to accurately calibrate wind tunnel temperatures. For routinetesting, a singletemperature probe can be relatedto the average stilling chamber temperaturevia calibration in order to eliminate the disturbing effects of anunnecessary thermocouple grid.

.L "Robson (Ref. 5) statesthat the copper-constantan thermocouples used in this probeare generally considered to have a systematicerror of +2.2"C (24°F).

-8. I I. I. Temperatures inthe 1/3-MeterTransonic Cryogenic Tunnel at NASA Langley havealso, been surveyed using a grid of thermocoupleprobes, Ref. 6.

65 Additionalinformation on thedesign and calibrationof total temperature probescan be found in Refs.7-10. Also, Bate (Ref. 11) has reviewedthe problem of errors in thermocouple measurements basedon experiencein the DFVLR windtunnels in WestGermany.

66 I

I I I. 6. References

1. Stickney, T. M.: "Recovery and Time-Response Characteristicsof Six ThermocoupleProbes in Subsonicand Supersonic Flow," NACA TN 3455, July 1955.

2. Pope, A. and Goin, K. L.:High-speed Wind TunnelTesting, Wiley, New York, 1965.

3. AmericanSociety for Testing and Materials (ASTM), Committee E20: Manual onthe Use of Thermocouples in Temperature Measurement, ASTM Special TechnicalPublication No. 470, Philadelphia, Pa., 1974.

4. The Omega Temperature Measurement Handbook, Omega Engineering,Inc., Stamford, Conn., 1975.

5. Robson, G. D.: "TestSection Temperature Calibratlon of the AEDC PWT 16-Ft TransonicTunnel atStagnation Temperaturesfrom -30 to 3OoF," AEDC-TR-69-2, Feb. 1969.

6. Polhamus, E. C. ; Ki lgore, R. A. ; Adcock,J. 6. ; and Ray, E. J. : "The Langley HighReynolds Number Wind-Tunnel Program," Astro. 5 Aero.,Oct. 1974.

7. Benedict, R. P.: Fundamentals of Temperature,Pressure, and Flow Measurements, Wiley, New York, 1969.

8. Baker, H. D.; Ryder, E.A.; and Baker, FI. H.: Temperature Measurement in Engineering,Vol. II, Wiley, New York, 1961.

9. Volluz, R. J.: "Handbook of SupersonicAerodynamics, Section 20, Wind TunnelInstrumentation and Design," NAVORD Rept. 1488 (Vol.6), 1961.

10. Dean, R. C., Jr.:Aerodynamic Measurements,MIT Gas Turbine Lab., Eagle Enterprises, New York, 1953.

11. Bate, J.: "Temperature Measurements in Wind Tunnels," RAE Lib.Transl. No. 1736, AD 922 120, Farnborough,Hants, England, June 1974.

67 1II.C. PITOT PRESSURES

Use of Pitot Pressures for Calibration As describedin Section II.C.l, when M > 1.6 the uncertainty in calculated testsection Mach number isless if thecalculation is based on Pitotpressure ratherthan freestream static pressure. Thus, mostsupersonic tunnels have been calibratedvia Pitot probe surveys and assumingan isentropicexpansion from thestilling chamber. Inthe past, investigators such as Hill (Ref. 1) and Hill, etal. (Ref.2) have reportedthat measurements insmall, supersonic tunnels (<0.5 m) of theratio of total pressure in the test section to reservoir pressure exhibit a range of 0.998 -+ 0.003. This type of resul t leads to the con- elusion thatnonisentropic expansion effects are negligible at normaloperating temperatures and pressuresin a properlydesigned supersonic tunnel, i.e., one in whichthe empty tunnelis free of shocks. However, largecontinuous tunnels areoften operated at relatively high humidity levels in order to increase the operatingtime prior to dryer saturation. For example,Maxwell and Hartley (Ref. 3) found in the AEDC-PWT 165 Tunnel that when thehumidity was 0.002 gm H20/gm of dry air, the average totalpressure of the test section was 2 to 6% lowerthan the reservoir pressure. This loss was reduced 50%by decreasing t tunnelhumidity to 0.001.

Inaddition to water vapor condensation, oblique shocks and real gas effects cancause a lossin total pressure. Also, large tunnels can havenon- uniformities** in total pressure causedby incomplete mixing in the stilling chamber, and smalltunnels can have losses caused by axialvelocity gradients (e.g., Ref. 4). Withthis number ofpossible causes oftotal pressure variation, it is recommended thatoperators of bothtransonic and supersonictunnels make calibration measurements tovalidate the assumption of uniform total pressure. This can be accomplished in subsonicflow via a Pitot probe,since it can be used directly to compare testsection total pressure with reservoir pressure. Insupersonic flow, another independent pressure must be measured suchas free- streamstatic, surface pressure on,a cone or wedge, or Pitotpressure behind an oblique shock. Once a choiceis made, the two pressures and theratio of specific

?t 2 2 These results wereobtained with 2 .o < M < 4.75, 3 .1 Wcm < Hs < 9.1 N/cm , 55 OC < To < 78 OC. A* This can be determinedduring sett ling chamber cal ibretion,Section 1II.A. 68 heatscan be used tocalculate the test-section Machnumber and tota pressure.

Barry(Ref. 5) has discussedin detail the errors thatoccur in computed Mach numberwhen usingpressures measured withtotal, static, conica , and wedge probes. A significantconclusion, obtained by Barry(Ref. 5) and Thompson and Holder(Ref. 6), is thatthe Mach number computed from a pressureratio involvingthe isentropic stagnation pressure is lesssensitive to measurement errors. For this reason,an isentropic stagnation pressure probe has been designed and tested by Goodyer (Ref. 7). The probeconsists of a Pitottube mountedon thesurface of a curvedcylinder of circular cross section. The Pitot tubesenses the impact pressure of a streamtube which has beenslowed tosubsonic speedby isentropiccompression along the leading edge of the curvedcylinder. A sketchof the probe is shown InFigure 3.C.l. The independent experimentalresults of Couch (Ref. 8) indicatethat this type of probe permits measurements of absolutestagnation pressures with an accuracy of 99.8 percent in aMach number range of 1.4 to 2.2. BeyondaMach number of 2.2, thepressure recoverydecays and theprobe ceases to offer any advar,tqeover a conventional probe. However, for Mach numbers lessthan 2.2 thestagnation pressure probe canbe used fordirect measurement of totalpressure loss. Also, for the case of equaluncertainty in measured pressures and 1.6 M < 2.2, theanalysis of Barry(Ref. 5) indicatesthe most accuratecalibration of Mach number would be obtained by usingthe isentropic probe in conjunction with a conventionalpitot probe.

If a supersonictunnel engineer elects not to use a Goodyer probebecause of its limited Mach number range,the next most accuratetunnel calibration procedure is to measure Pitotpressures in the freestream and on awedge, Ref. 5. Thisprocedure has beenused at anumber offacilities with success. Perhaps the most sophisticated use of this method has been developed atthe AEDC Propulsion Wind Tunnel,Ref. 3. A variableangle wedge with a movable pitot tubenear each surface is used inthe 16s facility. The purpose ofthe variableangle feature is tooptimize the wedge angle and therebyeliminate uncertaintyin effective angle caused by changes in boundary layergrowth. Ineffect, this feature eliminates wedge angleuncertainty in the calculation of Mach number. The complete Mach number probeincludes two conventional Pitot Dimensions In Centimeters

Straight Section

Dia .

Figure 3 .C .l. ISENTROPIC STAGNATION PRESSURE PROBE, Ref. 8 probeslocated outboard and alignedwith the leading edge ofthe wedge. A planview of this Mach number probe is shown on theright in Fig. 3.C.2. One of theseprobes has been usedon a sting to calibrate the empty test section of 165. These same data have beenused tocalibrate a retractableversion of theprobe which is mounted inthe ceiling of the test section. When fully extended,the wedge center1ine is 58.4 cm (23 in.) from theceiling. This permits routine Mach number measurements without the uncertainty of an isentropic expansionassumption. The interestedreader may refer to Reference 3 foraddi- tionaldetails.

A second type of Mach number probehas been employed in Tunnel A at the AEDC Von Karman Facility.This probe measures staticpressure onthe surface of a retractabledisk. The supporting arm 1s a 15 deg included-angle wedge and hasa smallPitot probe mounted belowthe disk. A schematic ofthis probe is shown onthe leftin Figure 3.C.2. Althoughthis probe is susceptible to leading edgeand angleof attack errors, it may be calibrated by conventional,sting 3: mounted probes and has theimportant feature of simplicity.

Inthe case ofintermittent tunnels, e.g., a Ludwieg Tube, a differenttype of Machnumber probe isrequired because ofthe short run-time and thepossi- bility of rapid changes intest-section flow. The AGARD TechnicalWorking Group, which isresponsible for selection and design of theLarge European High-Reynolds- Number TransonicWindtunnel (LEHRT), has recommended theprobe shown inFig. 3.C.3. As reportedby ROSS and Hartzuiker(Ref. g), thisminiature probe utilizes highfrequency-responsepressure transducers for measurement ofboth Pitot and static pressures and is designed to be used inthe small-scale pilot LEHRT facilities. The primarypurpose of this probe is tomonitor temporal changes in mean Mach number. Dougherty (AEDC) has pointedout to the present authors that measurements of static pressurefluctuations with this probe will have a limitedfrequency response and thusshould not be used to calibrate static pressure fluctuations associated with noiseand/or turbulence. However, thefluctuating Pitot pressures can be used for this purpose; see Section 1II.F. forfurther discussion of measurements of unsteady f 1 ow d is turbances.

Pitot Probesfor Freestream Calibration

Although a wide variety of Pitot nosegeometries have been used, thesimple cylindricaltube with square-cut nose is adequate forfreestream calibrations. For an internal to externaldiameter ratio of 0.125, thetests of Gracey(Ref. IO) R Compared tothe other two Machnumber probes, thistype of probe hasan additional disadvantage.Barry’s analysis (Ref. 5) shows theuncertainty in calculated Mach number isgreater when thereis equal uncertaintyin measured pressures. 71 Tunnel A Tunnel 16 S .051 orifice t 7-A 1 0'

II.'Pitot 31 75

1. Disknot presently Cali- bra ted w i th C, probe data.

1. Variableangle (10'-26') pitot wedge probe mounted on ceiling;extends center of wedge 58.4 cm from ceiling. 2. Calibratedwith identical wedge located at tunnel G, . Dimensions In Centimeters,

Figure 3.C .2. AEDC SUPERSON IC MACH NUMBER PROBES Scale 4:l Pitot and Static Measuring Transducers are Kullte CQL-062-50. Type, Mounted in Silastimer Compound in Probe.

,!. ,!. Static Holes 0.5 dia

d 1 r .- I -" I ? I t 3"

-Scale 1:l f-

-18d

U W demonstratedthat total pressure measurements withthis probe will be in error by 0.01q' at anangle of attack of 11 degrees and M,= 0.26 and1.62. 0 This same accuracy was attained at a = 223 , M = 0.26 and a = -+29O, M = 1.62 by increasingthe diameter ratio to 0;96. The angleof attack range was increasedeven moreby usinginternal.bevelling to increase the diameter ratio to near one. However, sinceflow angularity in empty testsections seldomexceeds 1 or 2 degrees, a tubewith a straight impactopening and a diameterratio of 0.5, or more, will provideimpact pressures with negligible error. (Of course, this assumes theprobe isfree of burrs.)

Effectsof Various Parameters on Pitot Probes -Size: Earlyexperiments with Pitot probes showed measuredpressures to be independent of probesize, e.g., Ref. 11. Thus, sizingis usually guided by the size of facility and Machnumber at which the probe is to be used. When totalpressure gradients are present, a Pitot probesenses an impact pressurecorresponding to a displacementtowards the higher pressure, Ref. 12. This effect decreases with probesize and withincreasina wall thickness. However, Livesey(Ref. 13) foundthat a conical nose Pitot,with a sharp edge atthe opening, isbest for use in a transversepressure gradient since it exhibits a negligibledisplacement error. Butsince cones cannot be used veryclose to walls, two dimensionalboundary layer measurements areusually made with Pitot probeshaving very small, flattened-oval openings anda square-cutnose, e.g., Refs. 11 and 12.

Analysisof data for the simple, circular Pitot tube indicates the measured pressure is independent of Reynolds number (based on insiderad ius of theopening) when it is greaterthan 100, Ref. 12.

Mach number: In dry air, the Pitottube has aenerally been found to be insensitiveto Machnumber and wil 1 reliablyprov idethe freestream stagnation pressure at subsonic speedsand thestagnation pressurebehind a normalshock at superson ic speeds.

Tu rbu1 ence : The incompressibleanalysis of Becker and Brown (Ref. 14) indicatesthat a circulartube Pitot with square-cut nose is relatively insensitive to turbulence. However, theseauthors suggest that the length of the constant diameter

74 opening be atleast three diameters. The intent is to eliminatesurging of the flow, in response toturbulence, and thusassure the existence of stagnation conditionsprior to changes ininternal geometry. When. theresults of Becker and Brown are viewed inlight of the data obtained byGracey (Ref. lo), which demonstrates decreasing flow angle sensitivity of Pitot probes with increasing Mach number, one may concludethat circular tube Pitot probesare unaffected by low levelsof turbulence. Since the turbulence intensity in most empty tunnels is lessthan two percent,the recommended Pitot probes(i.e., circular tubes with internaljexternaldiameter ratios > 0.5) can be used withconfidence to calibrate transonic and supersonicwind tunnels.

Rakes, Arrays and Supports,:

Insubsonic flows, impact pressure can be successfully measured with an orifice in a circularcylinder mounted normal tothe flow, e.g., Ref. 12. Thus, Pitot probesare generally considered to be insensitiveto support arrangements. However, near Mach one the bow shockgenerated by a supportcould conceivably interferewith the probe shock. The resultingpressure measuredbehind the shock interactionswould be expected todiffer from thenormal shock pressure. Thus, attransonic speeds the nose of the Pitot probeshould extend far enough forward toavoid this problem. A tubelength at least 12 timesthe support thickness is recommended. At subsonicspeeds, Dudtini ski and Krause(Ref. 15) haveobserved thatthe effect of proximity of a transversecylindrical supporting strut is negligible if the strut is two or more strutdiameters downstream from thePitot tubetip. For supersonic applications, Pope and Goin(Ref. 16, p. 353) notethat thePitot tube length is usually I5 to 20 tubediameters.

When severalPitot probes are used in a rakeor an array,the measured pressures may be affected by interactions between the bow waves on adjacenttubes. Bryer andPankhurst (Ref. 12) note that experiments at H = 1.6 indicatethe gap between Pitot probes may be assmall as one diameterwithout causing significant error. AsMach number decreasestoward one, theseparation distance must be increased. In subsonicflow, the spacing of Pitot probes is generally not considered to be critical, e.g., Ref. 17.

75 II .C. References

1. Hill, J. A. F.: "On the Cal Aero. Sci ., Vol . 22, No. 6, 2. Hill, J. A. F.; Baron, J. R. Schindel', L. H.: "MachNumber Measurements in High-speed Wind Tunnels,"HIT Naval Supersonic Laboratory Tech. Rept. 145, Jan. 1956 (Alsoavailable as AGARDograph 22,Oct. 1956).

3. Maxwell, H. and Hartley, M. S. : "Aerodynamic CalibrationResults for the

AEDC-PWT 16-Ft.Supersonic Tunnel at MachNumbers from 1.50 to 4.75 ,I' AEDC-TR-69- 102, May 1969.

4. Murphy, J. S.: "Evidences of an InherentError in Measurement of Total-Head Pressure at Supersonic Speeds," Aero.Engr. Rev.,Nov. 1953.

5. Barry, F. W.: "Determination of MachNumber From Pressure Measurements,'' Trans. ASME, April 1956.

6. Thompson, J. S. and Holder, D. W. : "Notes on Wind TunnelPressure Measure- mentsfrom the Operator's Point of View," R.A.E. Tech. NoteAero. 2547, Feb. 1958.

7. Goodyer, M. J.: "A New Probe forthe Direct Measurement of Stagnation Pressure in Supersonic Flow," R.A.E. Tech. Rept. 73122, March 1974.

8. Couch,L. H.: "Effectsof Geometric Variables on thePerformance of a Probe for Direct Measurement of Free-StreamStagnation Pressure in Super- sonic Flow," NASA TN 0-7887, May 1975.

9. Ross, R. and Hartzuiker, J. P.: "Recommended Flow Quality Measurements in LEHRT Pilot Facilities," AGARD TechnicalWorking Group AC/243 (PG.7/WG. 1) Action 61, June 2, 1975.

10. Gracey, W.: "Wind-Tunnel Invest igation of a Number of Total-Pressure Tubes at HighAngles of Attack," NACA Rept. 1303, Jan. 1956.

11. Folson, R. G.: "Review ofthe Pitot Tube," Trans ASME, Oct. 1956.

12. Bryer, D. W. and Pankhurst, R. C.: Pressure-Probe Methods forDetermining WindSpeed and FlowDirection, National Physical Laboratory, Her Majesty's StationeryOffice, London, 1971.

76 13. Livesey, J. L.: "The Behavior of Transverse Cy1 indrical and ForwardFacing TotalPressure Probes in TransverseTotal Pressure Gradients," Jour. Aero. Sci., Vol. 23, p. 949, Oct. 1956.

14. Becker, H. A. and Brown, A. P. G.: "Response of Pitot Probes in Turbulent Streams,'' Jour. Fluid Mech., Vol. 62, Part 1, 8 Jan.. 1974.

15. Dudziniski, T. J. and Krause, L; N.: "Effect of Inlet Geometry.onFlow-Angle Characteristics of Miniature Total-Pressure Tubes," NASA TN 0-6406, July 1971.

16. Pope, A. and Goin, R. L.: High-speed Wind TunnelTesting, Wiley, 1965.

17. Chew, W. L., Jr.: "Calibration of FiveTotal Pressure and Temperature Survey Rakes at Speeds from M = 0.2 to 1.0," AEDC TN-59-37, May 1959.

77 I I I.D. TESTSECTION STATIC PRESSURES

As discussed in Section II.C.l, measurement ofstatic pressure is

fundamental totransonic wind tunnel calibration. It iscurrently standard

practice to measure settling chamber pressure and empty-tunnel static to

calibratesubsonic and transonictunnels when M < 1.6. Typically, anaverage ofstatic pressure data, measured alongthe centerline, is used to calibrate a referencepressure. in order to avoid interference, it is generally considered good practice not to measure thisreference pressure with a probe

permanently mounted inthe test section. Thus, thereference pressure is usually measured eitherin the plenum chamber or atsidewall orifices located

inthe forward portion of the test section. Once calibrated,the reference

pressureis used tocontrol Machnumber duringroutine operation.

The bestlocation to measure thereference static pressure appears to be a matterof opinion. All of thelarger tunnels (> 2.4 m), whichresponded

tothe questionnaire, use plenum chamber measurements.The surveyindicated smallertunnels use either upstream orifices or plenum chamber measurements.

A totalof the responsesindicated a majorityof approximately 2:l preferred to use plenum chamber measurements.

Advantages ofusinq plenum chamber dataare: (I) it isrelatively

insensitiveto location at which the pressure is measured, and (2) it avoids havingto contend with erosion and/or contamination of orifices. However, experiencewith tunnel wall piezometer rings in a number oftunnels has proven orificedeterioration is not a significant problem. At supersonic Mach numbers, the plenum chamber pressure is generallylower than freestream static pressure, and thedifference increases with increasing tlach number,becoming increasingly more significantat Mach numbers exceeding 1.4. Incontrast, test-section-wall

78 pressuresare generally higher than freestream static pressure at supersonic Mach numbers. Inthe VoughtHigh Speed Wind Tunnel (withwalls slightly converged), tunnel-wallpressure Is closer to freestream static than plenum pressure when

1 < M < 1.6. Thus, thistunnel is calibrated using tunnel-wall pressures because smallerdepartures from freestream static offer the possibility of greater accuracy In Machnumber cal ibrat ion. In general, a more accurate tunnel cal ibra- tion may beexpected when thereference pressure is closer to freestream static pressure.

Inthe case of subsonic Mach numbers, test-section-wall and plenum pres- suresgenerally agree very closely. A possibleexception to this general conclusionis that models, withlarge blockage ratios (i.e., > 2%), may reduce plenum chamber pressurebelow the calibrated, empty-tunnel values at high subsonic Mach number, e.g., Parker(Ref. 57). As iswell known, inclined holes, suchas used in the AEDC transonictunnels, are designed to inhibit excessiveinflow from the plenum tothe test section, but ventilated tunnels with slots or normalholes are more vulnerableto this type of departure from empty-tunnelcalibration. In addition, the plenum chamber pressure may lag freestreampressure during rapid changes in model orientation.

III.D.l. TransonicSurvey Pipes

Responses tothe questionnaire indicate that 31 outof 53 transonictunnels haveused long pipes tosurvey centerline static pressure. The resultsof the centerlinestatic pressure survey are usually averaged over one or more lengths ofthe test section and used tocalibrate a referencepressure. In routine tests, a calibratedlength is selected which most closely matches thelocation and lengthof a particular model. An alternateprocedure is to construct a number of calibrationcurves to relate the reference pressure to several stationsalong the centerline. By locatingthe aerodynamiccenter of a model at

79 I1 I

a station which has been calibrated, the local Mach number at that station A can be used in data reduction. This method is used in transonic tests at f* NASA Ames And is considered to be important for measurements of Mach number at which a model encounters transonic drag rise. In low supersonic tests (M < l.6), the nose of the model is usually located at one of the calibrated stations for more accurate wavedrag measurements. In either case

(i.e., calibrations of the reference static pressure withan averaae along the centerline or with pressures measured at particular locations), buoyancy corrections are usually appliedby using centerline pressure measurements obtained in the empty tunnel.

Guidelines for the installation of a long survey pipe are presentedin

Reference 1. Some rules of thumb are:

1. The nose of the pipe shouldbe a sma.11 angle cone or ogive and

should be located well upstream in t,he subsonic portion of the

tunnel nozzle, e.g., in the 11-ft. Transonic Tunnel at NASA Ames

the nose of the pipe extends into the settling chamberand is

supported under tension.

2. In order to minimize pipe sway, the pipe should be loaded with a

large tensile force, and if appropriate, an upward moment should

be applied at the downstream support.

* In cases where a measured-average is used to calibrate a tunnel, a variation on this procedure would be to account for- local departures from the average. .L J. .. n Private communication, Mr. F. b!. Steinle, NASA Ames.

80 3. Inthe case of very long pipes, three or four cables should be

attachedto further control pipe sagand vibration.

4. All supportcables should be free of obstructions, and allturn-

buckles and cableattachment fixtures should be located behind the

tunnelwalls.

5. Cablesnear or withinthe test section should be sweptback at

anangle of approximately 30 deg to the centerline.

Although a number of tunnel calibrations have been conducted with the nose ofthe pipe located ,in the test section near the beginning of uniform ventila-

tion (e.g.,Refs. 1 and 2), thepreferred arrangement iswith the nose well upstream so that it is always in subsonicflow, Ref. 3. Thisarrangement minimizesdisturbances caused by the nose (e.g., no bow shock) and assures that no transonic shockpasses over the orifices.

A properly-designed,static-pressure survey pipe requires no transonic * calibrationcurve and suppliessimultaneous data throughout the length of thetest section. In transonic tunnels, boundary layer growth on the pipe does notusually induce any longitudinal Mach-number gradients because of ** theventilated walls feature. In contrast, the disadvantages of the long pipe are:

1. sag cancause the pipe to be inclinedto the flow which

in turn cancause erroneous staticpressure data,

2. vibration caninduce errors, see Appendix Ill,

* Thisconclusion is often sustained bydemonstration that a plotof pipe measured staticpressure versus plenum chamber pressure is smooth through ** transonic Mach numbers. However, as a rule of thumb, the blockage ratio of the pipe should be kept lessthan 0.5%, Ref. 6.

81 3. disturbancesgenerated by supports may introduceerrors, and * 4. itsbulk makes it difficult to use for surveys offcenterline

Orifices are a source of error, which is often overlooked when us ing a longfixed pipe. An example ofsignificant orifice-induced error hasbeen discussedby lsaacs (Ref. 5). Thistunnel calibration probe is shown in

Fig. 3.D.l. The orifices havea diameter of 0.076 cm (0.030 in.) and were ground and deburred with particular care to preventchamfer of the opening.

However, staticpressure variations as large as 0.3% ofthe dynamicpres- surewere observed. The factthat the errors wereindeed orificeerrors was ascertainedby moving the probe along the tunnel centerline. A repeatable patternin the variation of measured staticpressure was observed.Figure

3.D.2 shows a comparison ofdata obtained at two differenttunnel locations with M = 0.74.

This example illustrates the need forcaution when using a fixed, static

Pressuresurvey pipe; particularly when thepressure at a giventunnel stationis obtained with only one orifice. It is suggested thattunnel operators, who usesuch pipes,check the orifice problem by translating the pipefor at least one highsubsonic andone supersonic Mach number. If a problem isdetected, this source of error may be reducedby manifolding four or more orificestogether at a givenstation. A second alternativeis to translate the pipe either forward or rearward and takeseveral measurements at a givenstation with different orifices. Either of these procedures wouldimprove the accuracy ofstatic pressure calibrations. Also, it is

* A second pipe, mountedon thefloor, hasbeen used for subsonicCalibration measurements in the 11-FootTransonic Wind Tunnel at NASA Ames,

82 Dimensions In Centimeters

QUADRANT

STINGFAIRING 25 STATICPRESSURE HOLES AT :Oa.'076'ID, 5.1 CM SPACING

STATICPRESSURE PROBE MOUNTED ON CALIBRATION GEAR STING (FAC I L I TY FOR TRANSLATI ON ALONG TUNNELCENTER LINE OVER -229 CM)

Figure 3.D.1. R.A.E.SUBSONIC STATIC-PRESSURE PROBE

Q) W I Hole No. 2 P-Pw Hs-Pn Hole No. 2 -0.005

t I I I I -0.010 L I I -127 -102 -76 -51 -25 0 25 x - cm, distance downstreamfrom tunnel datum (a) PROBEDATUM53 CM UPSTREAM OF TUNNELDATUM

P -Pw Hs-Pw

-0.010 I 1 I i I I I -178 -152 -127 -102 -76 -51 -25 x - cm, distance downstream fromtunnel datum (b) PROBEDATUM 104m UPSTREAM OF TUNNELDATUM

Figure 3.D.2. TYPICALPRESSURE DISTRIBUTIONS ALONGPROBE AT TWO LOATIONS ON TUNNELCENTERLINE, M = 0.74 (choked) , R/R =19-7 x los PERMETER generallyconsidered good practice not to place orifices directly in line with each other, since a disturbance at an upstream orifice canpropagate downstreamand induce errors at adownstream orifice.

As notedby Pope and Goin(Ref. 6), thestatic pipe is seldom

used tocalibrate closed-wall supersonic tunnels. The pipenot only alters the Machnumber because of the reducedarea ratio but also inter- * fereswith the expansion pattern which is required for uniformflow.

A However,a static pipe hasbeen used quitesuccessfully for Mach numbers up to two at AEDC inthe Aerodynamic Wind Tunnel(4T), Ref. 7. Forexample, at M = 1.6 the 2 (I variation in measured centerline Mach numbers was only .007 and at M = 1.99 was 0.008. Thisapplication of a static pipe was made pos- sibleby the unique features of this tunnel, viz., adjustable porosity (0-10%), wallangle, andplenum pumping.

85 lll.D.2. TransonicStatic Pressure Probes

In caseswhere variations of Mach number transverse to the flow hadbeen

calibrated,respondents to the questionnaire indicated that such variations areoften larger than longitudinal variations along the centerline. These

data were obtainedwith conventional probes which, as discussed later, are

subjectto asymmetrical wall interference when moved off centerline and

M exceeds 0.85. Althoughthese transverse variations may be partlythe result

oftransonic wall-probe interference, the calibration of such variations is

obviouslyimportant, particularly for testing wingedmodels. Inthe past,

transonictunnel operators have traditionally concludedthat if (1) thetunnel

wallparameters are set to minimize Mach number variationsalong the centerline

and (2) theaverage ofthe centerline pressures agrees closely with the plenam

chamber pressure(for M < 11, thenthe transverse variations in Yach number

arenegligible. This conclusion is basedon thecomparison between two averages,

and, ingeneral, does notjustify the assumption of negligible transverse

gradients. Thus, windtunnel calibration should include off-centerline measure-

mentsas a standardpart of the calibration procedure. For this reason, one of

theprimary advantages of conventionalstatic pressure probes is mobility as

contrastedto the long,static pressure,survey pipe.

Questionnaireresults also reveal that the most populartransonic static

pressureprobe is a 10 deg apex-anglecone-cylinder with orifices located ten

or more cylinderdiameters (calibres) downstream ofthe shoulder. This criterion

for orifice location appears to have originatedwith the tests conducted by

Holder,et al. (Ref. 8). These investigatorsconducted a systematic,experimentalstudy of the effects of nosegeometries and orifice location on static

pressure measurement at !I = 1.6. A summary ofthese data is presented in Fig. 3.D.3.

The conclusionis that the measured staticpressure is within 0.5% of the refer-

ence pressure when 1 > 10d. As notedin the figure, the reference pressure, 0- 36 External Diameter d =O.ZO~CM

Four 0.0LSCI)I Diameter at 90' Intervals Collar Soldered eo Static Tuk

(a) Long Ogive (b) Short Oglve

10 p . Messured static Prusure - po = Static prusure mlssured by '40 kbc dl [ H = Total Head of Prcr Sirearn L,. Distance of stscic holes Dshind smulder - -1 -1 d = Diameter of static tuoe b*

0 5 I5

Figure 3.D.3 VARIATION OF STATIC-PRESSURE READING WITH POSITION OF STATIC HOLES AND NOSE SHAPEAT H - 1.6, Ref. 8 which wasassumed to be true freestream stat ic, was obtained with the long ogivalnose and orifices located 40 calibres downstream.

Highsubsonic (H = 0.6 to 0.9) dataind icate probe pressure generally returns to freestreamlevels (within 0.5% of q) atlo/d values of 4 to 6 calibres, Ref. 9. The exactlocation is dependenton nose geometry. For example, transonicdata presented by Ritchie (Ref. IO) for a staticpres- sureprobe, with a nosecorresponding tothe long ogive shown in Fig. 3.0.3,

indicate negligible measurement error when orifices are located only two calibres downstream ofthe nose-cylinder juncture.

However, sincethe overexpansion at the shoulder extends farther downstream inthe supersonic case,Gracey (kef. 9) concludedorifices located 10 or more diametersdownstream would sense freestream pressure with

"small-error" at bothsubsonic and low supersonic speeds.The consensus of transonictunnel operators seems toagree with Gracey.

As noted by Davis and Graham (Ref. 111, inthe past the data obtained byEstabrooks (Ref. 12) for a cone-cy1 inder hasbeen used,almost universally, as a standardfor transonic interference-free data. Although the purpose of these measurements was toinvestigate wall effects,the results are also pertinentto probe design and performance.Estabrooks obtained data on a 20° apex anglecone-cylinder in the AEDC-PWT 16T tunnelwith a model blockage ratio of 0.008% and M = 0.7 to 1.4. A cursoryexamination of these data indicates orificeslocated seven calibres downstream ofthe cone-cylinder juncture, will allowaccurate measurements offreestream static pressure throughout the transonic speed regime. The datawere unaffected by varying freestream Reynolds 6 6 number permeter from 4.5 to 12.7 mill ion (1.36 x IO < Re/ft < 3.87 X 10 ).

The factthat freestream static pressure cannot bemeasured at anyone location, as M + 1.0, will beestablished in the following discussion.

88 When the same cone-cylinder model was tested in the AEDC-PWT 1T tunnel,blockage was 28, and in this caseconsiderable wall interference was observed. Inthe Machnumber range of 0.95 to 1 .OS, Estabrboks concludedthe dominant wall interference effect was reflections of super- sonicexpansion waves (originatingat the shoulder) back to the model as compression waves. In orderto explain this phenomena, a discussion was givenconcerning too low a resistanceto inflow from the plenum to the testsecti,on. Unfortunately, Estrabrooks appeared to be unaware of the transonicshock which forms near the shoulder of this type of body and moves.rearward withincreasing Mach number. Althoughcompression waves

(generatedby inflow) may have been present,interpretation of this data mustaccount for passage of a transonicshock when 0.90 M 1.05. Thus, the effects of this shockon the measured pressure distributions weremis- interpreted as solelywall interference. For example, theexistence of a shockimnediately aftof the shoulder is clearly indicated (Fig. 3.D.4) for the 20 deg cone-cylinderat M - 0.95. Data forthis configuration at M = 0.975 are less definitive possibly because of a bifurcated shock or boundarylayer separation, either of which reduces the pressure gradient f: produced by the shock. As the Mach number is increased to one, the shock moves rearward and offthe instrumented portion of the cylinder. This type

* Althoughthe Reynolds number based onwetted length is approximately 2.55 x 6 10 nearx/d = 4, it isnot clear that the boundary layeris turbulent because ofthe shoulder expansion which thins and stabilizes the boundary layer. However, even if theboundary layer isturbulent, the transonic F shockcan cause separation if thelocal Mach number exceeds 1.3 (e.g., Refs. 13 and 14). The measured pressureratio at the shoulder doesindeed indicate a local Machnumber near 1.3. It is alsorelevant to here note thatHsieh (Ref. 15) foundthe. laminar boundary on a. hemisphere-cylinder separatednear M = 0.80.

89 H, 0.950 a - 0.0

0.975 - 0.0

1 .ooo 0 - 0.0

0 2 4 6 8 10 x/d, Distance from Nose inCalibers

Figure 3.0.4 TRANSONIC PRESSURE DISTRIBUTIONS ON A 20 DEG CONE-CYLINDER WITH 0.008% BLOCKAGE, Ref. 12

90 of phenomena may be clearly seen inthe sctilleren photographs presented

by Page (Ref. 16) for a 14 deg apex cone-cy1inder with a blockage rat lo

of 0.005%. A rathergeneral conclusion is that the transonic shock will

move off a cylindrical probeas M +l,and merge with the sting and/orsup-

port shocks,provided there is no wallinterference.

Movement of a transonicshock on axisymmetrlc bodies can now be cal-

culatedvia the computerprogram of South and Jameson (Ref.17). This

programprovides a solutionto the complete potential equatlons for steady

transonlcflow. Thus, inorder for thesesolutions to be applicable to

realbodies there must be noboundary layer separation (e.g., Ref. 15).

and in the case of windtunnel probes, the body must be free of wall

interference.

The existence of boundarylayer separation on cone-cylinders in

transonic flow has been investigatedby Robertson and Chevalier(Ref. 18).

Cone-cylinder mode 1s with a blockage ratio of 0.5% and 1.2% weretested

with M = 0.5 to 1. 17 inthe AEDC-PWT 1T tunnel. These investigators

foundthat the boundary layer separated at the cone-cy1 inder juncture when

the cone apex angle was 40 deg or more and tl < 0.85. Ingeneral, as cone

apex angleincreased,the Mach number for boundarylayer attachment increased

Althoughsurface pressures were only measured a distance of four calibres

downstream ofthe shoulder, the freestream static pressure was attained in mostcases withinless than four calibres. This was found to be true for

both sizes of models with apex anglesranging from 20 to 60 deg.

The primaryexception to thisobservation occurs when a transonic

shock locatesnear an orifice. If the shock isforward of the orifice,

91 the measured staticpressure will tendto be higherthan freestream. The exact amountdepends on strength of the shock and distance from orifice.

Correspondingly,the pressure will be low if the shock islocated near, but downstream of,the orifice. Since the transonic shock moves rearward with increasing Mach number, all stations along a probe's stem areaffected.

For example, schlierenphotographs obtained by Robertsonand Chevalier show thetra'nsonic shock ona 20 deg cone-cy1inder model initially forms near theshoulder at M = 0.8, and increasesin strength and moves rearward as Machnumber increases. The rateof movement of this shock isaffected bymodel blockage and theextent of the supersonic zone ona given model.

The effect of windtunnel blockage on movement ofthe transonic shock may be seen inthe M = 1 dataof Estabrooks (Ref. 12). By varyingthe sizeof cone-cylinder models to obtainwind tunnel blockage ratios from 0.5% to 4%, thetransonic shock moved forwardfrom x/d = 5 toless than 4 at

M =i 1. Thiseffect of blockage may also beseen inthe schlieren pictures of Page (Ref. 16) which compare thetransonic shock locations on the same model in two differentslotted-wall tunnels with blockage ratios of 0.25% and 0.005%. Furtherevidence of this phenomena may also be found inthe dataof Capone and Coates(Ref. 19) and Couch and Brooks(Ref. 20). For example, thesurface pressure data of Capone and Coatesprovide an excellent illustrationof transonic shock movement on a 20 deg cone-cylinder.In this case,the model was testedin the Langley 16-Foot Transonic Tunnel andhad a blockageratio of 0.198%. A sample ofthis data is reproduced in Fig. 3.0.5.

Here we see thetransonic shock move fromx/d = 3.25 at M = 0.90 back to x/d 3 10.5 at M = 1.025. Measurements atthe next higher Mach number, M - 1.04, * indicatedthe shock had moved pastthe instrumented portion of the cylinder. * Reflection of the bow shockback onto the cylinder was notobserved until M 2 1.10.

92 M = 0.90 0

M = 0.95 0

M = 1.00 0

M = 1.0250

x/d

Figure 3.0.5 TRANSONICPRESSURE DISTRIBUTIONS ON A 20 DEG CONE-CYLINDER,REF. 20.

93 ...... " ......

Similar measurements for a 40 degcone-cy1 inder, with the same blockage,

showed theshock was at x/d = 10.5 when M =I 1 .Ob. Thus, thelarger expansion

atthe cone-cylinder juncture and thecorrespondingly larger pocket of super-

sonicflow resulted in a retarded movement ofthe transonic shock.

Based onthese various results, it appearsthat a 20 degapex cone-

cylinder must have a cross-sectionalarea less than 0.01% ofthe tunnel area

inorder to avoid retarding the rearward movement of a transonicshock with

increasing Mach number. Thisconclusion is based on measurements made onlyat tunnelcenterlines. When theprobe is moved offcenterline, closer to a wall,

f even smallersizes would be necessary toavoid wall interference. Also, some

asymmetry ofthe shock may be expected. Thus,a non-perturbingflow measure-

menttechnique, such asa laserDoppler velocimeter, appears to be very

desirablefor tunnel calibrations in the range 0.95 < M < 1.05.

Recently,Neman and Klunker(Ref. 21) and South and Keller (Ref. 22)

haveperformed calculations for transonic flows about airfoils and axisym-

metricbodies which include wind tunnel walls in the boundary conditions. These

calculations show theshock moves forward when theopen-jet boundary condition

isapplied, i.e., Pw - Poo . Similarly,the shock moves rearward, compared

tothe free-air solution, when thesolid wall boundary conditionis applied.

