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NASA/TM-2003-210738

In-Flight Capability for Evaluating Skin-Friction Gages and Other Near-Wall Flow Sensors

Trong T. Bui and Brett J. PipitoneNASA Dryden Flight Research CenterEdwards, California

Keith L. KrakeSpiral Technology, Inc.Edwards, California

February 2003

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NASA/TM-2003-210738

In-Flight Capability for Evaluating Skin-Friction Gages and Other Near-Wall Flow Sensors

Trong T. Bui and Brett J. PipitoneNASA Dryden Flight Research CenterEdwards, California

Keith L. KrakeSpiral Technology, Inc.Edwards, California

February 2003

National Aeronautics andSpace Administration

Dryden Flight Research CenterEdwards, California 93523-0273

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*Aerospace Engineer†Engineering Student Trainee

‡Instrumentation Engineer

§

Note that use of trade names or names of manufacturers in thisdocument does not constitute an official endorsement of such productsor manufacturers, either expressed or implied, by the NationalAeronautics and Space Administration.

1

IN-FLIGHT CAPABILITY FOR EVALUATING SKIN-FRICTION GAGES AND OTHER NEAR-WALL FLOW SENSORS

§

Trong T. Bui

*

and Brett J. Pipitone

†

NASA Dryden Flight Research CenterEdwards, California

Keith L. Krake

‡

Spiral Technology, Inc.Edwards, California

Abstract

An 8-in.-square boundary-layer sensor panel has beendeveloped for in-flight evaluation of skin-friction gagesand other near-wall flow sensors on the NASA DrydenFlight Research Center F-15B/Flight Test Fixture (FTF).Instrumentation on the sensor panel includes aboundary-layer rake, temperature sensors, staticpressure taps, and a Preston tube. Space is also availablefor skin-friction gages or other near-wall flow sensors.Pretest analysis of previous F-15B/FTF flight data hasidentified flight conditions suitable for evaluatingskin-friction gages. At subsonic Mach numbers, theboundary layer over the sensor panel closelyapproximates the two-dimensional (2D),law-of-the-wall turbulent boundary layer, andskin-friction estimates from the Preston tube and therake (using the Clauser plot method) can be used toevaluate skin-friction gages. At supersonic Machnumbers, the boundary layer over the sensor panelbecomes complex, and other means of measuring skinfriction are needed to evaluate the accuracy of newskin-friction gages. Results from the flight test of a newrubber-damped skin-friction gage confirm that atsubsonic Mach numbers, nearly 2D, law-of-the-wallturbulent boundary layers exist over the sensor panel.Sensor panel data also show that this new skin-frictiongage prototype does not work in flight.

Nomenclature

Acronyms

FTF Flight Test Fixture

rms root mean square

RTD resistance temperature detector

Symbols

coefficient of skin friction transformed into the incompressible plane by the van Driest II transformation

g

gravitational acceleration

p

static pressure

T

temperature

u

streamwise flow velocity

u

+

velocity in wall units,

u

eq

van Driest effective velocity

u

τ

friction velocity,

x

streamwise distance from the leading edge

y

normal distance from the wall

y

+

distance in wall units,

δ

boundary-layer thickness

∆

pressure gradient parameter,

µ

viscosity coefficient

kinematic viscosity,

ρ

density

τ

shear stress

Cf*

ueq

uτ--------

uττw

ρw------=

y+ ρwuτy

µw----------------=

∆νw

ρwuτ3

------------dpe

dx--------=

ν µρ---

2

Subscripts

aw adiabatic wall

e edge of boundary layer

w wall

Introduction

Surface skin-friction drag is an important forceaffecting supersonic and hypersonic flight vehicles andvehicle propulsion systems. During the last 40 years, theNASA Dryden Flight Research Center (Edwards,California) has made a significant contribution toskin-friction research for high-speed flight. As early as1967, Garringer and Saltzman

1

demonstrated thesuccessful operation of a small, commercially availableskin-friction gage on the X-15 aircraft (North AmericanAviation, Inc., Los Angeles, California) to Mach 4.9.They found that the influence of the wall-to-recoverytemperature ratio on measured turbulent skin-frictionvalues was not as large as expected. In 1969, Quinn andOlinger

2

extended skin-friction measurement on theX-15 aircraft to Mach 5.25 and found that theexperimentally determined Reynolds analogy factor wassignificantly higher than the theoretically predictedvalues. In 1973, Fisher and Saltzman

3

measured the skinfriction and boundary-layer velocity profiles at variouslocations on the XB-70-1 aircraft (North AmericanAviation, Inc., Los Angeles, California) to Mach 2.5. Theskin friction was measured using a skin-friction forcebalance, a Preston tube, and a boundary-layer rake(using the Clauser plot method). Good agreement wasobtained with the Karman-Schoenherr correlation ataircraft locations that had approximatelytwo-dimensional (2D) flows. In 1980, Quinn and Gong

4

measured skin friction, heat transfer, and boundary-layervelocity profiles on a hollow cylinder. The cylinder wasmounted beneath a YF-12A aircraft (Lockheed MartinAeronautics Company, Palmdale, California) atMach 3.0. Quinn and Gong’s results showed goodwind-tunnel-to-flight skin-friction correlation.

Current hypersonic flight research efforts at NASADryden, primarily the X-43A research vehicle (MicroCraft, Inc., Tullahoma, Tennessee),

5

have made accurateskin-friction measurement even more critical than it hasbeen in the past. In addition to flying at comparativelyhigher Mach numbers of 7.0 to 10.0, where externalaerodynamic skin-friction drag is expected to be severe,the X-43A uses a scramjet engine through which flowremains supersonic. As a result, the skin-friction draginside the X-43A scramjet engine can be significant aswell.

Measuring skin friction in flight poses uniquechallenges in addition to those encountered inwind-tunnel testing. The gages and signal-conditioningsystems must be compact to fit into tight spaces in thevolume-limited flight vehicle. For ease of integrationinto the flight instrumentation system, the gages shouldnot have any signal-conditioning requirement other thanthose supported by the flight vehicle. The gages andsignal-conditioning systems are exposed to a widevariation of ambient pressures and temperatures in flight,and the test time in flight often is longer than in awind-tunnel test. Most importantly, the gages must berobust to survive extreme conditions encountered inflight, such as stage-separation shock,

g

loads, vibration,electromagnetic interference, and engine-firing heatloads, and still provide accurate measurements.