Inlight of the foregoing discussion, this implies that either slotted or

perforatedwalls act more likeopen-jets as thesize of transonic models is

increased.This phenomenon isapparently a resultof larger pockets of super-

sonicflow which impress lower pressures at the walls. This in turn draws in

more air fromthe plenum chamber and apparentlyshifts the model flowpattern

towardthe open-jet boundary condition. * This may partially explain the transverse Mach number gradientsreported by some ofthe questionnaire respondents.

94 As regardsstatic pressure probe design, small angle conescan be used tominimize strength of the transonic shock. Inaddition, the smaller ex- pansionangle will generateless wall interference for a givenprobe diameter and the boundary layer will remainattached at the cone-cylinder juncture.

A separatedboundary layer isundesirable because it introducesdisturbances whichare convected downstream a-nd cancause additional errors in static pressure measurements (see Eq. 3.D.1, p. 98). Thus, inthe past the IO deg apex anglecone-cylinder probe has served as a convenientcompromise between optimum transonic performance and ease of construct ion.;:

The problem of orifice-transonic shock interference,which is characteristic of cone-cylinder probes, may be avoidedby locating orifices on very small-angle cones. This was demonstrated by Sutton(Ref. 24) who compared thetransonic surface pressures on a 3 deg included-angle cone with a conventional IO deg cone-cylinderhaving 0.021% tunnelblockage. These probes a re shown inFig. 3.D.6. The correspond ingsurface pressures are shown in F ig. 3.D.7. for 0.9 c M < 1.02. The oscillation in calculated Mach number,based on the " cylindersurface pressure, is caused bypassage ofthe transonic shock. However , it isrelevant to note that transonic shock effects can be confinedto a srna 1 1

Machnumber range(e.g., AM = 0.02) when using a carefullydesigned 10 deg cone- cylinderprobe at * tunnelcenterline. In contrast, the 3 deg cone provides a monotonicallyincreasing pressure and decreasing Mach number throughoutthe transonicrange.

Unfortunately,the 3 deg cone isreported by Gracey(Ref. 9) to be sensitive to flow misalignment. For angles of attack between -+I deg, pressure * The AGARD needleprobe described in Reference 23 hasa cone apex angle of approximately 12 degand four 0.3 mm orifices located 11.3 calibres downstream of the shoulder. 95 4 HOLES, 0.061 DIAMETER

d = 0.305

(a) CONE-CYLINDERSTATIC HEAD

3 HOLES 0.036 1.3' 3" DIA.

TIP DIAMETER -

0g015 DIMENSIONS ARE tN (XN'TmERS

(b) 3" CONE STAT1C HEAD

Figure 3.0.6. DIMENSIONS OF THER.A.E. STATIC PRESSUREPROBES +o .02 - ' 0'.90 0.b2 0.44 0.96 0.'98 1.601.b2 = MACH NUMBER '' DEDUCED FROM AM - "- MR PLENUM CHAMBER 0- - STAT I C PRESSURE = M -M HC 6 155 cm FROM Re = 15.7 x 10 per meter THROAT. -0.02 I I I I I I I (a) 10' CONE-CYLINDER STATIC HEAD = MACHNUMBER MH DEDUCED FROM STATIC HEAD +o .02 1 1 I I I I I PRESSURE 0.90 0.92 0.94 0.96 0.98 1.00 1.02 I An = fl -M HC 6 Re = 10.6 x 10 per meter I I I I I I I -0.02 (b) 3" CONE STATIC HEAD

Figure 3.D.7. TRANSONICCHARACTERISTICS OFTHE TWO R.A.E. PROBES measurementson a 3 deg cone Inthe Langley 8-ft. TransonicTunnel indicate variationsof approximately 0.02 qinear M = 1. These results were the same for a 0.033 cm (0.013 in.)orifice located either 12.7 or 17.8 cm (5 or 7 in.) fromthe tip. Thus,a small-angle conecan be used tocalibrate wind tunnels near M = 1, but its sensitivity to flow angularity can make it difficult to resolve Mach number variationsas small as 0.001.n

Effectsof Various Parameters on Static Pressure Probes

Size:

As foundby Couch and Brooks(Ref. 20), cone-cy1inder bodies with a tunnel blockage ratio of only 0.03% canhave cylindersurface pressures which depart significantlyfrom freestream static near M = 1. Therefore,for accurate tunnelcalibrations along the centerline and near M = 1, staticpressure probes with blockageratios larger than 0.01% arenot recommended. The wall-interference- freeperformance of suchprobes can be caluclatedusing the South and Jameson computerprogram (Ref. 17). If for some reason a largerprobe is required, centerlineblockage effects can be estimated by theoreticalanalysis, e.g.,

Refs. 21,22 and 26. Wallreflected disturbances can also be detected by having orifices at more thanone station andchecking monotonicity of the data with increasingdistance from the nose. Pope and Goin(Ref. 6) suggestmoving the model off centerline and/orusing a schlieren system.

Once theprobe diameter is selected, at least four orifices witha diametevof 0.051 cm (.02 in.)should be located IO or more calibres downstream of thenose-cylinder juncture. Finally, errors induced by orifice s ize are discussed inSection lll.D.4.

A Additional transonic measurements of flow angularity with a 3 deg conical probeare reported by Wright,et al. (Ref. 25).

98 Reynolds .Number:

'At Reynolds numbers characteristicof transonic tunnels (i.e., Re/m > 5

million),static pressure probes are usually unaffected by this variable,

providedthe probe boundary layer is attached.

Tu rbu1 ence:

In contrast to the Pitot probe, staticpressure probedata can be

affectedby turbulence. The desiredquantity 1s the static pressure asso-

ciatedwith the mean flow.Bryer and Pankhurst(Ref. 27) notethat when the

turbulencescale is large, compared tothe probe dimensions, the measured

staticpressure will tendto be low and isproportional to the dynamic

pressureof the turbulence normal to an orifice. If theturbulence scale is

small,the measured staticpressure will be high. The followingequation

has been derived by Siddon(Ref. 28) torelate the error in measured,mean

staticpressure to the dynamic pressuregenerated by a turbulentflow. - 2 Pm - Pt = A p(U + (ut)')+ (3.D. 1)

If theorifices are located so there are no noseand/or supportstem-induced

errors,then A = 0. B is ameasure ofthe crossflow-induced error in measured

staticpressure. When a probe isinclined at an angle a in a steadyflow,

Eq. (3.0.1) maybe written as

2 2 2 Pm - Pt = 2q(Acos a + B sin a)

This type of measurement was performed in subsonic flow by Siddon with

a standard,classical probe. The probe had: (1) anellipsoidal nose, (2) a diameter of 0.305 cm (0.12 in.), and (3) sixorifices located 8 1/2diam-

eters downstream from the nose. For this probe, B was found to be -0.55.

99 A specialprobe, designed to measure unsteadycrossflow and fluctuating staticpressure, was alsotested by Siddon. This probe,which hada circumferential slit forsensing static pressure, was found .to havea B valueof

-0.23. Accordingto Siddon, the difference in B values isprimarily causedby thedifference in orifice arrangement.Additional measurements in a rotating- inclfnednozzle, a turbulentchannel flow, anda roundturbulent jet indicated

B variedover the range -0.46 to -0.35. Thus, themagnitude of crossflow- inducederrors varies with probe design and turbulencescale and intensity.

Bryer and Pankhurst(Ref. 27, p. 43) suggestthat probes be calibratedin flowswith turbulence closely matching those in which the probe will beused.

This is obviously anarea which needs furtherresearch.

-Yaw: When severalorifices (4 or more) arelocated around the probe clrcum- ference, flow angularity causes the measured staticpressure to below. This can be readily seen from thepressure distribution about a circular cylinder normal to a flow (e.g.,see Appendix I I I). Bryer and Pankhurst(Ref. 27) notethat the yaw-induced static pressure error of this type of probe is typically 0.01 PaD whenyawed 3 deg. The error,in a particular case, is dependenton nose geometry and orificelocation. Ritchie (kef. 10) reported yaw-induced errorsgenerally increase with Mach number.Thus, for a given allowederror in measured staticpressure, the permissible variation in flow misalignmentdecreases. Gracey (Ref. 91, Ritchie (Ref. 101, and Rittenhouse

(Ref. 29) have reported on staticpressure probes designed to minimize yaw sensitivity. These probes uti1 izeonly two orifices located 30 to 40 degrees from thewindward-meridian. Although these probes havesmall errors at transonic speedsand yaw angles up to 28 deg, theyrequire knowledge of the streamdirection. Since this is not usually known duringwind tunnel calibrations,this type of probe cannot be recommended. Hence, theconventional

100 probe with four or more orificesis preferred. Particularly since the flow angularity in the central core ofmany contemporary transonic and supersonic tunnels is less than one degree, the conventional probe will usually have negligible error due to yaw.* For example, Ritchie (Ref. IO) found a two degree angle of attack caused less than 0.2% error in measured static pressure through-out the transonic Mach number range. (These results were obtained

= 1/2 with a probe having an ogival nose (f r 12) and orifices located 12 calibres from the nose.) However, this is another reason for minimizing flow angularity in the empty tunnel, i.e., not only will model testing results be more representative of free-flight phenomena, but the tunnel calibration will be more accurate.

Lakes and Support Interference: The effects of a support flare on base pressures have been investigated by Chevalier (Ref. 30) over a Mach number range of 0.70 to 1.60. Based on the experimental results at M = 1, a flare located approximately I5 flare diameters downstream of the orifices will not interfere with static pressure probe readings. An II deg flare (semiangle) located at this distance from the orifices will have negligible effect on the flow at the orifices throughout the transonic speed regime. The required distance for no interference decreases with smaller flare anglesand increasing Mach number.

Interference caused by a cylindrical strut normal to the probe axishas been investigated by Krause and Gettelman (Ref. 31) for M = 0.3 to 0.9.

These authors found that a distance of 14 strut d iameters between static probe orifices and the strut was required for negliqible interference at these speeds. * It is conceivable that a static pressure probe couldbe calibrated for yaw errors. Static pressure readings could then be corrected for measured flow angles. However, most tunnels do not have sufficient flow angularity to warrant this procedure, and few operatorswould consider it practical. 101 Perhaps the most stringentrequirements for distance between oriflces anda strut have been reportedby Nichols (Ref. 32). Forthe case of a static pressure probe mountedon a double wedge strut support, the transonic measurements of Nichols indicate orifices should belocated 32 strutdlameters ahead ofthe strut. These dataare shown inFig. 3.0.8.

General criteriafor probe survey rakes have been suggestedby Gray

(Ref. 33) and arepresented in Fig. 3.D.9. The separationcriterion for adjacentstatic and pitot probes insubsonic flow is based onthe data of

Krause and Gettelman(Ref. 31). Insupersonic flows, Gray reconmends spacing adjacentprobes so thatthe neighboring bow shock intersects a staticpres- sureprobe 15 probediameters downstream ofthe orifices. The objective is toprevent disturbances caused by shock-wave/laminar-boundary-layer inter- actionfrom affecting the pressure measured atthe orifices. Since the criterion becomes impractical when M 2, Gray reconmends theflow deflection acrossthe Pitot shock be keptless than 3 deg at its intersection with the static probe, and furthermore,the probes should be spaced so thelnter- section is 5 or more static probediameters downstream of the orifices. il For Mach numbers between 0.9 and 1.2, rakesmust be used with caution because ofincreased blockage and near-normalshock waves. Also,rakes are notoriousfor inducing crossflow in the plane of the rake at high subsonic

Mach numbers, see Section II1.E. Inthis Mach number rangethe following alternativeis recommended: employ a single,static pressure probe (or combinationPitot-static probe* for validatingthe isentropic expansion assumption)with a slenderogival (L = ad) or verysmall angle conical n (x 5 deg)nose anda stingtype support which satisfies the criteria suggestedby Gray.

.L Here it isrelevant to note Bryer and Pankhurst (Ref. 27, p. 41) are of the opinionthat combination probes are in general less accurate than single- purposeinstruments.

102 0.4

0.3

p -P -w- H q 0 -m

0.2 a - 0.90

A - 1.00

0.1

-0.1 8 12 16 20 24 28 32 36

I /S, DISTANCE FROM ORIFICE To STRUT SHOULDER d 0 W Figure 3.D.8 EFFECT OF ORIFICELOCATION UTZIZIUG A DOUBLE- WEDGE SUPFORT STRUT, REF. 32 .. ,. ..- ._...... I I _,, , , . ..I . .I,. -

lo 11°16 d for M > le61.6

Cylindrical support

For negligible support interference: l/D -> 14, all valuesof M

For negligible adjacent probe interference: M < 0.9: dd1> 6 / M -> 1.2: Determined by the intersection of bow wave with static probe

Figure 3. D. 9 GENERAL CRITERIA FORPROBE SURVEY RAKES, Ref. 33

104 lll.D.3. SupersonicStatic Pressure Probes

AlthoughBarry (Ref.34) has shown thatsupersonic Mach number calcula-

tions basedon Pitot and freestream static pressure are not as accurateas

the Mach probesdescribed inSection III.C, this approach is often used

because of its familiar use in subsonic flows and its ease of construction.

In addition, it does provide amethod for calculating H wh ich does not depend

on theassumption of an isentropicexpansion from stilling chamber to test

section.

Walter and Redman (Ref. 35) measured pressuredistributions on a 7 deg

includedangle cone-cy1 inder at Mach numbers 1.55 and 2.87." These data

indicatethe surface pressure on the cylinder returns to freestream static

beyond IO calibresfrom the shoulder. In general, as Mach numher increases

theoverexpansion increases and longerdistances from the shoulder are re-

quired.Increasing cone angle hasa similareffect.

Pressure distribution data on cone-cylinder-flare configurations at

speeds up to M = 4.5 have been reported by Washington and Humphrey (Ref.36).

Data obtained on a blunt nosed, 10.3 deg cone-cylinderat M = 4.5 and zero

yaw indicatethe surface pressure returns to within two percent of freestream

static at nine calibres downstream ofthe shoulder.

In the case of yaw, the data of Reference 36 show the circumferentially

averagedpressure is belowfreestream static. For a given yaw angle,the

averagepressure decreases further as Machnumber increases.In general,

increasing the number of orifices aboutthe circumference will decrease yaw

sensitivity. However, Gray (P.ef. 37) and others have notedthat surface

* Thisdata is also presented in Reference 6. 105 pressure on long cy1 inders, measured approximately 240 deg circumferential ly

from the windwardlocation, provides a closeapproximatfon to freestream

static for M lessthan 4.0. Thus, thistype of probe canbe used to eliminate

yaw-induced errorsin static pressure.

As discussedpreviously, since most tunnels have small flow angularity

inthe empty testsection, it isunlikely that yaw induced errors will be

significant. But this must be determinedby the user., If staticpressure

is being used to calibrate Mach number andan accuracyof 0.1% is desired,

thensmall yaw-induced errors may be important.In which case, a static

probe with two orifices located circumferentially 70 - 80 deg apart canprovide

amore accurate measurement. Thiscan be accomplished by rotating the probe

tolocate the windwardgenerator (highest pressure) and then rotating the

probe until the two orifices agree."

Inorder to avoid support-interference, Gray (Ref. 37) recomendsthe

cylinderdiameter be constant for at least 8 diametersdownstream of the

orifices. Any subsequentenlargement indiameter should be restrictedto

no more than a 10 deg flare (semiangle).Additional criteria for rake

arrangementsare given inFig. 3.D.9.

Inthe past, Pitot pressures and surfacepressures on conicalprobes

have frequently been used insupersonic flows to calculate Mach number,

e.g., Refs. 38 and 39. Also, an extremelyaccurate conical static pressure

probe is briefly discussed by Pope and Goin(Ref. 6). Thisprobe design

has a short (1.78 cm) 8 deg conicaltip followed by a long(16.26 cm) 1 deg

included-anglecone. Orifices encircle the 1 deg cone at threelocations.

" This assumes no orifice-induced errors.

106 The error in measured static pressure is reported to be of the order of 0.1%

oftrue freestream pressure when M = 1.8 to 3.5. Thus, thesedata represent

some of the most accurate static-pressure measurements in supersonicflow.

However,Gray (Ref. 37) has recentlyreviewed the merits and limitations

ofsupersonic static-pressure probes.Cone-cy1 inder,sharp cone,and planar

probeswere considered. Based on theeffects of Mach number, angleof attack,

and Reynolds number, Grayconcluded that the cone-cylinder probe is, in general,

superior for use at Mach numbers below 4.

In this Mach number range, the Reynolds numbers of most supersonictunnels

arelarge enough forviscous corrections to be negligible.This effect canbe estimatedby calculating the equivalent inviscid pressure from anexpression suggestedby Gray (Ref. 37) forcone-cyl i ndersprobes in flows with Mach number

lessthan 5.

(’w’meas = 1 + 0.25 , (3.0.3) (‘w) inviscid

where -3x E M / (Rel /C) 1 /2 0

C (p /p ) (T /T ), Chapman-Rubesin viscosityparameter. Z weew

For a givenconfiguration, the viscous interaction coefficient decreases with

increasing Mach number. For example, thehypersonic experiments of Peterson * andGeorge (Ref. 40) indicate a coefficientof 0.08 isappropriate for a

20 deg cone-cylinderprobe at M = 7.2 and 14.0.

* These investigators, among others, ,havenoted that static-pressure probes shouldnot beused in flows with large axial or transversepressure gradients. Inorder to simultaneouslyminimize the effects of viscous interaction

and noseoverexpansion at supersonic speeds, it is recommended that orifices

be locatedat least 16 calibres downstream of theshoulder. In cases where

correct ion is judged to be necessary,the interested reader may refer ,to

Gray'sdiscussi on(Ref. 37) of a procedure for correctingthe measured

pressure for vi scous interaction and obtaining a betterestimate of the

inviscid,stati c pressure.

Sincecone-cylinder probes are relatively long, they not only havesmall

rigiditybut also cannotbe used inpressure gradients. In addition, they

aresensitive to yaw. Forthis reason, shorter supersonic static pressure

probeshave been investigated, e.g., Refs. 41, 42, and 43. Thiswork has

focused on theidea of designing a probe to have at least one station where

thecircumferentially averaged surface pressure remains a constantfraction of freestream static regardless of Machnumber or angleof incidence.

The probesdesigned by Donaldson and Richardson(Ref. 41) and Pinckney

(Ref. 42) areconventiona,l bodies of revolution, whereas theprobes of Smith andBauer (Ref. 43) have noncircularcross-sections. Measurements at M = 0.2

indicatethe noncircular probes of Smith and 3auer arecompletely insensitive

to flow angles of -+6 deg.Beyond this angle of attack ranqe,boundary layer separation becomesa factor,anderrors increase rapidly. Since the yaw sensitivityof conventional, circular probes increase with !lath number, theseprobes may or may notoffer anadvantage forsupersonic applications. Pinckney reportsfreestream static pressure canbe determined with his calibrated probe to within 2 percentfor incidence angles of 27 deg and M = 2.5 and 4.0. Better results have been reported by Dona ldson and Richardson.Using a 50 deg includedangle cone-cylinder probe w '4th 24 or ificeslocated 0.88 diameters

108 downstream of theshoulder, these investigators found the probe measured

0.793 POD,and thefreestream static pressure could be determined to within

1.2 percentover the Machnumber range 1.1 to 2.5. These results were obtained for zero yaw.By adjustingtheir calibration factor to 0.763, free-

streamstatic could be calculated to within 3 percent for incidenceangles

up to 18 deg in anyplane. This represents the smallest yaw sensitivity of

anysupersonic static pressure probe known tothe authors.*

As is well known, theoreticalsolutions for surface pressure distribu-

tionson probes can assist the placement of orifices to measure freestream

staticpressure. Based oncomparisons of measuredand predictedsurface

pressureson a hemisphere-cylinderprobe, Hsieh (Ref. 44) concludedthe

South and Jameson program(Ref. 17) canbe used successfully up to M = 1.3.

Beyond this Mach number, a method of characteristics algorithm is recommended

foraxisymmetric probes. A ratherlarge number of suchprograms arecurrently

available. For examples, theinterested reader may referto the paper

byHsieh (Ref. 44). Fornon-axisymmetric probes, anumber offinite dif-

ferencesolutions for three-dimensional, supersonic, inviscid flows are

available, e.g.,Marconi, etal. (Ref. 45).

Finally,the use ofmultiple probes in rake arrangements forsurveying

supersonictunnels is well known.Rakes can be successfully employed to

calibratesupersonic tunnels by applyingthe design criteria of Gray, Fig.

3.0.9, and avoidingreflections of how shock waves offthe tunnel walls.

* Donaldsonand Richardson also found a conventional,single bore, internal plenum provided less yaw sensitivitythan an annular plenum.

109 lll.D.4. Orifice-InducedStatic Pressure Errors

Errorsin static pressure measurements,caused by variations in orifice geometry,have been investigatedin a number ofstudies, Refs. 46-52. The relevantgeometric variables are: (1) holediameter, (2) ratioof hole depth todiameter, (3) therelative size of the cavity or tube connecting to the hole, (4) inclinationof hole axis relative to the surface normal, (5) the conditionof the hole entry, i.e.,whether the edges are square, rounded, chamfered, or have burrs.

Ideally,the measuring hole should be infinitesimallysmall so as to notdisturb the adjacent flow. Shaw (Ref. 47) notedthat the basic error caused by finite-sizedorifices consisted of three contributions. Firstly, dipping of the streamlines into the orifice causes a divergenceof stream- lineswhich results in a higherpressure in subsonic flow. Secondly, an eddy (orsystem ofeddies) is generated within the hole. (An approximate analysis by Nestler(Ref. 51)has shown how theturning of such aneddy cangenerate increased pressures.) And finally, a Pitoteffect occurs at the downstream edge ofthe hole. These three phenomena causethe measured pressureto be toohigh. Although the severity of these phenomena decrease with holesize, Rainbird (Ref. 49) observedthat holes with diameters less than

0.038 cm (.015in.) are difficult to produce with sharp edgesand negligible burrs.Also in short duration tunnels, the time required for pressure equilizationin typical measurement systems becomes excessive.

Based ona studyof orifice errors in turbulent pipe flow, Shaw (Ref. 47) and Franklin and Wallace(Ref. 50) have verifiedthat the effect of hole size scaleswith the local wall shear stress (T~)and fluiddensity (p) and viscosity (p), viz.,

110 - 'meas 'true = do , ?w ) - (3.0.4) T P W

In add ition, theactual magnitude of errors are a funct ion of anumber of

otherparameters. For example, Shaw (Ref. 47) and Livesey, etal. (Ref. 48)

foundthe relative depth of the, hole to also be a significantparameter. In

general,the measured pressuredecreases towards the true value as the ratio

of holelength to diameterdecreases. However, as holelength/diameter

decreasesbelow 2, Livesey,et al. noted a relativelylarge Cavity (14 do) * behindthe hole caused a negativeerror in measured staticpressure. In con-

trast,Rainbird (Ref. 49) and Shaw both used a cavitybehind the hole with a

diameterof only 2d0. Withthis arrangement, Shawls dataindicate a decreasing

statlcpressure error as thelength of the hole decreases from 1.5 do to ** 0.5 do.Shaw alsosystematically studied the effects of burrs and dis-

coveredthat burrs of the order of do/127can cause errors as large as occur

withvariatlons in smooth holesize, i.e., the solid curve in Fig. 3.0.10.

Rayle(Ref. 46) found,as expected, incliningthe axls of holes toward

the oncoming flowincreases the measured pressure. By incliningthe hole

downstream, a reducedpressure is measured. Raylealso studied the effects

of varying edges of an orifice.In general, rounding of the edges resulted

inhigher pressure; whereas,chamfering produced small negative errors. As

observedby Benedict (Ref. 53), theflow over a roundededge does not

imedlatelyseparate but instead is guidedinto the hole with a resulting

recovery of part of the dynamicpressure. Inthe case of a chamfered or

countersunkhole, the flow will separateat the upstream sharp edge, but * Similarly,Livesey, et al. also note that a contractionin tubing diameter (e do) will cause a higherpressure. ** Rainbird used a fixed ratio of hole length to orificediameter of 3.

111 it also acceleratesalong the sloping downstreamedge which results in a reduced pressure.Rayle concluded a0.076 cm (0.030 in.) hole with a 0.038 cm

(0.015 in.) deep countersinkshould provide a staticpressure near the true va 1 ue. Finally, Rayle's experiments.with water and air flows covered a

Machnumber range of 0 to 0.8. Thisdata demonstrates orif ce-induced errorsincrease with Machnumber.

A summary of subsonic data obtained hy Frank1 in and Wa lace(Ref. 50) and thesupersonic data obtained on a 25' apex-anglecone by Rainbird (Ref. 49)

is presentedin Fig. 3.0.10. Rainbird suggestedthe scatter in his data could be attributed to variations in the ratio of holediameter to boundary layer displacementthickness. However, thesubsonic data of Franklin and Wallace failedto indicate any effectof this ratio. Furthermore, Nestler (Ref. 51) demonstratedorifice-induced errors could not be correlated by the ratio of hole diameterto boundary layer momentum thickness. The problem is compounded further by thehypersonic w ind tunneldata discussed byCassanto (Ref. 52).

In this case,the diameters of square-edgedand chamfered (60°) or ifices on an 18 degcone were varied, respectively,from 0.076 - 0.635 crn (0.03 - 0.25 in.) and 0.152 - 0.635 cm (0.06 0.25 in.).For a freestream Machnumber of 8

= 6.37),the measured pressure was insensitive to orifice diameter ('local (decreasingslightly - lessthan 3 percent - withincreasing diameter). Thus, the effects of Machnumber on orifice-induced errors in static pressure needs additionalresearch.

Questionnaireresults indicate static orifice diameters typically range from 0.025 cm (0.01 in.) onsmall-angle cones to 0.228 cm (0.09 in.)on wind tunnelwalls. In order to minimize static hole errors*, it is recommended t Of course, a flush-mountedpressure transducer ispreferable whenever possible.

112 I

0 I 0. 0 0 0 A d d d 0 0

7 W Franklin : M<@5 8s Wallace

W/V M of Rainbird '8 Cone Data

I that a square-edge orifice with a diameter of 0.051 cm (0.020 in.)be adopted asan industry standard. S ince Rainbird (Ref. 49) hasdemonstrated thissize of orifice canbe used with satisfactoryresults in ablow-down tunnel, it can be used.- in most facil it es. Ideally,the length of the hole should be restricted to the order of /2 of an orificediameter, Ref. 47. However, such thinwall orifices are fragile and, thus,subject to damage. Hence, Gray

(Ref. 33) recomnends thehole length be greaterthan two orifice diameters.

The diameterof a connectingline, behind the hole, should be restricted to the order of two orificediameters, Refs. 47, 49, 50 and 33.

Of course,these last two criteria presuppose an installation which is accessible from theback side, e.g., a tunnelwail. In the case of a long, staticpressure survey pipe, tubing of the appropriate size is swaged or sweat solderedin a receivinghole and thenground or machined down flush withthe outside surface of the pipe. Hereagain an orificediameter of

0.051 cm (0.020 in.)can be used, and largerdiameter connecting lines may be used to reduceresponse time.

Inthe case of a conventionalstatic-pressure probe, a ratio of hole depthto orifice diameter of less than one isnot only possible but is frequently the case. For example, a probewall thickness of 0.033 cm (0.013 in.)is typical for 0.318 cm (1/8 in.) OD stainlesssteel tubing. Therefore, the recommended orifice size wouldprovide a holelength to diameter ratio of 0.65. Also, staticpressure errors of probes may be reducedby designing them to have laminarflow at the orifices. Although the existing correlations of orifice errorsare for turbulentflows, it appearsprobable that a laminarflow will * Of course, a flush-mountedpressure transducer is preferable whenever possible. dip into an orifice less than a turbulent flow. A laminar flow probe can

be obtained by properly sizing the probeand polishing the external sur-

face to 0.25 microns (1011 in.). For example, a 0.318 cm (1/8 in.) diameter

probe, with orifices located 10 calibres downstream, would havea local

Reynolds number of 1.25 million for a freestream unit Reynolds number of

39.4 million per meter. In general, if noise data is available for a given facility, the correlation ofBenek.and High (Ref. 54) can be used to estimate Reynolds numbers at which boundary layer transition occursin order to judge whether a laminar flow probe is feas ible.

Since the data of Shaw (Ref. 47) indicate stat ic pressure measurements

are very sensitive to burrs, considerable care mustbe taken to assure a

smooth, sharp-edged orifice. This may be done by beginning the hole with

drill bits several sizes smaller than the desired final hole size and

progressively increasing the hole size. Also, short flute drill bits should

be used to minimize flexing and a drill guide (of the same metal) clamped over the orifice location can be of considerable help. Finally, slower rates of dri 11 feed will produce smaller burrs,and pressurizing the hole, during

final drilling, with compressed air will aid the removal of burrs. Finishing

of the orifice can be done with a drill shank and an appropriate polish.

The finished orifice should be inspected for burrs with a microscope, and when possible, measured objectively,e.g., a Talysurf instrument.

Consideration should be given to the possibility usingof an electrical

discharge machine or a laser to manufacture smooth orifices. To the authors'

knowledge, no comparative study of different processesfor production of orifices has been made.

115 lll.D.5. A GeneralPurpose StaticPressure Probe

As discussedin Section II I.D.l, thelong, static pressure survey pipe, with nose located in the settling chamber, ispreferred for centerlinecalibra- tionsof transonic tunnels. This arrangement not only provides a large amount of simultaneousdata but also prevents the passage ofa transonic shock over theorifices. However, thequestionnaire results indicate a large number of transonictunnel operators (primarily smaller facilities) continue to use conventionalprobes. As mentionedpreviously in Section 111.0.2,an advantage of inexpensive,classical probes is their mobility and theconsequent ease of performingflow surveys off centerline. For thebenefit of tunnel operators who wish to continueusing this type of probe, the following probe design is suggested for calibrating transonic and supersonictunnels.

The basicprobe design is presented inFig. 3.0.11. An ogive nose with aneffective fineness ratio of 12 is suggested for tworeasons: (1) overexpansion atthe nose isminimal (e.g.,see Fig. 3.0.3) whichalso minimizes theextent of the supersonic pocket at supercritical speedsand thuswall interference, (2) atsupersonic speeds, the bow shock isattenuated byan ogive noseshape (e.g.,Ref. 19);thus, this design also reduces wall interference at supersonic speeds.

It shouldalso be notedthat the nosedesign specifies a distributed roughness for boundarylayer tripping. The objectiveof this feature is to preventshock-induced, boundary layer separation at all speeds. An additional benefitis reduced sensitivity to Reynolds number.Examples of aboundary layer transition strip on thistype of probe may be found in the report by Ritchie

(Ref.10). The size of grit and length of striprequired for a particular application canbe designed via the criter ia of Braslow and Knox (Ref. 55) and Braslow, et a 1. (Ref. 56).

116 PROBE DIAHETER SHOULO BESELECTED TO OBTAIN A TUNNEL BLOCKAGE -< 0.005% FOR TRANSONICAPPLICATIONS RECOPAENOEO ORIFICESIZE - 0.051 CPl (0.020 IN.)

- e Dl STR I BUTED ROUGHNESS

LC/+PATTER&WE3VIWUED BACK TO 30d FROM OGIVE-CYLINDERJUNCTURE

30'

VIEW VIEW8-14 B-B VIEW C-C VIEW 0-0 O RIFICES SINGLE ORIFICESINGLE6 ORIFICES ORIFICE6SINGLE ORIFICES ROTATED 30' FROM VIEWA-A ROTATED 30' 60' APART FROM VIEW 8-8

Figure 3.0.11 VRAWSONIC/SUPERSONlC STATIC PRESSURE PROBE An orifice diameter of 0.051 cm (0.020 cm) is recomnended for static pres-

sure ports along the cylinder.* The probe is designed to obtain primary static

pressure data at stations having six orificesin order to average out the

effects of any probe asymmetries, orifice errors, and small flow inclinations.

The purpose of the single orifices is to assist in locating the position of

either a transonic shock and/or the reflectionof a bow shock (or any other

disturbances) back onto the probe. The additional data will aid determination of where surface pressure equals freestream static. This feature will allow

the probe to be used off centerline where wall interference increases.

Finally, the flare angle should be 10 deg or less in order to minimize

interference near Mach one. The effects of this flare, as we11 as the wall-

interference-free transonic performance of this probe, can be calculated via the South-Jameson computer code (Ref.17). In the Mach number range of

0.95 to 1.00, it is necessary to keep probe blockage

realize wall-interference-free performance ata tunnel centerline. If the probe is used with higher blockage and/oroff centerline, wall proximity effects on shock locationand surface pressure distribution canbe estimated using the computer program of South and Keller (Ref. 22). In the case of supersonic applications (H > 1.3). probe blockage can be two ordersof magni-

tude larger without any deleteriouseffects. It is only necessary to apply

the criteria of Gray (Fig. 3.0.3) and avoid wall reflections of bow shocks.

The interference-free performance canbe computed with a number of existing asisymmetric method of characteristics codes.

ij A hardened, stainless steel is recommended for durability and corrosion resistance in order to maintain orifice integrity and minimize long-term abrasion by particles in the flow. 1II.D. References

1. Dlck, R. 5.: "The Influence of SeveralCable-Type Supports Upon theStatic PressuresAlong the Centerline Tube in a Transonic Wind Tunnel,"AEDC-TN-54-26, Feb. 1955.

2. Jackson, F. M.: "Supplemental CalibrationResults for the AEDC Propulsion Wind Tunnel (16T) ," AEDC-TR-70-163,Aug. 1970.

3. Jacocks, J. L. and Hartley, M. 5.: "Calibrationof the AEDC-PWT 4-Ft. TransonicTunnel with Modified Wails," AEDC-TR-69-134, June 1969.

4. Jackson, F. M.: "Calibrationof the AEDC-PIJT 16-FtTransonic Tunnel at TestSection Uall Porosities of 2, 4, and 68." AEDC TR-76-13, Jan. 1976.

5. Isaacs. 0.: "Calibrationof the R. A. E. Bedford 8 ft. x 8 ft. Wind Tunnel at Subsonic Speeds, Including a Discussion of theCorrections Applied to the Measured Pressure Distribution to Allow for the Direct andBlockage Effects Due tothe Calibration Probe Shape." ARC R. t M. No. 3583, Feb.1968.

6. Pope, A. and Goin, K. L.: High-speed Wind TunnelTesting, Wiley, 1965.

7. Gunn, J. A. and Maxwell, H.: "Calibrationof the AEOC-PWT Aerodynamic Wlnd Tunnel(4T) MachNumbers 1.6 and 2.0 NozzleBlocks." AEDC-TR-72-Ii1, Sept. 1972.

8. Holder, 0. W.; North, R. J.; andChinneck, A.: "Experiments withStatic Tubes in a SupersonicAirstream, Parts i and 11," A.R.C. R. t M. No. 2782, July 1950.

9. Gracey, W.: "Measurement of StaticPressure on Aircraft," NACA TN 4184, Nov. 1957, and NACA Report 1364, 1958.

10. Ritchie, V. 5.: "Several Methods for AerodynamicReduction of Static- PressureSensing Errors for Aircraft at Subsonic, Near-Sonic and Low Supersonic Speeds," NASA TR R-18, Feb. 1959.

11. Davis, J. W. and Graham, R. F.: "blind-Tunnel blali InterferenceEffects for 20' Cone-Cylinders," AlAA Jour.Spacecraft and Rockets,Oct. 1973.

12. Estabrooks. B. B.: "Wall-InterferenceEffects on AxisymmetricBodies in Transonic Wlnd Tunnels withPerforated Wall Test Sections," AEDC-TR-59-12, June 1959.

1 I9 13. Gadd, G. E.: "Interactions Between Normal Shock Llaves and Turbulent Boundary Layers,'' ARC R & t4 No. 3262, Feb. 1961. 14. Albers, E. E.; Bacon, J. W.; and Nason, B. 5.: "An Experimental Inesti- gation of Turbulent Viscous-Inviscid Interactions," AIAA PaperNo. 71-565, June 1971.

15. Hsieh, T.: "Hemisphere-Cylinder in Transonic, M_ = 0.7-1.0," AlAA Jour., Oct. 1975.

16. Page, W. A.: "Experimental Study of the Equivalence of Transonic Flow about Slender Cone-Cylindersof Circular and Elliptic Cross Section," NACA TN 4233, Apri 1 1958.

17. South, J. C., Jr. and Jameson, A.: "Relaxation Solutions for Inviscid Axisymmetric Transonic Flow Over Eluntor Pointed Bodies," Proc. AlAA Computational Fluid Dynamics Conference, July 1973.

18. Robertson, J. E. and Chevalier, H. L.: "Characteristics of Steady-State Pressures on the Cylindrical Portlon of Cone-Cylinder Bodiesat Transonic Speeds," AEDC-TDR-63-104, Aug. 1963. 19. Capone, F. J. and Coates, E. M., Jr.: "Determination of Boundary-Reflected- Disturbance Lengths in the Langley 16-Foot Transonic Tunnel," NASATN 0-4153, Sept. 1967.

20. Couch, L. M. and Brooks, C. W., Jr.: "Effect of Blockage Ratio on Drag and Pressure Distributions for Bodiesof Revolution at Transonic Speeds," NASA TN 0-7331, NOV. 1973. 21. Nebman, P. A. and Klunker. E. E.: "Numerical Modelina of Tunnel-\.la11 and Body- Shape Effects on Transonic Flow Over Finite Lifting Wings," Aerodynamic Analyses Requiring Advanced Computers, Part 11, NASA SP-347, Mar. 1975. 22. South, J. C., Jr. and Keller. J. D.: "Axisymmetric Transonic Flow Including Wind-Tunnel Wall Effects,'' AerodynamicRnalyses Requiring Advanced Computers, Part 11, NASA SP-347, Mar. 1975. 23. Sieverdling, C.; Maretto, L.; Lehthaus, F.; and Lawaczeck: "Design and Calibration of Four Probes for Usein the TransonicTurbine Cascade Testlng," Von Karman Institute for Fluid Dynamics, Tech. Note 100, May1974.

120 I

24. Sutton, E. P.: "The Development of Slotted Working-Section Liners for Transonic Operation of theR.A.E. Bedford 3-Ft. Wind Tunnel," A.R.C. R. E; M. No. 3085, Mar. 1955.

25. Wright, R. H.; Ritchie, V. S.; and Pearson, A. 0.: "Characteristics of the Langley 8-Ft. Transonic Tunnel with Slotted Test Section,''NACA Report 1389, July 1958. 26. Keller, J. D. and Wright, R. H.: "A Numerical Method of Calculating the Boundary-Induced Interference in Slotted or Perforated Wind Tunnels of Rectangular Cross Section,'' NASA TR R-379, Nov. 1971.

27. Bryer, D. W. and Pankhurst, R. C.: Pressure-Probe Methods for Determining Wind Speed and Flow Direction, National Physical Laboratory, Her Majesty's Stationery Office, London, 1971. 28. Siddon, T. E.: "On the Response of Pressure Measuring Instrumentation in Unsteady Flow," UTIAS Report No. 136 (AD-682 2961, Jan. 1969. 29. Rittenhouse, L. E.: "Transonic Wind Tunnel Results for Five Pressure Probes Designed to Minimize Static-PressureSensing Errors," AEDC-TDR-62-48, March 1962.

30. Chevalier, H. L.: "Calibration of the PWT 16-Ft. Transonic Circuit with a Modified Model Support Systemand Test Section," AEDC TN-60-164, Aug. 1960.

31. Krause, L. N. and Gettelman, C. C.: "Effect of Interaction Among Probes, Supports, Duct Walls and Jet Boundaries on Pressure Measurementsin Ducts

and Jets ,'I I. S.A. Jour., Vol . 9, Sept. 1953. 32. Nichols, J. H.: "Rake Interference Studies at Transonic Speeds,'' AEDC-TH- 56-1, Feb. 1956. 33. Gray, J. D.: "A Compendium of Flow Measurement Methods and Techniques," AEDC von Karman Gas Dynamics Facility, Notes preparedfor a short course at the University of Tennessee Space Institute, Tullhaoma,Tenn., Nov. 1973.

34. Barry, F. W.: 88Determinationof Mach Number from Pressure Measurements," ASME Trans., Apr i i 1956. 35. Walter, L. W. and Redman, E. J.: "Needle Static-Pressure Probes Insensitive to Flow Inclination in a Supersonic Stream,'' NAVORD Report 3694, March 1954.

121

I 36. Washington, W. D. and Humphrey, J. A.: "Pressure Measurements onFour Cone-Cy1 inder Flare Configurations at Supersonic Speeds,"RD-TM-69-11 (AD 699 3591, Oct. 1969. 37. Gray, J. D.: "Eva1uation.of Probes for MeasuringStatic.Pressure in Supersonic and Hypersonic Flow," AEDC-TR-71-265, Jan. 1972.

38. Norris, J. D.: "Calfbrationof Conical Pressure Probes forDetermination of LocalFlow Conditions at MachNumbers from 3 to 6," NASA TN 0-3076, Nov. 1965.

39. Vahl, W. A. and Weirich, R. L.: "Calibrationof 30° Included-Angle Cone forDetermining Local Flow Conditions in MachNumber Range of 1.51 to 3.51.," NASA TN 0-4679, Aug. 1968.

40. Peterson, C. W. andGeorge, 0. L.: "WindTunnel Pressure Probes: New Calibrations for New Geometries and FlowEnvironments," AIAA Jour., Vol. 13, No. 10, Oct. 1975.

41. Donaldson, 1. S. and Richardson, D. J.: "A ShortStatic Probe with Good IncidenceCharacteristics at Supersonic Speed,'' A.R.C. Current Paper No. 1099, 1970.

42. Pinckney, S. Z.: "A ShortStatic-Pressure Probe Design for Supersonic Flow," NASA TU 0-7978, July 1975.

43. Smith, A. H. 0. and Bauer, A. D.: "Static-Pressure Probes thatare TheoreticallyInsensitive to Pitch, Yaw and MachNumber," Jour.Fluid Mech., Vol. 44, Part 3, pp. 513-528, 1970.

44. Hsieh, T.: "Hemisphere-Cy1 inderin Low Supersonic Flow," AlAA Jour., Dec. 1975.

45. Marconi, F.; Yaeger, L. and Hamilton, H.:H. "Computation of High- Speed Inviscid FlowsAbout Real Configurations," NASA SP-347,Mar. 1975.

46. Rayle, R. E.: "Influenceof Orifice Geometry on StaticPressure Measure- ments," ASME Paper No. 59-A-234, Dec. 1959.

47. Shaw, R.: "The Influenceof Hole Dimensionson StaticPressure Measure- ments," Jour. Fluid Hech., Vol. 7, Pt. 4, April 1960.

122 48. Livesey, J. L.; Jackson, J. D.; and Southern, C. J.: "The StaticHole Error Problem: An ExperimentalInvestlgation of Errorsfor Holes of VaryingDiameters andDepths,'' Aircraft Engr., Yo1 34, Feb. 1962. 49. Rainbird, W. J.: "Errorsin Measurement of Mean StaticPressure of a Moving Fluid Due to Pressure Holes," DME/NAE Quarterly Bulletln No. 1967 (3) , Nat ' 1. Res.Counc. Canada, Oct. 1967.

50 ,Franklin, R. E. and Wallace, J. M.: "Absolute Measurements of Static-'Hole Error UsingFlush Transducers," Jour. Fluid Mech., Vol 42, Pt. 1, June 1970.

Nestler, D. E.: "StaticPressure Port Errors in Hypersonfc Turbulent Flow," AlAA Paper No. 71-270, Mar.1971.

Cassanto, J. M.: "An Assessment ofPressure Port Erosion Effects," AlAA Paper No. 75-150, Jan. 1975.

Benedict, R. P.: Fundamentals of Temperature,Pressure, and Flow Measurements, Wiley, New York, 1969.

Benek, J. A. and High, M. D.: "A Method Forthe Prediction of the Effects

of Free-Stream Disturbances on Boundary-Layer Transit ion,I' AEDC-TR-73-158, Oct. 1973, also AlAA Jour., P. 1425,Oct. 1974.

55. Braslow, A. L. and Knox, E. C.: "Simp1 ified Method forDetermination of CriticalHeight of Distributed Roughness Particlesfor Boundary-Layer Transition at MachNumbers from 0 to 5," NACA TN 4363, 1958.

56. Braslow, A. L.; Hicks, R. M. and Harris, R. V., Jr.: "Use ofGrit-Type

Boundary-Layer-Trans it ion Trips on Wind-Tunnel Model s ,I1 NASA TN 0-3579, 1966.

57 Parker, R. L., Jr.: ltFl0w GeneratipnProperties of Five Transonic Wind funnelTest Section Mal 1 Configurations,'' AEOC-TR-75-73, Aug- 1975. .. Ill. E. MEASUREMENT OF FLOWANGULARITY

The term yawmeter wi 1 1. beused herein' to denote probes designed to measure flowangularity in either two orthree dimensional flows. A widevariety of yawmetershave been used overthe years for different applications. The ones r' whishhave been used intransonic tunnels may be dividedinto three general types:(1) pressure probes, (2) hotwire or film probes,(3) force models I instrumentedwith a verysensitive force balance." A discussionof pressure probes is given first and isfollowed by a brief description of twoexamples

of thelatter types of yawmeter. A recentreview of the various types of pressure probes and a discussion ofthe pros andcons of each aregiven by Bryer and Pankhurst(Ref. 2). These authorsclassify pressure probe yawmeters into two categories:(1) those consistingof an arrangement of open-ended tubes, and (2) thosehaving a body withpressure sensing orifices. These may be subdividedinto probes designed to measure flowangles in one plane (2-D) or two(3-D).

III.E.l. DifferentialPressure Yawmeters: 2-D

Forflow direction measurements in oneplane, three types of yawmeter geometriesare most common, viz., twopressure taps on the surfaces of either awedge or a circularcylinder and twotubes withslanted inlets. A large varietyof suchprobes are available from commercial manufacturers. The circularcylinder yawmeters arenot recommended fortransonic flows because oftheir comparatively large interference with the flow and considerable sensitivityto Mach number, Ref. 2. Boththe cylinder andwedge have,greater susceptibility to error in thepresence of velocitygradients because of the largerseparation of orifices. Also, Sieverding, et al. (Ref. 3) havefound that a 30° (totalangle) wedge-shapedyawmeter is Reynolds number dependent inthe Machnumber range 0.8 to 2.2. In contrast,the two-tube type of yawmeter,generally referred to as a Conradprobe, provides: (1) minimum flow disturbance,(2) adequate sensitivity which is relatively free of Machnumber and Reynolds number effects (Refs. 4 and 5). and (3) orificeswhich are close togetherfor nearly point measurement of flow angularity.

124 References 3 and 5 containcalibration results for small tube type yawmeterswhich were designed to investigate the flow out of transonic turbine cascades and compressors. The objectiveof Sieverding, et al. k. (Ref. 3) was toinvestigate several probe geometries and arrangementswhich i couldbe used tosimultaneously measure total, static, and directional . .. pressures. The two-tube yawmeter, shown inFig. 3.E.1, was mounted on a rake ofrectangular cross section (2.3 mn thicknessnormal to flow and 6.0 m para1le1 to flow). The yawmeter was arrangedto measure flow angularity normal tothe plane of the rake. A smalldiameter (1.7 mm), truncated,2S0'.' total apex-anglecone probe was also mountedon therake. The separation distance between the yawmeter and theconical probe was 16 mm and bothnose tips werelocated 22 mm ahead of therake. The probes, withthis arrangement,were calibrated in the DFVLR/AVA Transonic Wind Tunnel (Im x lm). The resultingsensitivity * of the yawmeter is presented inFig. 3.E.l. 0.8 M 2.2. "

A combinationPitot probe andyawmeter was also calibrated with a similar arrangement,but two different companion probeswere used. In one case,the companion probe was a 150 coneneedle probe (1.5 mm OD) for measuring static pressure, and the secondcompanion probe was a 30' cone probe (1.5 mm OD). The sensitivityof the needle probe was found to be linearover the largest range of angles of yaw (%loo). The corresponding sensitivitydata are also shown inFig. 3.E.1. The differencein the angle sensitivity of the two-tube yawmeterand thecombination probe may be attributedto the difference in the inlet angle.

Sieverdlng,et al. (Ref. 3) also tested the yameter and needleprobe combinationin the small (135 rnm x 50 mm) VKI High Speed Cascade Tunnel C-2. Staticpressure measurements alongthe wall of this facility, with and with- outthe probes, indicated significant blockage whenMach number exceeded 0.30; whereas,a single AGARD needleprobe (12.3O cone and 1.5 mm OD) with a standard elbowtype support (3 mn OD) located 48 probediameters downstream showed . negligibleblockage. Hence, theseauthors conclude: if yawmetera (and/or otherprobes) are to be used intransonic tunnels, the probes and support

Althoughthese authors did not state the inside diameter of thetubing, .the ratio of I.D. to 0.0. should be keptgreater than 0.6 inorder to minimize loss or change ofsensitivity with increasing flow angularity, e.g., p. 19 of Ref. 2.

125 0.06

0.05

1J 1- Probe,Ref.3Combination 0.04 a

S :t Y 3 Tubes 0.8 mm OD 0.03 .a,.\J

0.02

\ Two Tube Probe, Ref. 3

0.01 - Tubing - 1.575 mrn OD x 0.254 rnm Wall Lp=j 6 = 80" 2 Tubes 1 .O mm OD

0 I I I I I I I I I 0.4 0.6 0.8 0.6 0.4 1 .o 1.2 1.4 1.6 1.8 2.0 2.2 M Figure 3. E. 1 . TWO D I MENS I ONAL YA\JMETERS mechanism_sh.oy.jd.~ . - ... ..be.c,a!ib_r_ated ~ in a tunnel of size similar to which it is to be used. And most importantlyfor tunnel calibrations, extreme care mustbe

taken. . ...~ ,in designing" ya.wmeter supports in order to avoid inducing extraneous flow angularity.

The flowangle sensitivity data of Buzzell (Ref. 5), obtainedwith a combi- nationPitot probe and yawmeter, is also shown inFig. 3.E.1 for M = 0.4, 0.6, and 0.8. These datawere obtained with a cobra-styled*rake with three cornbina- tion probesalternating with four single Pitot tubes. Again, this investigator followedthe accepted practice of locating the probe to measure anglesnormal to theplane of the rake. Mach number sensitivity checks indicatednegligible change in yawmeter sensitivityfor 0.41 5 M 5 0.81. Afterthis was ascertained,eight rakeswere tested at a meanMach number of0.6. The resultingspread in data are shown inFig. 3.E.1.

Vidal,et al. (Ref.6) have recently reported using a two-tube yawmeter to measure flowangularity in asmal 1 (30.5 cm) transonicwind tunnel. The objectiveof these tests is to compare measurements offlow angularity and staticpressure with calculated interference-free transonic flow solutions for a givenmodel, and therebyprovide a criterionfor adjusting tunnel wall porosity toattain wall-interference-free flow. A planar yawmeter was usedbecause an airfoil, which spans thetunnel, is being used for developmentaltesting. The yawmeter is similarto the one shown inthe lower right portion of Fig. 3.E.l.and was constructedof 0.0635 cm (0.025 in.) OD tubingwith each inlet chamfered at 45". These authorsclaim that their flow angle sensitivity is such thatwith a pressureresolution of 0.0007 N/cm2 (0.001 psi)they can "inprinciple" *+ resolveangles to within 0.03" in the Mach number range 0.55 to 0.725.

As noted by Bryer and Pankhurst(Ref. 2), yawmeter sensitivityin low speed flows has traditionally been defined as

A S" a(AP/q)/a$ (3.E.1) Y where AP isthe pressure difference across the yawmeter. However, these

+These rakeswere designed toplace six of them CirCUmferentiallYin the ** statordischarge plane of a compressor. Since the original writing of this section, the calibration anduse of a new yawmeter design has come toour attention, Lind (Ref. 28). This 2-D yameter consists of a two-hole, differentialpressure probe placed at the vertex of a forward-sweptwing. A yametersensitivity of 0.163 and anaccuracy of 0.01 deg is claimed for low-speedflows (M, -< 0.17). 127 authorsnote that, for compressible flows, less variation in sensitivity with Mach number is obtainedby using the following definition.

An example of this may befound inthe paper by Spaid, etal. (Ref. 7). In theirexperiments with a miniature (1 mm total span)combination Pitot probe and yawmeter, sensitivity, as definedin Eq. (3.E.21, didnot vary with 0.8 5 Ms 1.0 and -30" 5 3, 5 30". Inaddition, the supersonic data in Fig. 3.E.1 show much lessvariation with Machnumber when S is used.Thus, thecompressible Y definitionof yawmeter sensitivity, Eq. (3.E.2), is preferredfor transonic and supersonicapplications.

!ll.€.2. DifferentialPressure Yawmeters: 3-D

Forthe general case of flowangularity calibration in anempty tunnel(tran- sonicand/or supersonic), a pyramid yawmeter is recommended. The pyramidgeometry has two primaryadvantages compared to a conical or hemispherical yawmeter. ~irstl~, thepyramid probe performance is lesssensitive to positioning of the orifices. Secondly, it is relativelyfree of interference betweensimultaneous measurements ofpitch and yaw. Inaddition, the incompressible flow measurements of d. Bryer,et al. (Ref. 8) indicate S" for a pyramidprobe is comparativelyinsen- Y sitiveto Reynolds number and increasesslightly ("6%) with increased turbu- 1 ence.

A typicalpyramid yawmeter is shown inFig. 3.E.2.Here, theratio of orifice diameterto probe stem diameter is 0.16. Ingeneral , it is recommended that thisratio be keptless than 0.20. An additional,suggested constraint is thatthe diameter of the orifices be no smallerthan 0.508-mm (0.02in.) in -9. -9. orderto avoid clogging and excessivetime lag problems. The apex angle was chosen to conformwith the recommendation ofBryer andPankhurst (Ref. 2). These authorssuggest that a yawmeterbe designed so thatthe bow shock wave remainsdetached orattached throughout the range of Mach numbers withinwhich measurements areto bemade. Thus, an apexangle of 66" will maintain a detachedshock fortransonic Mach numbers up to 1.6. (A Mach number of 1.6

*Smallerorifices (0.25 mm) have been used at low speeds (Ref. 9) andsupersonic speeds (Ref. 2, P. 57). However, no significant improvement inperfor- mance is gained, and clogging and decreasedresponse time can make theprobe more expensiveto use.

128 D = 0.3175crn - ,/ 330 // &

I d = 0.16 D b16D"I

Figure 3.E.2. PYRAMID YAWMETER Is chosenbecause thisrepresents the upper limit ofoperation for the maJorlty oftransonic tunnels.) Beyond this speed regime,the bow shock will be attached, and theprobe can be used insupersonic flows. The purposeof this design featureis to avoid the sudden changes that canoccur inthe pressure response of sharp-nosedprobes near the shock attachment Mach number. Furthermore, Bryer and Pankhursthave noted that the maximum sensitivity of most pointed ymeters is obtained with an apex angle between 60 and 70 degrees.

Flnally,for use in transonic flows the probe stemshould extend down- streamfor a distanceof at least 16 diameters. Downstream of this station, thestem can safely be enlargedby a IO0 conicalflare to mate withthe avail- ableprobe support. Provided a massive,transverse probe support is not used, thisdesign wlll avoidinterference betweenprobe and support at transonic speeds.

The flow anglesensitivity, a(AP/HS)/aY, ofconical and hemispherical-nose probeshave been found toincrease with Mach number and reach a maximum ofabout 0.025 at M - 1.5 (e.9..Ref. IO andRef. 2).Further increases in Mach number resultin decreasing sensltivity. For example, datafor a hemispherical yawmeter withorifices located 45' from thenose indicate a 50% loss in sensitivity at H = 2.7, Fig. 35 of Ref. 2. Similarly,theoretical calculations * for a 60' conicalyameter indicate a 70% loss at M = 3.5, Fig. 3.E.3. If we assume similarbehavior for thepyramid probe and a maximum sensitivity of 0.025, the smallest change inflow angle which can be detected by a pressuremeasuring system with a resolutionof 3.45 x N/cmZ (0.005 psi) is

Inthe case of a transonictunnel with T = 37.8OC. Re/m = 19.7 x 106 , and S 2 H = 1.0, thesettling chamber pressure is 13.79 N/cm (20 psia).Substituting thisvalue for HS in Eq. (3). we findthat a flowangle of 0.01 degreecan theoretically be resolved.In practice, the effects of probe andfor support deflections,nonidentical internal geometry ofthe tubing and passageswhich

* For this reason,Barry (Ref. 11) and Zumwalt (Ref. 12) exploredthe use of Pitot probeslocated near the surface of wedges and cones toprovide increased sensitivityat high supersonic Mach numbers.However, for most applications, theconventional surface pressure yameters provide adequate sensitivity up to H = 3.5.

I30 .030 I I I I I I I I I

.025 -I P a EXPERIMENTALDATA (Ref. 25 / ,020 \ - CURVE FROM Ref. 10 -I

.005

0 0 I .o 2.0 3.0 4.0 5.0

MACHNUMBER

Figure 3.E.3. SENSITIVITY OF 60" CONICALYAWMETER connect two orifices to a differentialpressure transducer, vibration, turbulence, etc., may preventthe attainment of suchaccuracy. However, thepyramid probe can provideadequate anyle resolution for calibration of mostwind tunnels. Especiallyin light of the fact that the majority of tunnel operators are satisfied with a calibrationof (tunnel-empty) flow angles accurate to within 0.1degree.

If less-accurate,flow angularity measurements aresatisfactory and a

simultaneous measurement of Pitotpressure is desired, the nose of the pyramid yawmeter may be truncatedand an orifice placed in the center of the nose,;: e.g.,Ref. 9. Inthe case of subsonic tunnels, this would provide a convenientcheck on the uniformityof total pressure. In the case ofsupersonic tunnels,this permits Machnumber to be determinedsimultaneously, e.g., Refs. 14 and 15. Such a probenot only minimizes calibration time but also eliminates any uncertaintyin local Mach number atwhich flow angles are measured.

In summary, thesuggested dimensional criteria for a pyramid yawmeter shouldresult in a probewhich:

I. has a flow anglesensitivity which is relatively insensitlve to extraneous flow variablessuch as Reynolds number and turbulence,

2. is small enough to map flow angularityin most tunnelswith high resolution and minimum interference,""

3. has fast enough pressureresponse for most applications,

4. hasadequate structuralstiffness, and

5. can be used to calibrateboth transonic and supersonictunnels.

*It is recommended thatthe lip thickness be keptthin ( 0.005 cm) andthe orifice be beveled at an angleof 15" or more inorder to minimize sensitivity of the Pitot probe to flowangularity. ;:::In largetunnels &3 ft) where highresolution of theflow angularity field isnot required, the recommended pyramidprobe may be scaled up tolarger sizes.In any event,probe blockage should be lessthan 0.1 percentfor generalcalibrations of transonic tunnels. However, near M = I values an orderof magnitude smaller may be necessary in order to avoidprobe-wall interference, seeRef. 13. YaKmgJer Calibration, Rake Interfer-ence, and Blockage . . """ . .

As notedby Pope and Goin(Ref. 10, P. 134), a1 1 real yawmetershave asymmetriesand imperfectionswhich cause theprobe to indicate a nonzero AP at $ = 0. Thus, a yawmeter shouldalways be calibrated in flow conditions . similarto those in which it is to be used. The calibrationprocedure for differentialpressure yawmeters is describedin Refs. 10, 14, and 15 and will not berepeated here. However, it isrelevant to sound a note of cautionhere. When calibrating a yawmeter, thecenter of rotation should be at the nose tip.

Also,careful measurements ofthe angles bet W eenyawmeter axis and tunnelaxis areessential since these must be subtracted fromthe flow angles relative to theprobe inorder to determine flow angular i tywith respect to the tunnel center1 ine.

Inorder to reduce tunnel calibration t I me, most operatorsprefer to useprobes in rakes orarrays. However, Hartley and Nichols(Ref. 16) conductedtests in the AEDC 16T Tunnel withfive 7.62 cm (3 in.)diameter hemispherical yawmetersmounted on a 2.44 m.(8 ft) widerake. The rake consistedof a 22" (total-angle) wedge with a 7.62 cm (3 in.)wide base and was center mountedon a stingsupport. The yawmeterswere mounted 0.61 m (2 ft)apart with the nose locatedapproximately four diameters ahead ofthe leading edge ofthe wedge. The totalwind-tunnel blockage of the rake was approximately 1%. These investigatorsfound that the rake induced significant outflowtoward the tips of the rake. The inducedflow angularity, at the tips

ofthe rake, increased from about 0.5" at M = 0.6 to over 1" at M = 1.1. As Mach number increased from 1.1 to 1.2,the induced flow angularity decreased sharply and exhibitednear-interference-free characteristics for 1.2 5 M 5 1.5. A wall-mountedstrut support, with the same wedge angleof 22", was alsotested and found to induceeven larger outflow from the wall towardthe tip.

In bothcases, the support-induced flow angularity was ascertainedby mounting a singleprobe on a longsting. The firstsection of the sting had a diameterof 5.715 cm (2.25in.) and a lengthof approximately 16 probe diameters. The second section of sting had a diameter of 7.62 cm (3.0in.) and a lengthof over 20 probediameters which subsequently joined a conical flare and the rest of the st ingsupport mechanism. The stingsupport system enabledvertical traverses w ith thesting at zero angle of attack.In addition, theblockage of thesingle p robe was only 0.013%. Thus, the arrangementassured

133

I as near-interference-free,flow angularity measurementsas can be expected in a windtunnel.

An additionalconclusion reached by Hartley and Nichols (Ref. 16) isthat therake had negligibleeffect on flowangles normal tothe plane of the rake (;.e., with the rake vertical, the yaw datawere Val id and with therake horizontal,the pitch data wereVal id). Thus,based on these and other similar results, it is possibleto useyawmeters in a carefullydesigned rake arrangement to make twodimenslonal measurements."However, forgreatest accuracy, a singleprobe joined to a longsting with a supportwhich is symmetrical about thetunnel centerline is recommended.""

Finally,with regard to wind tunnel blockage at Mach numbers near 1 .O, thedata of Couch and Brooks(Ref. 13) indicatethat even withextremely smallvalues of model blockage (<0.0003) wallinterference occurs. Thus, yawmeters for measurements near Mach onemust be designedwith the utmost care, viz.,small probes and longstings, and theresulting data should be scrutinized for anysudden or unexpectedvariations around M = 1.0.

III.E.3. Hot-Wire/FilmYameters

Two hot-wiresinclined at anangle with respect to each other and the mean flow havelong been used in low speed flows to measure flowangularity (Ref.17). Three-wire probes have also been used extensively to simultaneously measure pitch and yaw inthree dimensional flows. In the past, hot-wires have not been used intransonic flows becausethey are so easilybroken. However, Hortsman and Rose (Ref. 18) have recentlydemonstrated that low aspect ratio (,f/d-lOO) tungstenwire probes can be used intransonic flows without a prohibitive breakageproblem. Also recently, Johnsonand Rose (Ref. 19) have reportedusing an X-arrayhot-wire to measureReynolds stressin a supersonic boundary layer and in a shock-wave/boundary-layer interaction (Ref. 20). Thus, although matching the sensitivities of two or more wiresfor accurate

;bDudzinskiand Krause(Ref. 4) pointout that a rakewith circular arms is unaffected by anglesof attack. Whereaswhen nonzero yaw anglesexist in subsonicand/or transonic flows, noncircular arms caninduce larger flow angularityat the nose of the yawmeter and alsocreate undesirable side forces on therake. **Another alternative wouldbe to calibrate a rakefollowing the procedure of Ref. 16. * man flow measurements canbe a problem, thereappears to be no prohibitive reason why two and three-wireprobes cannot be used asyawmeters. Inorder to avoidthe problem of matching sensitivities of more than one wire, Rosenberg (Ref.22) has successfully used a singlewire probe mounted in a rotatableholder. By rotatlng an inclined-wireabout the axis of the probe’s stem and takingdata at two distinct orientations, the three components of velocity and mass flux canbe determined at a point in a generalthree-dimensional flow.

In a study of the effects of contouring slotted walls to reduce transonic- wall- Interference, Weeks (Ref.23) has used a hot-f i lm probe for accurate measurement of flowangularity. This work involved the use of airfoil modelswhich spanned thetest section of the AFFDL Trisonic Gasdynamics Facility. The required planar measurements of flow angularity were obtained wlth a split-film, 22O total-angle wedge which was manufactured by Thermo Systems, Inc.,according to an AFFDL design.This probe consists of a quartzrod 0.152 cm (0.06 in.)in diameterwith the tip ground to a symmetrical wedge. As indicatedin Fig. 3.E.4, therod extends forward 1.27 cm (0.5 in.)from a0.3175 cm (0.125 in.)diameter supporttube. A pair of platinum films, 0.102 cm (0.04 in.)long, are deposited on each sideof the apex ofthe wedge. Four gold-filmleads are used to complete the twoseparate anemometer bridgecircuits. The calibrated yaw sensitivity

of this probe is shown in Fig. 3.E.4 for 0.85 ”< M < 0.95. Weeks claims that this probe will resolveflow angles to within 42 minutes of arc (0.03O). The primarylimitation is stated to be probevibration which was determinedexperi- mentally to induce errors 5 0.5 minutes of arc.

In summary, sincehot-film probes are (1) less delicate, (2) lesssuscep- tible to contamination because oftheir larger size, and (3) canhave corrosion resistantcoatings, hot-film yawmeters aresuperior to hot-wire yawmeters.Both hot-wires and hot-films require specialized data processing equipmentwhich may be considered a disadvantageby potential users. However,based on theresults obtainedby Weeks, hot-film yawmetersappear to offer a viable alternative to differentialpressure yawmeters.””

The uses ofhot-wires and hot-filmsare discussed further in Appendix 1. * Reference 21 alsodiscusses the fact that the calibration of a hot-wireis susceptible to change with time because ofcontamination and corrosion.This ** may require frequent calibration checks. Although the authors are not aware of any transient measurements of flow angles with a hot-wire/film,there is no inherent reason why this type of yawmeter can not be used in a continuous-traverse mode. Thiswould provide the advantage of reducedtunnel calibration time, e.g.,see discussion of forcebalance yawmeters. 135 To Separate Anemometer D I MENS IONS IN CENT IMETERS Bridge Circuits

Figure 3.E.4.SPLIT HOT FILM, 20" WEDGE PROBE CALIBRATIONBRIDGE VOLTAGE DIFFERENCE vs FLOW ANGLE, REF. 23 136 ForceBalance Yaweters The basicprocedure ofrunning a windtunnel force model upright and invertedto determine the average pitch angle is well known and is standard practicein professional wind tunnel testing. However, theuse of a small wedge mounted on a sensitive force balance to obtain a measure of tunnel-empty flowangles is new.Maxwe1.1 and Luchuk(Ref. 25) have recentlyreported the results of transonictests with this type of yawmeter.The probeconsists of a 20' included-angle wedge, with a 14.73 cm (5.80 in.) span,mounted on a specially-designed, two-component forcebalance, Fig. 3.E.5. The force balance was designed to measurenormal force and pitching moment with very thin,strain-gauged sections for maximum sensitivity.

The probe was testedin the AEDC Aerodynamic Wind Tunnel (4T) overthe Machnumber range of 0.6 to 1.3. The calibratedflow angle sensitivities, based on variationsof normal force and pitching moment, are shown inFig. 3.E.6. Althoughthe yaw sensitivity of thepitching-moment mode isapproximately 50% larger, Maxwell and Luchukfound thatflow angle measurements obtained from either mode were of equalaccuracy.

The wedge was supported by the 4T six-degree-of-freedom,captive trajectory system. Thispermitted the probe to be moved continuouslywith a variety of movements. Maxwell and Luchukconclude that flow direction data can be obtained"with an absolute accuracy that is little different from theaccuracy withwhich the probe is aligned." Furthermore, based oncomparisons ofdata obtainedwith the probe at rest and in motion,reliable and rapid measurements can be made withthe probe moving continuously with combined linear and rolling motion. The estimated rms deviationsfrom a mean value of flow angle was

I0.023O at all measured points. However, 24' sweeps in pitch and yaw produced larger variations, viz., -+O.O8O and -+0.25', respectively. These data were obtained with pitch and yaw rateswhich varied, respectively, from 1.16 to 1.28 and 1.01 to 1.36 deg/sec.

137 t L DIMENSIONSIN CENTIMETERS

1.67 Dia. \ 7I 14.73 I' I """""""

" "- - I -L """_ I"""_ $""_ \ ? 2-COMPONENTBALANCE INSTALLED HERE J

LOCATION OF GAGE SECTIONS

I 20°

Figure 3.E.5. GEOMETRY OF AEDCFORCE BALANCE YAWMETER 0.5 0.14

I CJ W n -0.12 U z V .I irr n W 0 n. A C 0.3 i0:lO f CNa J - LL U LL 8 ~0.08 V W rr0 0 LL z- 0.1 I 0 k n

0 0.04 L 5 0.6 0.7 0.8 0.9 1 .o 1,1 1.2 1.3

FREESTREAM MACH NUMBER, M

Figure 3.E.6. SENSITIVITY OF THE AEDCFORCE BALANCE YAmETER Unfortunately,the Reynolds number dependence of this yawmeter was not investigated. However,.some variationof total pressure at M = 0.6 indicated a decreasingsensitivity with increasing unit Reynolds number. At thispoint, there'ader may recallthat Sieverding, et al. (Ref. 3) alsoreported their wedge shapedyawmeter exhibited a Reynolds numberdependence.* Thisimplies that wedge shapedyawmeters should be calibrated as a function of Machnumber andReynolds number.Hence, thistype of yawmeter will bemore tediousto use. Additionaldisadvantages of the wedge forcebalance yawmeter is higher initial costs and pitch and yaw mustbe measured separately.** However, a forcebalance canbe calibratedto relate angular deflection of modeland supportto changes inloading. This provides a distinct advantageover differentialpressure and hot-wire/film yawmeters. Inconclusion, the force balanceyawmeter's advantage of rapid, continuous measurements offlow angularity appears tooutweigh the disadvantages.

* Neither of these yawmeters hada boundarylayer transition strip. Thus, dependence of wedge-yawmeter sensitivity on Reynolds number couldconceivably ** bereduced by utilizing a grit-type, boundarylayer trip. A double,intersecting wedge probe with four component forcebalance has more recently been constructed and testedat AEDC (Summers, Ref. 26). This yameterenables pitch and yaw datato be obtainedsimultaneously in even less time. An improveddesign, which isless sensitive to unsteadytransonic flow,specifies a small,symmetrical centerbody with flat plate wingsattached inorthogonal planes, Ref. 27. Recentexperience with this type of yawmeter in the AEDC 4T Tunnel indicatessimultaneous measurements of pitch and yaw canbe obtained at 225 pointsin less than six minutes and with anaccuracy of 0.01 degree.

140 Ill. E. References

1. Lennert, A. E.; Hornkohl, J. 0. and Kalb, H. T.: "Application of Laser Velocimetersfor Flow Measurements," Instrumentationfor AirbreathingPropulsion, Progress in Astronautics and Aeronautics, AIM Vol. 34, MIT Press, 1972.

2. Bryer, D. W. and Pankhurst, R. C.: Pressure-Probes Methods forDetermining Wind Speed and Flow Direction, Her Majesty'sStationery Office, London, 1971.

3. Sieverding, C.; Maretto,L; Lehthaus, F. and Lawaczeck, 0.: "Design and Calibrationfor FourProbes for Use inthe Transonic Turbine

Cascade Testing ,I' VKI TN 100 (AD 922 286) , May 1974.

4. Dudzlnski, T. J. and Krause,L. N.: "Flow-Direction Measurement with Fixed-Position Probes," NASA TM X-1904, Oct. 1969.

5. Buzzell, W. A.: "CalibrationResults for Stationary Pressure Rakes Sensing Yaw Angle Downstream of an Axial Compressor Stage," ARL TR 75-0104, Apri 1 1975.

6. Vidal, R. J.; Erickson, J. C. and Catlin, P. A.: "Experiments with a Self-

Correcting Wind Tunnel," Windtunnel"." . Deslgn and Testing Techniques, ~~ AGARD-CP- 174, Oct. 1975.

7. Spaid, F. W.; Hurley, F. X. and Hellman, T. H.: "MiniatureProbe for TransonicFlow Direction Measurements," AlAA Jour.Vol. 13, No. 2, Feb. 1975, P. 253.

8. Bryer, D. W.; Walshe, D. E.; and Garner, H. C.: "PressureProbes Selected for Three-DimensionalFlow Measurement," R.&M. No. 3037, 1958.

9. Schulze, W. M.; Ashby, Jr., G. C.; and Erwin, J. R.: "SeveralCombination Probes forSurveying Static and TotalPressure and FlowDirection," NACA TN 2830, Nov. 1952.

10. Pope, A.; and Goin, K. L: ."High-speed - ~ ~ " Wind Tunnel Testing, Wiley, 1965.

11. Barry, F. W.: "Comparison ofFlow-Direction Probes at Supersonic Speeds," J. Aero. Sci., Sept. 1961, P. 750-752.

141 12. Zumwalt, G. W.: "ConicalProbes forDetermination of Local MachNumbers and Flow Directionin Supersonic Wing Tunnels," SCTM 355-60(71), SandlaCorp., Nov. 1960.

14. Vahl, W. A. and Weirich, R. L.:"Calibration of 30"Included-Angle Cone forDetermining Local Flow Conditions In MachNumber Range of 1.51 to 3.51," NASA TN D-4679, August 1968.

15. Norris,J. D.: "Calibrationof Conical Pressure Probes for Determlnation of LocalFlow Cond i t ions at MachNumbers from 3 to 6," NASA TN D-3076, Nov. 1965.

16. Hartley, M. S.; and Nichols, J. H.: "Effectsof Rake Blockageon Flow Angularity Measurements atTransonic MachNumbers in the AEDC-PWT 16-FootTransonic Tunnel," Twenty-Fifth Supersonic Tunnel Assoc. Meeting, NASA LangleyResearch Center, May 1966 (referenced with, author'spermission).

19. Johnson, D. A. and Rose, W. C.: "LaserVelocimeter and Hot-wire Anemometer Comparison In a SupersonicBoundary Layer," AiAA Jour.(Tech. Notes), Vol. 13, No. 4, April 1975.

20. Rose, W. C. and Johnson, D. A.: "Turbulence in Shock-Wavea Boundary-Layer Interaction," AiAA Jour.,Vol. 13, No. 7, July 195 21. HotWire-Hot Film-Ion Anemometer Systems, CAT/FORM 6560375,Thermo-Systems inc., 1975.

22. Rosenberg, R. E.: "A ThreeDimensional Hot-wire Anemometry Technique Employing a SingleWire Probe," ARL 71-0039, March 1971.

23. Weeks, T. M.: "Reduction ofTransonic Slotted Wall interference by Means of SlatContouring," AFFDL-TR-74-139, March 1975.

142 24. Maxwell, H. andLuchuk, W.: "Evaluation of a Wedge on a ForceBalance as aFlow Angle Probe," AEDC-TR-74-110,Feb. 1975.

25.Raney, D. J.: "Flow Direction Measurements in Supersonic Wind Tunnels," Aero. Res.Coun. Lond. , Current Papers No. 262, 1956.

26. Summers, W. E.; personalcomnunication, AEDC, Feb. 1976.

27. Luchuk, W.: "Flow Angle Measurements Using a 2-Inch Span Cruciform-Wing Force Model," presentedat 45th Semi-Annual STA meeting,Albuquerque, N.M., April 1976 (referencedwith author's permission).

28. Lind, I. A. : "A SensitiveFlow Transition Probe," KTH Aero Memo FI 175, TRITA-FPT-019, lnstitutlonenfor Flygteknik Stockholm, Sweden, July 1975. 1II.F. MEASUREMENT OF UNSTEADY FLOW DISTURBANCES

Theneed for measurements of flowunsteadiness was briefly reviewed in Section ll.C.6. The primaryobjective of noise calibration of a windtunnel is to obtain a measure of the fluctuations in static pressure and flow angu- laritythat exist in the empty testsection. Here a review will be given ofthe instrumentation that hasbeen used to obtainthis type of data. How- ever,before discussing sensors, it is germain tonote the amplitudes and frequencies of unsteadystatic pressure which characterize transonic tunnels.

Inthe center of transonic test sections the fluctuating pressure coefficient,defined as

- AC = - x 100 percent, Pq may range from 0.5% to 5% dependingon the tunnel configuration, Mach number, and Reynolds number. Dougherty, etal. (Ref. 1) havenoted a value of 0.45% corresponds to a level of sound which istyp.icaIly radiated from turbulent boundarylayers on solid test section walls. However, Hartzuiker,et al. (Ref. 2) have pointedout that AC actuallydecreases with increasing Reynolds numberand P 6 8 rangesfrom 0.5% at Re = 5.7 x IO to anestimated 0.2% at Rex = 1.1 1.7 x IO , X - see Fig. 3.F. I. McCanless andBoone (Ref. 3) havereviewed noise measurements made inboth perforated and slottedtest sections. These authorsnote that in perforated-walltunnels center1 ine measurements of AC tend to be lower (40-609;) P thanwall measurements;whereas, theopposite trend is found inslotted-wall tunnels.Generally, the edgetones generated by perforated-wall tunnels* tend to make thesetunnels noisier than slotted-wall tunnels. For example, AC datafor P twelveperforated-wall tunnels range from 1% to 7.4%; whereas, datafor five slotted-walltunnels showa range of 0.5% to 2%) Ref. 3. A peak in AC is P usually measuredbetween M = 0.70and 0.80 forboth perforated and slotted-wall tunnels. The frequencyspectra of noise at M = 0.80 is presented inFig. 3.F.2 fornine different, continuous tunnels, Ref:2. These dataare representative ofsolid, perforated, and slotted-walltunnels and arepresented in terms of the Mabey spectrumparameter which is discussednext.