To address the challenges of measuring skin friction inflight and to support the development of skin-frictiongages and other near-wall flow sensors for flightresearch, a boundary-layer sensor panel has beendeveloped for use on the NASA Dryden F-15B(McDonnell Douglas Corporation, St. Louis,Missouri)/Flight Test Fixture (FTF). This reportdescribes the boundary-layer sensor panel and flightsignal-conditioning system, evaluates flow quality overthe sensor panel for both subsonic and supersonic flightconditions using previous F-15B/FTF flight data, anddiscusses results from a recent in-flight evaluation of anew rubber-damped skin-friction gage.

Flight Facility Description

The F-15B/FTF is an aerodynamics and fluiddynamics research test bed at NASA Dryden.

6

Figure 1shows the F-15B/FTF in flight, carrying theboundary-layer sensor panel. The FTF is the black,vertical, fin-shaped object mounted on the centerline ofthe F-15B lower fuselage. Primarily made of compositematerials, the FTF was designed for flight research atMach numbers to a maximum of Mach 2.0. Without theaft fairing, the FTF is 107 in. long, 32 in. high, and 8 in.wide. To improve the flow quality aft of the FTF, the aftfairing was used for the flight discussed in this report.The aft fairing adds an additional 18.8 in. to the lengthof the FTF, as shown in figure 2. The boundary-layersensor panel is the small white aluminum panel locatedtoward the aft end of the FTF. As shown in figure 2b, theflow at the proposed sensor panel location is relativelystraight and uniform. The FTF noseboom provides localincoming flow properties in addition to the airdataprovided by the aircraft noseboom. Signal-conditioningsystems for the experiment are mounted inside the FTF.

3

Figure 1. NASA Dryden F-15B/FTF in flight with the boundary-layer sensorpanel.

4

(a) FTF without aft fairing.

Aft fairing

020586

Attachment jointProposed sensorpanel location

(b) FTF with aft fairing.

Figure 2. Tuft flow visualization of the F-15B/FTF at Mach 0.7, 45,000 ft (from Richwine

6

).

5

Boundary-Layer Sensor Panel Description

Figure 3 shows a closeup view of the boundary-layersensor panel. The sensor panel is an 8.00-in.-square,0.75-in.-thick aluminum plate designed to fit intoexisting 8.00-in. hatches on either side of the FTF,which facilitates joint flight testing with other FTFexperiments. Detailed survey of the boundary layer overthe sensor panel is made possible by a high-resolutionboundary-layer rake and a Preston tube. Resistancetemperature detectors (RTDs) and Micro-Foil

®

(RdFCorporation, Hudson, New Hampshire) heat-fluxsensors have been installed on both the front and backside of the sensor panel. In addition, three rows of staticpressure taps (indicated by the triangles in figure 3) arepresent to provide local-wall static pressures on thesurface of the sensor panel. In the top row, two staticpressure taps are located in front of the Preston tube. Inthe middle row, two static pressure taps are placedacross the skin-friction gage, and two static pressuretaps are located in front of the rake in the bottom row.Space is available on the sensor panel to accommodate

two skin-friction gages or other near-wall flow sensors.As figure 3 shows, only one skin-friction gage has beeninstalled at location No. 1. The mounting hole for thesecond skin-friction gage has been filled with amatching aluminum blank plug at location No. 2.

As figure 3 shows, limited space on the sensor panelnecessitates the mounting of the boundary-layer rakeand the Preston tube 1.5 in. downstream of theskin-friction gages. Boundary-layer analysis for theF-15B/FTF flight conditions shows that this locationmismatch should result in a skin-friction coefficientchange of only 0.25 percent. This small change is wellwithin the accuracy of direct skin-friction measurementgages, which is ±5 percent based on past skin-frictionmeasurements.

7

The location mismatch, therefore, is notexpected to present problems in evaluating skin-frictiongages. Furthermore, mounting the rake and the Prestontube downstream of the skin-friction gages reduces anyinterference effects that these intrusive instrumentsmight have on the gages.

Figure 3. Closeup view of the boundary-layer sensor panel installation on the F-15B/FTF in flight.

6

The boundary-layer rake and the Preston tube used onthe sensor panel have been designed and builtspecifically for the F-15B/FTF flight conditions.

8

Asfigure 4 shows, the rake has a curved body, which allowsthe pitot tubes to be more densely clustered in thenear-wall region than conventional rakes allow. For theFTF boundary layer, the law-of-the-wall region extendsapproximately 0.5 in. above the surface of the FTF, andthe pitot tubes are spaced on the curved portion of therake such that approximately ten tubes are located insidethis region. This number of tubes should be sufficientfor computing the skin-friction shear stress using theClauser plot method. The rake total height is 2.94 in.,which allows it to span the entire F-15B/FTF boundarylayer over the expected flight envelope. The centerlineof the first pitot tube is approximately 0.04 in. from the

wall. This boundary-layer rake has been found to giveaccurate measurements in a wind-tunnel test to amaximum of Mach 2.0.

8

Hopkins and Keener

9

discussed a method for sizingPreston tubes. Using their method and a representativeF-15B/FTF flight condition at Mach 0.8 and an altitudeof 30,000 ft, analysis has shown that the maximumPreston tube diameter for which a single calibrationcurve would be expected to be applicable is 0.312 in.,and the minimum Preston tube diameter is 0.012 in.Essentially the same range of Preston tube sizes isobtained using the sizing formulas from Allen.

10

Thecurrent Preston tube outer diameter of 0.125 in. fallswell within this allowable range. This Preston tube alsohas been found to perform well in a wind-tunnel test.

8

18.9°

0.936 in.

0.250 in.

0.030 in. 0.042 in.

0.027 in.

Detail AScale 8:1

26°- - - - - - - - - -

- - - - - - - - - -

18.9°020523

A

R 1.625 in.

0.188 in.0.165 in.

R 1.375 in.

3.000 in.

2.500 in.

1.750 in.

3.030 in.

2.536 in.

2.940 in.

Figure 4. Detailed three-view drawing of the F-15B/FTF curved rake design.

7

Signal Conditioning System for the Boundary-Layer Sensor Panel

The various flow sensors and pressure taps on theboundary-layer sensor panel require an onboardsignal-conditioning system. A compact andhigh-performance signal-conditioning system wasconfigured for the boundary-layer sensor panel.Conditioned signals from each of the boundary-layersensor panel measurements are added to the standardFTF data acquisition system, which is described byRichwine.

6

Analog signal conditioning for the low-output-levelskin-friction gages and heat-flux gages was provided bya multiple-purpose design containing bridgecompletion, gain, offset, and active three-poleButterworth filtering. Each channel was configuredindividually for each sensor. Signal conditioning islocated in close proximity to the boundary-layer sensorpanel.

Pressure measurements, including all boundary-layerrake ports, the Preston tube, and surface static ports, areobtained using two temperature-controlled,electronically scanned pressure transducers sampled at arate of 25 samples/sec. The surface static ports, Prestontube, and six boundary-layer rake ports nearest thesurface are measured with a ±5-lbf/in

2

differentialtransducer. Estimated uncertainty in thesemeasurements is ±0.018 lbf/in

2

. The remainingboundary-layer rake ports are measured with a±10-lbf/in

2

differential transducer. Estimateduncertainty in these measurements is ±0.036 lbf/in

2

.Both pressure transducers are referenced to the FTFnoseboom static pressure with estimated uncertaintywithin ±0.023 lbf/in

2

. Transducer temperatures aremonitored and available in the data stream.

The RTD, configured as one arm of a Wheatstonebridge with a three-wire hookup, measures walltemperature. Measurement uncertainty is estimated at±0.4 °F. Signal conditioning, including cold-junctioncompensation, is supplied for both Type-K and Type-Tthermocouples installed in the skin-friction andheat-flux gages, respectively. In addition to theuncertainties associated with the specific thermocoupletypes, the measurement uncertainty from the onboarddata system is estimated to be ±2.0 °F.

Pre-Test Boundary-Layer Analysis for Flows Near the Boundary-Layer Sensor Panel

During the design phase of the boundary-layer sensorpanel, analysis of boundary-layer velocity profilescollected during previous F-15B/FTF flight tests

6

wasconducted to determine the flow quality in the vicinityof the proposed sensor panel. An understanding ofboundary-layer properties is important to determine thecorrect instrumentation size and location. Also, thePreston tube and Clauser plot methods assume a 2D,fully turbulent boundary layer. Therefore, theidentification of flight conditions that produce 2D, fullyturbulent boundary layers is important so that anevaluation of the skin-friction gage accuracy can bemade.

Richwine published boundary-layer data for the entirematrix of flight conditions conducted in his experiment,covering a significant portion of the F-15B/FTF flightenvelope.

6

Boundary-layer velocity profiles near thecurrent boundary-layer sensor panel location areavailable for altitudes of 15,000, 30,000, and 45,000 ft,at Mach numbers ranging from 0.39 to 2.0. Anunderstanding of the local boundary-layer propertiescan be gained from analyzing this comprehensivedatabase.

The van Driest effective velocity concept, asdescribed by White,

11

can be used to collapse theboundary-layer velocity profiles at different free-streamMach numbers into the well-known incompressible lawof the wall. Figure 5 compares the turbulentboundary-layer velocity profiles obtained at an altitudeof 15,000 ft for aircraft Mach numbers ranging from0.39 to 0.98. The velocity profiles at different Machnumbers all collapse into the ordinary incompressiblelaw of the wall, and the agreement is good. The rakeused in Richwine's flight experiment is a canted rakethat spans approximately 5 in. across the flow. Becausethe velocity profiles do not have discontinuities,Richwine's rake data show that the flow isapproximately 2D at this location.

Figure 6 compares the turbulent boundary-layervelocity profiles obtained at an altitude of 30,000 ft foraircraft Mach numbers ranging from 0.51 to 1.48. Onlythe data from subsonic flights agree with the law of thewall. At supersonic Mach numbers, the velocity profilesare below the law of the wall, and the disagreementbecomes larger at greater Mach numbers.

8

Figure 7 compares the turbulent boundary-layervelocity profiles obtained at an altitude of 45,000 ft foraircraft Mach numbers ranging from 0.7 to 2.0. Again,only the data from subsonic flights agree with the law ofthe wall, a result similar to the data obtained at analtitude of 30,000 ft. At supersonic Mach numbers, thevelocity profiles are significantly below the law of thewall. Starting at Mach 1.06, the velocity profiles moveaway from the law of the wall until Mach 1.39; then atMach numbers greater than 1.39, the velocity profilesmove back toward the law of the wall.

At supersonic Mach numbers, the disagreementbetween the velocity profiles and the law of the wallappears to be caused by shock formation over the FTFsurface, as observed by Richwine.

6

At transonic andsupersonic Mach numbers, wall pressure data from hisexperiment indicated the presence of shocks over theFTF. Richwine observed that the shocks weakened atMach numbers greater than 1.39.

6

Figure 7 shows thatthe velocity profiles move back toward the law of thewall at Mach numbers greater than 1.39, a trend thatcorresponds with Richwine’s observation.

The skin friction can be calculated from Richwine'sboundary-layer velocity profiles using the Clauser plot

method and the Fenter-Stalmach law of the wall, asdescribed by Allen and Tudor.

12

Figure 8 presents theskin-friction results for altitudes of 15,000, 30,000, and45,000 ft. Comparison with the incompressibleKarman-Schoenherr correlation is possible, because thecompressible skin-friction values have been transformedinto the incompressible plane using the van Driest IIcorrelation, described by Hopkins and Inouye.

7

Adifferent Karman-Schoenherr curve exists for eachaltitude, because the Reynolds number changes withaltitude.

For the three altitudes considered, the agreement withtheory is good at subsonic Mach numbers, and theagreement worsens as the flight Mach number isincreased toward Mach 1.0. At supersonic Machnumbers, the rake results diverge from theory, whichindicates complex boundary layers. At an altitude of45,000 ft, for aircraft Mach numbers between 1.16 and1.69, the computation of skin-friction values using theClauser plot method is not possible. At these particularconditions, the velocity profiles could not be made to fitthe log-law profile, regardless of the skin-frictionvalues.

15

20

10

25

u+

30

35

100 1,000 10,000 100,000

020524y+

Ln(y+)/0.41 + 5.00

Mach 0.39

Mach 0.52

Mach 0.61

Mach 0.70

Mach 0.80

Mach 0.90

Mach 0.98

Figure 5. F-15B/FTF boundary-layer velocity profiles at 15,000 ft.