* Recentresearch indicates there are anumber of ways to reducethe level of noiseqenerated by edgetones, Refs. 1 and 4..

1 44 6

.OM

Fig. 3.F.1FREQUENCY SPECTRA OF NOISE FROMTURBULENTA BOUNDARYLAYER ON A SOLIDWALL, Ref. 2

,014

,010

.w1

.ffl

Fig. 3.F.2 NOISEFREQUENCY SPECTRA FOR SOME EXISTING CONTINUOUSWINDTUNNELS AT HaD = 0.80, Ref. 2 * As noted by Harttuiker, et al. (Ref. 2) , flow qual ity can af.fect the results of tests on: (1)dynamic stability, (2) staticforces and moments, (3) buffet, and (4) flutter.Tests of these quantities generally involve increasinglyhigher frequencies In theorder listed. Apparently, little work hasbeen done to measure fluctuations in flow angularity and correlate thesewith pressure fluctuations. However, Mabey (Refs. 5-7) and Harttuiker, et al. (Ref.2) have found the effectson model testsof both pressure and incidence fluctuations can be correlated by using a spectrum function defined as follows:

Here AC2 is the mean-squared value of thefluctuating statlc pressure coef- P 2 ficient, and F(n) isthe contribution to AC perunit bandwidth at the reduced P frequency n.Mabey hassuggested

canbe used to measure windtunnelflow unsteadiness if the reducedfrequency is chosen tocorrespond to a naturalfrequency of the model, e.g., fundamental wingbending mode, torsion mode, etc. A variety of tests have shown thet greater model excitationfollows increases in nF(n). Thus, criteria for acceptable flow quality canbe established for various types of tests bysuc- cessivelyreducing nF(n) until the results approach anasymptote and cease to varysignificantly with tunnel flow quality. For example, analysesof buffet measurements on aircraft models with different natural frequencies led Mabey (Ref. 7) to concludethat an acceptablelevel of flow unsteadiness for the detectionoflight buffeting is e 0.002. Testsatother facilities haveconfirmed the usefulness of Mabey'sspectrum parameter for correlating a variety of dynamicmodel tests, e.g., Hartzuiker,et al. (Ref. 2). Unfortunately, a preciseboundary cannot bedrawn toseparate acceptable from unacceptable levelsof flow unsteadiness. Rather, there is a "grayregion" separating flow qualities that are either acceptable or not for a given type of test.

* Reference 2 is mainly concerned with estimating the flow quality that will be necessary to make the LEHRT cost-effective in light of the planned 10 sec run t ime. 146 However, the utility of including fluctuating pressure measurementsin tunnel calibration is now well established.

Condenser microphone measurements on the AEDCIO deg transition cone,in six different transonic tunnels, indicate98 percent of the energyof background pressure fluctuations are contained within0-20 KHz, Ref. 8. However, since there is presently no criterion for an upper limiton frequency beyond which boundary layer transition is unaffected, Westley (Ref. 9) recommends that the frequency rangeof noise measurements extendat least up to 30 KHz.

Thus, acoustic calibration of transonic tunnels requires instrumentation that can measure dynamic pressures with these ranges of amplitudes and frequencies. The sensors employed should also be relatively insensitive to vibrations of the mounting surface and durable enoughnot to be easily damagedby either particles in the ,flow or overloading.

III.F.1. Dynamic Pressure Measurements A rather wide variety of instrumentation has been used to measure unsteady flow disturbances in wind tunnels. Condenser microphones, strain gage,and piezoelectric dynamic pressure transducers have been employed for noise measurements in stilling and plenum chambers, dlffusers, and on test section walls, and models and probes located on the centerline, Refs. 10-15. In addition, hot-wire anemometers have beenused to measure flow disturbancesin stilling chambers (e.g., Ref. 13) and in the test section of transonic tunnels(e.g., Refs. 9 and 16) and supersonic tunnels (e.g., Ref. 17). Also, laser Doppler velocimeters (LDV) are being used to measure turbulence by an ever increasing number of tunnel operators, Ref. 9.

Unfortunately, this lack of standardization makesit difficult to compare measured levels of flow disturbances. For example, Lewis andDods (Ref. 18) noted significant variations in the frequency responseof 12 different microphones and dynamic pressure transducers. In general, Lewis and Dods found small diameter transducers (0.1 to 0.3 cm) gave higher power-spectral-density values, at all frequencies, than larger diameter transducers (0.5 to 1 cm). Also, the high frequency portion of the spectrumof pressure fluctuations varles with the particular sensor,and as is well known, the rms valueswill be under- estimated when a significant portion of thehigh frequencies are attenuated. In

1 47 addition, when measuring vorticitywith hot-wires, hot-films, or an LDV, the data may alsovary because of differences in frequency response.

The comparisonproblem is compounded furtherby the choice of sensor mounting. Ideally,acoustic pressures should be measured with no relative motion between sensor and thetest medium, Ref. 19. However, sincethis is neitherpractical nor relevant to windtunnel model testing, some representative location on aprobe, model or tunnelwall must be selected.

The first measurements ofwall pressure fluctuations beneath turbulent boundary layersin awind tunnel were reported by Willmarthin 1956 (Ref.20). Willmarth (Ref.21) has recentlyreviewed the problems of dynamic pressure measurements attunnel walls and notesthat most ofthese measurementshave been made withflush-mounted transducers. Hanly (Ref. 22) has recentlystudied the effect of sensorflushness on fluctuating-surface-pressure measurements at M = 1.68, 2.0, and 2.5. These tests show spectralpressure measurements are extremelysensitive to flushness with protrusion causing greater error than submergence. Hanlyconcludes that more repeatabledata canbe obtainedwith transducers mounted approximately 0.0254 cm (0.0l"in.).beneath a surface orifice. Thus, it is recmendedthat acoustic measurements attunnel walls conform to this criterion.

Also, it isrelevant to note here that two or more wall-mountedtransducers canbe used todetermine useful correlations between disturbanceswhich may existin atunnel, e.g., Refs. 11 and 25. Inaddition, Boone andMcCanless (Ref. 1 I) have notedthat wall data canbe used toextrapolate, through M = 1, measurements obtained at the centerline with probes or modelswhich may be subject to oscillating shocksand/or othermodel-induced unsteadiness near Mach one. However, experience(Ref. 8) with microphone measurementson a 10 degcone indicates (1) thisis not aproblem and (2)the real advantage of awall-mounted sensor is that it can be calibratedwith respect to centerline measurementsand usedas a permanent monitorfor assessing anysubsequent changes in tunnelflow unsteadiness.

Concerninginstallation of sensors inperforated walls, Credle and Shadow (Ref.24) mounted a 0.635 cm (1/4 in.)piezoelectric microphone inthe center of anarea which was filled andsanded smooth. The radiusof the area was

148 approximately 40 microphonediameters. These investigatorsstated:

"Thisinstallation technique precluded the measurement of purely near- field influence of the mostadjacent upstream holes and allowedfor the measurement of what might be consideredas the radially integrated averagevalue of pressure fluctuations at the wall surface."

Credle and Shadow alsoinstalled an identical,but shielded, microphone in the wallin order to monitor microphone response to wall vibration. In general, it is considered good testing practlce to ascertain the component of a microphone'soutput which is due tovibration. Finally, questionnaire results indicatedthat strain-gage transducers are most often used for acoustic measurements at windtunnel walls. Presumably, this is because theinstrumentation requiredto process the signal from a strain gage is readily available at most wind tunnel s.

Inaddition to tunnel wall measurements, tunnelnoise calibrations require dynamic pressuredata near the center of the test section. Some of the first such measurements in a transonictunnel were reported by Chevalier and Todd (Ref.25). Inthese initial tests, dynamic pressuretransducers were mounted ona wedge, a wingprobe, and an ogive-cylinder.Later acoustic measurements in the AEDC-PWT 16T and 165 tunnels wereperformed with condensermicrophones and strain gage transducers mounted on a 10 deg included-anglecone, Ref. 10.

A varietyof other probe geometries have also been used.For example, an ogive-cylinder anda flat plate have been used in some of the NASA Ames tunnels,Ref. 23. A 10 deg cone-cylinderprobe hasbeen used inthe 8 x 6-ft. Supersonic Wind Tunnel at NASA Lewis,Ref. 26. As partof a reviewof probes foracoustic calibration, Boone and McCanless (Ref. 11) consideredslender cones, wedges, flatplates, hemispheres, and sharp-tipped,flow-thru, circular cylinders. These authors recommended a 10 deg apex-angleconical probe for wind tunnelacoustic measurements because:

1. cones arenot as sensitiveto tip flow as wedges and flat plates, 2. thetransonic Machnumber range,where unstable shock waves occur, issmaller for slender cones thanfor flat plates, flow- thru cy1 i nders , and hemi spheres, 3. a slendercone introduces minimal disturbance to the flow.

As noted by Credle and Shadow (Ref. 24), a smallerangle cone ispreferred, but a 10deg cone is aboutthe minimum anglewhich will allow installation of instrumentationunder a laminarboundary layer. These same authorsalso reportedthat by 1970 the IO deg cone had become a standarddevice at AEDC for calibratingwind tunnel flow disturbances.

Up tothis time the AEDC acousticcalibration cone had two flats,located 180 deg apart,for flush mounting of sensors. By 1970 experiencewith this cone indicated satisfactory noise measurements couldprobably be made with a completely symmetricalcone and two 0.635 cm (1/4 in.)condenser microphones. Furthermore, * thesurface of the cone was polishedto an rms finishof 3 microns anda traversingPitot probe mechanism was mounted aft of the cone for boundarylayer transitionstudies, Ref. 27. Thissubsequently becameknown asthe AEDC transi- t ion cone andhas now been used to measuredynamic flow qua 1 ity in over I8 domestic and foreigntunnels, Refs. 8, 28 and 29.

The work of Dougherty with the AEDC transition cone and the correlations of Pate and Schueler (Ref. 30) and Benek and High(Ref. 31) have established a directrelationship between AC and boundarylayer transition. In addition P toproviding a common measure of dynamic flowquality, Treon, et al. (Ref.32) reporteddata from the AEDC transition coneenabled better agreement to be obtainedin a seriesof tests on a transportaircraft model. Inthese tests, the same model was testedin the AEDC-PWT 16T, the NASA Ames 11-ft. Transonic Wind Tunnel, and theCalspan 8-ft. TransonicTunnel, and differencesin drag coefficients, measured atzero-normal-force, were found to be less when a correctionfactor was used to accountfor relative Reynolds number effects between facilities. An effective Reynolds number forthe Calspan and Ames tunnels,relative to the AEDC 16T tunnel, was defined on thebasis of a common, boundary-layer-transitionlength. Thus, theutility of a standardacoustic calibrationdevice hasbeen demonstrated.Subsequent to a planned flight calibration,this device will also be usefulin correlating tunnel and free-flight conditions.

Because ofthe demonstrated utility of the AEDC transition cone and its pastuse in a number ofmajor facilities, Westley (Ref. 9) recommends noise and turbulencelevels in transonic tunnels bemeasured with two 10 deg cones fitted, respectively,with: * Apparently,the surface finish was later improved to 0.25 microns,Ref. 8.

I50 1. skin-friction gages todetermine transition Reynolds numbers and flush-mountedmicrophones to measure noise levels onthe test section centerline,

2. a crossedhot-wire anemometer mountedon thetip (an ONERA design) . The proposalto * eliminate the traversing probe mechanism will reduceprobe inducednoise and windtunnel blockage. Although Westley's recommendation is notspecific, it is assumed that he isnot recommending floating-elementskin- friction balancesbut rather heat transfer devices suchas thin-films or thermocouples for transitiondetection, e.g., Refs. 11, 34 and 35. Withregard to hot-wire measurements,Westley expresses a concensus thatthese instruments areideal for measuring disturbances in a windtunnel test section. Because of this importance and thefact that recent research has demonstratedhot-wires can be used effectively in some transonictest sections (Ref. 161, theyare discussedseparately in Appendix 1.

Alternate Acoustic Calibration Probes -=.---

Not onlywould the AEDC transition cone be expensiveto reproduce, but noise measurementson it aresusceptible to a number of probe-inducederrors. Credle and Shadow (Ref. 24) observedthat pressure gradients exist on a cone at subsonic and transonic speeds, andhence acoustic measurements are influenced byboth static and totalpressure gradients. A laminarboundary layer may modulateacoustic disturbances which passthrough it to an underlyingsensor, Ref. 9. Also,Siddon (Ref. 36) has shown thatboth axial and lateralfluctua- tions cancause errorsin measurements of AC with probes. Hence, otherprobe P alternatives may offer some advantages for acoustic calibration of wind tunnels.

The followingconcise summary ofSiddon's work (Ref. 36) isextracted from Willmarth'sarticle (Ref.21). ---" Siddon has reportedconstruction of an excellentprobe for unsteady static-pressure measurements, and hehas calibrated it invarious contrived flows to remwe the errors causedby theinteraction

* Credle(Ref. 33) notedearlier that the traversing probe support structure appeared togenerate additional noise, basedon comparisons withacoustic dataobtained on an ogive-cylinder and the AEDC 10 deg cone with flats.

151 of the body of the probe with streanwise and cross-flow velocity fluctuations. The pressure is transmitted to thediaphragm of a miniaturecondensor microphone(0.25 dm diam.) insidethe probe (0.305 cm diam.) throughan annular slit approximately 2 diametersdownstream from the tip of thebalsa wood, venose. A 0.318 cm (1/8 in.)collar around the probe, downstream of the sl t, was carefully positioned to make thesteady pressure at the slit equal to thefree-stream static pressure when there is nocross flow. The probe- co lar compensation was checked at zero angle of attack in a flow with sinu- so dalaxial velocity fluctuations and was foundadequate.

"Siddon'sunique achievement is the development of a compensation scheme to cancelthe pressure fluctuations produced by cross-flow fluctuations. His scheme is based uponan earliertransducer in which piezoelectric force-sensing elementswere used to measure lift fluctuations of a smallairfoil, which are proportionalto velocity fluctuations normal to the airfoil when theangle-of- attackfluctuations are small. In the present caseSiddon used an arrangement of fourpiezoelectric Bimorph plateelements in anI-beam configurationto measure theorthogonal bending moments producedby cross-flow-induced transverse forces on the nose of theprobe. If quasi-steady,slender-body, aerodynamic theoryis applicable, the transverse force will be proportionalto the instan- taneoustransverse velocity at the nose. The two electricalsignals representing theorthogonal components of transverse velocity wereeach squared and summed withanalog-computer elements to obtain a signalproportional to thesquare of thetransverse velocity. A fractionof this signal was added tothe pressure measuredby thecondensor microphone togive (approximately) the true static- pressurefluctuations that would have existedin the absence ofthe probe.

"Siddon calibratedthe probe in a flowproduced by a jet of air passed through a rotatinginclined nozzle. The probe slit was placedat the point where thestatic pressure is constant (theintersection of the nozzle axis and theaxis of rotation). At thatpoint any fluctuatingpressure signals from the condensormicrophone were assumed to beproduced by cross-flow interaction with theprobe. These signals werecancelled by additionof the proper fraction of thesquare of the transverse velocity. "Siddonconcluded that the error in the pressure produced by cross-flow interactionin turbulence was lessthan 20 percent. Thus, reasonablyaccurate measurements offluctuating pressure could be made as long as theassumption of quasi-steadyflow was notviolated and thetime lag between thetransverse-force signalfrom the probe noseand thepressure measured at the annular slit was notimportant. generally, th.is requ:ires that the spatial scale of the pressure

"fluct-uations be much largerthan the probe dimensions.* As a result of his workSiddon was ableto conclude that in many practicalcircumstances where only root-mean-squarepressure fluctuations were measured withprobes 1 ike Strasberg's(Ref. 37), the.correction for cross-flow interaction is 1 kely to be small. Owing todifferences in the corrected and uncorrected wave forms, one mustuse thecorrected pressure when instantaneousvalues are desi red even when thecorrected and uncorrectedroot-mean-square pressures are the same."

Thus, for anumber of aerodynamics and acousticreasons, as well as simplicity and economy, Westley(Ref. 9) notesthat anumber of opera ors of transonic and supersonictunnels are measuring fluctuating static and pi tot pressures. He recommends thatthese types of acoustic probes bedeve opedand standardized. Dynamic staticpressure probes should be developed for high speed flows and probablyshould follow the Siddon type design.

Inthe case of dynamic Pitot probes, a seriesof design studies at NASA Langley has culminated in a designwhich appears to be satisfactoryfor acousticcalibration of wind tunnels, Refs.38, 35 and 14. A schematic ofthe probereported in Ref. 14 is shown inFig. 3.F.3. Briefly, it consistsof two0.318 cm (1/8in.) diameter piezoelectric transducers mounted in tandem. The probediameter of 0.635 cm (1/4 in.) was selected because mean pressure measurements indicatethe pressure is nearly constant across the center of a flat-faceddisk in supersonic flow. The diaphragm ofthe exposedtransducer is coveredwith a thin coating of RTV rubber to reduce vulnerability to damage by particlesin the flow. The purpose ofthe shielded rear transducer is to serve as an acceleration(or vibration) monitor. The signalfrom this transducer issubtracted from the exposedtransducer inorder to account for the effects nThe underlining was inserted by thepresent authors to point out that the size of the AEDC transition cone may induceerrors in noise measurements.

153 AUDimensions in Centimeters

rEmsed tnnrduccr

Figure 3 .F. 3. SMALL PIEZOF&EC!i'RIC DYNAMIC PRESSURE PROBE, Ref. 14 of probevibration.* In order for this technique to be valid, Anders, etal. (Ref.14)noted that the two transducers mustbe matched to give identical outputsfor given acceleration levels.

In conjunction with the probe, a hot wire was also used in the stilling chamber and test section of a small Mach 5 windtunnel at NASA Langley. Assuming purelyacoustical disturbances, the following relation was used to relate the fluctuating static pressures obtained from the hot wire to fluc- tuatingPitot pressures.

Here isthe rms fluctuatingtotal pressure behind a normalshock, is the rms fluctuating,freestream static pressure, andu isthe S sound-source velocitydetected by thehot wire. The resultsfrom this equation gave excellent agreement withthe fluctuating Pitot probe data, e.g., see Appendix I, Fig. 7. It isrelevant to notehere the conclusion reached by Anders, etal. (Ref.14).

"The hotwire and Pitot probegenerally indicate the same trend and levelwith respect to the Reynolds number. This agreement isof great practicalimportance since the piezoelectric Pitot probe is a much more ruggedinstrument with simpler data reduction procedures than the hot- wireprobe. For diagnostic studies, the Pitot probe ca,n provideessen- tially the same information as thehot-wire probe .with much less effort. However, thehot wire does have one particular advantage inthe present investigation.That is, in a pure sound fieldthe hot wire can distinguish betweenmoving sources and fixed sources."

A similar comparison of hot wire data with fluctuating Pitot probe data hasbeen reported byGrande and Oates(Ref. 39). However,a 1.78 rnm (0.070 in.) diameter strain gage transducer was employed and datawere obtained for 1.1 < M < 2.25. Nondimensionalized power spectraldensities obtained in a turbulent boundary layer and inthe freestream were found to agree remarkably

*Dougherty and Steinle(Ref. 8) reportthat the accelerometer used inthe AEDC transition cone has notdetected any significant vibration effects during testsin a number oftunnels. However, Stainback,et al. (Ref. 35) didreport significant probe vibration effects.

155 we1 1. These authorsconclude that "the frequency response characteristics of thestagnation pressure sensor are identical to those of thehot wire, i.e., it canbe considered an ideal point sensor for fluctuations with spatial scales somewhat largerthan the probe diameter.''

To summarize theadvantages and disadvantages of fluctuating Pitot probes, thefollowing points are noted.

D isadvan tages :

1. cannotseparate the three possible flow disturbance modes of entropy,vorticity and pressure.

2. cannot deduce whetherdisturbance sources are stationary or movi ng.

3. shockmodulation of disturbances may be unknown in some cases (e.g.,see Ref. 39 ).

Advantages :

1. relativelyinexpensive and off-the-shelf, commercialtransducers are readi ly available.

2. speedand ease of measurement.

3. simplerdata reduction.

4. durable,;.e., farless susceptible to particle damage compared to hot wires.

5. highsignal to noise ratio.

6. reduced influenceof flow perturbations associated with finite probesize (compared withthe AEDC transitioncone). 7. minimum wall-probeinterference. 8. easily moved about to surveyentire test section.

Inconclusion, recent research with fluctuating Pitot probes indicates theseinstruments may be adequate forinitial calibration of flow disturbances intransonic and supersonictunnels. This type of measurement couldserve as aconvenient and inexpensivestandard to compare tunnelnoise levels. It 1s alsorelevant to note that Doughertyof thePropulsion Wind Tunnelgroup at AEDC plansto use afluctuating Pitot probe to monitor freestream disturbance levelsduring flight tests with the AEDC transition cone.

1 57 1II.F.References

1. Dougherty, N. S. ; Anderson, C. F.; and Parker, R. L. : "An Experimental Investigation of Techniques to SuppressEdgetones from Perforated Wind

Tunnel Wa 1 1 s ,I' AEDC-TR-75-88, Aug . 1975. 2. Hartzuiker, J. P. ; Pugh, P. G. ; Lorenz-Meyer, W. ; and Fasso, G. E. : "On theFlow Quality Necessary for theLarge European High-Reynolds-Number TransonicWindtunnel LEHRT," AGARD-R-644, March 1976.

3. McCanless, G. F. and Boone, J. R.: "NoiseReduction in Transonic Wind

Tunnels ,I1 J. Acoust. SOC. Am., Vol . 56, No. 5, Nov. 1974.

4. Schutzenhofer, L. A. andHoward, P. W. : "Suppression of BackgroundNoise in aTransonic Wind-Tunnel Test Section,'' AlAA Jour., Nov. 1975.

5. Mabey, D. G.: "FlowUnsteadiness andModel Vibrationin Wind Tunnels at Subsonic and Transonic Speeds," C.P. No. 1155, Aero. Res.Coun., 1971.

6. Habey, D. G.: '#AnHypothesis for the Prediction of Flight Penetration of Wing Buffeting from Dynamic Testson Wind TunnelModels," RAE TR 70189, Oct. 1970.

7. Mabey, D. G.: "The Influence of FlowUnsteadiness on WindtunnelMeasure- ments atTransonic Speeds," RAE Tech. Memo. Aero. 1473, 1973.

8. Dougherty, N. S. and Steinle, F. W.: "TransitionReynolds Number Comparisons in SeveralMajor Transonic Tunnels," AIAA Paper No. 74-627, July 1974.

9. Westley, R.: "Problems of Noise Measurement in Ground-Based Facilities

with Forward-Speed Simulation(High-speed Wind TunnelNoise) ,I1 Appendix 5 of AGARD-AR-83, Sept. 1975.

IO. Riddle, C. D.: "Investigationof Free-StreamFluctuating Pressures in the 16-Ft. Tunnels ofthe Propulsion Wind Tunnel Facility," AEDC-TR-67-167, Aug. 1967.

11. Boone, J. R. and McCanless, G. F.: "Evaluationof the Acoustic Sources of BackgroundNoise in Wind Tunnel Facilities,'# NASA CR-98155. 12. Boone, J. R. andMcCanless, G. F.: "Application of the Techniques for Evaluatingthe Acoustic Sources of Background Noise in Wind Tunnel Facilities," Tech. Rept. HSM-R05-69, ChryslerHuntsville Operations, Mar. 1969.

13. Credle, 0. P.: "An Evaluationof the Fluctuating Airborne Environments in the AEDC-PWT 4-Ft.Transonic Tunnel," AEDC-TR-69-236,Nov. 1969.

14. Anders, J. R. ; Stainback, P. C. ; Keefe, L. R. ; and Beckwith, I. E. : "Sound and FluctuatingDisturbance Measurements in the Settling Chamber and Test Section of a Small, Mach 5 Wind Tunnel," ICIASF '75, Int'l Congress on Instrumentationin Aerospace SimulationFacilities, Ottawa, Canada, Sept. 22-24, 1975, publishedby IEEE, 345 E. 47th Street, New York.

15. Anders, J. B.; Stainback, P. C.; Keefe, L. R.; and Beckwith, 1. E.: "FluctuatingOisturbances in a Mach 5 Wind Tunnel,"Proc. AIAA 9th Aerodynamics TestingConference, June 1976.

16. Horstman, C. C. and Rose, bl. C.: "Hot Wire Anemometry inTransonic Flow," NASA TM X-62495,Dec. 1975.

17. Donaldson, J. C. and Wallace, J. P.: "Flow Fluctuation Measurements at

MachNumber 4 inthe Test Section of the12-Inch Supersonic Tunnel (D) ,I' AEDC TR-71-143, Aug. 1971.

18. Lewis, T. L. andDods, J. B.: "Wind-TunnelMeasurements of Surface-Pressure Fluctuationsat MachNumbers of 1.6, 2.0, and 2.5 Using 12 Different Transducers," NASA TN D-7087, Oct. 1972.

19. Fuchs, H. V.: "Measurement of PressureFluctuations Within Subsonic TurbulentJets," J. Sound & Vib., Vol. 22, No. 3, 8 June 1972.

20. Wi 1 Imarth, W.W. : "PressureFluctuations Beneath Turbulent BoundaryLayers," AnnualReview of Fluid Mechanics,Vol. 7, Annual Review Inc.,Palo Alto, Cal if., 1975.

21. Wi 1 Imarth, GI. W. : "Unsteady Force and Pressure Measurements ,I' Annual Review of Fluid Mechanics, Vol. 3, AnnualReview Inc., PaloAlto, Calif., 1971.

22. Hanly, R. D.: "Effectsof Transducer Flushness on Fluctuating Surface Pressure Measurements, AIAA Paper No. 75-534, Mar. 1975.

23. Dods, J. B. and Hanly, R. D.: "Evaluation of Transonic and Supersonic Wind- TunnelBackground Noise and Effectsof Surface Pressure Fluctuation Measurements," AlAA Paper No. 72-1004, Sept. 1972. 159 24. Credle, 0. P. and Shadow,T. 0.: "Evaluation of theOverall Root-Mean-

Square Fluctuating Pressure Levels in the AEDC PWT 16-Ft.Transonic Tunnel ,I' AEDC-TR-70-7,Feb. 1970.

25 Chevalier, H. L. andTodd, H. E.: "Measurement of thePressure Fluctuations in the Test Section of the 16-FootTransonic Circuit in the Frequency Range from 5 to 1000 cps," AEDC-TN-61-51 (AD 255 7631, May 1961.

26. brabinus, R. J. and Sanders, B. W. : "Measurements ofFluctuating Pressures in 8-by6-Foot Supersonic Wind Tunnel for MachNumber Range of 0.56 to 2.07,'' NASA TM X-2009, May 1970.

27 Credle, 0. P. and Carleton, W. E.: "Determination of Transition Reynolds Number inthe Transonic MachNumber Range," AEDC-TR-70-218, Oct. 1970.

28. Dougherty, N. S., Jr.: 'IPrepared Comment on Cone Transition Reynolds 1 Number Data Correlation Study,'l AGARD-CP-187, June 1975.

29. Whitfield, J. and Dougherty, N. S., Jr.: "A Survey ofTransition Reynolds Number Work at AEDC," to be presentedat AGARD Fluid Dynamics Panel Symposium onLaminar-Turbulent Transition, Copenhagen, Denmark, May 1977.

30 Pate, S. R. and Schueler, C. J.: "RadiatedAerodynamic Noise Effects on

Boundary LayerTransition in Supersonic and HypersonicMind Tunnels ,'I AlAA Jour.,Vol. 7, No. 7, Mar. 1969.

31 Benek, J. A. and High, M. D.: "A Method forthe Prediction of theEffects of Free-Stream Disturbances on Boundary-Layer Transit ion," AEDC-TR-73-158, Oct. 1973, also AlAA Jour., P. 1425, Oct. 1974.

32 Treon, S. L.; Steinle, F. M.; Hagerman,J. R.; Black, J. A., and Buffington, R. J.:"Further Correlation of Data fromInvestigations of aHigh-Subsonic- Speed TransportAircraft Model in ThreeMajor Transonic Wind Tunnels," AlAA PaperNo.71-291, Mar. 1971.

33 Credle, 0. P.: "PerforatedWall Noise inthe AEDC-PWT 16-Ft. and 4-Ft.

Transonic Tunnel s ,'I AEDC-TR-7 1-21 6, Oct . 197 1 . 34. Heller, H. H. andClemente, A. R.: "FluctuatingSurface - Pressure Characteristics onSlender Cones in Subsonic,Supersonic, and Hypersonic Mach-Number Flow," NASA CR-2449, Oct. 1974.

160 35. Stainback, P. C. ; Wagner, R. D. ; Owen, F. K. ; andHorstman, C. C. : "ExperimentalStudies of HypersonicBoundary-Layer Transition and Effects of Wind-Tunnel Disturbances," NASA TN D-7453, Mar.1974.

36. Siddon, T. E.: "On the Response of PressureMeasuring Instrumentation in Unsteady Flow," UTlAS Rept. No. 136, Institute for AerospaceStudies, Univ. of Toronto, Jan. 1969.

37. Strasberg, M.: "Measurements ofthe Fluctuating Static and Total-Head Pressures in a Turbulent Wake," DavidTaylor Model Basin Rept. 1779, (AD 428 7001, Dec. 1963 (a1 so AGARD Rept.464).

38. Stainback, P. C. and Wagner, R. D.: "A Comparison ofDisturbance Levels Measured in HypersonicTunnels Using a Hot-wire Anemometer anda Pitot Pressure Probe,'' AlAA Paper No. 72-1003,Sept. 1972.

39. Grande, E. and Oates, G. C.: I'Response ofMiniature Pressure Transducers toFluctuations in Supersonic Flow," Instrumentationfor Airbreathing

."' ."' Propulsion Progress in Astronauticsand Aeronautics, AlAA Vol. 34, MIT Press, 1972.

161. .I.. .I.. , ...... - .. . .. _... .._

111.6. TRANSONIC TUNNEL BOUNDARY CONDITIONS AND WALL INTERFERENCE

I I I.G.l. Conventional Ventilated Walls

The history of development of ventilated walls for transonic tunnels has been reviewed by Goethert (Ref. 1). The three primary milestones in this development wereas follows: 1. Theoretical analyses in Germany, Italy, and Japan during the 1930's indicated tunnel walls with proper arrangement of longitudinal slots would provide wall-interference-freeflow simulation. This work wasinterrupted by World War 11. 2. During 1946, Wright and Ward (Ref. 2) developed a "subsonic theory for solid-blockage interference in circular wind tunnels with walls slotted in the direction of flow." Subsequently, a 12-in. diameter tunnel was designed with ten evenly spaced slots providing a total openness ratio of one-eighth. The tunnel was put into operation in 1947. This design did indeed prevent choking and enabled testing thru Mach one of a model with 8.5% blockage. 3. Unfortunately, the solid slats in slotted tunnels were found to cause significant reflecticns of bow shocksand expansion waves at supersonic speeds. Thus, around 1950 theoretical analyses at Cornel1 Aeronautical Laboratory* indicated better shock wave cancel lation could be achieved with small-grain porous walls, Goodman (Ref. 3). Unfortunately, exploratory testsshowed such walls clog easily, and even worse, the porosity needed to vary with each change in Mach number and/or shock strength. As an outgrowth of thiswork, the now familiar perforated wall was selectedas a convenient compromise.

The early mathematical models of tunnel-wall-interference werebased on the governing differential equation for perturbation velocity potential in subsonic, compressible flow, e.g. ,'Baldwin, et el. (Ref. 4).

A The current name of this facilityis Calspan.

162 If the wa 11 is so lid, the boundary conditionfor no flow through the wall is

-=a4 solid wall. an 0 at solid (3-6.2)

In the case of an open-jet, there can be no pressure difference across the boundary; thus

-=a4 0 at open boundary. ax (3.G.3)

The corrections to measured valuesof model lift and pitching moment, which result when solving Eq. (3.6.1) with either solid or open-wall boundary conditions, are discussed in detail by Garner, et al. (Ref. 5). The theoretical results generally agree with experiments. In order to facilitate applications ofthls type of boundary-induced corrections, Heyson (Ref.6) has compiled solutions in the form of curvesand charts.

In the case of ventilated walls, the boundary conditions become more complex. In fact, the centralproblem of theoretical analysis of transonic-wall-interference is selection of the appropriate boundary conditions to use withEq. (3.G.l). This is still an area of active research,and only a brief review of boundaryconditions for ventilated tunnels wi 11 be given here.

An approximate boundary condition for porous walls was derivedby Goodman (Ref. 3, Part II), viz.,

+ -1 0 R at ideal, porous wall.

This boundary condition was derivedby assuming the average velocity normal to the wall is proportional to the pressure drop through the wall(a linearized approximation to viscous flow through the wall),and that the pressure outslde the wall is equal to freestream static. The value of R, for a given wall, is usually determined experimentally by measuring pressure drop and the associated mass flow through a wall sample (e.g. Ref. 71, i.e. , An approximate,uniform boundary condition for slotted walls was derived byBaldwin, etal. (Ref. 4).

-a4+K- a20 = 0 atideal, slotted wall ax axan (3. G. 6) where K is related to slot geometry by

DS wS K = In {CSC (E -lr -11 , DS and

Ds = distance between slotcenters,

Ws = widthof slots.

In anattempt to account forviscous effects, Baldwin, et al. suggested addingthe porous boundary condition to Eq. (3.6.6)and measuring R forthe slotof interest.

-a+ + K -& + -1 -a4 = 0 at viscous, slotted wall ax axan Rs an (3.6.8)

Keller (Ref. 8) has recentlysuggested this boundary condition be extended to includevarying slot width by replacing l/RS with l/Rs + aK/ax.

After more thantwo decades of testing and comparisons of theory with experimentalresults, it is now generallyrecognized that application of the linear boundaryconditions, with constant v aluesof K and/or R, is inadequate. As an example ofthis discrepancy, a brief summary of a typical case is presentedhere.

Lowe (Ref. 9) measured thewall porosi ty parameter for a wall with 22.5% porosity and normalholes. The standardporosity parameter, R, was determined experimentally by measuringthe static pressure drop and mass flowacross a

164 9-inchby 21-inch section of a sidewall of theGeneral Dynamics 4-footHigh Speed Wind Tunnel.Data were obtained for Mach numbers of 0.5, 0.7, and 0.9 anda correspondingunit Reynolds number range of 19.7 to 37.7 mill ion(per meter),Using the measured valuesof R and theresults of 1 inearperturbation theoryobtained byLo and Oliver(Ref. IO), the upwash and stream1ine curva- turecorrections indicated the wall did not have the characteristics of an open jet.This contradicted the results of tests with an aircraftforce model which,when corrected for open-boundaryinterference,agreed with data obtained I withthe same model inthe Langley 8' TransonicPressure Tunnel (Ref. ll)*. ! Thus, theresults of Lowe, as well asothers, indicate the measurement of R and I use ofthe classical, linear perturbation theory is not very useful for calibrat- ingthe effects of transonic wind tunnel walls. If,in genera1,this approach to correctingfor wall interference had provensuccessful, measurements of R for porous and slottedtunnels would have become a standardpart of transonic tunnel calibration.

The current consensus is:the true, transonic-tunnel boundary conditions are dependent on thelocal flow conditions near the wall. This, in turn, means adependence onboth tunnel operating conditions @the particular model configuration, e.g., Newman and Klunker (Ref. 14). Recent effortsto obtain improvedboundary conditions for fixed (passive) wall conditions include the studyof variations in R between top and bottomperforated walls, Ref. 15, and nonlinearboundary conditions for walls with normal holes, Ref. 16, and slotted walls, e.g.,Refs. 17 and 18. Of course,the basic objective of these studies is toattain data correction procedures which can reliably account for theeffects of real, ventilated walls.

111.6.2. AdaptiveWall Studies

The difficulties in applying transonic wall corrections, which do not reduce to simplemodifications of speed and angle of attack, are well known. Also, one ofthe conclusions obtained with the conventional, linear theory of walleffects is the impossibility of using uniform porosity to simultaneously - * The recent,supercritical airfoil tests of Elackwell and Pounds (Ref. 12) indicatethe actual boundary conditionshifts towardthe free-jet as porosity increases,i.e., the transonic shock moves forwardfor a given Mach number. This same trend was alsoobserved asa resultof increased blockage in the supercriticalcone-cylinder tests of Page (Ref. 13).

165 eliminatethe effects of wall interference on both normal force and pitching moment, Ref. 16.

Forthese reasons, other procedures have been suggested and are currently beinginvestigated. The theorydeveloped by Ferri and Baronti(Ref. 19) and Sears(Ref. 20) seems to offerthe promise of being more correct. These authors suggested that the pressure distribution and the streamline deflection angle be measured alongthe tunnel walls (outside the boundary layer) with amodel in place. The scheme theninvolves calculation of (I) theflow-deflection angles whichcorrespond to the measured pressuredata and (2) thepressure distribution correspondingto the measured flow deflectionangles. The difference between the measuredand calculated pressure distributions andstreamline deflections arethen used to accomplish one of the following:

1. determinethe wall porosity which eliminates wall interferencefor a givenexternal pressure distribution, 2. providethe correct pressure distribution outside ofthe porous wall for a given porosity distribution, 3. determinewall contours to conform with free-air stream1 ines, or 4. calculatethe wall corrections to be appliedto the experimentalresults.

One of the advantages ofthis procedure is that it onlyrequires the linearizedperturbation equations to be valid nearthe wall. This means the procedure may be valid as long as supersonicpockets do notextend to the tunnelwalls. The primaryadvantage ofthis procedure is that it uses data toestablish the appropriate boundaryconditions. However, asnoted by Ferri and Baronti,the primary disadvantages are the requirements for "accurate measurements of flowdeflections and pressurevariations at several angular positions and at many stationsalong the test section."

Inconjunction with the theory of Ferri and Baronti,an experimental program was begun in the 15" Trisonic Gasdynamics Facility at the Air ForceFlight Dynamics Laboratory.Since Streamline angles can bemeasured more accurately away from thewall, the theory was subsequentlymodified to use flow angles and staticpressures measured at an intermediate"midfield1' location between the modeland wall.For angularity measurements, a new hot-film, 20 deg-wedge probe

166 was designed and fabricated.Calibration tests show it to be capableof resolving flow angles to within +2- minutes of arc (Ref.21). This probe, togetherwith a conventional, cone-cylinder,static-pressure probe,provides theinput required bythe Ferri and Barontitheory.

Results of this workhave demonstrated the feasibility of changing slotted-wallcontour to minimize transonic-wall-interference with the flow over 6% thickbiconvex airfoils at zero angle-of-attack. As expected,the resultsfor nonzero angles-of-attack indicate the wall contour will need to be changed as changes in lift and/or model configurationare made. The studyof lifting airfoil models iscontinuing. However,enough resultsare now availableto conclude that this approach offers a decidedadvantage over the pre- viousapproach of measuringpressure drop and mass flow through a wall sample and then trying to use linear boundaryconditions toestimate wa1l-model interferencefactors.

Work is also underway atthe University of Southampton(England), Ref. 22, and ONERA (France),Ref. 23, on usingadjustable, solid walls to conform with free-airstreamlines.

Vidal,et al. (Ref. 24) have reportedon recent progress with the Calspan one-foot,self-correcting tunnel. The conclusionsregarding transonic cross- flowcharacteristics of perforated walls are quite Interesting. The following is quotedfrom their paper. "The usualtheoretical approach is to assume thatthe normal velocity component inthe inviscid stream is linearly related to the velocity throughthe wall, which is linearly related to the pressure drop across thewall. Our results show thatneither linear relation is applicable and thatthe wall boundarylayer amplifies the normal velocity in the inviscidstream by a factorranging at least from 1.15 to 6. It does not appear to be feasible to calibrate this boundary layeramplifica- tion because the latter will depend, in part, on theupstream history ofthe boundarylayer. The mainadvantage tothe flowmeter technique is that it isnonintrusive anddoes notproduce disturbances in the flowfield. This one advantage is outweighed by thedisadvantages cited above. Consequently,the flowmeter technique for inferring the normal velocity componenthas been discarded, and we are now usingconventional pitch probes for this determination." Thus, this is another case which shows the linear boundary condition at ventilated, transonic walls is basically incorrect.

The basic technique used to correct wall porosity is as follows. First, theoretical estimates of the unconfined, longitudinal, disturbance-velocities, are made at a chosen distance from the tunnel wall. The wall porosities are initially set to provide these distributionsby monitoring the local static pressures with a long survey pipe. Second, the normal velocity components, at this same distance, are measured with small pitch probesand used as input for solutions of thc transonic, small-disturbance equation which assume unconfined flow. The resulting solutions provide new approximations for the longitudinal velocity distributions. The wall porosities and/or plenum pressures are then adjusted to provide this new velocity distribution.” Next, the normal components are again measured,and the process continuesuntil the differences between all the normal velocity components, measuredat successive iterations, are less than 0.0005 Vm. At this point, unconfined flow conditions are assumed to be achieved.

Experience with this iterative procedurehas shown that the convergence criterion is unnecessarily stringent, and a better criterion is being considered. However, for the case of an NACA 0012 airfoil at M = 0.55, a = 4O and 6O, and M = 0.725, a = 2O convergence was obtainedin five to seven iterations. The significant result was the measured airfoil pressure distribution, obtained in the one-foot tunnel with wall control, agreed very well with data obtained with the same airfoilin the 8-foot tunnel.

Although the Calspan results are encouraging, there arestill a large number of problems to overcome before three dimensional models becan similarly tested, i.e., adequate theoretical models of 30 transonic flows and porosity adjustment of all four walls.

* The perforated walls are dividedinto ten segments on the top and eight on the bottom. The four central segments in the top wall and the two central segments in the bottom wall are designed to provide an adjustable porosity with linear variation in the streamwise direction. Each segment has a separate plenum for individual control of suction or blowing.

168 Recently, Kemp (Ref. 25) has suggested a hybrid scheme. He hasproposed usinglimited adaptive-wall control to reduce interference to analytically

, correctablelevels. In sumnary, reductionof transonic-wall-interference and improveddata-correction methods are areas of active research in the USA, Canada, and Europe. Considerableprogress isanticipated in the near future.

111.6.3. BoundaryLayers and WallGenerated Noise

As notedby Pindzola, et al. (Ref. 161, slotted-walltunnels generate less noisethan do perforatedwalls. An illustrationof the importance of this phenomena hasbeen given by Cumning and Lowe (Ref. 11). Intheir tests, anF-111 aircraft model was tested in the same tunnel with bothporous and slotted walls. Near-interference-free data and minimum dragwere obtained over a Machnumber of 0.60 to 0.80 withthe slotted walls. With transition free, this corresponded to anobserved rearward movement of boundary layer transition compared to the porouswall tests. Thus, thisprovides a specific case ofwall-generated noise f affecting boundarylayer transition onan aircraft model.

In thesupersonic and hypersonic speed regimesPate (Ref. 27) and Dougherty (Ref.28) have developed correlationsto relate tunnel wall boundarylayer propertiesto radiated noise."" And inthe transonic range, the recent tests ofVidal, et al. (Ref. 24) and Starr (Ref.29) reaffirmthe essential role the boundary layerplays in determining wall crossflow characteristics. These tests, among others,have also demonstrated that model-induced pressure gradients can significantlyalter wall-boundary-layers in transonic tunnels. This means empty-tunnelboundary-layer surveys must be supplemented by takingadditional surveyswith models inplace (particularly for high lift configurations).In summary, wall boundarylayer surveys are a necessarypart of calibrating both transonic and supersonictunnels.

A The new National Transonic Facility at NASA Langley will have slotted walls becausethey generate less noise and interferenceat subsonic speeds,Ref. 26. Parker(Ref. 30) also found slottedwalls, asopposed toperforated walls, provided amore uniform centerline Machnumber distribution up to M - 1.1. ** Tunnelnoise measurements arediscussed in greater detail in Section I1I.F.

169 A review of various means for measuringboundary layer profiles has been given byKenner and Hopkins (Ref. 31). These investigatorsobtained boundary layer measurementson a supersonictunnel wail (2.4 < HOD,<3.4) with a single traversing probe,three different rakes, and a 12 deg.wedge with orifices in theleading edge. The interestedreader may consultthis reference for details of boundary layerprobe designs anda discussion of the results that can be expected.Also, Allen (Ref. 32)has givenageneral discussion of the effects of Machnumber on Pitot probe measurement errors in turbulent boundarylayers. 111.6. References

1. Goethert, E. H. : Transonic Wind Tunnel Testing, Pergamon Press, 1961.

2. Wright, R. H. and Ward, V. G.: "NACA Transonic Wind-Tunnel Test Sections," NACA Report 1231 , June 1955 (supercedesNACA RH L8J06, 1948). 3. Goadman, T.. R.: "The Porous Wall Wind Tunnel: Part I1 - Interference Effect on a Cylindrical Body in a Two;Dimensional Tunnel at Subsonic Speed," Rept. No. AD-594-A-3, 1950, "Part 111 - Reflection and Absorption of Shock Waves at Supersonic Speeds,I' Rept. No. AD-706-A-l , Nov. 1950, "Part IV - Subsonic interference Problems in a Circular Tunnel," Rept. No. AD-706-A-2, Aug. 1951, Cornel1 Aero. Lab., inc. 4. Baldwin, B. S.; Turner, J. 8.; and Knechtel, E. D.: "Wall Interference in Wind Tunnels with Slottedand Porous Boundaries at Subsonic Speeds," NACA TN 3176, May 1954. 5. Garner, H. C. ; Rogers, E. W. E. ; Acum, W. E. A. ; and Maskel 1, E. C. : "Sub- sonic Wind Tunnel Wall Corrections," AGARDograph 109, Oct. 1966. 6. Heyson, H. H.: "Rapid Estimation of Wind-Tunnel Corrections with Application to Wind-Tunnel and Model Design," NASA TN 0-6416, Sept. 1971. 7. Pindzola, M. and Chew, W. L.: "A Summary of Perforated Wall Wind Tunnel Studies at the Arnold Engineering Development Center,'' AEDC-TR-60-9, August 1960. 8. Keller, J. D.: "Numerical Calculation of Boundary-induced Interference in Slotted or Perforated Wind Tunnels Including Viscous Effects in Slots," NASA TN 0-687 1 , Aug. 1 972. 9. Lowe, W. H.: "Subsonic Crossflow Calibration of a 22.5 Percent Perforated Wall," HST-TR-355-3, General Dynamics, presented at39th STA Meeting, Mar. 1973. 10. Lo, C. F. and Oliver, R. H.: "Boundary Interference in a Rectangular Wind

Tunnel with Perforated \Jal 1s,I@ AEDC-TR-70-67, Apr 11 1970.

11. Cumming, D. P. and Lowe, W. H.: "Experimental Wall Interference Studies in a Transonic Wind Tunnel," AlAA Paper No. 71-292, Mar. 1971.

171 12. Blackwell, J. A., Jr. andPounds, G. A.: "Wind TunnelWall Interference Effects ona Supercritical Airfoil at Transonic Speeds,'' Paper No.1, Proc. AIM 9th Aerodynamic TestingConference, June 1976.

13. Page, W. A.: "ExperimentalStudy ofthe Equivalence of Transonic Flow aboutSlender Cone-Cylinders of Circular and Elliptic CrossSection," NACA TN 4233, Apri 1 1958.

14. Newman, P. A. and Klunker, E. B.: "NumericalModeling of Tunnel-Wall andBody-Shape Effects on TransonicFlow Over Finite Lifting Wings," Part 11 of Aerodynamic AnalysisRequiring AdvancedComputers, NASA SP- 347, Mar. 1975.

15. Mokry, M.; Peake, 0. J.; andBowker, A. J. : "Wall Interference on Two- DimensionalSupercritical Airfoils, Using Wall Pressure Measurements to Determinethe Porosity Factors for Tunnel Floor and Ceiling," NRC-13894, Nat'l. Aero.Estab., Ottawa, Canada,Feb. 1974.

16. Pindzola, M. ; Binion, T. W.; and Chevallier, J. P.: "Design ofTransonic WorkingSections," App. 8 of A Further Review of Current"Research Aimed at theDesign and Operation of Large WindTunnel-s., AGARD-AR-83, Sept. 1975.

17. Berndt, S. B. andSorensen, H.: "Flow Properties of SlottedWalls for TransonicTest Sections," Paper No. 17, WindtunnelDesign and Testing Techniques, AGARD-CP-174, Mar. 1976.

18. Barnwell, R. W. : "Improvements inthe Slotted-Wall Boundary Conditions,'' Paper No. 3, Proc. AlAA 9th Aerodynamic Test-ing-Conference, June 1976.

19. Ferri, A. and Baronti, P.: "A Method forTransonic Wind-Tunnel Corrections,'' AIM Jour., Jan. 973.

20. Sears, W. R. : "Se f Correcting Wind Tunnels," AeronauticalJour., Vol. 78, Feb.-Mar. 1974.

21.Weeks, T. M. : "Re duction of TransonicSlotted Wall Interference by Means of SlatContouring,'' AFFDL-TR-74-139, Mar. 1975.

22. Goodyer, M. J.: "The LowSpeed SelfStreamlining Windtunnel," Paper No. 13, WindtunnelDesign and TestingTechniques, AGARD-CP-174, Mar. 1976- 23 Chevallier, J. P. : "SoufflerieTranssonique A ParoisAuto-Adaptable," Paper No, 12, \Jindtunnel~Design and TestingTechniques, AGARD-CP-174, ~ ~ ~- Mar. 1976.

24.

25 - Kemp, W. B., Jr.: "Toward theCorrectable Interference Transonic Wind

Tunnel ,'I Proc. AlAA~~ 9th Aerodynamic-~ - "" Testing Conference, June 1976.

26. HcKinney, L. W. andHowell, R. R.: "The Characteristicsof the Planned NationalTransonic Facility,'' Proc. "- AlAA 9th Aerodynamic TestingConference, June 1976.

27. Pate, S. R.: "Measurementsand Correlations of Transition Reynolds Numbers onSharp Slender Cones atHigh Speeds," AlAA Jour.,Sept. 1971.

28. Dougherty, N. S.: "Correlation of TransitionReynolds Number with Aero- dynamicNoise Levels in a Wind Tunnel at MachNumbers 2.0 - 3.0," AlAA Jour., Dec. 1975. .

29. Starr, R. F.: "Experiments to Assess theInfluence of Changes inthe Tunnel WallBoundary Layer on Transonic Wall Crossflow Characteristics,'' WE TunnelDesign and Testing Techniques, AGARD-CP-174, Mar; 1976-

30 Parker, R. L., Jr.: "Fiow GenerationProperties of FiveTransonic Wind TunnelTest Section Wall Configurations,'' AEDC-TR-75-88, Aug. 1975.

31. Keener, E. R. and Hopkins, E. J.:"Accuracy of Pitot-Pressure Rakes for TurbulentBoundary-Layer Measurements in SupersonicFlow," NASA TN D-6229, Mar. 1971.

32. Allen, J. M.: "Effectsof MachNumber on Pitot-ProbeDisplacement in a

Turbulent BoundaryLayer ,I' NASA TN D-7466, June 1974 (also see AlAA Jour., PP. 949-590, July 1975).

173 1II.H. STANDARDMODELS

III.H.l. AGARD ForceModels

Theneed for standardmodels was recognizedearly by the AGARD Wind Tunnel andModel TestingPanel. In 1952, thispanel adopted AGARD Models A and B for thepurpose of building and testing calibration models in supersonictunnels, Ref. 1. It was thoughtthat this would be"extremely useful in establishing standards of comparisonbetween wind tunnels." It wouldalso be usefulin studyingthe effects of changes in Reynolds number, turbulence, model size and model fabricationtolerances.

AGARD Model A was an existingrocket body withfins which had been designed by NACA and had theprior designation of RH-IO. AGARD Model B is awing-body combinationwhich consists of an ogive-cylinder anda delta wing with a symmetrical, 4% circular-arcairfoil. In 1954, the AGARD Model B was modified to includevertical and horizontaltail surfaces. This configuration was designated AGARD Model C and was designed"primarily for testing and calibration in thetransonic speedrange.'' The purpose of the tail was to have a model which wouldbe more sensitive to flow curvature and wall reflections of shockand/or expansion waves.

The geometricalspecifications for the various AGARD models aregiven in References 1, 2 and 3. The associatedwind tunnel data is presented in Reference 4. Goethert(Ref. 5) alsodiscusses some ofthe early AEDC dataobtained for AGARD Models B and C. The followingconclusion was derivedfrom these early tests. Based oncomparisons of datafor models having 1.15% and 0.01% blockage inthe PWT 16T TransonicTunnel, it was concludedthat satisfactory results could beachieved intransonic tunnels if aircraft models didnot exceedabout 1% blockageratio. Responses tothe questionnaire indicate this rule of thumb has been adoptedalmost universally. However, forprecision testing, Goethert (Ref. 5) recommended blockageratios be keptas small as 0.5% and with awing span notexceeding half the tunnel width.

These earlyconclusions werebased on the results of testing force models. The more recenttesting of the past Few years hasemployed models designed to measure pressuredistributions. It is now known thataircraft models with 1% tunnelblockage can experience considerable wall interference, especially near

174 Machone, even in thebest ventilated tunnels. Thus, currentstudies of transonicwall interference require the use of pressuremodels to providethe necessary data.

lll.H.2. TransonicPressure Models: 2-D

Inthe past, anumber of airfoils have been used in studies of transonic wallinterference. What followsis a brief list of airfoils whichhave been employed recently.

Weeks (Ref. 6) hasused a symnetrical, 6% circular-arcairfoil to study wallinterference * in a contoured,slotted-wall tunnel. In France, an NACA 64 A010 airfoil has beenused at ONERA (Ref. 7) fortwo-dimensional studies of solid,adjustable walls in transonic tunnels. Whereas, Calspanstudies of walls withadjustable porosity have utilizedthe symmetrical NACA 0012 airfoil, Ref. 9. A 15.2 cm (6 in.) chord modelhas been testedin the Calspan 8-ft Tunnel to providebaseline force and pressuredata which are wall-interference-free. Also, this airfoil hasbeen found to be lesssensitive to Reynolds number and tunnel flow quality. Thus, Pindzola,et al. (Ref. 10) have recommended the NACA 0012 airfoil beadopted as a standard 2-D model inorder that transonic wall development work have a cmon basis.

I I l.H.3. TransonicPressure Models: 3-0

A 20deg cone-cylinder hasbeen used in anumber of transonic facilities (particularly for M > 1) toselect operational values of wall porosity, wall angle and plenumpumping. Davis and Graham (Ref. 11) havedescribed a typical casewhich illustrates this procedure. They have alsoreviewed the wall- interference-free,transonic data which is available for this model geometry.

At onetime, it was thoughtthat if thewall parameters were selected to give minimum interference on this model,through Mach one, this wouldbe satisfactory for testingall types of models.However, anumber of aircraft model tests have shown thisis not the case.For example, Davis' (Ref. 12) transonic testswith an AGARD Model B indicatedbetter agreement withthe AGARD reference datacould be obtained with different wall settings. Thus, care must beexercised in selecting a calibration model. Inpartic- ular, it is now recognizedthat a calibration model mustbe '%imilarll to models whichare to be tested.Unfortunately, precise criteria for how similar have notyet been defined. All that may be saidat this time is: more than one type ofstandard calibration model is necessary for valid testing of missile, airfoil, and aircraft models inexisting transonic tunnels.

The testsof Treon, etal. (Refs. 13 and 14) establishedthe need for identical models and instrumentation when comparing resultsfrom different tunnels.In this study, a 0.0226-scale model ofthe Lockheed C-5A was testedin theCalspan 8-ft., the NASA Ames 11-by 11-ft., and the AEDC 16-ft. (16T) transonicwind tunnels. The same combinationof mdel-support sting and internal forcebalance was used in each of thetunnels. This allowed analyses of smalldifferences in blockage,buoyancy and Reynolds number effectswhich would not have been possible if different models hadbeen used. Inaddition to forces and moments, seven orifices onthe fuselage were used to measure localstatic pressure.This enabled comparisons of buoyancyand model-induced changes in effectivefreestream Mach number. The resultingcorrections for relative buoyancy and effectiveReynolds numberk reducedthe spread inaxial force by 75 percent for Mach number belowthe drag rise value? Finally, these tests permitted estimatesof the "best expectancy agreement" between data obtained inthe three facilities.

The utility of the AEDC transition cone(Refs. 15-17) has been discussed inSection 1II.F. and will not be repeatedhere. However, it should be noted thatthis study of the effects of tunnel environment onboundary layer transition was also basedon thefundamental premise employed by Treon, etal. Namely, the same model, instrumentation, and support mechanism areessential for meaningful results.

A simplified,but versatile, aircraft modelhas been testedin the AEDC 16T and 4T transonictunnels by Binion(Ref. 8). Themodel consistsof twogeo- metricallysimilar, centerbodies with rectangular-planform wings. The center- bodieshave pointed, ogive-type noses and thewings have the NACA 63 A006 airfoil prof i le. The smal ler body servedas a tai 1 and was mountedon a separateforce * Seep. 150. A* Subsequent tothis work,Binion and Lo (Ref. 15) showed, in some cases,wall interference canovershadow theeffects of Reynolds number variations. balance and sting. Four different model arrangementswere tested in both tunnels, viz., the wing by itself, the tall by itself, and the wing with tail mounted close behind and at a more aft position.

After the force tests, the tests were repeatedand pressure distributions were measured on the centerbodiesand the wings. Angles of attack were repeated by duplicating the pressure difference across the model forebody which was initially calibrated as a functlon of a in the 16T. The conclusions reached by Binion include the following. 1. Flow angulari.tY can be induced into the tunnel flaw which is a function of model configuration, model attitude, and tunnel configuration. This flow angularity is distinct from theusual upwash correctionand varies nonlinearly with Mach numberand model incldence. No existing theoretical corrections can account forthis phenomena.

2. The movable tail feature confirmed the expected dependence of wall interference on model configuration in the transonic regime. Also, the more-aft tail position encountered wall-reflected disturbances at supersonic Mach numbers. 3. The attainment of an interference-free value oflift does not ensure an interference-free flow field. 4. There is no value of porosity, with the presentAEDC 4T walls, which will yield interference-free pressure distributions forthis aircraft model (0.9% blockage) when extensive regionsof supercritical flaw exists. The magnitude of wall interference appears tobe a function of size and extent of supersonic pocketsand the model-induced pressure gradient at the wall.

A transonic transport model has been designed and developed at ONERA and has been offered as a standard for transonic tunnel calibrations. A family of five different sizeshas been fabricated so that an appropriate size is available for even small tunnels. However, only the largest model provides for measurements of wing pressure distributlons. An equivalent body of revolution is also available for thelarge model. A description of the model geometry may be foundin Reference 19.

177

-. Two sizes of this model (large M5 and 1/4 scalesmaller H3) have recently been tested in the AEDC 16T and 4T tunnels and the NASA Ames 11-ft tunnel as part of a cooperativeprogram with ONERA. The statedpurposes of thisstudy were ''to provide an experimentaldata base for (1) the evaluation of theoretical or empiricalwall-interference correction factors and (2)the establishment of guidelinesto allow reasonable selection of wind-tunnel-to-model size ratios inthe transonic speed regime.'' The testresults and evaluationare reported by Binion (Ref.19).

An unexpected result of these tests was theobserved sensitivity to Reynolds number. Infact, the modelswere found to bemore sensitiveto Reynolds number when boundary layertransition was fixedthan with free transition. Also, greatervariation of the data from tunnel-to-tunnel occurred with fixed transition. Wing pressuredata from thelarger model showed thesedifferences were causedby differencesin shock-boundarylayer interactions and trailing edge separation.Finally, even thoughstate-of-the-art manufacturing tolerances wereused tofabricate, the models, there appears to be smalldifferences in tail incidence between the twomodels. Thisprecluded useful model-to-model comparisons ofpitching moment. In summary, the ONERA modelswere found to be overly sensitive to Reynolds number and tunnelflow quality and exhibited insufficient model similarityfor accurate studies of wall interference. Thus, the objectives of these tests were notachieved.

Based onexperience with the AEDC simplified aircraft modeland the ONERA models, the following criteria haveevolved for amodel tostudy transonic wall- interferenceproblems.*

1. The aircraft model shouldhave a smallcylindrical centerbody with an ogive nose (thecenterbody musthouse a forcebalance and provide a passage for surface pressure 1 ines). 2. Surfacepressures on thecenterbody should be selected and calibrated to directly measure Mach number and angle of attack. 3. The wingshould have an NACA 0012 airfoil,zero taper, and should be alignedwith the centerbody axis. A variable sweep featurewould be desirable in order to study the effect of lift onaxia1,interference gradients.

~ ~~ ~~~ ~ ~ * Binion, T. W., Jr., personalcommunication, AEDC, March 1977.

178 4. The horizontal tail should be separately ins.trumcnted andgeometrically similar to the wing. 5. Standardization of instrunentationand sting configuration is essential. 6. Both model forces and pressure distributionson wings and centerbody should be measured.

Work is continuing at AEDC to develop a model with these features. In sunwnary, a satlsfactory aircraftmodel for calibrating transonicwind tunnels does not yet exist. Until wall interference effects are clearly defined and separable from Reynolds number and flow quality effects,a simplified aircraft model is required. Once this objective is realized, more realistic aircraft models, e.g., the ONERA transport models, canbe utilfzed much more effectively for tunnel-to-tunnel comparisons.

179 1II.H. References

1. "AGARD Wind Tunnel Calibration Models,!' AGARD Specification 2, Sept. 1958. ..

2. Fail, R. and Garner, H. C.: "Calibratton Models for Dynamic Stabi 1 i ty Tests ,I' AGARD Report 563,: 1968.

3. Curry, W. H., ed.: "The FirstFifteen' Years ofthe Supersoni c Tunnel Association,''Sandia Laboratories, Sept. 1969.

4. Hills, R., ed.: "A.Review of Measurementson AGARD Calibration Model s ,'I AGARDograph 64, Nov. 1961.

5. Goethert, B. H.: Transonic WindTunne.1 Testing, Pergamon, 1961.

6. Weeks, T. M.: "Reduction of TransonicSlotted Wall Interference by Means ofSlat Contouring," AFFDL-TR-74-139, March 1975.

7. Chevallier, J. P. : "SoufflerieTranssonique a' ParoisAuto-Adaptables," Wind TunnelDesign and Testing~Techniques, AGARD-CP-174, March 1976.

8. Binion, T. W., Jr.: "An Investigationof Three-DimensionalWall Interference

in a Variable Porosity Transonic Wind Tunnel ,'I AEDC-TR-74-74, Oct. 1974.

9. Vidal, R. J.; Erickson, J. C., Jr.; and Catlin, P. A.: "Experiments with a Self-Correcting Wind Tunnel,'' Wind TunnelDesign and TestingTechniques, AGARD-CP- 174, March 1976.

10. Pindzola, M.; Binion, T. W., Jr.; and Chevallier, J. P. : "Design of Tran- sonicWorking Sections," A Further Review of Current Research Aimed at theDesign and Operation of Large Wind Tunnels, AGARD-AR-83, Sept. 1975

11. Davis,J. W. andGraham, R. F.: "Wind-TunnelWall InterferenceEffects for 20° Cone-Cylinders," AlAA Jour.Spacecraft 5 Rockets, Oct. 1973 (a 1 so see NASA TN 0-7432).

12. Davis, J. W.: "AGARD Model B TransonicBlockage Investigation," NASA Marchall Space FlightCenter, Huntsville, Ala., presented at 39th Semi- Annual STA Meeting, Mar. 1973 (referencedwith author's permission).

13. Treon, S. L.; Steinle, F. W.; Hofstetter, W. R.; and Hagerman, J. R.: "Data Correlationfrom Investigations of aHigh-Subsonic Speed Transport Aircraft Model in ThreeMajor Transonic Wind Tunnels," AlAA Paper No. 69-794, July 1969. 180 14. Treon, S. L. ; Steinle, F. W. ; Hagerman, J. R. ; Black, J. A. ; and Buffington, R. J.: "Further Correlation of Data From Investigations of a High-Subsonic- Speed Transport Aircraft Model in Three Major Transonic Wind Tunnels," AIM Paper No. 71-291, March 1971. 15. Binion, T. W., Jr. and Lo, C. F.: "Appl fcation of Wall Corrections to Transonic Wind Tunnel Data," AlAA Paper No. 72-1009, Sept. 1972.

16. Dougherty, N. S., Jr. and Steinle, F. W.: "Transition Reynolds Number

Ccnnpari sons in Several Ma Jor Transonic Tunnels,I' AlAA Paper No. 74-627, July 1974. 17. Dougherty, N. S., Jr.: "Prepared Comment on Cone Transition Reynolds Number Data Correlation Study,'' AGARD-CP-187, June 1975. 18. Whitfield, J. and Dougherty, N. S., Jr.: "A Survey of Transition Reynolds Number Work at AEDC," to be presented as Paper No. 25 at AGARD Fluid Dynamics Panel Symposium on Laminar-Turbulent Transition, Copenhagen, Denmark, May 2-4, 1977. 19. Binion, T. W., Jr.: "Tests of the ONERA Calibration Models in Three Transonic Wind Tunnels," AEDC-TR-76-133, Novo 1976-

181 111.1. OPTICAL METHODS

111.1.1. SupersonicTunnels

The use of schlieren and shadowgraph flow-visualization-methods to detect unwantedshocks inan empty test-section* is well known (Ref. 1) and, infact, may be designated a classicaltechnique . Obviously,the observance of a shock inthe empty-tunnelindicates corrective action is necessary. These means of flow visualization are also helpful in assessing the performance af probesand rakesand theirinteraction with nearby boundaries. High quality pictures also enableflow separation on probes to be observed and thereby provide additional datato guide improved designs.

A third,classical method forflow visualization is the Mach-Zehnder interferometer. However, theseinstruments are seldom used forwind tunnel calibration because of their cost and hypersensitivity to vibration and align- ment errors.

Detaileddiscussions of these methods may befound in a ratherlarge number ofreferences. References 2 thru 4 arerepresentative of both older and newer literature which deals with these three methods of flow visualization. 111.1.2. TransonicTunnels

As previouslydisucssed in Section III.D., movement of a transonicshock ona staticpressure probe is strongly affected byblockage and wallcharacter- istics. For example, theschlieren photographs of Page (Ref. 5) arequite instructive as tothe effects ona probecaused by varyingtunnel blockage from 0.25% to 0.005%.

In thecase of supercriticalflow about a hemisphere-cylinderprobe, the shadowgraphs of Hsieh(Ref. 6) werevery helpful in detecting boundarylayer separation and interpretingthe measuredpressure distributions.

* Most ofthe respondents to the questionnaire indicated they routinely used one or both of thesetechniques.

182 At lowsupersonic speeds, schlierens andshadowgraphs arevery useful in studiesof the shock-cancellation properties of ventilated walls, e.g., Ref. 7. Also,Dougherty, etal. (Ref. 8) havevery effectively used schlieren photographs tostudy the sound fieldgenerated by perforatedwalls when ex- posed tohigh-subsonic flows.

111.1.3. Newer Methods

Newer optical methods forflow visualization include laser Doppler velocimeters (LDV) , holographicvelocimeters (HV) , and holographicinterferometry (HI) fordensity measurements. The primaryadvantage of LDV'sand HV's is their potential to measurethree-dimensional flow fields without disturbing the flow. Aswe have seen, thisis particularly important in transonic tunnels near Mach one.

The currentstate-of-the-art of LDV'sand theirapplication to tunnel calibrationis reviewed in Appendix II. Sincethe accuracy of current LDV systems isapproximately 0.4 - 0.5%, theyare not yet superior to conventional probeswhich provide comparable accuracies of 0.1%.

The fundamentals ofholographic velocimetry are reviewed by Shofner, et al. (Ref. 9). A verycomprehensive review ofthe use of holography in windtunnel testing hasbeen compiledby Havener (Ref. IO). Progress(up to 1975) inauto- mating HI datareduction is reportedby Hannah andHavener (Ref.11). Since this is still a developingtechnology, applications of holography to empty-tunnel calibration appears to be inthe future.

Sparksand Ezekiel (Ref. 12) have recentlydemonstrated the usefulness of LaserStreak Velocimetry (LSV) for quantitative measurements of low-speed velocity fields nearmodels. This technique has theadvantage ofproviding, simultaneously, velocities on a planeas opposed tothe point-by-point measurements requiredwith LDV's. However, theaccuracy of LSV's iscurrently insufficient for empty-tunnel calibrations.Finally, Sedney, etal. (Ref. 13) have given a reviewof flow tra- certechniques and theirapplications in supersonic-flowfield diagnostics.

183 111.1. References

1. pope,A.and Goin, K. L.: High-speed Wind TunnelTesting., Wiley, 1965.

2. Ladenburg, R. W.; Lewis, B.; Pease, R. N.; and Taylor, H. S. (editors): Physical Measurements in Gas Dynamicand Combustion, PrincetonUniversity Press, 1954.

3. Vasil'er, A. L.: Schlieren Methods, Transl.by A. Baruch, Israel Program forScientific Translations, Jerusalem, New York andLondon, 1971.

4. Merzkirch, \J.: Flow Visualization, Academic Press, 1974.

5. Page, W. A.: "ExperimentalStudy of theEquivalence of Transonic Flow AboutSlender Cone-Cylinders of Circular and Elliptic CrossSection,'' NACA TN 4233,Apr i 1 1958.

6. Hsieh, T.: "Hemisphere-Cylinder in TransonicFlow, Mm = 0.71.0," AlAA Jour.,Oct. 1975 (also AlAA Paper No. 75-83, Jan. 1975).

7. Goethert, B. H.: Transonic Wind TunnelTesting, Pergamon, 1961.

8. Dougherty, N. S., Jr.; Anderson, C. F.; and Parker, R. L., Jr.: "An ExperimentalInvestigation of Techniques to SuppressEdgetones From Perforated Wind Tunnel Mal Is," AEDC-TR-75-88, Aug. 1975.

9. Shofner, F. M.; Menzel, R.; and Russell, T. G.: "Fundamentals of Holographic Velocimetry," AFFDL-TR-68-140, Nov. 1968.

10. Havener, A. G.: "A UsersGuild on PulseLaser Holography for Wind Tunnel Testing," ARL TR 75-0213,June 1975.

11. Hannah, B. W. and Havener, A. G.: "Applicationsof Automated Holographic Interferometry,"Int'l Congress on Instrumentationin AerospaceSimulation Faci 1 ities, I EEE Pub. 75 CHO 993-6 AES , Sept. 1975.

12. Sparks, G. W., Jr. and Ezekiel, S.: "LaserStreak Velocimeter for Two- DimensionalFlows in Gases," AlAA Jour.,Vol. 15, No. 1, Jan. 1977.

13. Sedney, R.; Kitchens, C. W., Jr.; and Bush, C.:C. "The Marriage of Optical, Tracer and SurfaceIndicator Techniques in Flow Visualization," BRL R-1763, USA Ball istic ResearchLaboratories, Feb. 1975.

184 1II.J. Humidity Measurements

The effects of moisturecondensation and the necessity for air drying have been discussed inSection Il.C.7. Measurement and monitor ing of the moisturecontent of the tunnel flow 1s therefore an essential part of tunnel calibration and operation.

The moisturecontent of a gas is expressed in a number of ways; relative humidity(ratio of moisture partial pressure to saturation pressure), dew point (or frostpoint) temperature at atmospheric pressure, specific humidity

(MSS of waterper mass of dry gas), and volume ratio (parts of water vapor permillion parts of air). The dew pointor ice-point at atmospheric pressure is the most commonly usedform ofexpression for wind tunnel operations.

A number of measurement systems are utilized by facilities responding tothe questionnaire, extending from the visual observation of fog in the tunnelflow to completelyautomatic, continuously recording dew point system5.

All dew point measurement instruments can be calssifiedaccording to the bas ic principles used.

One of the more basicnon-continuous dew point measurement instruments operateson the principle of allowing a hand-pump pressurized sample of gas, at known pressure and temperature, to expand to room temperature(Ref. 1). If theexpansion reduces the sample temperature to or belowthe dew point temperatures,fog is created which may beobserved visuallythrough a viewing window.

A trial and errorprocedure is required to determine the initial samplepres- surewhich will expand tocreate a just-visiblefog- Since a known relationship exists between pressure and temperatureratios, the dew point canbe determinedusing ambient as the final pressure. These instrumentsare low in cost,reasonably accurate and are widely used both as primary dew point monitors

185 and for monitoringthe accuracy of lessbasic instruments. They provideonly

periodic or spot checks and are therefore not satlsfactory for facilities

where relatively sudd,en changes in dew point canoccur. A1 1 readings must

be manuallyrecorded. Heasurements below about -4OC are difficult to make.

Continuousindicating and recordinghumidity sensors include the Dunmre

typewhich changes resistance in a non-linear fashion with relative humidity.

A modifiedform respdnds to dew po'int(Ref.' 2). Each sensor has a1 imited

range, so thatseveral are required if thehumidlty range is large. The range

extends downward to about-40 C.

An electrolytichumidity sensor is also available. These sensors utilize

an elementwhich electrolyzes water into hydrogen andoxygen, causing an

electrolysiscurrent to flow. The electrolyticinstrument is usually calibrated

in parts per million, with fullscale ranges aslow as 0 to 100 parts per million.

Thisinstrument, like the resistance device, canbe configuredto both indicate

humidity and provide an electricalsignal for an externalrecording device.

Dew point temperaturescan be determined by controlling automatically

the temperature of a polished metal mirror to the point that a trace film of

condensation(or frost) is maintained. Several instruments based on this

principle have been developed. More recenttypes are simplified in that the

therm-electric cooling effect is used to chill the mirror (with auxiliary

refrigeration If necessary). The condensation film isautomatically maintained byfeedback control of the mirror temperature, utilizing an optical source

reflectry light from the cooled mirror toa pair of photo-detectors forming a bridge circuit. The dew pointtemperature is measuredby athermocouple or

186 resistance-temperaturedetector attached directly to the mirror. The dew point is indicated bya meter orother Indicator, and thetemperature sensor output may also berecorded, supplied to the wind tunnel data system, etc.

The rangecan extend to aslow as 200 OK (-100 OF).

A continuous-recording, dew point monitor has obviousadvantages both withregard to monitoring and controlof-tunnel ineasurements.They can also provide Information onthe performance of dryersand.other tunnel equipment. I I I. J.References

1. Pope, A. ; and Goin, K. L.: High SpeedWind TunnelTesTing., pp. 223-226,

John Wiley and Sons, New York.

2. Doeblin, E. 0.: MeasurementSystems; Application and Design, pp. 596-'598,

McGraw-Hi 11 , New York.

3. Fraade, I). J.: "Measuring Moisturein Gases," Instruments and Control

Systems, Apri 1 1963.

188 I

IV. ERROR AND UNCERTAINTY IN CALIBRATION MEASUREMENTS

Treatment of accuracyand sources oferror in experimental data involves

principlesof statistics and probability.Unfortunately, eachbranch of

sciencetends to developspecialized terminology, which impedes understanding

and communication in comparing measurement results. An attempt will there-

fore be made to define and recmendbasic terminology which may be used to

advantage inevaluating, defining and comnunicatingcalibration accuracy.

As a first step, a definition and classification of various types of errors will be stated.

1V.A. Random Error

Errors may be classifiedin two generalcategories: random and fixed.

Random erroris frequently referred to by theengineer in less precise terms

as"scatter," "noise," etc., allimplying that repeated measurements do not

yieldthe same value.Most processes are such that if a sufficientlylarqe

number of measurements are made and thefrequency with which each value is

measured isplotted against the measured value,the resulting plot (the proba-

bility density function) will approachthe familiar bell-shaped normal distri-

butioncurve. In this case, the arithmetic mean value,or average,

N xi - (4.A i=l N .1)

occursat the peak ofthe curve. When plottedin normalized form, the area

underthe curve isunity. The precision,which is a measure ofthe scatter

or random error,is specified bythe standard deviation,

(4.A.2) a= N-1 i=l 189 if,the distribution is based on a sufficiently large number of measurements, 68 percentof the measurements will lie within the range -+1 u , 95.4 percent within -+2 u and 99.7 percent within 23 u . A wide!. flat distribution there- forecorresponds to measurements with a largestandard deviation, a large amount of scatter, a large random error, or a lack of precision, all of whichrefer to the same characteristicsof the measurement.The random error

isquantatively stated in terms ofthe standard deviation and errorstatements shouldalways be specified as 1 u , 2 u , etc.

1V.B. Fixed Error

A second formof measurement error is referred to assystematic error, fixed error or bias.This component of error will be the same in each of many repeated measurements. The magnitude and signof the bias may not be known a priorisince thesecan be determinedonly by comparison with the true value of the measured quantity. As one example, an undetected change inthe calibration of an instrument suchas a pressuretransducer will introduce a fixed bias of unknown magnitude and sign. Upon detection,this bias or fixederror can beremoved by recalibration.Since unknown fixederrors are not correctable, unless detected, theirminimization depends (1) upon careful mon toringof results, (2) routine pre-andpost-test calibrations of instruments, n place, (3) end-to-end calibrations of instrumentsprior to and duringtests, etc.This same philosophycan be appliedto a basicinstrument suchas a presS uretransducer or to the tunnel- flowcalibration. The objectiveshould be toeliminate all large, unknown fixederrors.