9

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020525y+

Ln(y+)/0.41 + 5.00

Mach 0.51

Mach 0.62

Mach 0.70

Mach 0.79

Mach 0.89

Mach 1.00

Mach 1.05

Mach 1.15

Mach 1.28

Mach 1.39

Mach 1.48


Ln(y+)/0.41 + 5.00

Mach 0.70

Mach 0.79

Mach 0.90

Mach 0.98

Mach 1.06

Mach 1.16

Mach 1.28

Mach 1.39

Mach 1.49

Mach 1.69

Mach 1.80

Mach 1.90

Mach 2.00

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020526y+


10

The results from this pretest analysis indicate that atsubsonic aircraft Mach numbers, at any altitude in theF-15B/FTF flight envelope, the evaluation ofskin-friction gages in flight is possible using theboundary-layer rake and the Preston tube. At transonicand supersonic Mach numbers, however, the complexboundary layers over the FTF preclude accuratecomputation of skin-friction values using the Clauserplot and Preston tube methods. Other methods, such as acalibrated direct-measuring skin-friction gage, must beused to estimate skin-friction values. This type ofskin-friction gage, however, is currently not availablefor the F-15B/FTF.

Accuracy of the Control Skin-Friction Measurements

The boundary-layer rake and Preston tube in thesensor panel provide the control skin-frictionmeasurements used to evaluate new skin-friction gages,and a review of the accuracy of these approaches isimportant. Skin friction can be calculated from theboundary-layer rake data using two different methods:skin-friction theory and the Clauser plot method. Thetheoretical skin-friction value is calculated from the vanDriest II transformation and the Karman-Schoenherrcorrelation, the method recommended by Hopkins andInouye. They reported that this method can predict theskin-friction value to within 10 percent of the data

considered.

7

Allen

13

compared the accuracy of severaldifferent Clauser plot methods and concluded that theFenter-Stalmach and Baronti-Libby methods give thebest results with accuracy of approximately ±5 percentof the data considered. Allen preferred theFenter-Stalmach method, because it is simpler than theBaronti-Libby formulation. The Fenter-StalmachClauser plot method is used in this report.

Finally, skin friction can be calculated from thePreston tube data. A popular Preston tube method forcompressible turbulent boundary layers is theBradshaw-Unsworth method.

14

Allen

15

found that thismethod is more accurate at comparatively higher valuesof calibration parameters defined in reference 16 than atlower levels. The F-15B/FTF flight conditions generallyresult in higher values of the Allen calibrationparameters, and data published in reference 15 showthat the Bradshaw-Unsworth method is expected to beaccurate to approximately ±10 percent.

In summary, the control skin-friction measurementsprovided by the boundary-layer rake and Preston tube inthe sensor panel are expected to be accurate to within±10 percent. If more accuracy is desired, then a bettermethod of control skin-friction measurement must beused. No other method, however, is presently availablefor flight research applications.

.0020

.0015

.0025

Cf*

0 .5 1.0 1.5 2.0 2.5Mach number

Karman-Schoenherrcorrelation,15,000 ft

Flight data,15,000 ft





020527

Figure 8. F-15B/FTF skin-friction results and comparison with theory.

11

Flight Test of a New Rubber-Damped Skin-Friction Gage

A new rubber-damped skin-friction gage wasevaluated in flight using the boundary-layer sensorpanel on the F-15B/FTF. Magill, et al,

17

described anearly design of the gage. Sang and Schetz

18

discussed asignificantly improved and ruggedized version of thisgage, which was specifically made for flight testing onthe F-15B/FTF. Figure 9 shows a diagram of the gage,which uses a cantilever-beam, non-nulling approach,with a rubber sheet covering the top of the sensing disk.

Environmental Test Results

Before it is accepted for use in NASA Drydenaircraft, new flight hardware is normally subjected to aseries of rigorous environmental tests, includingvibration, altitude (pressure), and temperature tests. Inaddition to qualifying the equipment for flight, thesetests provide an opportunity to evaluate the performanceof new sensors and equipment in the controlledenvironment of a ground laboratory. The specificationsfor the environmental testing of F-15B/FTF equipmentare provided in reference 6.

For the vibration test on the new gage, the gage wassecurely mounted on a vibration table, and randomvibration measuring 8

g

rms was applied for 20 min ineach of the 3 normal directions. The frequency for thevibration test ranged from 15 to 2000 Hz. Severalmodifications were done before the gage passed the

vibration test. The first gages had glycerin fill tubes andcaps that were loosely attached to the gage housing. Forthe gage to withstand the vibration tests, the tubes andcaps were replaced by set screws. Later, when glycerinwas not used to fill the gage cavity, the glycerin fillholes were left open.

Problems with strain-gage wiring also werediscovered during the ground vibration tests. Theseskin-friction gage prototypes use semiconductor straingages that are extremely delicate and small, the size ofsingle strands of hair. In the first skin-friction gagesdelivered to NASA Dryden for this project, therelatively large electrical wires on the outside electricalconnector were soldered directly to these delicatesemiconductor strain gages, allowing forces on theexternal wiring connector to be transmitted directly ontothe strain gages. Consequently, a slight movement of thewiring connector on the skin-friction gage housingwould produce significant erroneous gage outputs.When the vibration test was first attempted, and theskin-friction gage was being mounted to the vibrationtable, the solder joints to the semiconductor strain gagesbroke, causing the skin-friction gage to becomeinoperative before any vibration was applied.

To solve this problem, aircraft-quality electricalconnectors and wiring harnesses were used onsubsequent skin-friction gages. To relieve the stress onthe delicate semiconductor strain gages, an extrasoldering pad was used inside the skin-friction gagehousing. Small electrical wires were used between theinternal soldering pad and the strain gages, and largerelectrical wires were used to connect the soldering padto the outside electrical connector. In addition,stress-relieving wiring loops were used inside theskin-friction gage housing to further isolate the delicatestrain gages from forces on the electrical connector.After these modifications, the skin-friction gage passedthe ground vibration test.

For the pressure and temperature environmental tests,the skin-friction gage was placed inside a sealed testchamber. The pressures and temperatures inside thechamber were varied to simulate conditions at altitudewhile the gage output was monitored. For the altitude(pressure) test, the gage was left at ambient temperature,and the pressures were varied between ambient and thatof an altitude at 50,000 ft. For the temperature test, thetemperatures were varied between -50 and 100 °F, andthe pressure was held at ambient value. Several differentgage configurations were tested. The first skin-frictiongages had significant sensitivity to changes in bothpressure and temperature. With no glycerin fill (the

Sensorhead

Wall shear, τw

Cantileverbeam

Rubbersheet

Gagehousing

Straingages

020528

Emptycavity

Figure 9. Diagram of a rubber-damped F-15Bskin-friction gage.

12

configuration that was flown on the F-15B/FTF), thegage was found to be relatively insensitive to changes inambient pressure. The sensitivity to temperaturechanges, however, was significant, and it remains aserious flaw of this gage design.

In the first gage prototypes, only half of a Wheatstonebridge was used for the strain gages inside theskin-friction gage. Bridge completion was done throughthe signal-conditioning unit. This bridge arrangementcaused serious, incomprehensible temperaturesensitivity in the skin-friction gage, because half of theWheatstone bridge was with the skin-friction gageinside the temperature test chamber, and the other halfwas in the signal-conditioning unit outside the testchamber. The use of a full Wheatstone bridge inside theskin-friction gage improved the temperature sensitivityby making it more repeatable. The use of an aluminumcantilever beam instead of a plastic cantilever beaminside the skin-friction gage further improvedtemperature sensitivity and repeatability. In addition, thealuminum cantilever beam provides a better bondingsurface for the semiconductor strain gages than theplastic beam provides. Figure 10 shows the temperaturesensitivity of the gage compared to the total calibratedgage output range for the F-15B/FTF skin-friction gageconfiguration flown. Significant temperature sensitivity

still exists. The variation in the gage output over theexpected temperature range in flight is approximatelysix times the total calibrated gage output. In otherwords, the gage output caused by temperature changesis many times larger than the expected gage outputcaused by skin friction. In addition, the data in figure 10has a very wide hysteresis band.

In an attempt to correct the temperature sensitivity ofthe gage, a quadratic fit of the data (shown in figure 10)was made. This fit was used together with the rubbersheet temperature on the gage to subtract out the gageoutput caused by temperature alone. Although thisapproach does not eliminate the uncertainty caused bythe wide hysteresis band, it removes the overall effectsof temperature on the gage output. Figure 11 shows theresults of the temperature correction scheme on theskin-friction gage laboratory calibration data. Thesecalibrations were performed with rubber sheettemperatures ranging from 83 to 87 °F. Although this isa narrow range of ambient temperature, it is wideenough to cause the uncorrected gage-calibration data toscatter. With the application of the temperaturecorrection scheme, all of the temperature corrected datacollapses into a single calibration line.

–60

5

020529

–40 –20 0 20 40 60 80 100 120Rubber sheet temperature, °F

4

3

2

1

0

Change ingage output(volt) from

70 °F

–1

–2

–3

Total calibratedgage output

All temperature test dataAll temperature test data

quadralic fit

Figure 10. Temperature sensitivity of the rubber damped skin-friction gage.

13

Flight Test Results

In-flight evaluation of a non-glycerin-filledrubber-damped skin-friction gage was performed usingthe F-15B/FTF. On December 20, 2000, a dedicatedskin-friction flight was conducted on an F-15B aircraft,NASA Dryden tail number 836. An FTF aft fairing wasused for this flight. Figure 12 shows the profile for thisflight. During the first portion of the flight, a sweep ofthree altitudes (15,000, 30,000, and 45,000 ft) wasconducted. As the aircraft descended from 45,000 ft, thesame sweep was conducted in reverse order to confirmdata repeatability. At each altitude, four subsonic Machtest points (Mach 0.6, 0.7, 0.8, and 0.9) were obtained.At each test point, the F-15B aircraft maintained straightand level flight for approximately 2 min to maintainhigh flow quality and keep acceleration loads frominfluencing the skin-friction gage reading. Each of thetest points shown in figure 12 was carefully chosen toassure a 2D, law-of-the-wall boundary layer over thesensor panel, as discussed previously.

When reducing flight data, care was taken to averageonly the data in the time intervals during which all theflight conditions, including Mach number, altitude,angle of attack, and angle of sideslip, were

simultaneously constant, indicating steady-stateconditions had been reached. The heat-flux sensorsmalfunctioned during this flight; therefore nomeaningful heat-flux data were collected. The walltemperature measurements, however, provide a goodindication of wall heat-transfer rates. Figure 13 shows aplot of the ratio of the wall temperature as measured bythe outside wall RTD to the calculated adiabatic walltemperature ( ) for all the flight conditions. Thewall temperature is close to adiabatic wall temperaturefor most of the flight conditions, with rangesfrom 0.88 to 1.16. An “adiabatic wall–seeking”temperature variation is generally observed. As theairplane climbs to higher altitudes, the warm wall coolstowards the adiabatic wall temperature. Conversely, asthe airplane descends, the cool wall heats up towards theadiabatic wall temperature. The in-flight variation of

T

w

/

T

aw

is not expected to influence the accuracy ofvarious instruments in the sensor panel. Hopkins andInouye

7

found that skin-friction theories can accuratelypredict the skin friction at

T

w

/

T

aw

ratios greater than 0.3.Also, Allen

16

found that the compressible Preston tubecalibrations agree very well with the data at

T

w

/

T

aw

ratios as low as 0.32 to 0.51.

Tw Taw⁄

Tw Taw⁄

8

020530

–.2

7

6

5

4

3

2

1

0

–10 .2 .4 .6 .8

Change in gage output, V

Changein weight,

g

1.0 1.2 1.4 1.6

X-dir, uncompensated

X-dir, compensated

Linear fit of temperaturecompensated data

Figure 11. Temperature correction to the ground calibration data at different ambienttemperatures.

14

Altitude,ft

Machnumber

50,000

45,000

40,000

35,000

30,000

25,000

20,000

15,000

10,000

5,000

0 10 20 30Flight time, min

020531

40 50 60 70

1.0Altitude, ftMach .9

.8

.7

.6

.5

.4

.3

.2

.1

0

Figure 12. Flight profile for skin-friction gage evaluation.

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020532Flight conditions

1.2

1.0

1.4

.8

.6

.4

.2

0

Figure 13. Wall temperature ratio at the sensor plate.

15

Data from the wall static pressure taps indicate thatvery small streamwise pressure gradients exist in thevicinity of the sensor panel. The pressure gradientparameter

,

as discussed by Patel,

19

is plotted in figure 14. In thisfigure, pressure gradients for the rake and Preston tubewere computed using two wall pressure taps in front ofthe rake and two in front the Preston tube. Pressuregradients for the skin-friction gage were computedusing one wall pressure tap upstream and onedownstream of the skin-friction gage. Because theskin-friction gage pressure taps are located beside theboundary-layer rake, these taps also provide thestreamwise pressure gradient beside the rake.Interestingly, the small adverse pressure gradient infront of the rake is almost exactly mirrored by the smallfavorable pressure gradient beside the rake, whichshows that the rake is producing a small local pressuredisturbance on the sensor plate in a manner similar tothat of a small airfoil. The Preston tube produces lessdisturbance than the rake. The pressure gradient in frontof the Preston tube is close to zero for all conditions.