Some typesof unknown fixed errors cannot be readilyeliminated by cal i bration. An example might be the drag of a standardmodel, where no "true" valueis known. Facility-to-facility comparisonsallow only an estimateof the probable maximum magnitude ofthe bias. Correction may bepossible only to the extentthat the comparison tests allow determination of and correction for the cause(or causes) of the bias (or a portionthereof).

The fixed error limit, whichnormally must be estimated,is the upper limit on the fixed error or bias, and may be symmetrical or non-symmetrical, i.e., it may be 0, + or 0, - rather than -+.

1V.C. Uncertainty

The totaluncertainty interval for ameasurement representsthe largest, reasonably-expectederror (i.e., the true value should fall in the uncertainty interval) and is a combination ofthe precision (standard deviation) and the estimatedbias.

A method described by Abernathy, etal. (Ref. 1) and recornendedby the

National Bureau of Standardsexpresses the uncertainty as the range centered aboutthe mean value and defined as

Up+- (B+tg5u) (4.C.l)

Where U isthe uncertainty, B thebias or fixederror 1 imit, and t isthe 95 95thpercentile point for the Student "t" distribution. The valueof t depends onthe number ofvalues used in computing u ; for a large number of measurements theStudent t distributionis identical to the normal distribution. The use ofthe t factorincreases the uncertainty limit when small 95 samples are used tocalculate (or, more accurately,to estimate) d . Abernathy, etal. recommended that a valueof be used for t for 31 or moresamples 2.0 95 (compared to 1.96 for an infinite number).Reference (1) and most statistics

texts (e.g., Ref. 2) containtables of Student's t distributionsfrom which tg5 can be obtainedfor less than 30 samples. Statistical methods employing the t distribution are frequently called small-sample methods forobvious

191 reasons.Although this example Issimplified, it canbe extended toInclude othererror terms.

An additional problem in accurate determination of measurement error, not discussedabove, is that the measured propertiesnormally have small- amplitudevariations with time. In addition to obtaining a statistically adequate number of samples, the sample interval must span at least one complete cycle of thelowest-frequency component of tunnelunsteadiness, as discussed by Huhlstein and Coe, (Ref. 3).

1V.D. ErrorPropogation

In essentially all cases, calibrationparameters are determined from basicproperties which are measured and a known function relating the measured quantities and thedesired parameter. An obviousexample would be thedetermina- tion of Machnumber inthe test section from measured pressures. Random error sourceswould include the precision (standard deviations) of the pressure measurements. Staticpressure probe uncertainty limits may be estimated as a fixedbias in the absence of a calibration.Another fixed bias could be theestimated uncertainty in y . As an illustrative example, the random error in Mach number can be calculatedfrom

(4.D.1)

where thevariations in H and P aretaken to be uncorrelated. The fixed S error or bias limit can similarly be calculated from

(4.D .2)

192 where and Bp arethe estfmated uncertainty 1 imitsfor the ratio of Y specificheats and forthe static probe error, respectively.

The results can be combined accordingto Eq. (4.C.1) todetermine the

totaluncertainty interval for a specific(point) Mach number measurement.

(4.D.3).

The uncertaintyinterval of an individualproperty measurement, suchas ,. a pressure,can also be estimated as above;where individualerror sources such as thestandard deviation of thetransducer, the excitation power supply,the

instrumentationamplifier and theanalog-to-digital converter are all taken

into account.Normally, however, the calibration is performed end-to-end utilizing all components so that all of the above factorsare taken into accountand attributedto the pressure transducer.

193 IV. Ref erences

1. Abernathy, R. B., etal and Thompson, J. W., Jr.: "Handbook; Uncertainty

in Gas Turbine Heasuremcnts," AEDC-TR-73-5, February 1973 (also CPlA No. 180).

2. bel, P. G.: Introductionto Hathematical Statistics, pp402-403, John

Wiley and Sons, Inc., New York, 1966.

3. Huhlstein, L. Jr., and Coe, C. F.: "Integration TimeRequired tG Extract

Accurate Static and Dynamic Data From Transonic Wind TunnelTests," AlAA

Paper 75-142, Pasadena, Cal if., Jan. 1975.

194 V. CONCLUSIONS AND RECOMMENDATIONS

V.A. Summary ofState-of-the-Art of Transonic and Supersonic Wind Tunnel Calibration

Reference hasbeen made throughoutthe previous sections of this report to information ,obtained from thequestionnaire survey, literature search and, personalcontacts. A condensat.ion of thisinformation hasbeen presented where appropriate.Primarily, information hasbeen summarized in anattempt

to define "best state-of-the-art" calibration accuracy. , Attention has been

focusedon the primary problems which were considered to bemeasurements of , stagnation and staticpressure and caiculation of Mach number. A concluding sumnary ofthe questionnaire results is presented here for convenience.

Based ona judgment evaluationof data reported in the questionnaires, thebest, current, pressure-measurementaccuracy (on the basis of standard deviation)ranges from 0.025 to 0.10 percent. These accuracies were reported forboth blowdownand continuoustunnels.

The survey showed thatapproximately twice as many transonictunnels use plenumchamber pressurefor a referenceto monitor Mach number, asopposed to test-section wall pressure. However, bothtypes of measurements are used and bothrequire a calibration(s)to relate the associated data to static pressure measurements along the centerline.

The most popularstatic-pressure-probe is a 10 deg cone-cylinderwith static orifices located ten or more cylinder diameters downstream of the shoulder. The ten-deg-coneappears to bea trade-off between therequirements to minimize disturbanceof the flow and, simultaneously, be easily fabricated and durable enough to beused repeatedlyin a windtunnel environment. Although athigh subsonic speedsa shock forms on the cylinder and accurate measurement of the static pressure requires orifices at severalstations, only afew of theprobe designssubmitted with the questionnaires have thisfeature. A smaller-angle cone not only hasa lower,shock-attachment Mach number but it alsogenerates a

195 weaker transonicshock on the cylinder and thussmaller deviations from free- streamconditions. Of thevarious static probe designs described in response tothe questionnaire, a two-degree (total-includedangle) cone was thesmallest.

An additionalsource of error in calibrating transonic tunnels is the neglectof variations transverse to the flow. Almost without exception, in caseswhere measurements had been made, t,he questionnairesindicated greater Mach number gradientsoccur across the flow than along the tunnel centerline. This may be most significant.i.n the determination of drag divergence and/or buffetonset for transonic aircraft models. However, thepresent state-of-the- art of wind'-tunneltesting is to use off-centerlinedata exclusively as a diag- nostictool to detect unacceptably large variations. In which case nozzle and/or test section configurations are altered.

Themost popularflow-angularity-probes appear to be the 30-deg-cone for simultaneousmeasurements of pitch and yaw.Wedges ofvarious angles are often used for planar measurements. It appearsfeasible to design probes of this type (i.e., differential-pressure) which can resolve flow angles to -+0.01 degree. (This objective was proposed in 1970 bythe ad hoc Air Force-NASA Committeeon TransonicTesting Techniques.) The quotedaccuracy for flowangle measurements ranged from 0.01 deg to 0.04 deg. A spatialvariation of +1/4 deg was frequently mentioned.

Quotedstagnat ion-temperature accuracy usual ly ranged from 1 to 2 OC.

The majority of reporting facilities do notcontinuously monitor humidity. In order to achieve a Machnumber accuracyof 0.001, humidity must be monitored continuously.

Nearly 50% of the tunnels have made noise measurements in either the stilling chamber, thetes section and/or the plenum chamber. In mostcases, either miniature strain gauge transducers or condensermicrophones were used to measure thenoise data. The followingtechniques have beenemployed to measure freestreamdisturbances in transonic and/or supersonlc wind tunnels.

1. High-frequency-responsepressure transducers mounted near the tip of cones to measure fluctuating static pressures beneath a laminarboundary layer.

II I II 2. Pressuretransducers mounted on wedges withthe measurement surface aligned with the flow. 3. Pressuretransducers mounted onthe cylindrical portion of ogive-cylinders. . 4. Pressuretransducers mounted inPitot probes to measure fluctuationsin Pitot pressure. 5. Hot-wire and hot-f i lm measurements.

Approximately 25% ofthe tunnels reported having made hot-wireor hot- film measurements ofturbulence. However, inthe majority of cases onlyvery limited centerline and/or wall boundarylayer measurements have been made. Only afew' tunnelsreported measurements of fluctuating Pitot pressure.

The majority of the tunnels reported that Schlieren systemswere of valuein detecting unwanted disturbancesin the test section. When combined withhigh speed photography, this method alsoprovides data on flowunsteadiness.

Most ofthe responding facilities have used one or more standardforce models duringcalibration. However, comparisons withreference data are usually only qualitative and are of limited use inpinpointing undesirable flow characteristics.

The well known rule of thumb thatthe model crosssection should not exceed 1% ofthe tunnel area for transonic testing appears to be universally accepted. A consequence ofthis criterion is that very fewtunnel operators attemptto correct for wall interference. This also reflects the lack of an acceptedtheory for correcting for transonic wall interference.

Finally,the consensuson frequency of windtunnel calibration is that a tunnelshould be recalibratedor at least spot checkedwhenever:

1. tunnelconfiguration changesoccur,

2. significantinstrumentation modifications are made,

3. erroneousdata is beingobtained, or

4. inthe absence of any ofthe above,once each year.

Static-pressureorifices should also be inspectedbefore recalibrating.

197 V.B. TRANSONIC TUNNELS

The goal of Machnumber calibrations in transonictunnels should be to achievean accuracy of +O.OOl, particularly in the transonic drag rise regime: 0.75.- < M < 1.0. Foran air test medium, thisrequires the following constraint " ontotal errors in total and stat ic pressure measurements.

However,changes in Reynolds number (Section 11.8.2) or humiditylevel (Section Il.C.7) can easily cause Machnumber variationsseveral times larger than 0.001. For this reason,considerable care should betaken to calibratean empty-tunnel overthe entire range of Reynolds numberand humiditylevels normally encountered duringroutine operations. Once thetunnel is calibrated for typical,operating humiditylevels, a continuousmonitoring of humidity is preferred, e.g.,a strip recorderfor subsequentreference. Inaddition, excessive spatial variations of total pressure (i.e. , AHS/HS > 0.001) across the sti 11ing chamber may require correctiveaction, e.g., additionalscreens, honeycombs, etc.Finally, the assumption of an isentropicexpansion from the stilling chamber to the test sectionshould be evaluated by direct measurements inthe test section, both on and off-center1 ine.

The long,static-pressure, survey pipe iswell established as thestandard forobtaining centerline measurements. Bestresults are achieved with the nose ofthe pipe located well upstream in the settling chamber; this is necessary in orderto prevent passage of a transonic shockover the length of the pipe. In all cases,the resulting data should be carefullyinspected for orifice-induced errors. Once thecenterline data i'sdetermined to befree of orifice errors (SectionIII.D.l.), standard procedure is to use the static pressure data to calibrateeither plenum chamber pressure or wallstatic taps. If plenumpres- sure is used for Machnumber control, the possibi 1 ity of departures fromempty- tunnelcalibrations should be carefully examined in cases of (1) large (> 1%) lifting models athigh-subsonic speeds and (2) rapidlyvarying Machnumber conditions suchas occur during rapid changes in model orientation. If wall statictaps are used for Machnumber control,at least oneshould be located on

198 each wall, ahead ofthe model location, and averaged with a "piezometerring"; this average ispreferable to using only a singlewall static pressure.

Since a long pipe is difficult to move aboutthe test section and off- centerlinedata is important for aircraft-model testing, it is recommended thatsupplementary, off-centerline data be obtainedwith a conventionalstatic pressureprobe or rake of suchprobes, when M is not near one (Sec. I I I .D. 2). Sincethe wing span of most transonic aircraft models is restricted to 60 percent,or less, of the tunnel width, both Mach number and flowangularity data shouldbe obtained over this span in the vertical and horizontalplanes.

Off-centerline measurements of flow angularity are sensitive indicators oferrors causedby nozzlecontour, wall settings, seal leaks, etc. The most accurate measurements of flow angularity can be obtained in the least amount of timewith a probeconsisting of two,orthogonal, symmetrical wings and a force balance housed in a smallcenterbody."

In addition to measurinq Machnumber and flow angularity off-centerline, datashould be taken atrepresentative forward, center, and aftstations in theuseable test section. It is suggested thatthe resulting data, at a givenstation, be expressed in terms ofstandard deviation from the mean of thecenterline measurements. Thistype of data will provide more complete informationon flow quality and should be considered when selectingwall porosity, wall angle or amount of plenumevacuation.

Unfortunately, no generalconstraints exist as to what areacceptable off-centerlinevariations. Jackson (AEDC) has suggestedthe following criteria for ''good" un iformi ty in center1 ine Machnumber :

2 aM 5 0.005 for M < 1 ,

2 UM 20.01 for M > 1

Inthe past, a criterion for acceptable flow angularity along the centerline has not beenneeded because a given model isusually run upright and inverted inorder to establish the effective angle of incidence. This is a valid and * Acceptableaccuracy can also be obtained with conventional, differential- pressureprobes, see Section 1II.E.

199 well-establishedtesting practice; however, flowangularity in the yaw plane isfrequently ignored. To summarize, standard criteriafor flow uniformity need to be developed for representative(preferably standard) models in variouskinds of tests, =.g., force,buffet, flutter, etc. These criteria shouldinclude standards for acceptablevariations in Machnumber and flow angularity,both on and off-centerline.

Unsteadydisturbance measurements intransonic tunnels should be a standardpart of tunnelcalibration. Recentuse ofhot-wires in the test sectionof a transonictunnel at NASA Ames indicatesthese may be usefulfor unsteady-flowcalibration in at least some tunnels. However, asdynamic pres- sureincreases hot-wires become more vulnerable to breakage and probably will be impracticalfor use in the new high-Reynolds-number faci 1 ities(except in thesettling chamber). Based on extensiveexperience with the AEDC transition cone in twenty-onemajor windtunnels of the United States andWestern Europe, thisdevice has become an "unofficial"tunnel acoustic calibration modeland is currentlyconsidered to be thebest available disturbance calibration instrument However, thereis a definite need for a lessexpensive instrument which canbe easilyreproduced and used in all sizesof facilities. The development offluctu- atingPitot probes appears to meet this need (Section I1I.F). Thistype of instrumentcan be used to measure centerline noise and to calibratewall-mounted, dynamic pressuretransducers. A walltransducer(s) can then beused as a permanent monitorof tunnel noise. The walltransducer(s) should bemounted approximately 0.025 cm (0.01 in.)below the plane of the tunnel wall and, preferably, shouldhave a frequencyresponse out to 30 kHz. By using two or more wall- mounted transducers,the direction of propagation of disturbancescan be ascer- ta ined .

Althoughthere are no generalcriteria for acceptable levels of flow unsteadiness, Mabey (RAE) has developed some noiseguidelines for interference- freeflutter and buffettests (Section I I1.F.). The goalof transonic tunnel, noise-reductionresearch is to reduce unsteadiness to levels characteristic of turbulent boundarylayers on solid walls, i.e., AC < 0.005. P An accepted measure of wave-cancellation characteristics of ventilated walls is to obtain pressure distribution data on a 20 deg cone-cylinder and com- parethe results with wall-interference-free data. It is now apparentthat the

200 traditionalassumption of a linear boundary condition at ventilated, transonic wallsis erroneous. Thus, thisexplains the failure of pastattempts to theo- reticallycalculate the effects of wallinterference on madel testing. Measure- ments oftest-section-wall boundarylayers, both with and without models in situ, arebeing made in efforts to gain a betterunderstanding of ventilated walls and theircross-flow characteristics. Current research on transonicwall inter- ferenceis focusing on three areas: (I) thederivation of more exactboundary conditions, (2) the development of a self-correctingwind tunnel wlth automatic control of longitudinallyvarying ventilation, and (3) varyingwall contours to attain wall-interference-free flow aboutmodels.

The NACA 0012 airfoil is currently being usedas a standardpressure model in two-dimensionaltests. The ONERA transportaircraft modelshave been tested in anumber oftransonic tunnels and have been found to be extremelysensitive to flowquality. AGARD has notyet adopted a standard,transport-aircraft model. There is currently a genuine need for a standard,aircraft, pressure model (or models) toaid transonic wind tunnel calibration and datacomparisons between tunnel s.

LaserDoppler Velocimeters are an importantaddition to the tools availab le forwind tunnel calibration. The obviousadvantage of an LDV is it does not perturbthe flow. At Mach numbers nearone, this is an importantadvantage. In addition,an LDV canbe used not only for mean-flow velocity measurements but alsocan measure flow angularity and turbulenceintensities greater than about one. LDV measurements of velocity and flowangularity are currently only 1/4 to 115 as accurate as thebest conventional probes at high subsonic and low supersonic speeds. However, near Mach onean LDV is expected toprovide superior data. At thistime, the LDV isnot being used toroutinely calibrate empty- tunnels. However, we anticipate suchuse inthe future.

201

I V. C. SUPERSONIC TUNNELS

Mach number insupersonic tunnels should be calibrated bymeasuring two independentpressures inthe test section. Accurate results can be obtained bymeasuring Pitot pressures in the freestream andbehind the bow shock of a wedge. However, freestreamPitot and staticpressures are preferable to only Pitot data and theassumption of isentropic flow from the settling chamber.

Sincetransverse gradients in Mach number are typically larger than axial variations, it is consideredessential to calibrate both Mach number and flow angularityoff-centerline. At leastthree cross-sections should be surveyed nearthe forward, center and aftportions of the useable test section. This type of data can be obtained most easily with Pitot and static proberakes or arrays mounted on a traversingsting. This data should be used tocalibrate a permanent Machnumber probewhich should be installedin supersonic tunnels forfrequent, routine checks on calibration. :ngeneral, a calibrated Mach number accuracy of 0.5 to 1% isconsidered good. Industrystandards need to be developedwhich define acceptable flow qual ity for particular kinds of testing, e.g., force and pressuretests of missiles and aircraft models.

As intransonic tunnels, centerline noise measurements shouldbe obtained and used to calibrate one or more, wail-mounted,dynamic-pressure transducers. The AEDC transition cone is currently the only flow disturbance calibration devicewhich has been testedin a large number oftunnels. A smaller and less expensivenoise calibration device is needed toserve asa standard.Probes designed to measure fluctuatingPitot pressures should be consideredfor this purpose. TraditionalPitot surveys of tunnel-wal 1 boundarylayers not only establishthe size and geometry of the inviscid f low but also aid correlations of faci 1 i ty noise.

Inaddition to keeping the total temperature high enough toavoid lique- faction of the test gas, the effects of typical levels of water vapor in the test gasshould be carefullycalibrated. As iswell known, theprimary effect of watercondensation is a loss oftotal pressure anda risein static pressure. Also, variousoperators haveobserved thatpressure tests are more sensitive to humiditylevels than force tests.

The number of surveys of supersonic flow fields with a iaserDoppler veio- cimeteris increasing as some earlier problemshave been resolved.In the future this new tool may enabie more accurate calibrations of supersonictunnels.

202 APPENDIX I * Hot Wires and Hot Films

Introduction ' A hot-wire anemometer is a means of measuring fluctuations in localized areas of the flow at frequencies up to 200 KHz. The sensor may be a smalldiameter wire suspended between needle-like prongs a orthin metallic film on an insulative substrate that may be shaped in various geometries. It responds to cooling effectsand thus measures both kinematicand thermodynamic fluctuations of the flow.

The hot-wire has been a generally accepted standard for measuring fluctuations in wind tunnel flow since the work of Drydenand Kuethe in 1929 (Ref. l), Its use can be very tedious and thus has often been avoided. However, it has not been replaced because of its advantages that include: small sensor size, high frequency response and sensitivity to pressure, vorticity, and entropy fluctuations. Dr. Kovasznay (Ref. 2) opinioned in 1968 that the hot-wire has not been replaced by other methods becauseof its unique characteristics. Furthermore, significant developmentsin hot-wire methodology in the 1970's indicate continued use of this instrumentin both specialized experiments and in wind tunnel calibrations. Reference 3 is a recent textbook on hot-wire technology. References 4 and 5 provide further background and extended lists of references relative to measurements of fluctuating propertiesin wind tunnels. Reference 2 provides summaries of the early history and the technologyup to 1968.

Useful application of hot-wires to incompressible flow is commonly dated as 1929 (Ref. 1). Experiments with hot wires in supersonic flows began in the mid 1940's. However, equipment and analysis techniques were not considered adequate until the mid 1950's (Refs. 6, 7, and 8). Applications to transonic flows encountered particular difficultyin separating the components of the output signal (Refs. 3 and 8). Recent reports of progress (Refs. 9 and 10) have outlined approaches for practical applicationsin the high-subsonic and transonic test regime.

* This section has been contributed by C. J. Stalmach, Jr., Vought Corporation. Equ ipment Descr ipt ionand Operat-ion

The sensor is electrically heated to maintain either a constantcurrent or constanttemperature (resistance). In the case of constantcurrent, compen- sationfor the thermal lag of thewire is obtained by an output amplifier whose gainwith frequency is adjusted (during a square wave heatinginput) to compensate for decay ofthe output with increasing frequency. For largeaspect ratios (Ud -> 150) the wire exhibits a first order response that is simple to compensate electronically.Adjustment of the non-linear, compensating amplifier is required for eachchange in mean flow condition or sensor.

Inthe case of a constanttemperature anemometer, a high-gainfeedback system provides power tothe wire in response to fluctuations in cooling causedby the flow such that the wire resistance (temperature) remains essen- tiallyconstant. The square ofthe voltage required to maintain constant wire resistanceis a direct measure ofthe heat transfer between thesensor and its environment. The constanttemperature anemometer has severaladvantages compared to a constantcurrent system including:

1.' thermallag not a problemsince sensor temperature isconstant, 2. automaticadjustment to large changes in mean flowconditions whichreduces accidental burnouts and continuesdata acquisition during mean flow changes, 3. direct DC outputas a functionof mean velocity, 4. compatiblewith film and low R/d wiresensors that have complexfrequency response characteristics, and 5. output may be linearized and temperature compensated.

The constantcurrent approach was initially preferred because it provided higher-frequency-response and signal-to-electronic-noiseratio. Solid state electronics, however,have permittedthe constant temperature systems to have comparableperformance and arethe systems now generallypreferred. An exception is measurement oftemperature fluctuations independent of velocity and densityeffects. Here a minimum wiretemperature is required that is best achievedwith a constantcurrent operation. Moderncommercial unitsgenerally incorporateboth circuits. Reference 3 and literaturefrom commercialequip- ment suppliersprovide further details on the power systems and commonly used sensor sty1 es.

204 Response to Mean Flow

A wire or film sensorresponds to changes in flow conditions that affects theheat transfer of the sensor to its environment. For steady flow the sensor response may be expressed as

Nu -A+B & (1.1) or E2 - [C+D (PU)~] (Tw - Te) (1.2)

where x=0.5 forclassical analysis (King's law) of flow around heated cylinders. Figure A.I.l shows theresponse of a hotwire to the mean flow.For supersonic flow, theNussel t number is evaluatedbehind the normal shock(Ref. 6).

Response toFluctuations in the Flow

The simultaneousreaction of the heated sensor to density, velocity and temperature is thekey both to the advantages and difficulties of the hot-wire anemometer approach to measuringflow fluctuations. It is an advantage to haveone sensor measure bothkinematic and thermodynamic fluctuations.In comparison, a laserDoppler velocimeter can only measure thefluctuating velocities and a microphone orpressure sensor responds only to the net sound or pressurefluctuations. Separation of the modes composing theoutput of the heatedsensor isnot simple and, ingeneral, requires a priori knowledge ofthe flowcharacteristics being sampled.

The choiceof techniques for separating the modes of a fluctuatinghot- wiresignal is somewhat dependent on the Machnumber and Reynolds number of thetest. AS indicatedin the summary curve of Fig. A.I.l, wireresponse to mean flowis well defined for the incompressible case.For isothermal, incorn- pressibleflow, a hot-wireresponds only to velocity changes. The sensorOut- put is well behaved forsupersonic Mach numbers asindicated by the lower curve

Of Fig. A.I.1. The sensoroutput, however is Machnumber dependentbetween thesetwo bounds forthe lower range of wire Reynolds number.

Testequipment and data analysis techniques are sufficiently developed to permituseful applications for tunnel calibrations over the Mach range of interestin this report (0.4 5 3.5). Research to improveequipment and analysis shouldcontinue however. In particular, additional work needs to bedone in theareas of transonic flow applications and separationof the signal into its component modes.

205 (Ref. 3)

2.0 1.8 1.6

1.4

1.2

1 .o I t .8 .6 .4

0 0 .2 .4 .6 .8 1 .O

Figure A. 1.2 FLUCTUATIONDIAGRAM FOR 1 PERCENT MASS FLOW FLUCTUATIONS AND 1 PER CENT STAGNATION TEMPERATURE FLUCTUATIONSWITH VARYING DEGREES OF CORRELATION. (Ref. 7) 206 Separation ofModes in Fluctuating Flow In supersonic flow the fluctuating voltage aof heated I wi re placed normal to the flaw canbe expressed in terms of the fluctuating velocity, density and total temperature (Refs. 3, 7, 8, and lo), (using the notation of Ref. 10) :

The sensitivity coefficients for constant temperature sensor operation are:

Rn Nu a 1 T Rn Ret wr a

a Rn Nut 1 ” (I T aRn~ .5) wr

1 1 =- + - (K - 1 - nt) + rn + - sP) (I .6j ’Tt 2 Asw 2 tPs 2 (sU -

For supersonic flow (M > 1.2), the heat loss is insensitive to Mach number, and sensitivity to velocity and density are essentially equal (Refs.3, 6, 8).

:. su = s = s P PU Thus for supersonic flow, Eq.(1.3) may be simplified to (Ref. 7):

The root-mean-square of the sensor output may be expressed as:

where the correlation coefficientof mass flux and temperature is defined by: Sensor output,obtained at three different sensor temperatures, and sensor sensitivity,obtained from calibration, can provide solutions for thethree unknowns <(pu)*>, and RpuT . The normal practiceis to plot data, obtainedat several wire overheat ratios, in modal diagramsas developed by Kovasznay(Ref. 7) and Morkovin(Ref. 8). A fluctuationdiagram for varying degrees ofcorrelation is given in Fig. A.1.2.The characteristic modal diagrams of Kovasznay forfluctuations in velocity, temperature, and sound are shown in Figs. A.1.3, A.1.4, and A.1.5.

Independent fluctuationswhich characterize the flow field are the vorticity (turbulence),entropy (temperature spottiness), and pressure(noise or sound) modes.The vorticity,entropy, and pressuresensitivity coefficients are relatedto the measured density,velocity, and totaltemperature sensitivity coefficients as follows(Ref. 8):

Su + 8 ST (1.10) t

s =s +UST (1.11) UP t n X Su + (y-1) (l+nxM) ST (I.12) -M a t

Where nx is thedirection cosine of the normal to a plane sound wave front relativeto the flow direction. If two or more measureable sound sourceswith distinctorientation exist, then asound sensitivitycoefficient would be requiredfor each sound wave direction.

Themodal diagramtechnique is an accepted means ofdistinguishing the Pri- mary characteristicsof the sensor signal in supersonic and certainhypersonic flows. An importantapplication of the modal diagram to windtunnel calibration is to distinguish sound from a fixedsource (such as causedby roughness orholes in the wall) asopposed to a moving sound source(such as eminating from a turbulent boundarylayer on the tunnel wall). This technique * is effectiveonly for flows where thetemperature spottedness isnegligible. The diagram for a fixedsource of soundhas an origin-intercept(see Fig. A. 1.5) and thediagram for a movingsource of soundhas a positive ordinate-intercept similar to thetemperature diagram of Fig. A.1.4. * See Refs. 11, 12, 13, and 14 for examples.

208 //I

- STt

-1 0 1 2 'Tt

Figure A. I .3 FLUCTUATIONDIAGRAM FOR 1 PERCENT TURBULENT VELOCITYFLUCTUATIONS (VORTICITY MODE). (Ref. 7)

Figure A.1.4 FLUCTUATIONDIAGRAH FOR I PER CENTTEMPERATURE SPOTTINESS (ENTROPY MODE). (Ref. 7) 2

0 -I 0 I 2

~l~~~~A.I.~ FLUCTUATION DIAGRAM FORSOUND WAVES THAT AREALMOST MACH WhVES HAVING1 PERCENT PRESSURE FLUCTUATIONS (Ref. 7)

0 -1 -.5 0 .5 I

sPu’sT Figure A.1.6 FLUCTUATIONDIAGRAM FOR UNCORRELATED MODES AT M = 1.75; TEMPERATURE SPOTTINESS 0.1 PER CENT;TURBULENT VELOCITY FLUCTUATIONS 0.2 PER CENT; SOUNDWAVES (DETECTABLE) 0.1 PER CENT OF MASS FLOW FLllCTllATIONS.(DOTTED LINES SHOW SEPARATECONTRIBUTIONS.) (Ref. 7) 21 0 When thedominate mode is sound, thefoilcwing lsentroplc relations betweenpressure, density and temperatureare appropriate (Ref. 11):

(1.14)

The resultanthot-wire equation and calculation of the fluctuation quantities inthe freestream are given in References 11 and 12.

In Reference 15. an applicationof the diagram approach is shown which identifiestemperature fluctuations in the wake of a wedge in hypersonicflow where theadjoining freestream signal level is low. If bothtemperature fluctuations and moving sound sourcesare likely present (Ref. 161, theInterp- retation of thedlagrams becomes more difficult. An example of a hot-wire output,in a casewhere all three modes,are presentis shown In Figure A.1.6. It is readily seen that in the absence of onedominant mode, separationof the component modes canbe a problem.References 7. 8, and 12 may be consulted for more detailsconcerning the modal diagramtechnique.

TransonicFlows The mdal diagramapproach cannot be generally applied for compressible subsonicand transonic flows (see Fig. A.l.1) where thederivatives of the Nusselt number and recoveryfactor with respect to Mach number arenot zero, and su # s forail overheatratios (Refs. 3, 8, and IO). Transonicopera- P tion at high dynamicpressures also increases problems with wire breakage.

The aboveproblems helpexplain the limited usage ofhot-wire systems in transonicwlnd tunnels. Recent developments, however, provide examples for overcomingthese difficulties. The sensor failure problem may be alleviated withthe use of shorter wires (Ud - IOO), wireswith insulative backing or film sensors(Ref. IO, 17, and 18). Heat lossesto end supportsor substrate and possibleinterference effects necessitate that eachsensor be calibrated in a representativeflow environment. The sensitivitycoefficients Sp and Su

21 1 havebeen systematically measured intransonic flow by independently varying density and velocity (Refs. 9 and IO). These resultsestablish that Sp and Su areapproximately equal for all Mach numbers (includingthe troublesome transonicrange) if thewire overheat ratio is greater than 0.5 and thewire Reynolds number isgreater than 20. Operationwithin these restrictions again permitsthe use of the simplified expression of Eq. (1.7).For many transonic windtunnels the total temperature fluctuations are xgiigible relative to themass-flux term. In this casethe hot-wire directly senses thefluctua- tions of themass-flux. If thelevel of temperature fluctuation Is unknown In a facility (such as anew cryooenictunnel) the level may beascertained with a sensoroperated atconstant current and nearthe recovery temperature.

Reductionof mass-flux measurements in a transonicflow into its elements requiresfurther assumptions, e.g., possibleapplication of the modal diagrams or anindependent measurement with a laservelocimeter, pressure transducer or a special hot-film geometry.For boundary-layer flow. the pressure fluctuations cangenerally be neglected relative to vorticity.Operating within the above describedwire and flow domain,boundary layer profiles of velocity.density and Reynoldsshear stress weresuccessfully obtained in Ref. 10 for a nominal Hach number of 0.8. For operationin the transonic freestream, it appearsthat the principle of modal diagramscan be applied to high-Reynolds-number flow byobtaining data at several over-heat ratios, all being greater than 0.5. Straightline fairings OF thedata, extrapolated to theordinate, would then provideinformation on the dominate mode asin the supersonic case. Another approach for separation of thefluctuation modes is to employ a yet-to-be defined fi Im sensor that has two (or more) films and a geometrysuch that one film responds to themass-flux and a second film responds to pressure (geo- metricallyshielded from velocity). An application of specialsensor geometry, fordirect correlation measurements,hasbeen reported for hypersonic,boundarylayer flow (Ref. 17).

Comparison of Hot Wire to Other Systems

Hot-wiredata have, ingeneral, compared wellwith measurementsfrom otherdevices. Agreement of velocityprofiles obtained with hot wire and laser systems within a shock-wave/boundary-layer interaction(Ref. 19) gives credence toboth systems and tends to validatethe assumptions employed in

21 2 datareduction. In the freestream of many transonic and supersonicwind tunnelsthe sound mode generallydominates, and in suchcases good agree- menthas been obtained between hot-wire and pressuretransducer measurements(Refs. 14and 20). An examplecomparison of Pitot and hot-wire measurements at M = 5 is shown inFig. A.1.7. Fordiagnostic measurements of flowswith noise dnminated disturbances, a dynamic pressuretransducer may sensesuch fluctuations with much less effort than a hot-wire system. A dynamic Pitotpressure survey may satisfy many windtunn-1 calibration requirements.Another widely used Calibration model isthe AEDC developed, pressure-instrumented IOo cone. (See Section 1 I I .F. 1 .).

In summary, thehot-wire can provide more informationthan a dynamic pressuretransducer, e.g., it can distinguish betweenmoving and fixedsources of sound with a single sensor. The wirealso provides a higherfrequency re- sponseand, ingeneral, has a more omnidirectionalresponse to noise sources. The smallersensor of a hot-wire system isimportant (in addition to the frequencyaspect) if localfluctuating measurements in a shock or boundary layerare required. Further, the wire/film sensor may be locatedin areas hidden from theview of a laserDoppler velocimeter.

OtherData Analysis Techniques

Otherdata-reduction techniques for the fluctuatinq signal include time (autocorrelation) and spatial(location and direction)correlations of the signal.Spectral analysis provides the energy content at eachfrequency. These techniques and associatedelectronic equipment are fairly common and are described inthe literature. (Suggested referencesinclude 3, 12, 14, and 21).

Sensor Choice and Calibration Reouirements

Inthe past, the wire-breakage problem had encouraged minimizingcalibra- tion time. Minimum calibrationrequires maximum use ofcorrelations of exist- ingcalibration data and end-losscorrections to predict the performance of a wire (Refs. 8, 12,22, and 23). A more recentapproach, that promises improved datagathering efficiency and dataaccuracy, is the use of more durable,rugged sensors. A ruggedsensor may include a fine-diametershort wire (E/d - loo), a wire backedby an insulator,or a largerdiameter film sensor(Refs. IO, 17, 18 and24). The complexheat loss and possiblesupport interferences with such sensors requireindividual sensor calibrations and theuse of a constant temperature system. 213 Fully turbulent nozzle wall baundary layer

-+I

- -turbulen-First aTa*nce smt Oi

I t J

Figure A. I .7 COMPARISON OF PITOT PROBE AM, HOT-WIRE WSUREMENTS OF FREE-STREAM PRESSURE FLUCTUATIONS IN A CONVENTIONAL, MACH 5 NOZZLE, Ref. 14. 214 Calibration of a heatedsensor may, in general, be obtainedin situ by maintaining a constant Machnumber and temperature and varyingthe total pressure (and therebyReynolds number). For supersonic flow and for R > 20 intransonic flow, S = Su = S . Furtherdiscussions on such mean flow P PU approaches to calibration and more-refined dynamic calibrations are given in References 3, 8, 9, 14, 18, 25, and 26. Rose andHorstman have successfully used therugged class of tungsten Yk hot-wiresin a transonicflow for over 16 hourswithout breakage. Commercially available film sensors(such as 0.002 to 0.006-inchdiameter cylinders) should alsoreceive consideration for wind tunnel calibration because ofthe following advantages :

. superiorresistance to particle damage, . superiorresistance to surface contamination and straingaging, . highersensor Reynolds numbers (particularlyimportant in transonic flow).

Compared to the rugged classof tungsten wire sensor, the Platinum-film Sensor has comparable capabilitiesin maximum overheatratio (-one) and frequency response ( - 150K Hz).

* Privatecommunication Summary of Advantagesand Disadvantages of HotWire System

The advantagesand disadvantages of using a hot-wiresystem to measure thefluctuating flow properties during a windtunnel calibration are summarized as follows.

Advantages

1. Smallsensor size

2. Highfrequency response

3. Highsensitivity

4. Sensitiveto both kinematic and thermodynamic fluctuations

5. Distinguish betweenmoving and stationarynoise sources with a singlesensor (in flows where temperaturespottedness isnegligible)

6. Re1 iable systemsand sensors commercially available

7. Rugged sensorsavailable, particularly film type

D i sadvan tages"

1. Possiblefrequent breakage of fine-wire sensor (due to air loads, vibrations,particle impingement, burnout,oxidation, accidents, etc.)

2. Possiblefalse signal (usually apparentldue to strain guaging, contamination,or vibration of probe.

3. Calibration may be requiredin situ (facility time may be expensive).

4. Separation of signalinto independent modes requiresassumptions concerningflow characteristics or independent measurements

5. Analyses ofsignal particularly difficult for compressible subsonic ortransonic flow, unless restricted to higher Revnolds numbers and wiretemperatures.

f: Since the original writing of this section, a review of hot-wire anemometry by Comte-Bellot has been published(Ref. 27). Thisreference provides additionaldiscussion of problem areas.

216 NOMENCLATURE

constants a En R~ overheatparameters, 1/2 a En , wire d iameter

E wire voltage K2 time-averaged, totalpressure behind a normalshock h heat-transfer coefficient I wire current k d in Rw/d EnTw

Kn Knudsen number, n/d

R wirelength m d Rn p/d Rn Tw

M Mach number n d En k/d Rn Tw

directioncosine of normal to sound plane "X wave front relative to flow direction hd Nu Nusselt number, -k

staticpressure

resistance

Reynolds number, pudh

correlation coefficient of mass-flux and total temperature fluctuations,

sensor sensitivity coefficient

temperature

axialvelocity

exponent inequation 2 21 7 a

Y ratio of specific heats,1.40 used for air

n recovery factor, r'T t

x molecular mean free path U vi scos i ty

P dens i ty

'w r temperature overheat, CTW - Tr) Tr root mean square

Superscripts

( 1' fluctuating value

(7 time averaged

Subscripts

e environment

r recovery of adiabatic wall

t total or stagnation conditions

t2 stagnation correction behind normal shock T temperature

U velocity

W wire

U sound

P dens i ty

PU mass flux

0 entropy

T vorticity

218 A. REFERENCES1.

1. H. L. Dryden and A. M. Kuethe: "Effectof Turbulence in Wind Tunnel Measurements," NACA Tech, Rep. 342, 1930.

2. "Advances in HotWire Anemometry, Proceedings ofthe International Symposium onHot Wire Anemometry," Edited by W. L. Melnik and J. R. Weske, University of Maryland, AFOSRNo. 68 - 1492, USAF Office of Scientific Research, July, 1968.