The limiting values of

∆

, as recommended by Patelfor a maximum Preston tube error of 6 percent, areplotted in figure 14. The

∆

limits range from -0.007 to0.015. As figure 14 shows, all pressure gradients arewell within the Patel limits. Frei and Thomann

20

alsostudied the Preston tube error caused by local pressuregradients. They found that the error depends on both thelocal pressure gradient and the Reynolds number (basedon the local friction velocity and the Preston tubediameter) and published an empirical fit for the error.Using the Frei and Thomann empirical fit, the expectedPreston tube error can be computed for the current flightexperiment. Figure 14 shows that the local pressuregradient reaches a maximum of 0.003 at Mach 0.9, andan altitude of 45,000 ft. At this flight condition, theReynolds number is 419. The expected Preston tubeerror based on the Frei and Thomann fit isapproximately 0.7 percent. This error is well within theuncertainties of this instrument, and the local pressuregradient effects can be considered negligible for thisexperiment. Note that if the upper

∆

limit value of 0.015was used in the Frei and Thomann fit, then the errorbecomes approximately 6 percent, in agreement withPatel's findings.

The validity of both the Preston tube and Clauser plotmethods is based on the well-known log law. Patel

19

showed that constants in the incompressible log law can

be derived from the Preston-tube calibration curve.Allen

16

demonstrated that a number of popularcompressible Preston tube calibrations can be reducedto the incompressible log law at the limit of zerofree-stream Mach number. The findings pertaining to thePreston tube (discussed in the preceeding paragraph)can also apply to the Clauser plot method; therefore, thepressure gradient effects can be considered negligiblefor the Clauser plot method as well.

Figures 15, 16, and 17 show the boundary-layervelocity profiles obtained in flight. Figure 15a showsdata from the ascending portion of the flight, andfigure 15b shows data from the descending portion ofthe flight, both at altitudes of 15,000 ft. Analogouslabeling is used in figures 16 and 17 for the test points ataltitudes of 30,000 and 45,000 ft, respectively. NoMach-0.9, 45,000-ft test point exists for the descendingportion of the flight (figure 17b), because this test pointis at the top of the flight profile as shown in figure 12. Incalculating

u+ and y+, the skin-friction values used wereobtained from the actual local boundary-layermomentum thickness and the Karman-Schoenherrtheory. When skin friction was calculated this way,compressibility was accounted using the van Driest IItransformation. To compare data with theincompressible law of the wall, the van Driest effectivevelocity concept was used. As expected, goodagreement was obtained between the measured velocityprofiles and the law of the wall. Also, for mostconditions, very good repeatability of data was achievedbetween the ascending and descending portions of theflight. In the descending portion of the 45,000-ft flight,however, the measured velocity profiles are noticeablyabove the standard log-law profile.

During the flight, the pilot had difficulty maintainingsteadiness at the Mach-0.7, 45,000-ft condition, because0.7 is close to the low Mach number limit for the aircraftat this altitude. Figure 12 shows that as a result of thisdifficulty, very little time was spent at the Mach-0.7,45,000-ft condition. This difficulty might havecontributed to the large disagreement with the log-lawprofile shown in figure 17. Interestingly, results derivedfrom Richwine's data (figure 7) also show a worsecorrelation between skin friction and theory at a45,000-ft altitude. Apparently, the boundary-layerprofiles at a 45,000-ft altitude do not correlate with thelog-law profile as well as boundary-layer profiles atlower altitudes.

∆νw

ρwuτ3

------------dpedx--------=

16

In figures 18, 19, and 20, the boundary-layer profilesalso are compared with the 1/7th-power-law profile, apopular approximation for turbulent flat-plate boundarylayers.11 The measured boundary-layer profiles are inagreement with both the 1/7th-power-law profile andlog law, which indicates that a flat-plane condition hasbeen reached. At subsonic aircraft Mach numbers,Richwine's boundary-layer profiles also are in goodagreement with the 1/7th-power-law profile. Atsupersonic aircraft Mach numbers, however, Richwine'sprofiles do not agree with the power law.

Figure 21 compares the in-flight skin-friction gageoutput with results obtained using theKarman-Schoenherr theory, Clauser plot method, andPreston tube method. The Reynolds number, based onmomentum thickness, ranges from approximately 3,600(at Mach 0.7 and an altitude of 45,000 ft) to 30,000 (atMach 0.9 and an altitude of 15,000 ft), and theskin-friction shear stress from approximately 0.3 to 1.4psf, respectively. Flying low and fast generally results inhigh skin-friction levels and Reynolds numbers.

Both the Clauser plot and Preston tube results agreewith the Karman-Schoenherr theory (with thevan Driest II compressibility transformation). Thetemperature corrected and uncorrected skin-frictiongage measurements, however, erratically andunpredictably vary. The gage measurements do notcorrelate with the Preston tube values, Clauser plotvalues, or the Karman-Schoenherr theory. Figure 21clearly shows that this skin-friction gage design doesnot work in flight.

Sang and Schetz suggested “mildly 3D” flows as apossible reason for the disagreement between the gagemeasurements and other results,18 but this is not likelythe reason. The conditions for this flight test werecarefully chosen from the pretest analysis results. Onlyflight conditions that produce nearly 2D,law-of-the-wall turbulent boundary layers were used.The quality of the boundary layers obtained during theflight test is further validated by good agreement amongthe Preston tube results, Clauser plot results, and theory.As previously discussed, small local variations in walltemperature and streamwise pressure gradients arepresent; but they are not likely to cause the largedifferences between the skin-friction gagemeasurements and theory.

Sang and Schetz18 also suggested temperaturemismatch between the rubber sheet and the surroundingwall as a cause of the erratic gage outputs, but this is notlikely the cause either. As shown in figure 22,temperature differences exist between the rubber sheetand the wall, but the differences are small. The ratio ofrubber-sheet temperature to the wall temperature(mismatch temperature ratio) ranges from 0.94 to 1.07.The mismatch temperature ratio of 1.07, the largesttemperature mismatch during the entire flight, occurredduring the descending portion of the Mach-0.9,15,000-ft condition.