3. V. A. Sandborn: Resistance Temperature'fransducers, MetropologyPress, Fort Call ins, Colorado, 1972.

4. V. A. Sandborn: "A Review of Turbulence Measurements in Compressible - r Flow," NASA TH X-62,337, March 1974.

5. R. Westley:"Problems ofNoise Measurements In Ground-Based Facilities with Forward-SpeedSimulation," Appendlx 5 of "A Further Review of Current Research Aimed atthe Design and Operationof Large Wind Tunnels,'' Second Report ofthe Minilaws Working Group,'AGARD-AR-83, Sept. 1975.

6. L. S. G. Kovasznay: "The Hot Wire Anemometer in Supersonic Flow," Jour. Aero. Sci., Vol. 17, No. 9, September, 1950. 7. L. S. G. Kovasznay: "Turbulence in Supersonic Flow," Jour.Aero. Sci., Vol. 20, No. 10, October, 1953.

8. M. V. Morkovin:"Fluctuations andHot Wire Anemometry in Compressible Flows,'' AGARDograph,24, November, 1956.

9. W. C. Rose and E. P. McDaid: "Turbulence Measurement in Transonic Flow," Proc. AlAA 9th Aerodynamic TestingConference, June 1976.

10. C. C. Horstman and W. C. Rose: , "HotWire Anemometry in Transonic Flow," NASA TM X-62,495, December 1975. 11. J. Laufer: "Aerodynamic Noisein Supersonic Wind Tunnels,"Jour. Aero. Sci.,Vol. 28, No. 9, September 1961.

12. J. C. Donaldson and J. P. Wallace:"Flow Fluctuations Measurements at

Machnumber 4 inthe Test Section of the 12 InchSupersonic Tunnel (D) ,'I AEDC-TR-71-143, August 1971.

13. M. C. Fischer and R. D. Wagner: "Transition andHot Hire Measuremen.ts in Hypersonic He1 ium Flow," AlAA Journal,Vol. 10, No. 10, October 1972. 14. J. B. Anders, P. C. Stainback, L. R. Keefe;and 1. E. Beckwith: "Sound and FluctuatingDisturbance Mea'surements in the Settling Chamber and TestSection," ICIASF '75 Record, Int'l Congress,onInstrumentation in

Aerospace Simulation Faci 1 ities, Ottawa, Canada Sept.22-24,: 1975, , published by IEE, 345 E. 47thStreet, New York.

15. R. D. Wagner and L. M. Weinstein: "Hot Wire Anemometry in Hypersonic He1 iumFlow," NASA TN 0-7465,June 1974.

16. P. C. Stainback, etal: "ExperimentalStudies of Hypersonic Boundary - Layer Transition and Effects of Wind TunnelDisturbances," NASA TN D-7453, NASA LangleyResearch Center, March 1974.

17. V. Mikulla and C. C. Horstman: "TurbulenceStress Measurements in a Non- adiabaticHypersonic BoundaryLayer," AlAA Journal,Vol. 13, No. 12, December 1975.

18. W. C. Rose:"The Behavior of a CompressibleTurbulent BoundaryLayer in aShock-Wave-Induced AdversePressure Gradient,'' NASA TN D-7092, NASA Ames ResearchCenter, March 1974.

19. w. C. Rose and D. A. Johnson: "Turbulence in a Shock-Wave Boundary-Layer Interaction," AlAA Journal,Vol. 13, No. 7, July 1975.

20. E. Grandeand G. C. Oates: "Response ofMiniature Pressure Transducers toFluctuations in Supersonic Flow," Instrumentation for Airbreathing Propulsion, AIAA Series onProgress in Astronautics and Aeronautics, Vol. 34, 1972.

21. R. K. Otnes and L. Enochson: Digital Time SeriesAnalysis, Wiley, New York, 1972-

22. W. Behrens: "TotalTemperature Thermocouple Probe Based onRecovery Temperature of CircularCylinder," Int. J.Heat and Mass Transfer, Vol. 14, 1971.

23. C. F. Dewey, Jr.: "Hot Wire Measurements in Low Reynolds Number Hypersonic Flows," ARS Journal,Vol. 28, No. 12,December 1961.

24. E. L. Doughman: "Development of aHot Wire Anemometer forHypersonic Turbulent Flows,"The Review of ScientificInstruments, Vol. 43, No. 8, August 1972.

220 25. R. F. Rosenberg: "Some Aspectson Hot Wire Anemometry Leading to a SpecialCalibration Method for HotWire Probes," ARL 71-0038, USAF AerospaceResearch Laboratories, March 1971.

26. R. H. Klrchhoff and R. R. Safarik,"Turbulence calibration of a Hot Wire Anemometer," AlAA Journal,Vol. 12, No. 5, May 1974.

27. Comte-Bellot, G.: "Hot-Wire Anemometry," Annual Reviews ofFluid Mechanics,Vol. 8, Palo Alto, Calif. 1976.

22 1 APPEND IX I I LASERDOPPLER VELOCIMETER MEASUREMENTS

Thedevelopment of the laser was quickly folloyed by its application to the measurement of the velocity of a moving object by observing the Doppler shift in thefrequency of theincident laser light. Liquid flow velocities weremeasured by Yeh and Cummins (Ref. 1) in 1964, utilizing Dopplerradiation fromsmall particles entrained in the flow. In 1965, thetechnique was refined andmeasurements were made in seededgas flowby Foreman,George andLewis (Ref.2). Since thattime very significant advanceshave been made both in thetheoretical understanding of the technique, and in improved opticalarrangements and signalprocessors. The LaserDoppler Velocimeter (LDV) technique has been appliedto measurements of mean velocity,turbulent intensity and flow directionin a variety of flowfields and inboth liquids andgases. By 1970 thevelocity range over which measurementshad been made extended from 10’ 4 cm/sec to 1000 m/sec (Ref. 3). Thisrapid rate of developmentcontinues. Literature on thesubject is now extensive; a bibliography (Ref. 4) published in 1972 contains 190 references.

Laservelocimeter systemshave been operated inseveral wind tunnels, e.g., Refs. 3, 5, 6, 7, 8, and theirunique capabilities, especially for non-intrusivevelocity and turbulencesurveys around models in thetunnel, shouldinsure their continued development and application.Although to the present,the LDV has not been widelyapplied to basic tunnel calibration measurements, this capability hasbeen demonstrated and their availability to thewind tunnel operator will probablylead to this use.The discussionhere- in will be restricted primarily to consideration of the LDV techniquefor application to basic facility flow calibration measurementssuch as mean velocity (Mach number) distribution,turbulence intensity and flowangularity.

BasicPrinciples

Consideration of theadvantages and disadvantages of laser velocimeter techniquesrequire some discussion of the basic principles of the system.

The principle of the LDV can be descr,ibedby both the Doppler effect and byan interference-fringe model.Since both approaches yieldthe same basic equationsdiscussed inthis report, the fringe model will beused since it aidsvisualization of the physical principles involved. Several optical arrangements arepossible, but considerations will be limitedto the dual-

222 beam or differential-Doppler systemmost commonly used for windtunnel measurements.

For a single-component,dual-beam system, Fig. A.II.l, a laser beam is separatedinto two parallel beams of equalintensity separated by a distance A. These beams enter a lenswhich causes them tocross at the focal point of thelens where themeasuring or probe volume is formed. Inthis region, the wavefrontsinterfere constructively and destructively to form stationary, alternatedark and brightregions or fringes,Fig. A.11.2. A particle moving throughthe measuring volume causes variations in the intensity of the light scatteredby the particle. The scatteredlight is collected and focusedby a second lensonto a photodetector,usually a photomultipliertube. r The receivingoptics and photodetector may be located on the same side of themeasurinq volume as the laser and transmitting optics or may be located on theopposite side. If locatedon the same side,the svstem utilizeslight scatteredin the backward direction by particlesin the flow (backscatter mode). If laser and receivingoptics are on oppositesides of the measuring volume, the system utilizes 1 ightscattered in the forward direction (forward-scatter mode). Obviousoperational advantages are associated with the backscatter mode, partic- ularlyfor wind tunnel application. However, since much more light is scattered inthe forward direction, Fig. A.11.3, thesignal-to-noise ratio is significantly higherwith the forward-scatter system.

The photomultipliertube generates an electricalsignal at a frequency directlyproportional to the velocity and inverselyproportional to the fringe spacing,Fig. A.11.4, accordinqto the relation

where fdis the Doppler frequency, U thevelocity component normal to the fringes and perpendicularto the bisector of the beam angle, and 8f is the fringespacing.

The fringespacing can be determined from the systemgeometry and the wavelength ofthe laser as shown inFigs. A.II.l and 2, i.e.,

(I I .2) N N Ream &

Collecting

/ i\

L Processor imatomutiplier I Tube (Backscatter Mode)

Figure A. I 1.1. DUAL BEAM LASERDOPPLER VELOCIMETER, WITH CPPIOPIAL FORWARD AND MCKSCA!M!ER MODES A-

Figure A. I I .2 GENERATION OF INTERFERENCE FRINGES IN MEASURING VOLUME OF DUAL BEAM LASER - DOPPLER VELOCIMETER Forward Scatter Scatter Scatter

Electrical signal Incident Light _c c

Time

Figure A I I 3 LIGHT SCATTERED BY Figure A. I I. 4 LASER ANEMOMETER SIGNAL A SMALL PARTICLE FROEl PHGT0DE;TECTOR

. where X isthe wavelength of the laser and 0 is the angle between theinter- secting beams.From the above relations:

The measured velocity is the absolute value of the velocity of the particle which isnot necessarily equal to that of thefluid. The problems ofpartic'le lag will be discussedin more detaillater in this section. The ambiguity withregard to flow direction becomes a problem inturbulent flows at low, mean-component velocity where flowreversal may be encountered,but can be overcome by introducing an acousto-opticmodulator (Bragg cell) into one of thetwo parallel laser beams. The Bragg cellintroduces an accurately known frequencyshift into one ofthe beams, whichresults in a moving ratherthan a stationaryfringe system.Zero particlevelocity then corresponds to the frequencyshift; higher or lower velocities generate higher or lower output signalfrequencies. In this manner thedirectional ambiguity can be eliminated and reversingflows can be measured. Formain-stream, empty-test-section measurements in a windtunnel the directional ambiguity will notnormally be a problem,but the Bragg cell may still be usefulat high flow velocities to ' down-shiftthe signal frequency to a range that can be more readily measured by theelectronic signal processor.

The measuring orprobe volume is an ellipsoid which may bedefined by thecontour where the light intensity decreases tol/e times the maximum intensityat the center of theprobe volume, Ref. 9. The width and length of the volume Fig. A.11.2 may be defined by

where do isthe diameter of the intersecting laser beams atthe focal point of thetransmitting lens. The focuseddiameter,do,is related tothe initiallaser beam diameter, Do, by

227 (I 1.6)

where F isthe transmitting lens focal length. Other relations may be derived forthe dimensions of the probe volume,dependinq upon how the"effective" geometricboundaries are defined(Refs. 3, 10, 11). The above equation demon stratesthat a large,original (or expanded) laser-beamdiameter anda short- focal-lengthlens yield the minimum sizemeasuring or probe volume.The focal lengthis also determined by thedistance from thetransmitting lens to the point of the flow where measurements are desired.

The selection of the measuringvolume is a designproblem during which conflictingrequirements must be balanced. The initialrestrictions are usuallydetermined by the minimum number of fringesrequired for the signal processorto measure thefrequency with adequate accuracy. Also, the Doppler signalfrequency must not exceed the maximum whichcan be measuredby the processor:this frequently is a functionof the fringe spacing and themxi- mum velocityto be measured, asdefined by Eq. (11.1). The required minimum number of usablefringes canbe as low as eight with a countersystem, but a somewhat larger number isnormally preferred, The number of fringescan be determinedfrom the width of the measuring volume, wv, divided by the fringe

spacing,bf, as definedby Eqs. (I1.2), (I1.4) and (i1.6), i.e., .

0 8F sin (F) - (I 1.8)

TI Do

From theoptical configuration, Fig. A.II.1, it may be notedthat the angle (8/2) isnormally smal 1.

Here A isthe beam separationdistance at the transmitting lens. From Eqs. (II. 8) and (II. 9). the number of fringes can be expressed in terms of onlythe beam separationdistance and the initial laser-beamdiameter.

4A Nfr = - - (11.10) TI Do 228 Againnote that the number of fringes can be increased by reducingthe unfocused beam diameter, Do, butthis in turn increases the probe diameter. Increasingthe beam separation, A, for a fixedtransmitting lens focal length, is aneffective means ofincreasing the number of fringes.

Based on thepreceeding equations, it isof interest to calculate the dimensions of the measuringvolume and other characteristics of a velocimeter system.Consider an Argon laser with a wavelength of 514.5 nm andan input beam diameter of 1.5 mm. If thetransmitting lens focal length is 1.5 m and the beam spacing is 100 mm, the beam angleis 3.818". From Eq.' (11.6) the focused beam diameter is 0.65 mm which is alsothe maximum widthof the probe volume. The lengthof the probe volume is, from Eq. (11.5) , 19.7 mm. The fringespacing is 7.7 and the maximum number offringes is 84. At a velocity of 300m/sec., theDoppler frequency (Eq. (11.1) will be 38.96 MHz. Thissystem isnot necessarily typical, but it illustratesapplication of the system relationshipsdiscussed previously.

To thispoint the measurement of a singlevelocity component (normal to thefringe pattern) has been described. The fringepattern of a single- component LDV can be rotatedby rotating the transmitting optics. This permits the measurement of velocity components at two or more anglessuch as -+45 degrees tothe nominal tunnel centerline. From thetwo velocity component measurements, thevelocity vector in the tunnel plane normal tothe bisector ofthe intersecting beams canbe determined. Thus, flow angularitycan be measured inaddition to velocity and turbulenceintensity.

Opticalarrangements to yield simultaneous two-componentmeasurements can be realized by using two pairsof intersecting beams, usuallywith the plane defined by the second pairof beams normal tothat defined bythe first beam pair.Separation of the two measurementscan be achieved by splitting a laser beam into two pairsof beams, each pairpolarized 90" tothe other. Polarized filters on the two photodetectorsallow each detectorto see only thelight scattered from the fringes formed by two ofthe four beams atthe intersectionpoint. Two wavelengths oflaser light can also be used for two componentmeasurements. The use of an Argon ionlaser is particularly con- venientfor this purposesince two strongcolor hands, 488 nm (blue) and 514.5 nm (green)are available. Optical filters allow separation of the scatteredlight so eachphotodetector sees onlythe light of interest. A third velocity component, parallel to thebisector of the intersecting beams, can alsobe measured simultaneously.For example, Orloff and Logan (Ref. 12) havedescribed an LDV system for measuring all three velocity components whichemploys backscattering and a reference-beam method.

SignalProcessors

The outputsignal from the photomultiplier tube, shown on Fig. A.11.4, is a frequency'burstat the Doppler frequency with amplitude modulated accordingto the intensity distribution across the fringes. This amplitude-modulated envelope is commonly referredto asthe "pedestal" and must be removed byhigh- pass filtering or optical means beforeprocessing. The number ofcycles of thedoppler signal and themodulation intensity about the pedestal envelope will varyaccording to the location at which the particle crosses the probe volume, thesize of the particle and the number ofparticles present at one timewithin the probe volume. Signal bursts of measurable amplitude and the minimum required number of cyclesoccur at random timeintervals, and phase reversalsduring a singlesignal burst will occur when multipleparticles are present .

Several methods forprocessing the data from the photomultiplier have been used.These include:

spectrumanalyzers photoncorrelators filter banks

o frequencytrackers

0 counters

Onlythe last twotypes of processorsproduce essentially real-time velocity informationdirectly and arecurrently used for mostwind tunnel applications.

The frequencytracker, as the name implies,converts the Doppler frequencyreceived from the photodetector into a proportional,analog voltage. The trackercircuit is implementedusing phase orfrequency locked loops, or a combination ofthe two.Both types ofloops function by comparingthe outputfrequency of a voltage-controlled oscillator (VCO) or a voltage-to- frequencyconverter (V/F) tothe input signal frequency, and bothutilize the differencein frequency to modify or adjust the ac input voltage to the in- ternalfrequency generator. The dc voltage is thenproportional to the input frequency.Alternatively, the internally-generated frequency can be converted to a digital signal by means of a counter. Trackerprocessors are characterized by the maximum frequencyrange, capturebandwidth, dynamic range and slewrate. If the change invelocity from one particle to the next exceeds thecapture bandwidth or capture range,the tracker will loselock and nottrack the particle or other particles which are outside the capture range.

Where Dopplersignal tracking is interrupted bya drop-out,the last outputvoltage level is normally placed in "hold". If thesignal returns duringthe hold period, tracking is resumed. If thesignal is not recap- turedduring the hold period, the search or sweepmode is activated until thesignal is re-acquired. Signals can be providedto external data systems torecord mean velocity and ac or turbulent fluctuation signals, e.g.,Ref. 13. Dopplerfrequencies can bemeasured from 2kHz to 50 MHz. Velocity changes over a 200: 1 rangecan be followed, and inthe search mode the frequencyslew rate can be as high as 400 MHz/ms. Data-validationfeatures arealso normally incorporated. For example, one systemrequires tracking for 8 Dopplercycles and holdingfor 2 additionalcycles without drop-out in orderto be considered a validdata point. Thus, forheavily seeded flows, datarates up to 1 x 10 6 per second can be processed.

The counteror burst processor for laser-anemometer signalsaccurately measures thetime required for a particle in the flow totravel across a fixed number offringes in the measuringvolume, i.e., a known distance. From these two quantities, a counterdetermines the Doppler frequency and therebythe particle velocity. The counter may be inherently a digitalin- strument, so drift and calibration problemswhich may be associatedwith analogprocessors, such as trackers,are avoided. Counter processors are normallyconfigured to yield a directdigital output.

A counterfunctions by passing a lasersignal burst through a thresh- oldlevel detector which, when theadjustable amplitude-threshold is exceeded, enables a zero-crossingdetector suchas a Schmidt trigger. Those Dopplersignals above theamplitude-threshold level are then converted into a train of square waves, with a frequencyequal to the origianl signal frequency. Many electronicfrequency-counters function by counting each cycle of the unknown signalfor an accurately-fixedtime period, suchas 0.1, 1.0 or IO seconds.The readingsare then converted to the signal frequency

23 1 inhertz. Since the maximum number ofcycles available from the passage of a particleacross the measuring volume of a LDV isextremely small, the accuracyof the direct counting procedure would be totally inadequate. To avoid this problemcounter processors are period measurement devices,i.e., pulses from an accurate, high-frequency oscillator or clock are accumulated in a registerduring the time interval corresponding to a fixed number ofpulses from theSchmidt trigger. Alternatively, the time intervals canbe converted tovoltage amplitudes which can bemeasured digitally. The timeresolution forthe n periodranges from 2 to IO x 10-9 seconds,depending upon theclock frequency, e.g., 500 MHz corresponds to 2 x 10-9 seconds resolution. The counterprocessor also normally includes computational capabilities to convert theperiod information into either frequency or velocity units for digital display. The computationtime istypically about 1 x IOm6 sec so that, even athigh velocities where theDoppler frequency is IO to 20 MHz, thetotal acquisition and computationtime for one individual measurement can be asshort -6 as 2 to 3 x 10 sec.Data acquisitionrates of 100,000 readings/secare thereforetheoretically possible (but far from common) with moderateconcentra- tionof particles, Ref. 14.

Counterprocessors include several data-validation features to allow the rejectionof noise bursts, detect the loss of a bit or cycle during a process- ingcycle, reject signals from large particles, etc. The primarytechnique utilized to rejectdata sequences in which one or more cycles may be missing (cycleamplitude below threshold) consists of using two or mre registers. The clockpulses are gated into both a high and low register on the first cycle. The clockpulses to the low register are gated off after NL cycles,while the highregister accumulates NH cycles. A comparaterthen computes theratio of the twotime intervals, which should be equal,to the ratio NH/NL. If the erroris within pre-set limits, the measurement isvalidated.

Asher(Ref. 14) demonstratesthe advantage of an odd ratio, NL/NH, such as 5/8 or 10/16 overeven ratios. Odd ratios suchas 5/8 are commonlyused. A counterprocessor using three different registers hasbeen used at AEDC. None of thesesystems completely reject spurious signals, but they do greatly reducethe probability of such databeing considered valid. Large particles, which may be lagging the fluid flow to anexcessive degree, can be detected if thetotal input signal amplitude (pedestal plus Doppler frequency) exceeds a pre-set limit. The adjustablethreshold level (required to enable a zero- crossingdetector) can be used to reject signals with inadequate ~igna1-t~- noiseratio. Setting the threshold level high, on the other hand, canbias thedata acquired to largeparticles by rejectinglow-level signals fromsmall particles. Mostcounter processors also include functions suchas a digitalindication ofeither the data rate (validated data points per second) orpercent of total

datasignals processed that are validated. An output is normallyprovided to ' indicate eachtime a new datapoint is validated andprocessed. A digital outputcan be made availableto allow introduction of the data directly into a computer orother digital recording system. An analogvoltage output is also normally available.

Due tothe difference in operating principles, the counter is notlimited byslew rateor tracking rate performance. The counter is an extremelywide- band instrument,while the tracker is inherently narrow-band. Assuming the noisepresent is broadband,the noise rejection characteristics of thetracker is superiorto that of the counter. However, sincethe tracker operates in thefrequency domain, it is responsiveto Doppler frequency spectrum broadening resultingfrom the finite duration of the signal burst. This broadening is similarto the modulation sidebands generated when a carrierfrequency is amplitudemodulated. The presence of multipleparticles in the probe volume alsogenerates phase reversalswhich cause spectrum broadening. Since the standarddeviation about the mean velocity is a measure of turbulence and is relatedto the corresponding Doppler frequency deviation through Eq. (ll.l), spectralbroadening can interfere with turbulence measurements.

-Particle Size and Distribution Effects The size,size distribution, concentration, and physicalcharacteristics ofthe' particles in a transonicor supersonic tunnel flow field are of great importance,whether the particles are naturally present in the flow or the flowis seeded. Inthe presence of velocity gradients or turbulence, a significant velocity lag may exist between the fluid motion and theparticle motion. Lag effectsare most significant at highfrequencies and inregions of rapid fluid acceleration or deceleration, as across a shock or expansion wave or along a streamlineapproaching a stagnationpoint. These environments,of interestin flow field surveys, would not normally be encountered in empty test sectioncalibration measurements, but the ability of theparticles to follow relativelysmall mean velocityperturbations and respond to low-to-moderate levelsof turbulence are of interest. 233 The upper limit on particlesize is determined by inertia or lag effects. and thelower limit, forhigh speed:flows, may be determinedby the reduced amount of 1 ightscattered by the particle which results in an unacceptably lowsignal-to-noise ratio.

Flowseeding may be required to obtain 'anadequate number of particles of controlledsize, especially if a trackerprocessor is used.For measurements with a counterprocessor, naturally occuring particles in the flow may be utilized, Refs- 6 and 7. These particlesare normally sparsely concentrated and result in nomore than one particlein the measuring volume at the same time. However, seeding may be employed with a counterprocessor to control particlesize and increasethe data rate. Measurements inthe 16-foot transonic tunnelat NASA LangleyResearch Center have been made usingatomized oil to seed theflow, Ref. 5. Seeding in a continuouswind tunnel can create contam- inationproblems which encourage the use of measurement procedures applicable to unseeded flows.

The motionof a spherical particle in a fluid flow hasbeen reviewedby Hinze (Ref. IS). So0 (Ref. 16) alsoreviewed andsummarized theequations describingparticle motion. The completeequation, as givenby Hinze is:

+ 3 dp2VT& ltdt' [% - Fe dr7'%I+ t0 The subscript g refers to the gas andp to the particle; dp isthe diameter ofthe particle (assumed to be spherical), and t' is a dummy variable.

In this formthe term on the left of the equality sign is the force requiredto accelerate the particle. The first termon the right is the drag force basedon Stokes' law. The second termaccounts for thepressure gradient in the fluid aroundthe particle causedby acceleration of the fluid. The third term is theforce required to accelerate the apparent mass of the particle relative to the ambient fluid.

234 The fourthterm (designated the "Basset"term) accounts forthe deviation of theflow pattern from steadystate. The Vastterm, Fe, represents body forces due to gravity, Lorentz force on a charged particle in an electric field,laser photon pressure, etc. Base (Ref. 17) reviewsthe complete equation as given byHinze above but employs a morecomplex expressionfor drag(Oseen's law) to accommodate higherrelative Reynolds numbers*

According to Hinze,the second, third and fourthterms on theright of Eq. (11.11) may beneglected if thedensity of the fluid is significantly lessthan the density of the particle, which is normally true. In the range of speeds encountered intransonic and supersonictunnels, gravity effects arenegligible compared tothe drag force. Retaining only the Stokes' law drag term on theright side, Eq. (11.11) may be written

(I 1.12)

Thisdifferential equation, which agrees with that given by So0 (Ref. 16), may betransformed into the standard transfer function form

where S isthe Laplace operator,and T isthe time constant defined by P

(11.14)

The steadystate sinusoidal amplitude and phaseresponse of the particle with respect to the gas may beexpressed in thefrequency domainby substituting jw for S, where j = 6.

Herew isthe frequency of gas motionin radians per second. The phaseangle, 4, bywhich the particle lags the fluid motion is

(11.16)

235 The aboveequations based onStokes' law agree with those given by Feller and Meyers(Ref. 18) and others. Mazumder, Hoyle and Kirsch (Ref. 19) and Yantaand Gates (Ref. 20) use a timeconstant expression simi lar to Eq. (II. 14) exceptthat a correctionterm is applied to the Stokes' drag coefficient to extend its applicability to the range of flow conditions where the Knudsen a. number (Kn = - ) becomes appreciable.Epstein (Ref. 21) derived a correction dP termto Stokes' law from the kinetic theory viewpoint, as is discussed by Happeland Brenner (Ref. 22), butvarious forms and empiricalconstants have evolved. One ofthe simpler forms, usedby Yanta and Gates(Ref. 201, results in a timeconstant expressed as

where k is the Cunningham constant (1.8 for air), and 9. is the mean freepath. From Eq. (ll.l7),the effect of increasing Knudsen number is to increasethe timeconstant. The effect becomes significant (18%) for a ratioof mean free pathto particle diameter of 0.1. Sincethe time constant increase can be significantfor low densityflows, Eq. (11.17) is more accuratefor investi- gatingparticle response in these cases. This, of course, assumes Stokes' law to be valid. Thus, atthis point, it isappropriate to discuss some of thelimitations on theuse of Stokes' law for particle drag in compressible flows.

A more detaileddrag coefficient expression for large-differential Mach and Reynolds Number isgiven byWalsh (Ref. 23) and is based onexperimental data. Walsh(Ref. 24) hascompared resultsobtained using Stokes' drag coefficient equation with morecomplex expressions which account for a widerange of differential Reynolds andMach numbers. Flow fieldstudies included normalshocks insupersonic flow and velocitygradients up to 365 sec". He concludesthat the use of Stokes'law generally yields conservative results compared to the more accuratedrag coefficient expressions, and it overpredictsthe velocity lag by less than 10% for particlediameters less than 5 microns and velocitygradients up to 333 sec'l. Thisoverprediction decreasesas the initial gas velocityincreases, the velocity gradient decreases, and theparticle size decreases.Considering other uncertainties, suchas the shape of the particle, the use of Stokes'drag coefficient is considered to beadequate for thepurpose of this study; the exception being lowdensity flows for which Eq. (11.17) is recommended.

The timeconstant defined by Eq. (11.14) is an extremelyuseful quantity for studyingthe particle lag problem. As in Eq. (I1.151, it canbe used to definethe frequency response of theparticle. In the time domain, the time constant is a measure of particle transient response. A step change in velocitylag, for example, is reduced to I/e of its initial valuein onetime 2 constant,I/e in two time constants, etc. The relaxation 1 engthcan also be determinedby the product of the time constant and the gas velocity. As before, in onerelaxation length the particle lag will reduce to l/e of its initial velocity,etc.

Accordingto Eq. (Il.lh),the fidelity with which particle motion repre- sentsfluid motion in a specificflow condition (test gas viscosity known) can beimproved by reducing the particle diameter and density.Diameter reduction isparticularly effective since the time constant increases according to diameter squared. As discussedpreviously, however, the minimum particlediameter is limited bythe minimum acceptablesignal-to-noise ratio.

An average particle density to 1 gm/cm3 is commonly used forflow seeding at conditionstypical of those encountered in transonic and supersonicwind tunnels. Seedingagents include dioctyl phthalate (DOP), silicone oil, and polystyrene latex (Ref. 19). A reviewof the several types of generators for introducing seedingparticles of controlled size is given by Mazumder, Blevins and Kirsch (Ref.25). For seeding high temperature gas flows,particles with hlaher melting pointsare necessary. Zirconium dioxide (Zr Oz), which has a meltingpoint above 3000 K and a densitv of 5.9 gm/cm, and aluminum oxide (A12 03)have been used forseeding high temperature gas flows, e.g., Ref. 26.

To demonstratethe effects of particle diameters, the frequency response of particles with a density of 1 gm/cm3 and diametersranging from 0.5 pm to Io pm is shown inFig. A.11.5 for Mach one flow and a stagnationtemperature of 40 Celsius(I04 OF). Theseresponse data clearlydemonstrate the desirability of using particles with a diameter of approximately 0.5 urn. The timeconstants for part fcles with a dens i ty of 1 gm/cm3 are shown in Fig. A.11.6 as a function

237 N. w 0 Mach No. 1.0 T, = 40 Celsius Pp = 1 g/cc

1.0

0.5

0.1

Figure A.11.5 EFFECT OF FARTICLEDIAMETER ON FREQUENCYPXSXINSE of particlediameter and with Mach numbers ranging from 0.5 to 3.0,. From thisFigure it may be seen that, for the assumed constantstagnation temperature,the time constant does not change significantlywith Machnumber. More effectwould be apparent at Mach 2.0 and 3.0 for lowdensity flow, in accord- ance with Eq. (11.17).

Various criteria may bechosen todefine the required degree of fldelity offrequency response. If it is assumed that adequatemeasurements can be made when theparticle lags the fluid motion by no more than 5%, i.e., V /V = Pf

0.0523 e- (11.18) f0.95 T' P where fo 95 is the upperfrequency limit withoutrealizing more than 5% attenuat ion of the particle velocity response to fluid mot ion.In Fig. A.11.7 thisfrequency limit is shown as a functionof particle diameter andMach number, againfor a particledensity of 1 gm/cm3. InFigs. A.11.6 and 7, it is demonstratedthat diameters of the order of 1 Urn orless are required for measurements up toapproximately 10kHz, and thatdiameters less than 0.5 pm arenecessary to extend accurate measurements to 100 kHz. These guidelines are obviously dependenton flow conditions and the definition of anaccept- able amount ofparticle lag, but they are in general agreement withthe con- clusionsof several investigations, basedupon bothexperimental and analytical results.Pedigo and Stevenson(Ref. 27) statethat for a particleto follow transonicflows with reasonable accuracy, the diameter should be lessthan 1 vm. Asher(Ref. 14), Mazumder, Hoyle and Kirsch(Ref. 19)and Seasholtz (Ref. 29) reachsimilar conclusions for rather widely varying flow conditions.

With regard to turbulence measurements in boundarylayers, Yanta (Ref. 30) foundthat mean velocity and turbulence intensity distribution measurements with both 1 pm and 5 pm dioctyl phthalate particles were invery close agreement.These measurements were made in aMach 3 flowchannel operating at one atmospherestagnation pressure. Yanta points out that inturbulent boundary layer flow, dominatedby vortexmotion, the particles are moving withthe vortices andmust respond to changes in velocity in a frame ofreference moving with theflow (Lagrangian). A hot'wire mustrespond to changes invelocity withrespect to a fixed(Eulerian) frame of reference. Asa consequence, larger particles canbe used forturbulence measurements withoutparticle lag effects.

239 h) & 0

lx T

lx 1 30 0.1 1.0 10 Parkj.de Diameter, p m Particle Diameter, Pm

Figure A. 11.6 TIMECONSTANT AS A FUNCTION Figure A. I I .7 MAXIMUMFREQUENCY FOR NO OF PARTICLEDIAMETER FOR MORE THAN 5% ATTENUATION VARIOUS MACH NUMBERS,PARTICLE OF SINUSOIDALVELOCITY DENSITY = 1 gm/cc VARIATIONS,PARTICLE DENSITY = 1 gm/cc I --

DataAnalysis and Accuracy

The datanormally obtained with a laservelocimeter include the mean velocity and theturbulence intensity in one, two or three components. By specialanalytical techniques, the spectrum of theturbulence can also be derived .

An individual measurement obtainedfrom the velocimeter, Ui, is taken to representthe passage of an individualparticle through the measuring volume (individualrealization) since this occu'rence istypical of high-speedflows and counter-typeprocessors. The individual measurement can deviatefrom the true mean velocity due to turbulence,noise andsystem resolution. The acquisition of a large number of measurements aretherefore necessary to improve the accuracy of both the mean velocity and turbulence measurements.

The mean velocityas determined from a large number of individual measurements is N c ui -= i =1 (11.19) U N 9

where ti is the number of values measured for a singletest condition,and U. is I a single velocity measurement.

The standard deviation of the velocity probability distribution function is

( 1 I. 20)

If the effects of broadening of theDoppler frequency spectrum due to finite sample length andphase reversals(tracker processor) or, in general, apparentvelocity fluctuations due: tonoise, etc., are negligibly small, thestandard velocity deviation is equal to the root-mean-squareturbulence velocity

cu = u' ; (I 1.21) - U' where U = U + u',andthe turbulence intensity is 7 . U

24 1 Correction of the measured turbulence velocity for the effects of Doppler spectral broadening (as occurs witha tracker processor) is discussed, for example, by George in (Ref. 31)- The correction techniques for this type of bias are based on the white noise or broad-band characteristics of the intensity modulation of the Doppler frequency; whereas, the turbulenceis band-limited.

Flack and Thompson(Ref. 32) have identified ten different biases which influence individual-realization, velocimetry measurementsof mean velocity and turbulence. Magnitudes of the individual biases range from less than 0.1% to 31% for the turbulence componentand from 0.1% to about 12% for the mean velocity component. The larger errors are associated withhigh turbulence intensities. The largest bias is due to the probability, in a turbulent flow, that more high velocity particles will be measured than low velocity particles.

This bias occurs because the individual measurements are not randomly distributed. If the scattering particles are uniformly distributed in the flow, the rate at which particles pass through the measuring volumeis weighted linearly with velocity, Fig. A.11.8. This form of statistical bias is discussed by Barnett and Bentley (Ref. 33) and by HcLaughl in and Tiederman (Ref. 34).

N ” (I I .22) U -N -1 c ui i=l

Barnett and Bentley derive the correction to the biased (arithmetic] meanin terms of the turbulence intensity as

N (I 1.23)

The correction to the arithmetic mean velocityfor velocity bias is therefore significant only when the turbulent intensityis large; a turbulent intensity of 10% would result in a 1% correction to the mean velocity. It should be emphasized that Eqs. 11.22 and 11.23 are based on a one dimensional analysis and may not be generally applicable toall flows.

242 Biased Average of Individual Measurements f f Individual Measurements The work to date on individual realization or velocity biasing has also been restrictedto constant-density, velocity-fluctuation flows suchas turbulent boundarylayers. This work assumes uniformdensity of scattering particles. Furtherstudy is needed of other flow fields where theunsteadiness is dominated byunsteady shocks, acoustic sources, etc. In such cases, significant density variationsoccur, and density and velocityfluctuations may be correlated.For empty-test-sectionsurveys of mean velocityin windtunnels, turbulence is sufficiently low that the contribution of velocity biasing to the total measurement error wouldappear to be minor.

Yanta(Ref. 9) and Yanta and Smith,(Ref. 35) discussthe problem of determiningthe sample size, N, requiredto establish the uncertainty in a measured value, such as mean velocity, asa functionof the turbulence intensity. For example,

(11.24)

When Z is the number ofstandard deviations corresponding to a desiredconfi- dence level (1.645, 1.96 and2.58 for 90, 95 and 93-percentconfidence 1 imits,

" respectively), AU/U is the error in the mean value, and ul/rthe turbulence intensity.

The confidencelimits for the standard deviation of the velocity prob- abilitydistribution (the rms turbulence,ul) is given by

N= 22 (11.25) ~(Au'/u')~ WhereAu' isthe error in magnitude of the turbulence, and Z and N areas previously defined.

Inthe case ofwind tunnel calibrations, the measurement periodshould be sufficientlylong to averagethe lowest frequency component oftunnel flow unsteadiness.Therefore, the required measurement period will extendfrom one to morethan 10 seconds,which may be of the same orderof magnitude as the timerequired to obtain the necessary number oflaser velocimeter measurements,

24 4

L dependingon the data rate. Some measurements inthe AEDC Tunnel 1-T required severalminutes per station.

An evaluation of the agreementbetween laservelocimeter and conventional, wind-tunnel-calibration measurements can be made from testsin several wind tunnels.Meyers, et. al., in Ref. (5) foundthat freestream velocities measured in the 4.9 m (16-foot)Langley Transonic Tunnel compared tothe tunnel ' t calibration measurements within 22%;where the uncertainty of the velocity based onthe tunnel calibration is stated to be +I%. Althoughthe particles used to seed theflow ranged in size from 10 to 15 microns,the authors did notconsider particle lag to be a problem inthe test section; although significant lag was presentin the flow accelerating section of the tunnel.

Measurements inthe 0.3 m (I-foot) AEDC Tunnel 1T inthe Machnumber rangefrom 0.6 to 1.5 arereported by Smith etal. in Ref. (8). The error estimatefor the conventional calibration data rangesfrom approximately two percentat Mach0.6 to 0.5 percentat Mach 1.5. flultiple-pointaxial surveys of centerline calibration data, basedon pressuremeasurements, were obtained only at Mach numbers 0.6 and 0.8, with a single-point,pressure-calibration, velocity measurement (atthe same point onthe tunnel centerline) of 2.5 to 2.7 percent.Comparison ofthe axial distributions obtained byboth techniques at M = 0.6 showed thebest agreement with an averagedifference of about 1.4 percent for a mean velocity of 216rnps, and theuncertainty bands overlapped. At Mach 0.8, thedifferences ranged from about 0.3 to 2.0 percentat a mean velocity of approximately 280 mps. Interestingly,all of thevelocities meas- uredwith the laser velocimeter were higher than the velocities determined from pressure and temperature measurements. No explanation was offeredfor the difference,but obviously particle lag was not a factor. All measurements were made withnaturally-present, light-scattering particles in the flow.