Westkaemper examined the effects of temperaturemismatch on direct measurements of drag at Mach 5and mismatch temperature ratios from 0.92 to 1.09 on aflat plate.21 He found that “there was no apparentcorrelation of the drag variation with the conditions oftemperature mismatch.” The drag variation found inWestkaemper's experiment was within the repeatabilityof ±2 percent of his measurements; therefore, “thevariation in drag did not appear to be related to themismatch condition, but appeared rather to be randomin nature.” Voisinet22 studied the effects of temperaturemismatch for Mach numbers of 2.9 and 4.9 and largermismatch temperature ratios ranging from 1.1 to 2.7.Using an empirical correlation of his data, Voisinetfound that the skin-friction variation in Westkaemper'sexperiment was indeed negligible.

For the current F-15B/FTF experiment, the expectederror in skin-friction values caused by temperaturemismatch can be computed using Voisinet's empiricalcorrelation. At the Mach-0.9, 15,000-ft, descendingflight condition, the Reynolds number is

per ft or per m, which is in themiddle of the range of Voisinet's data. Figure 10 inVoisinet's report shows that at this Reynolds number, theexpected skin-friction error per degree of temperaturedifference is approximately 0.042 (N/m2)/K or0.0005 (lbf/ft2)/R. The temperature mismatch isapproximately 33 R, resulting in a skin-friction error of0.017 lbf/ft2. With a nominal skin friction value of1.47 lbf/ft2 as predicted by theory, the error caused bythe temperature mismatch is approximately 1 percent,which is well within the uncertainties of the currentapproaches. The temperature mismatch effects,therefore, can be considered negligible for thisexperiment.

3.98 106

× 1.3 107

×

17

Possible causes of the skin-friction gage failureduring the flight test include the temperature sensitivityof the semiconductor strain gages, faulty installation ofthe strain gages, mismatch of the strain gages, and thetemperature sensitivity of the rubber sheet. The rubbersheet can unevenly expand and contract in response tothe ambient temperature changes during flight, inducingextraneous forces on the sensing element of theskin-friction gage. Because of the plastic-elastic natureof the rubber sheet, the effect can be highly non-linear

and non-repeatable, which causes the failure of the

temperature compensating algorithm in flight. The

rubber sheet can modify the local wall surface. The

rubber sheet is not suitable for skin-friction

measurements in hot flows, such as a scramjet

combustor. Because of these problems, semiconductor

strain gages, as well as rubber or polymer materials,

should be avoided in the construction of skin-friction

gages used in flight test applications.

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Localpressuregradient

parameter,∆

020533Flight conditions

.020

.015

.010

.005

0

–.005

–.010

–.015

–.020

RakeSkin-friction gagePreston tubeUpper limit, 6-percent maximum errorLower limit, 6-percent maximum error

Figure 14. Local pressure gradient on the sensor plate.

18

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020534y+

Ln(y+)/0.41 + 5.00

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

(a) Ascending portion of the flight.

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020535y+

Ln(y+)/0.41 + 5.00

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

(b) Descending portion of the flight.

Figure 15. Boundary-layer velocity profiles at 15,000 ft, in wall units.

19

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020536y+

Ln(y+)/0.41 + 5.00

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90


10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020537y+

Ln(y+)/0.41 + 5.00

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90



20

10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020538y+

Ln(y+)/0.41 + 5.00

Mach 0.70

Mach 0.80

Mach 0.90


10

15

20

25

u+

30

35

100 1,000 10,000 100,000

020539y+

Ln(y+)/0.41 + 5.00

Mach 0.70

Mach 0.80



21

0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

1.0

020587

.8


0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

1.0

020588

.8


Figure 18. Boundary-layer velocity profiles at 15,000 ft.

22

0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

1.0

020589

.8


0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.60

Mach 0.70

Mach 0.80

Mach 0.90

1.0

020590

.8



23

0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.70

Mach 0.80

Mach 0.90

1.0

020591

.8


0

.2

.4

.6

y/δ

.8

1.0

.2 .4 .6u/ue

1/7th-power law

Mach 0.70

Mach 0.80

1.0

020592

.8



24

Karman-SchoenherrTheory

Clauser Plot(Fenter-Stalmach)

Preston Tube(Bradshaw-Unsworth)

Skin-friction gage(compensated)

Skin-friction gage(uncompensated)Shear,

psf

10

0

8

6

4

2

0

–2

–4

–6

–8

–10

020540

5,000 10,000 15,000 20,000Reynolds number based on momentum thickness

25,000 30,000 35,000

Figure 21. Comparison of skin-friction gage measurements in flight with other methods.

Gro

un

d

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15K

Ratio ofrubber-sheettemperature

to walltemperature

Flight conditions

1.5

1.4

1.3

1.2

1.1

1.0

.9

.8

.7

.6

.5

020593

Figure 22. Ratio of skin-friction gage rubber-sheet temperature to the surrounding wall temperature.

25

Conclusion

An 8-in.-square boundary-layer sensor panel has beendeveloped for in-flight evaluation of skin-friction gagesand other near-wall flow sensors on the NASA DrydenFlight Research Center F-15B/Flight Test Fixture (FTF).Instrumentation on the sensor panel includes aboundary-layer rake, temperature sensors, staticpressure taps, and a Preston tube. Space is also availablefor skin-friction gages or other near-wall flow sensors.

Pretest analysis of previous F-15B/FTF flight data hasidentified flight conditions suitable for evaluatingskin-friction gages. At subsonic Mach numbers, theboundary layer over the sensor panel closelyapproximates the two-dimensional (2D),law-of-the-wall turbulent boundary layer, andskin-friction estimates from the Preston tube and therake (using the Clauser plot method) can be used toevaluate skin-friction gages and other near-wall sensors.At supersonic Mach numbers, the boundary layer overthe sensor panel becomes complex, and other means ofmeasuring skin friction are needed to evaluate theaccuracy of new skin-friction gages.

Results from the flight test of a new rubber-dampedskin-friction gage confirm that at subsonic Machnumbers, nearly 2D, law-of-the-wall turbulent boundarylayers exist over the sensor panel location. Theboundary layers are in good agreement with both thelaw of the wall and the 1/7th power law. At subsonicMach numbers, the high quality of flows over the sensorpanel enables in-flight evaluation of skin-friction gagesand other near-wall sensors. In-flight evaluation of anew rubber-damped skin-friction gage prototypeshowed that the gage did not work in flight. Theskin-friction gage measurements did not agree with thecontrol measurements and theory. The gage failure inflight was probably caused by the temperaturesensitivity of both the rubber sheet and thesemiconductor strain gages used in the skin-frictiongage. Because of these problems, semiconductor straingages, as well as rubber or polymer materials, should beavoided in the construction of skin-friction gages usedin flight test applications.