Flow angularity measurements were also made with the 2-component laser velocimeter, which demonstrated the ability to make angularity measurements withdeviations ranging from 2.015degree to 2.25 degrees. The deviationsare a functionof the rms deviationsof the component velocities. No conventional angularity data werepresented for comparison. This test demonstrated the capability for measuring the flaw angularity andan average of the flow fluctuations in both magnitude and direction.

The two testsdescribed should be regarded as operational feasibility demonstrationsonly and shouldnot be regardedas the best accuracy currently available.Further, since the laser velocimeter data are compared to conventional calibration data with a stated accuracy of approximately 50.5 to -+2.0 percent, an absoluteevaluation of the laser velocimeter accuracy cannot be made fromthese results.

Measurementsby Johnson(Ref. 36), Johnson and Rose (Ref. 371, Yanta and Lee (Ref. 38) and Boutier and Lefevre(Ref. 391, provideadditional evaluations ofthe agreementbetween flow velocity measurements bythe laser velocimeter and by conventionalPitot and static probes. These freestreamvelocity measurements are somewhat limited in scope sincethey are obtained from boundarylayer velocity profile measurements; nevertheless,agreement within 0.5% is demonstrated.

The laservelocimeter yields a direct measurement of flow velocity, while the Mach number is requiredfor calibration of windtunnels. A second measurement, thestagnation temperature, is thereforerequired. Using the energy equation,the local static temperature canthen .be determinedfrom the measured velocity

Where Tw isthe test section temperature corresponding to Vm, To isthe stagnationtemperature, C isthe specific heat at constant pressure, R isthe P gas constant, and y isthe ratio of specific heats. The local Mach number can then be determined from

VW M= am

"m (I I .28)

246 Sensitivity coefficients for measurements of both velocity and stagnation temperatureare shown in Fig. A.II.9 for a stagnationtemperature of 40° Celsius. From thesedata, measurement of Mach numbers to anaccuracy of +_O.OOl requires a velocity measurement accuracyof about 0.1% in the transonic speed range.

Thisaccuracy isnot believed to be withinthe state-of-the-art at present. v i Conclusions

The advantages and disadvantages of thelaser velocimeter may besummarized as follows:

Advantages

1. No probeor other device introduced into the flow, i .e ., non- pertubing. 2. Providesdirect, linear measurement ofveloc i ty and ve locity fluctua- ti ns;no calibrationrequired. 3. Ab lity to measure reversingflows. 4. Ab lity to separate mean and fluctuatingvelocities into components. 5. Po nt measurements can be approached by propercontrol of measuring vo ume. 6. Ca be realized as an inherentlydigital instrument.

-D i sadvantages 1. Complex, expensiveequipment required 2. Measurement ofhigh-velocity flows in large tunnels with an air test medium presentsspecial problems with regard to single-to-noise ratio, frequencyresponse, and sensitivity of equipment totunnel vibration, temperature,etc. 3. Furtherdevelopment needed to improvesignal processor, particularly withregard to data validation features, rejection of largeparticles, etc. 4. Light-scattering particles of size needed for good signal may not follow flow in regions of rapid velocity change. 5. Signal-to-noise ratio may precludeaccurate turbulence intensity measurements at lowlevels, i.e., 0.1 to 1.0 percent. 6. Current accuracy attainable is not as good as with conventional techniques. 7. Takes excessivetime to make thesurveys required for wind tunnel calibration. 247 .WO

.018

,016 To = 3U°K I

.014

.012

.010 - -" 1/OK aTo .008

.006 .004

.002

0 0 1 2 3 4 Mach No. Figure A. I I -9 SENSITIVITY COEFFICIENTS FOR DETERMINATION OF MACH NUMBER FROM VELOCITY AND STAGNATION 248 TEMPERATURE MEASUREMENTS Nomenclature

aaD streamfree acoustic velocity specific heat at constant pressure =P initial diameter of laser beam DO d0 diameter of laser beams atthe focal point of the transmitting lens d particle. diameter P F focallength of tkansmitting lens

body forceson particle (Eq. 11.11) Fe Dopp 1 er frequency Hz fd , frequency at which particlemotion is attenuated 5% relativeto steady fo. 95 statesinusoidal fluid motion m

Knudsen number k Cunningham constant (Eq. I I. 17)

R mean free path

length of measuringvolume llV M Machnumber

N sample size

Nf number of fringes in measu'r ing volume r number ofpulses counted in the high.register of a counterprocessor NH number of pulsescounted in the register of a counterprocessor NL low R constant gas

S Lap1operaace tor

stagnationtemperature TO

249 time constant def inlng part icle response to variations in fluid flow : TP velocity, seconds .. To3 freestreamtemperature

t time,seconds

t' dumny variable (Eq. 11.11)

U velocity component nom1 to fringe pattern and to bisector of theangle formed by two Intersectinglaser beams - U mean velocity - mean velocity, corrected for velocity bias uC

U' rms turbulence velocity

Au I error in magnitude of turbulence velocity

V gas flow velocity 9 V particle velocity P "" f fees t ream ve 1oc ity

W width of measuring velocity V

2 number of standard deviations corresponding to a desired confidence level (Eq. I I .24). -Greek Y ratio of speciffc heats

A beam separation distance at the transmitting optics

fringespacing in measuring volume bf

0 angle between laser beams at the.ir intersectionto formthe measuringvolume

x wavelength of laser light

viscosity of gas

density of gas

density of particle pP

a standard deviation of velocity probability-distribution-function

250 .-; phase angle, degrees .. ..

angular velocity, radians per second

. .... I 25 1

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22. Hoppel, T. and Brenner, H: Low Reynolds tJumber Hydrodynamics with SpecialApplications to Particulate Media, pp 50-51, Prentice-Hall, EnciewiSod Cliffs, N. J., 1965. 23 - Walsh, M. J., "Drag CoeffIcientEquations for Small Particlesin High Speed Flows," AIAA Journal, V 13, No. 11, Flov., 1575.

24. Walsh, M. J.:"Influence of Drag CoefficientEquations on Particle MotionCalculations," Proceedings of the tlinnesota Symposium on Laser Anemometry, Universityof Minnesota, 1975.

25. Mazumder, M. K.; Blevins, C. W. and Kirsch, K. J.: I'Wind TunnelFlow Seeding for LaserVelocImeter Applications," Proceedings of the Minnesota Symposium asLaser Anemmetry, University of Minnesota, 1975.

26. Self, S. A., "Boundary Layer Measurements in HighVelocity High Tempera- ture MHD ChannelFlows," Proceedings ofthe Second International Workshop on LaserDoppler Velocimetry, Purdue UnIversity, 1974.

27 Pedfgo, M. K. and Stevenson, W. H., "The Design of a LaserDoppler Veloci- meter for TransonicFlows,'' Purdue University, Prepared for Army Missile Cmand, AD-774 302, October 1373.

28. Haertig, J., InformalPresentation, Proceedings of the Second International Workshop onLaser Doppler Velocimetry, Purdue University, 1974.

253 29 Seasholtz, R. G., "Laser Doppler Velocimeter Measurementsin a Turbine Stator Cascade Facility," Proceedings of the Second International Work- shop on Laser Doppler Velocimetry, Purdue University,1974.

30 Yanta, W. J., "Laser Doppler Velocimeter Measurementsof Turbulence Properties of a Mach 3 Turbulent Boundary Layer,'' Proceedings of the Second International Workshop on Laser Velocimetry, Purdue University, 1974. 31. George, W. K., "The Measurement of Turbulence Intensities Using Real-Time Laser Doppler Velocimetry," Proceedingsof the Second International Work- shop on Laser Velocimetry, Purdue University,1974.

32. Flack, R. D. and Thompson, H. D., "The LVD's Potential in Understanding Turbulent Structure,'' Proceedings of the Minnesota Symposiumon Laser Anemometry, University of Minnesota, 1975. 33. Barnett, D. 0. and Bent ley, H. T., "Statistical Bias of Individual Realization Laser Velocimeters," Proceedings of the Second International Workshop on Laser Velocimetry, Purdue University, 1974.

34. McLaughlin, D. K. and T iederman, W. G., "Biasing Correction for Individual Realization of Laser Anemometer Measurements in Turbulent Flows,ll The physics of Fluids, Vol. 16, 1973. 35 Yanta, W. J. and Smith, R. A., "Measurement of Turbulence Transport Proper- ties with a Laser Doppler Velocimeter,''AlAA Paper No. 73-169, Jan. 1973. 36. Johnson, D. A., "Turbulence Measurements in a Mach 2.9 Boundary Layer Using Laser Velocimetry," AIAA Journal V 12 No. 5, pp 711-714, May 1974. 37. Johnson, D. A. and Rose, GI. C., "Turbulence Measurements in a Transonic Boundary Layer and Free-Shear Flow Using Laser Velocimetry and Hot-wire Anemometry Techniques," AIAA 9th Fluid and Plasma Dynamics Conference, Paper No. 76-399, July 1976.

38. Yanta, W. J. and Lee, R. E., "Measurements of Mach 3 Turbulence Transport Properties on a Nozzle Wall ,Ii AlAA Journal, V.14, No. 6, pp 725-729, June 1976.

39. Boutier. H. and Lefevre, J.: "Some Applications of Laser Anemometry in Wind-Tunnels," The Accuracy of Flow MeasurementsBy Laser Doppler Methods, " Proceedings of LDA Symposium - Copenhagen, 1975.

254 APPENDIX 111 EFFECTS OF VIBRATION OF A CYLINDRICAL PROBE ON STATIC PRESSURE MEASUREMENTS

I In the process of collecting material for writing about mean static pressure measurements, thequestion arose as to how measured datawould be affectedby vibration of a probe. The problem of error in measured pressure causedby unsteady cross-flow has been previouslyconsidered by Siddon (Ref. 1). The followingdiscussion is taken from this reference.

Some insightinto the problem of a probe in an unsteadycross-flow can be obtained by use of an idealizedflow model thatignores viscosity. Consider a cylindricalpressure probe of diameter d, subjected to a uniform, unsteadycross-velocity Vn(t), see Fig. A.III.1. Any couplingeffect of the axialvelocity U(t) is neglected. Assuming theflow to be irrotational,the appropriatepotential function is:

d2 4 = Vn (r + r)COS n.

The unsteadyform of the Bernou lli equat iongives:

1 a4 p (r, t) Pt(t) = "p (vn2 - v - v 2, + p n, - 2 r n at where,

P (t)is the pressure which would have occurredat r = 0 inthe absence of t theprobe (i.e., the 'true' pressure). At thesurface of the probe (r = d/2) , thepressure equation becomes:

P (Tl,t) - Pt(d = -1 (1 - 4 Sin2 n) + ptnd Cos n - (111.1) 2 P 'n

The first termon the right hand sideis recognized as thepressure distribu- tionfor steady potential flow. The second termarises from unsteadiness. For a pressureprobe which registers the exact circumferential average of P(n,t), (e.g.,by means of a circumferentialslit) the error will be: P,(t) - Pt(t) = - $P vn 2 Y

.where P,(t) = measured unsteady staticpressure.

Inthis ideal situation the part of the pressure distribution associated withthe acceleration term Vn does not contribute to the error.

A realprobe will nottake an exactaverage over P(rl); therefore, an additionalerror proportional to inmay arise.

P,(t) - Pt(t) = - 1 p Vn2 + K p ind

Roughlyspeaking, the coefficient K (<1)represents the fractional inaccuracy ofthe average over P(n). If we regard Vn as sinusoidal,the error becomes \i n increasinglyimportant with frequency wVn). Nevertheless,for a (in - 0.318 cm (1/8in.) diameter probe with an averaginginaccuracy of and V 5% n 3.05 m/sec (10 ft/sec), K p ind is lessthan 5% of the V errorat 100 Hz. n Based onresponses tothe questionnaire, many investigators havelocated a single row of orifices on eitherthe top or side of a longpipe for tunnel calibrations.Byusing Eq. (Ill-I) with 11 = 0 (orifices on eithertop or bottom) andassuming the pipe oscillates sinusoidally, one can estimatethe error in measured, mean staticpressure, i.e.,

1 2 P(0, t) - P (t) = T p Vn (t) + p \ip(t)d. (I 11.2) t

Taking a timeaverage over one cycleresults in

- - T T P,(O) - P = - 1 V 2(t) dt + in(t)dt, (I 11.3) t 2TOn 0

where T = period of oscillation.

Before we canproceed any further, a characteristicfrequency and amplitude must be assumed. For calculationpurposes we here assumea frequency of 100Hz.

256 Since we have previously established thatit is desirable to measure mean static pressure to within 6.89 N/m2 (0.001 psi), the amplitude of pipe - = 2 oscillation wi 11 be calculated for the caseP m - Pt 6.89 N/m . Hence, if the pipe oscillation is represented by

Displacemetjt E D = A sin wt;

then '8

'. Vn = b = Aw coswt 2 in = D = -AU srnwt.

Substituting into Eq. (111.3), we have

22 2wt) PA w a*. 6.89 = - dt - (111.4)

It may be noted from Eq. (1 11-41 that the required amplitude increases as frequency decreases. For the particular case ofw = 21rf = 200~and 2 4 52 24 p = 1.71 N sec2/m4 (0.00332 lbf sec /ft ) [Po = 2.41 x 10 N/m (35 psia), T = 311°K (560"R), M = 11, the amplitude required for significant pressure 0 error is

Therefore, a static pressure survey pipe witha single row of orifices, located on either the top or bottom, would have tooscillate at a frequency of 100 Hz and an amplitude of 0.639 cm in order to cause the measured pressure 2 to be 6.89 N/m (0.001 psi) too high. Figure A. I I I. 1 PFESSURE DISTRIBWION ON A CIRCULAR CYLINDER IN CRoSSFIX)W, Ref. 1 I-

Goingthrough the same procedurewith n = 90" (orifices on side of pipe), an amplitudeof 0.368 cm (0.145 in.)is predicted to cause the measured, 2 staticpressure to be low by 6.89 N/m (0.001 psi).Since large tension or compressionloads are usually placed on staticpressure survey pipes to minimize sag, it appears unlikelythat pipe vibration is a significantsource of error.Although this conclusion is based onan inviscid,incompressible analysis, it isconsidered to be conservative(i.e., gives a lower bound estimateof A) because of the neglect of pneumatic damping intubing which connect orifices with pressure transducers.

A.III REFERENCES

1. Siddon, T. E., "On the Response of PressureMeasuring Instrumentation in Unsteady Flow," UTlAS Report No. 136, AD 682296, January 1969. APPENOIX IV: FACILITIES RESPONDING TO OUESTIONNAIRE TABLE I

Organlratlon Re/m x Productlon Testlng Faci 1i ty & ionLocat M @ H - 1.0 Began 4' Trisonlc 24.6 60.7 B 1owdown - 26" Transonic 0.66 m Octagon Farm1nadale.N.Y. 31 - 92 1957 16-Inch ETH Zurich, 0 0.95 1935 - Con t sq. (First. Cont. S.W.1 .Supersonic W Swltz. I. 2-2 (2.5) 7 - 7.5 . 10.40 m I NLR Amsterdam, 2 Cont. x m High Speed WT Hol land - .37 3.5 - 19 1.6 2.0 1959 (H = 1.2) 2 Continuous nsert.5-1.2 Bd 10.27 x 0.27 m 1960 Supersonic WT II 1.2 - 5.8 30 - 50 Run Tlme I II Pilot WT 0 - 1.0 14.1 Cont. 10.55 x 0.42 m21I 1956 vmuI e 2 II 1.2 x 1.2 Supersonic WT 1.2 - 4.0 21 - 54 Bd m n.s Y 1.7 nl2 1963 Supersonlc U.S. Army (M = 1.25) 2 Tunnel No. 1 Aberdeen, Md. 1.25 - 5.0 h - 71 Con t . 0.38 x 0.33 m 1956 USAF-FDL10.3 1.2 0.38 x 0.38 m: Trlsonlc G.F. - 3.3 27 Cont. WPAB, Ohio 1.5 - 4.76 - 0.61 x 0.61 m N.A. USAF ARL Mach 3 Hi-Re - 2.97 3.0: Bd WPAFB , Ohio - 33 - 359 . 1970 SandiaAlbuquer- 0.305 x2 12" Tr ison ic 2.5 16.4 Bd que, N.M. 0.5 - 0.305 m 1956 NASA Lewis R.C. 8' x 6' SwT o.36 2.1 15 Cleveland,Ohio -

II (M = 2.0) IO' x 10' SWT 2.0 - 3.5 1 - 11 Aero. Res. Unit 18- Inch Pretoria, S.Afr. o*6 - 4*2 12 - 30 . .. .- .. . . APPENDIX IV: FACILITIES RESPONDING TOOUESTIONNAIRE TABLE I

Organlzatlon I Re/m x Production Testlng Faci I i tv I Location IM @ M = 1.0 I Tvne,. Cross-Sect ion Began I .2 m HSWT BAC 2 Preston, Eng. 0.4 - 4.0 21 - 74 1.22 x 1.22 m 1960 2 30" x 16" NAE 0.4 - 2.0 15.1 Suction 0.76 x 0.41 m 1952 SuctlanW.T.da , i sonic 2 Tr II 1.52 x 1.52 m 1964 t3lowdown Wf 0.1 - 4.4 idBd II 2 HRN-PDT 0.15 - 0.95 - Bd 0.38 x 1.52 m 1969 l/3 m Trans. NASA Langley Res 0.05 1.2 13.45 341 Cont. 0.34 m Octagon 1974 Cryoqenic Tun. Ctr.Hampton, Va. - - High Speed (M = 0.9) II On2 - 2 7' x IO' Tun. 11.8 - 16.1 Cont. 2.0 x 2.92 m 1946 2 1.37 x 1.37 m 1948 6" x 28" 2 I' 0.3 - 1.20 32.8 - 98.4 Bd 15.2 x 72.4cm 1974 Transonic Wl , Transon ic (1.05-2.53) air '4.88 m x 4.88m II 1960 DynamicsTun. 0 - 1.2 (8.04-19.7)Freon Cont. i 0.61 m corner f i 1 lets Unitary II 1.47 - 2.86 2.6 - 28 (k1.5) Plan Wl - 4.63 1.6 - 20("2.3) 2.29 Conto 1.22 x 1.22In2 1955 I6l Transonic 11 0.2 - 1.3 I2 - 13.8 Con t. A.72 m Oct. 1950 TI I I I I 8' Transonic 2 II 0.2 - 1.3 1.3 15.3 Cont. '2.16 x 2.16 rn 1952 Pressure Tun. - Northrop 0.20 - 1.3 2' Trisonic 4 Hawthorne, Ca. 11.5 - 3.0 7 0.28 - 49 Bd 0.61 m sq. 1962 High Speed Picatinny Arsen- 0.2 - 0.76 (M = 0.6) Subsonlc WT al, Dover, N.J. 11.12 12.96 Induct ion '0.61 m Diarn. N.A. ". - .. . .. APPENDIX IV: FACILITIES RESPONDING TO OUESTIONNAIRE TABLE I

Productlon Testing TY Pe Cross-Section Began 2 Bd 0.42 x 0.41 m N.A. 2 Bd 0.71 x 0.51 m 1973

7 lndraft 10.2 x 12.7 cm" 1956 Bd 10.66 m sq. I 1965 kcDonnel1-Doug 1as Polysonic \-IT UIS. Mo. 1 Tunnel A ,G;:z ryz; Cont. 11.02 m sq. I 1958 I ac.AEDC/ARO I I I Tunnel D II

~ ~~ ~~ Ludwi eg 2 Pilot HlRT Tu beTu !18*6x23*2 cm 1N.A. (Research Only) Transon i c echn ion 1 Induction 0.8 x 0.6 m2 1968

30 cm WT i I Bd 0.30 m Sa. 1960 40 x 50 cm SWl I 2 1 0.4 x 0.5 m 1968 ockwel I Int. ?3kn i c 1 Sequndo. Ca. 2.13 I Bd m SQ. 1958

Transonic W bedford, Eng. 24' x 23' I II SWr I Cont. 0.69x0.76 m 2 1959 " . . , .. . " .. . . APPENnlX IV: FACILITIES RESPONDING TOOUESTIONNAIRE TABLE I

Organization Re/m x Production Testing Faci 1 ity & Location M @ M = 1.0 Type Cross-Section Began

Bd 0.305x0.406m2 1964 1956

WRE Salisbury, 0.4 1.0 - 3.3 - o.38 2 Cont. 0.38x0.38 m 1957 15-lnch SWT s. Aust. 1.4 - 2.8 I1 2.8 - 5.0 (M = 2.8) 2 s3 Bd Wl 16.4 - 65.6 Bd 17.8x15.2 cm 1966 ARLMe1 bourne, 2 3.28 - 6.56 Cont. 0.81x0.53 m 1957 2 VolvoTrollhatten o,5 - 1.5 23 - b9 Bd 0.5x 0.5 m 1952 I TransonwT i c WT ISweden Aust. 1 0.4 - 1.4 2 II 1.4 3.2 (M = 1.5) Bd 0.5 x0.5 m 1962 WT9 - 29 - 55 Boe ing Boe ing (M = 1.2) - 4*0 2 Supersonic WF. Seattle, Wash. 19.7 - 45.9 Bd 1.22x1.22 rn 1957 2 2D-TWT I1 0.2 - 1.25 26.25 - 77.1 Bd 0.305x0.91 m 1965 2.44 m x 3.668 Boe ing II 0. - 1.11 Cont. 1944 Tranyznlc WT 0.61 m corner Lockheed tr I let Lockheed 0.2 - 5.0 14.8 - 67.9 Bd 2 Triqn,nic Saugus, Calif. 1.22x1.22 m 19h0 Hawker-Siddeley 2 HS - TWT Hatfield.Enaland 0.5 - 1.1 10.2 - 11.5 Cont. 0.76x0.61 m 1954 Cont. 6' low 8' Transonic Calspan 1.3 - 18 0 - 1-34 Re; Interm. 2.44 m Sq. 1956 w-r Buffalo, N.Y. 28 - 42 - Variable Dia, zg'"0.81m-l.52m 1966 Tube, .. APPENDIX IV: FACILITIES RESPONDING TO OUESTIONNAIRE TABLE I

Organization Re/rn x Testing Production Faci 1 i ty @ M = 1.0 Type Cross-SectionBegan (M = 20'' SGrr 1.5) 2 0.47 - 23.6Cont. 3.51x0.46 rn 1950 VoughtCorp. High Speed Wl Dal las. Texas 10.5 - 5.0 Aerodyn. WT - Pro ulsion LIT Fac 1T AED!/ARO Tulla- 10.2 - 1.5 17 I Cont. 10.305 m Sq. I 1953 I ,ho-. Aerodyn. WT - II 0.1 - 1.3, 4T & 1.3 - 22.5 Cont. 1.22 rn Sq. 1968 1.6 7.0 I Propu 1 s ion Vf -16T It 0.2 - 1.6 1.3 - 23 Cont. 4.88 rn sq. 1957 Propulsion Wl II (M = 2.0) 16s 1.5 - 4.75 I961 1.3 - 8.4 Cont. 4-tuLa3.q- 14" Trisonic NASA Marshall SFC WT Huntsvi 1 le,Ala. Bd 0.356 rn Sq. 1956 12' Pressure NASA Ames RC 0.- 0.98 2.1 - 5.4 Cont. 3.44 m Sq. wr Moffett Field,Ca. I I I 1946 I 14' Transonic 2 w-r 12 - 16.5 Cont. 4.11x4.18 m 1956 11 I Transonic LIT 0.4 - 1.4 6.7 - 32 Cont. 3.35 m Sq. 1956

2' Transonic 1II w-r 0.2 - 1.4 12 - 23.5 Cont. 0.61 rn Sq. 1951 Injector-Driw In 0 - 1.0, II 16.7 - 167 Con t 15.24 crn Dia. Pilot Model Transonic Wl 1.2. 1.4 9' x Super 7' II (M = 2.0) sonic WT 1.55 - 2.5 7 - 71.7 Cont. 2.74 x 2.13 rn2 1956

8' x 7' Super I1 (M = 2.5) sonic WT 2.45 - 3.5 2.44 x 2.13 rn 2 3 - . .. 17,...... Con t .. - .. .. . - 1956 APPENDIX IV: FACILITIES RESPONDING TO OUESTIONNAIRE TABLE I

Organitat ion Faci 1 ity & Locat ion

61 x 61 Super-NASAAmes RC Mof- sonic WT fettField, Ca.

1' x 3' Super- II CWT Supersonic Uaval Surf.Weapon Tunnel No. 1 Zntr.,Silver Spg. Supersonic Tunnel No. 2 II

Boundary Layer II 1

NSWC Hyper- II sonic Tunnel Aero.Res.lnst. Tunnel FFA-Sb (FFA)Stockholrn,Sb Transon ic Tun- nel FFA-HT II I I Tr isonic Tonne; =FA-TVM 500 II

Tunnel FFA-S5 II

8 X 6-Ft TWT II Royal Airc.Est. x 4-Ft* '* Bedford.Enqland

DFVLR x TWT Gottingen,W.Ger. APPENDIX IV: FACILITI€S RESPONDING TO OUESTIONNAIRE TABLE I

Production Testing :ross-Section Began:ross-Section I 0.6 rn Sq. I 1966 0.5 rn Sq.(M>I 0.6x0.34 rn2

0.25-0.50 m

0.81 m Dla. 1970 APPENO I X IV (Cont Id) TABLE II: TESTSECTION CHARACTERISTICS

Venting of Plenum Facll fty Cross-Section Porosltv WallAnqle Chamber 1.27cm, 90 SW 20% sw NAL (India) 1.22 m Sq. Perforated loto -2O EJector Flaps 300 Top & Bot. 6%TsB GAC (NY) 0.66 m Octagon Slotted I 12% & 6% 0. oso -EjectorFlaps

~ ~~ ETH (Swi tz) 0.4 m Sq. Sol id

I NLR (HST 2 Slotted 1.6 x2 m 5 cm,40 cm 12% T E B 0.13' FixedEjector Slots .Hal land) - (T & B) 2 NLR (CSST) 0.27 x.27 m Solid I I 2 Slotted , NLR (PT) 0.55 x 0.42 m lO%T&B 0.22O FlxedEjector Slots (T & B) 0.525 cm, 5.25 cm ' 2 NLR (SST) 1.2 x 1.2 m U.S. Army 2 0.38x 0.33 m Sol id (SSTl) WPAFB (TGF) 0.38 x 0.38 mL ~ 4 - 12% O0 'EJectorFlaps WPAFB (M3HR) 20.3 x 20.3cm2 Solid I

II t-Sand ia (TWT) 0.305 x 0.305 m2 Perforated [ 0.318 cm,30' 6% lo EJectorFlaps 2.44 x 1.83 m2 -"--Perforated 2.54 cm,30' ' 6% O0 Aux i 1 iary Pumps 3.05 x 3.05 m2 Sol id I I 2 ~ ARU (SWT) 0.45x 0.'45 m Perforated 0'.'6O T E B lAuxiliary Pumps APPENDIX IV (Cont'd) TABLE II: TEST SECTION CHARACTERISTICS

Venting of Plenum Fac i Wall Anqle Chamber

BAC (HSWT) 1.22 x 1.22 m Perforated 1.6 cm,90' 19% O0 EjectorFlaps lJAE (Su WT) 0.76 x 0.41 m Slotted 0.58 cm, 4.7 cm lo -0.5' to-. 25' Ejector F 1 aps

2 Perforated NAE (2DT) 0.38 X 1-52 m 1.27cm, 90' 20.5% Vent Inq to Atm. (T B) TEB O0 9RA (TWT) 2.44 x 2.74 m2 Perforated cm,1.27 90' 22.5% O0 9ux i 1 iary Pumps

4fM (SWT) 0.305 x 0.406 m2 Solid 1 20.3 x 22.9 cm2 1 Perforated 0.185 cm,90' I 22% O0 EjectorFlaps

O0 Ejector F 1 aps

0.5' 2 Volvo (WTl) 0.5 x 0.5 m Slotted (T G B) ? 4%T&B 0.21O ElectorFlaps I I I Volvo (Wr9) 1 0.5 x 0.5 m2 I Sol id Boeing (SWT) 1 1.22 x 1.22 m2 1 Sol id - Boe 1 ng 2 Perforated 1.03 cm, 90° 0.305 X 0.91 m (2D-TWT) (T & B) 34.1% T&B O0 EjectorFlaps APPENDIX IV (Cont'd) TABLE II: TESTSECTION CHARACTERISTICS

HoleSize/Angle or Venting of Plenum Faci 1 i ty Cross-Section Wall Type Slot Width/Soaciu PorQsjtv WallAnqle Chamber 2.44 x 3*66 Slotted 7.45 cm,70.41 cm 11% - 3.5% (TWT) O0 Ejector Flaps p;fleTscorner tJithinsertsTdB 47.31 cm SW cm* 'O0 22% '-0.75' Auxiliary Pumps Lockheed (TWT) 1.22x 1.22 m2 Perforated '*'' Hawker Slotted T&B 2.54 cm, 10.80 cm Perf. InsertsO.10 cm, goo 3% -0.17' EjectorFlaps 0.067' SW 6.35 cm, 148.6 cm 4.4% 0.167OT~B EjectorFlaps

LRC (UPWT) 1.22 x 1.22 m Sol id

LRC (16' TT) 4.72 m Octagon Slotted .62 cm, 1.95 rn 3.9% 0' - 0.75' Auxi I iary Pumps - p 2.16 x 2.16 mz Slotted 3.18 cm, 43.20 cm T&E 0,083' SW LRC (8' TPT) 3-6% FixedEjector Slots no - o-,,h~TFR Tapered fromzero EjectorFlaps or LRC(1/3mTCT) 0.34 m Octagon Slotted to 1.727 cm 14 cm 0-125; 0.083' Vent Ins to Atm. Auxiliary Pumps & LRC(HST) 2.0 x 2.92 m Slotted 4.60 cm, 73.03 cm 4.8% TF.R O0 EjectorFlaps

LRC (4' SPT) I. 37 x 1.37 m Solid

LRC (6" TWT) 15.2 x 72.4 cm2 Slotted 3.476 cm, 3.81 cm 12.5% O0 FixedSlots TL R Slotted 3.318 x 2.54 cm, 0.61 m Sq. ' NC (2l TWT) Holes 1.91 cm 10% -2/3' to +lo Ejector Flaps PA(HSSWT) 10.61 m Dia. I Solid I I IO0 10' I I Va 1 ve Cont ro1 s PA (''I' Tbn) 0.42 x 0.41 m* Perforated 0.48 cm,30' 0.25 - 8% NA Diffuser Pumping h, 4 0 APPENDIX IV (Cont Id) TABLE II: TEST SECTION CHARACTERISTICS

Venting of Plenum Faci 1 i ty Cross-Section Wall Type Mal 1 Anqle C harnber 2 L-G (CFF) 0.71 x 0.51 m Perforated 0.635 cm, 30' 0- 10% 0.25' T&B EjectorFlaps

UM (4" x 5" 10.2 x 12.7 cm Sol id

1Perforated 0.635 cm, 90° 23 oo to +.3O EjectorFlaps MD (PSWT) 1.22 m Sq. Perforated 0.953 cm, 90' 25 -0.75' to Oo Auxi 1 iary Pumps I I - Sol id

AEDC (Tunnel ol 0.305 m Sq. Sol id AEDC (Pi lot 2 Ejector Flaps and/ x 23.2 cm Perforated 0- 10% HIRT) 18.6 .305 cm, 30' O0 or Exhaust to Atm. Techn ion 2 0.8 x 0.6 rn Perforated .635 cm, YOo 21% TbB +0.5' EjectorFlaps (TIWT) 1 I I Techn ion n0.3 m Sq. Sol id Techn ion 2 &haAm" 0.4 x 0.5 m Sol id RI (TU) 2.13 m Sq. Perforated 0.635 cm, 90° 19.7% O0 Ejector F I.aps Auxiliary Pumps E Perforated to 0*670 1.27 cm, 90° 22.7% O0 Venting to Atm.

Perforated 2 JPL (SWT) 0.51 x 0.46 m Sol id APPENDIX IV (Cont'd) TABLE It: TESTSECTION CHARACTERISTICS

HoleSize/Angle 01 Venting of Plenum Facl 1 i ty Cross-Section Wall Type Slot Width/SDacint Porositv WallAnqle Chamber VC (HSWT) 1.22 m Sq. Perforated 1.04cm, 90' 22.5% 80.33: T&B Ejector Flaps Q-h' cw 6% and to 0.5' EjectorFlaps & AEDC (AWT- 1T) 3.305 m, Sq. Perforated 0.318 cm,30' -0.67' 0 -10% TEB Steam Eiector SYS. 1 to 0.55' EjectorFlaps & AEDC (Am-4T) 1.22 m Sq. Perforated 1.27cm, 30' 0 - 10% - I T&B Aux I llary Pumps EjectorFlaps E AEDC (PWT- 16T) 4.88 m Sq. - Perforated I 1.905cm, 30' 6% -1' to O.jOT&B Auxi 1 iary Pumps AEDC (PWT-16s) 4.88 rn Sq. Sol id NASA Marshall 0.356 Sq. 03.4% Auxi iary Pumps (14'' TWT) m 1

ARC (12' PWT) 3.44 m Sq. Sol id 2 ARC (14' TWT) 4.11 x 4.18 m 'lottedbl.74 cm,26.4cm ).18O T&B EjectorFlaps withinsert 5.6% Ejector Flaps & ARC (1 I' TWT) 3.35 m Sq. 5.8% 1.19O sw Auxiliary Pumps ARC (2' TWT) 0.61 m Sq. 22% with Throttle bo to 0.35' Ejector Flaps -Bars - ARC ( I -D TWT) 0.152 m Dia. Sol id 2 4RC (7 'x7 ' SW) 2.74 x 2.13 m

2 ARC(8Ix7' SWT) 2.44 x2.13 m Sol id

N ARC (6'x6' SWT 'lottedT&B11.03 cm,28 cm 5.1% O0 Auxi 1 iary Pumps v, 1.83 m Sq. with insert N 4 N APPEND I X I V (Cont Id) TABLE II: TESTSECT ION CHARACTERIST ICs

Venting of Plenum Faci 11ty Cross-Section

I NSWC (ST #I) I 0.4 m Sq. I NSWC(ST #2) I 0.4 m Sq. TransonicJozzle used in ST.I#l isavail Solid 1 0.86' NSWC (HyT) 0.41 x 0.41 m Sol id I FFA-Sb 0.92 x 0.90 m Slotted I 2.5 cm, 30 cm I 6% TES 0.15' FlapsEjector - 0.89 x 0.89 mL FFA-HT Octaaona 1 Slotted 3.4 cm, 37 cm 9.2% O0 ' EjectorFlaps d

FFA-TVM 500 0.5 m Sq. Perforated 0.5 cm,30' 6% -6' to 00 Auxi 1 iary Pumps

FFA-SS 0.46 x 0.48 m2 Slotted 0.21 cm, 5.3 cm 4% T&B 0.15' Ejector Flaps

RAE(8Ix6' TWT) 2.44 x 1.83 m2 -0.2' to 0.45' Auxiliary Pumps RAE(3Ix4' SWr)l 0.91 x 1.22 m2 Sol id - 1 Slotted 0.99 cm, 8.9 cm 9.75% -0.4' to 0.9 Ejector Flaps vAuxi iary Pumps 1 Perforatedl 1 cm, 30' 1 6% 0' to 0.5' DFVLR (TT Exhausted to W1) 0.6 m Sq. - Perforated 0.6 cm, 30' 6% APPENDIX IV (Cont'd) TABLE I I : TESTSECTION CHARACTER ISTICS

HoleSite/Angle or Venting of Plenum Facll Ity Cross-section Wail Type Slot Width/- Porosftv Wall Anqlc Chamber 0.5 m sq. (M > 1) cm, 6.8 cm 1 10% to 0.1' DFVLR (T-s WT) 0.6 x 0.34 rn2 Slotted 0.8 - 0.05' Exhausted to Atm. 0.50 m 0.25 - so, id DFVLR(HGK) x 0.30 .m NASA Marshal 1 Varies 1-1 0:: 0.81 Dia. Perf. 00 Ejector Flaps (HRNTT) m 1.27 cm, 30' along axis 1. Rmort No. 2. Gobwnmmr Accr*on No. 3. Rripient's cltabg No. NASA CR-2920 4. Title rd Subtitle 5. Rem13 Date November 7 7 "Calibration of Transonic and Supersonic Wind Tunnels" 19 6. F'erforming Orgmnization Coda -

T.D. Reed, T.C. Pope and J.M. Cooksey I . 10. Work Unit No. 9. MaminQOrp.nilrtion Narm ud Ad&- - Vought Corporation "11. Contract or Grant No. Dallas, Texas NAS 2-8606 - 13. Type of Repon and Period Covered 12. Spotnoring Apncv Nlnu md Address Contractor Report - National Aeronautics E Space Administration 14. Smmroring Agmcy code Washington, D.C. 20546 1 I - 15. Supplemntarv Notes

- 16. Abstraa State-of-the art instrumentation and procedures for calibrating transonic (0.6 C M < 1.4) and supersonic (M 5 3.5) wind tunnels are reviewed and evaluated. Major emphasis is given to transonic tunnels. Background information was obtained via a literature search, personal contacts and a questionnaire which was sent to 106 domestic and foreign facilities. Completed questionnaires were received for 88 tunnels dnd included government, industry and university-owned facilities. Continuous, blowdown and intermittent tunnels are considered. The required measurements of pressure, temperature, flow angularity, noise and humidity are discussed, and the effects of measurement uncertainties are summarized. Included is a comprehensive review of instrumentation currently used to calibrate empty-tunnel flow conditions. The recent results of relevant research are noted and reconmendations for achieving improved data accuracy are made where appropriate. lt is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation. provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 M < 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline. With regard to procedures. tunnel operators are cautioned to: (1) verify by measurements that expansions from a settling chamber to a test section are indeed isentropic, and (2) obtain calibrations over the entire range of reynolds number and humidity levels. Also, it is suggested that calibration data should include off-centerline measurements of Mach number and flow angularity. Finally, three problem areas for transonic tunnels are identified and discussed, viz. (1) the lack of standard criteria for flow uniformity and unsteadiness, (2) the undesirable noise generated by ventilated walls, and (3) wall interference. - Wind tunnels, Calibration, Testing I UNCLASSIFIED-UNLIMITED I I STAR Category 09 - 19. Scurity Clrit. (of this vtt 20. SMiw aaif. (of chis pet 21. NO. of p.pa 22. Rice' UNCLASSIFIED UNCLASSIFIED 287 $9.25 - 'For ule bv thc Nniwl Tuhniul InfarmtionSmvii. Spriqtidd, Virginia 22161

*U.S. GOVEKtB.!ENT PRINTIKG OFFICE: 1977 - 735-078143