References

1Garringer, Darwin J., and Edwin J. Saltzman, FlightDemonstration of a Skin-Friction Gage to a Local MachNumber of 4.9, NASA TN D-3830, Feb. 1967.

2Quinn, Robert D., and Frank V. Olinger,Flight-Measured Heat Transfer and Skin Friction at a

Mach Number of 5.25 and at Low Wall Temperatures,NASA TM X-1921, Nov. 1969.

3Fisher, David F., and Edwin J. Saltzman, Local SkinFriction Coefficients and Boundary-Layer ProfilesObtained in Flight From the XB-70-1 Airplane at MachNumbers up to 2.5, NASA TN D-7220, June 1973.

4Quinn, Robert D., and Leslie Gong, In-FlightBoundary-Layer Measurements on a Hollow Cylinder ata Mach Number of 3.0, NASA TP-1764, 1980.

5Rausch, V., C. McClinton, and J. Sitz, “Hyper-XProgram Overview,” ISABE 99-7213, Sept. 1999.

6Richwine, David M., F-15B/Flight Test Fixture II: ATest Bed for Flight Research, NASA TM-4782, Dec.1996.

7Hopkins, Edward J., and Mamoru Inouye, “AnEvaluation of Theories for Predicting Turbulent SkinFriction and Heat Transfer on Flat Plates at Supersonicand Hypersonic Mach Numbers,” AIAA Journal, vol. 9,no. 6, June 1971, pp. 993-1003.

8Bui, Trong T., David L. Oates, and Jose C. Gonsalez,Design and Evaluation of a New Boundary-Layer Rakefor Flight Testing, NASA TM 2000-209014, Jan. 2000.

9Hopkins, Edward J., and Earl R. Keener, Study ofSurface Pitots for Measuring Turbulent Skin Friction atSupersonic Mach Numbers - Adiabatic Wall, NASA TND-3478, July 1966.

10Allen, Jerry M., “Critical Preston-Tube Sizes,”Journal of Aircraft, vol. 7, no. 3, May-June, 1970, pp.285-287.

11White, Frank M., Viscous Fluid Flow, 2nd ed.,McGraw-Hill, Inc., Boston, Massachusetts, 1991.

12Allen, Jerry M., and Dorothy H. Tudor, Charts forInterpolation of Local Skin Friction From ExperimentalTurbulent Velocity Profiles, NASA SP-3048, 1969.

13Allen, Jerry M., Use of Baronti-LibbyTransformation and Preston Tube Calibrations toDetermine Skin Friction from Turbulent VelocityProfiles, NASA TN D-4853, Nov. 1968.

14Bradshaw, P. and K. Unsworth, “Comment onEvaluation of Preston Tube Calibration Equations inSupersonic Flow,” AIAA Journal, vol. 12, no. 9, Sept.1974, pp. 1293-1295.

26

15Allen, Jerry M., “Reply by Author to P. Bradshawand K. Unsworth,” AIAA Journal, vol. 12, no. 9, Sept.1974, pp. 1295-1296.

16Allen, Jerry M., Evaluation of Compressible-FlowPreston Tube Calibrations, NASA TN D-7190, May1973.

17Magill, Samantha, Matthew MacLean, JosephSchetz, Rakesh Kapania, Alexander Sang, and WadePulliam, “Study of Direct-Measuring Skin-FrictionGauge with Rubber Sheet for Damping,” AIAA Journal,vol. 40, no. 1, Jan. 2002, pp. 50-57.

18Sang, Alexander K., and Joseph A. Schetz, “Studyof Rubber Damped Skin Friction Gages for TransonicFlight Testing,” 40th Aerospace Sciences Meeting andExhibit, AIAA-2002-0533, Jan. 2002.

19Patel, V. C., “Calibration of the Preston tube andlimitations on its use in pressure gradients,” Journal ofFluid Mechanics, vol. 23, part 1, pp. 185-208, 1965.

20Frei, D. and H. Thomann, “Direct measurements ofskin friction in a turbulent boundary layer with a strongadverse pressure gradient,” Journal of Fluid Mechanics,vol. 101, part 1, pp. 79-95, 1980.

21Westkaemper, John C., “Step-Temperature Effectson Direct Measurements of Drag,” AIAA Journal, vol. 1,no. 7, Jul. 1963, pp. 1708-1710.

22Voisinet, Robert L. P., “Temperature Step Effects onDirect Measurement of Skin-Friction Drag,”International Aerodynamics Testing Conference,AIAA-78-779, April 1978.

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NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18298-102

In-Flight Capability for Evaluating Skin-Friction Gages and OtherNear-Wall Flow Sensors

WU 710-55-04-RR-00-000

Trong T. Bui, Brett J. Pipitone, and Keith L. Krake

NASA Dryden Flight Research CenterP.O. Box 273Edwards, California 93523-0273

H-2518

National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA/TM-2003-210738

An 8-in.-square boundary-layer sensor panel has been developed for in-flight evaluation of skin-friction gages and othernear-wall flow sensors on the NASA Dryden Flight Research Center F-15B/Flight Test Fixture (FTF). Instrumentation onthe sensor panel includes a boundary-layer rake, temperature sensors, static pressure taps, and a Preston tube. Space is alsoavailable for skin-friction gages or other near-wall flow sensors. Pretest analysis of previous F-15B/FTF flight data hasidentified flight conditions suitable for evaluating skin-friction gages. At subsonic Mach numbers, the boundary layer overthe sensor panel closely approximates the two-dimensional (2D), law-of-the-wall turbulent boundary layer, andskin-friction estimates from the Preston tube and the rake (using the Clauser plot method) can be used to evaluateskin-friction gages. At supersonic Mach numbers, the boundary layer over the sensor panel becomes complex, and othermeans of measuring skin friction are needed to evaluate the accuracy of new skin-friction gages. Results from the flight testof a new rubber-damped skin-friction gage confirm that at subsonic Mach numbers, nearly 2D, law-of-the-wall turbulentboundary layers exist over the sensor panel. Sensor panel data also show that this new skin-friction gage prototype does notwork in flight.

Boundary layer, Flight test, Flow sensors, F-15B, Skin friction

A03

31

Unclassified Unclassified Unclassified Unlimited

February 2003 Technical Memorandum

Presented at the 41st AIAA Aerospace Sciences Meeting and Exhibit, Jan. 6-9, 2003, Reno NV asAIAA-2003-0741

Unclassified—UnlimitedSubject Category 34

This report is available at http://www.dfrc.nasa.gov/DTRS/

in-flight capability for evaluating skin-friction gages and other … · 2013-06-27 ·...

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